Patrick George and Raymond Beach

Glenn Research Center, Cleveland, Ohio

Beamed-Energy Propulsion (BEP) Study

NASA/TM—2012-217014

February 2012

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Patrick George and Raymond Beach

Glenn Research Center, Cleveland, Ohio

Beamed-Energy Propulsion (BEP) Study

NASA/TM—2012-217014

February 2012

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Approved for Public Release, Distribution Unlimited

Notice for Copyrighted Information 

This manuscript is a joint work of employees of the National Aeronautics and Space Administration and employees of 

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Aeronautics and Space Administration. The United States Government may prepare derivative works, publish or reproduce 

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NASA/TM—2012-217014 iii 

Contents 

1.0 EXECUTIVE SUMMARY .................................................................................................................. 1

2.0 STUDY SYNOPSIS ............................................................................................................................. 2

2.1 Study Background ..................................................................................................................... 2

2.1.1 Rationale ....................................................................................................................... 2

2.1.2 Previous Component Demonstrations .......................................................................... 3

2.1.3 Study Scope .................................................................................................................. 4

2.1.4 End Goals of Study ...................................................................................................... 4

2.2 Study Approach ......................................................................................................................... 4

2.3 Study Findings ........................................................................................................................... 6

2.3.1 Launch DRMs .............................................................................................................. 6

2.3.2 Space Design Reference Missions (DRMs) ............................................................... 19

2.4 Unexpected Findings ............................................................................................................... 26

2.4.1 Spacecraft Remote Operation ..................................................................................... 26

2.4.2 Propulsion Engine ...................................................................................................... 26

2.5 BEP Development Direction ................................................................................................... 27

2.6 Future Potential of BEP ........................................................................................................... 30

3.0 STUDY REPORTS ............................................................................................................................ 31

3.1 Launch DRMs ......................................................................................................................... 31

3.1.1 Laser Optical Collaborative Modeling for Parametric Assessment of Space 

Systems (COMPASS) Report .................................................................................... 32

3.1.2 Laser Optical Consultant Report—Design Reference Mission 1–A ........................ 112

3.1.3 Laser Optical Cost Analysis ..................................................................................... 188

3.1.4 Laser Thermal COMPASS Report ........................................................................... 194

3.1.5 Laser Thermal Consultant Report—Design Reference Mission 1–B ...................... 228

3.1.6 Laser Thermal Cost .................................................................................................. 285

3.1.7 Millimeter-Wave Thermal Collaborative Modeling for Parametric 

Assessment of Space Systems (COMPASS) Report ............................................... 291

3.1.8 Millimeter-Wave Thermal Consultant Report—Design Reference Mission 

1–C ........................................................................................................................... 324

3.1.9 Millimeter-Wave Thermal Cost Estimates ............................................................... 370

3.2 Space Design Reference Missions......................................................................................... 376

3.2.1 Low-Earth-Orbit to Geosynchronous-Earth-Orbit Collaborative Modeling 

for Parametric Assessment of Space Systems (COMPASS) Report ........................ 376

3.2.2 Low-Earth-Orbit to Geosynchronous-Earth-Orbit Consultant Reports—

Design Reference Mission 2–A ............................................................................... 433

3.2.3 Low-Earth-Orbit to Geosynchronous-Earth-Orbit Consultant Reports—

Design Reference Mission 2–B ................................................................................ 442

3.2.4 Low-Earth-Orbit to Geosynchronous-Earth-Orbit Cost ........................................... 445

3.2.5 Deep Space Mission Design Lab (MDL) Report ..................................................... 447

3.2.6 Deep Space Consultant Report—Design Reference Mission 3 ................................ 505

4.0 CONCLUDING REMARKS ........................................................................................................... 515

APPENDIX A.—ACRONYMS AND DEFINITIONS ............................................................................ 517

APPENDIX B.—CONSULTANT RESUMES ........................................................................................ 521

APPENDIX C.—THRUST AUGMENTATION OF AN IN-SPACE PROPULSION SYSTEM ........... 569

APPENDIX D.—MICROWAVE ENERGY ISSUES ............................................................................. 575

APPENDIX E.—ABL EXPERIMENT .................................................................................................... 593

APPENDIX F.—ENVIRONMENTAL CONSIDERATIONS FOR A GROUND-BASED BEP 

STATION ......................................................................................................................................... 597

APPENDIX G.—EFFECTS OF BEAMING ENERGY THROUGH THE ATMOSPHERE .................. 619
NASA/TM—2012-217014 iv 

G.9.2.1 Continuous Wave Source ......................................................................................... 642

G.9.2.2 Pulsed Source ........................................................................................................... 644

APPENDIX H.—SUMMARY PRESENTATION ................................................................................... 649

List of Tables 

TABLE 2.1.—OTHER FOMS FOR THE DRMS ...................................................................................... 10

TABLE 2.2.—NONRECURRING AND RECURRING COSTS FOR LAUNCH VEHICLES .............. 15

TABLE 2.3.—LAUNCH VEHICLE AND INFRASTRUCTURE COSTSA ............................................. 18

TABLE 2.4.—LEO–GEO TUG CONOPS ................................................................................................. 21

TABLE 2.5.—RECURRING AND NONRECURRING COSTS FOR LEO–GEO DRM ........................ 22

TABLE 2.6.—RECURRING AND NONRECURRING COSTS FOR DEEP SPACE DRM ................... 25

TABLE 3.1.—DRM 1–A SUMMARY OF REQUIREMENTS AND ASSUMPTIONS ........................ 117

TABLE 3.2.—MASS BREAKDOWN FOR 120-CM DRM 1–A SPACECRAFT ................................. 127

TABLE 3.3.—TRADE STUDY—DATA RATE AND COVERAGE VERSUS BEAM WIDTH 

TO GATEWAY .................................................................................................................................. 131

TABLE 3.4.—TRADE STUDY—DATA RATE AND COVERAGE VERSUS BEAM WIDTH 

TO USER .......................................................................................................................................... 131

TABLE 3.5.—PEL FOR 120-CM LIGHTCRAFT LAUNCHER ............................................................ 132

TABLE 3.6.—THERMAL ANALYSIS ASSUMPTIONS—SPACECRAFT INSULATION ................ 138

TABLE 3.7.—MEL FOR 120-CM LIGHTCRAFT LAUNCHER FOR SIX CUBESATS, 

DEPLOYED FROM DUAL P–POD ON ORBIT ............................................................................ 142

TABLE 3.8.—MEL MASS BREAKDOWN FOR 35-CM VERSUS 120-CM BEPS 

LIGHTCRAFT ................................................................................................................................. 143

TABLE 3.9.—MEL BREAKDOWN FOR AFRL 35-CM NANOSAT LAUNCHER (REF. 5) ............ 144

TABLE 3.10.—EXISTING SOLID STATE, DIODE-PUMPED LASERS ............................................ 145

TABLE 3.11.—ISO 14644-1 CLEANROOM STANDARDS; (CLASS 1 IS PROPOSED FOR 

THE VAB) ....................................................................................................................................... 155

TABLE 3.12.—SSDP LASER LAUNCH SYSTEM SPECIFICATIONS AND COST ......................... 163

TABLE 3.13.—LIFE-CYCLE COST MODEL FOR 35-CM LIGHTCRAFT LAUNCHED WITH 

10-MW SOLID-STATE LASER (REF. 19) .................................................................................... 165

TABLE 3.14.—MISSION ∆V SUMMARY FOR 1.2-M LTD (DRM 1–A) USING OTIS .................... 171

TABLE 3.15.—MASS, ∆V, AND PROPELLANT SUMMARY FOR 1.2-M LTD ............................... 171

TABLE 3.16.—HX PARAMETERS FOR CURRENT (DRM–1B) DESIGN ........................................ 231

TABLE 3.17.—ISSUES FOR LASER HX .............................................................................................. 232

TABLE 3.18.—PROPERTIES AND MASS BREAKDOWN OF CONCEPTUAL 100-MW HX 

LAUNCH VEHICLE AND DRM–1B LASER HX LAUNCH VEHICLE ..................................... 236

TABLE 3.19.—DELTAS IN MASS FOR NON-DROP-TANK VERSION OF COMPASS 

DESIGN ........................................................................................................................................... 238

TABLE 3.20.—CHARACTERISTICS OF CHEVROLET VOLT BATTERY PACK ........................... 249

TABLE 3.21.—LAUNCH VEHICLE COST PER KILOGRAM DRY MASS ....................................... 261

TABLE 3.22.—CAPITAL COST FOR A NOMINAL 60-KW MAIN-LASER BEAM MODULE ...... 266

TABLE 3.23.—CAPITAL COST FOR MAIN LASER SITE WITH 2500 BEAM MODULES ............ 266

TABLE 3.24.—CAPITAL COST FOR A NOMINAL 200-KW BOOST-LASER BEAM 

MODULE ......................................................................................................................................... 267

TABLE 3.25.—CAPITAL COST FOR BOOST LASER SITE WITH 1200 BEAM MODULES .......... 267

TABLE 3.26.—CAPITAL COST FOR LAUNCH SITE ......................................................................... 268

TABLE 3.27.—COST PER LAUNCH FOR 1000 LAUNCHES/YEAR, 80-KG PAYLOAD................ 269

TABLE 3.28.—TRL LEVELS AND BASIS FOR HX LAUNCH SYSTEMS ....................................... 270

TABLE 3.29.—THE THREE DESIGN POINTS ASSOCIATED WITH DRM–1C ............................... 326
NASA/TM—2012-217014 v 

TABLE 3.30.—FIGURES OF MERIT FOR THE MICROWAVE THERMAL ROCKET 

VERSUS OTHER CONCEPTS FOR COMPARISON [RASCAL, RP1, ROCKET 

PROPELLANT 1.] ........................................................................................................................... 331

TABLE 3.31.—MARGINS OF OTHER CONCEPT STUDIES ............................................................. 332

TABLE 3.32.—TECHNOLOGY MARGINS PRIORITIZED BY GREATEST POTENTIAL 

IMPACT FIRST ............................................................................................................................... 333

TABLE 3.33.—PAYLOAD DESCRIPTION FOR EACH OF THE DESIGN POINTS 

ASSOCIATED WITH DRM–1C ..................................................................................................... 336

TABLE 3.34.—BEAM FACILITY POINT DESIGNS ........................................................................... 340

TABLE 3.35.—COMPARISON OF APERTURE PARAMETERS ....................................................... 342

TABLE 3.36.—CONOPS FOR ORBCOMM LAUNCH ......................................................................... 344

TABLE 3.37.—ASSUMPTIONS USED IN THE BEAM FACILITY COST CALCULATION ........... 345

TABLE 3.38.—ASSUMPTIONS USED IN THE BEAM FACILITY COST CALCULATION ........... 347

TABLE 3.39.—KEY FACILITIES AND MILESTONES ....................................................................... 351

TABLE 3.40.—PRELIMINARY EXPERIMENT ROSTER FOR THE GYROTRON BEAM 

FACILITY ........................................................................................................................................ 354

TABLE 3.41.—WEATHER AT THE CONUS SITE EXAMPLES GIVEN IN FIGURE 3.131) ........... 362

TABLE 3.42.—CONOPS FOR LEO–GEO LASER TRANSPORT ....................................................... 438

TABLE 3.43.—TECHNOLOGY READINESS ...................................................................................... 440

TABLE 3.44.—CONOPS FOR LEO–GEO LASER TRANSPORT 2–B ................................................ 443

TABLE 3.45.—TECHNOLOGY READINESS FOR DRM 2–B ............................................................ 444

TABLE C.2.— AUGMENTED THRUSTER SYSTEM PERFORMANCE ........................................... 571

TABLE C.3.—THRUSTER SYSTEM PARAMETERS ......................................................................... 571

TABLE C.4.—MISSION PARAMETERS FOR AUGMENTED AND AUGMENTED 

THRUSTER SYSTEMS .................................................................................................................. 573

TABLE D.1.—POWER BEAMING REQUIREMENTS, VARIABLES, AND ASSUMPTIONS ......... 577

TABLE D.2.—REQUIRED TRANSMITTER EIRPS TO ACHIEVE A MINIMUM OF 30 

MW/M2 AT THE MTR .................................................................................................................... 579

TABLE D.3.—PHASED ARRAY ANTENNA ADVANTAGES AND DISADVANTAGES ............... 579

TABLE D.4.—PHASED ARRAY ANTENNA INPUT PARAMETERS ............................................... 589

TABLE F.1.—EXPOSURE LIMITS BY WAVELENGTH AND EXPOSURE TYPE 

(REPRODUCED FROM SOURCE 3) ............................................................................................. 601

TABLE G.1.—THRESHOLD POWER LEVELS BEYOND WHICH PROPAGATION 

EFFECTS OCCUR AT INFRARED AND MILLIMETER WAVELENGTHS FOR 

VARIOUS OUTPUT APERTURE RADII, REFF ............................................................................. 623

TABLE G.2.—BEAM RADII AT TARGETS IN LOW EARTH ORBIT (LEO) AND 

GEOSYNCHRONOUS ORBIT (GEO) FOR VARIOUS OUTPUT APERTURE SIZES. ............ 627

List of Figures 

FIGURE 2.1.—BEAMED THERMAL SYSTEM EFFICIENCY—LAUNCHING PAYLOADS 

AND SPACE PROPULSION. DDCU, DC–DC CONVERSION UNIT; PPU, POWER 

PROCESSING UNIT. .......................................................................................................................... 5

FIGURE 2.2.—TWO SIZES OF LASER OPTICAL VEHICLES STUDIED. ............................................ 6

FIGURE 2.3.—FUNCTION OF EXTERNAL SURFACES IN AIR-BREATHING 

PROPULSION MODE......................................................................................................................... 7

FIGURE 2.4.—LASER THERMAL VEHICLE. LASER THERMAL ROCKET 

CONFIGURATION SHOWING THE LASER RECEIVER HX MOUNTED ON ITS SIDE. .......... 8

FIGURE 2.5.—MILLIMETER-WAVE THERMAL DRM. MILLIMETER-WAVE THERMAL 

ROCKET CONFIGURATION SHOWING A CIRCUMFERENTIAL RECEIVER HX. .................. 9
NASA/TM—2012-217014 vi 

FIGURE 2.6.—ORBCOMM SPACECRAFT. COURTESY OF ORBITAL SCIENCES 

CORPORATION; USED WITH PERMISSION. ............................................................................. 10

FIGURE 2.7.—PEGASUS XL. .................................................................................................................. 10

FIGURE 2.8.—LAUNCH VEHICLES COMPARED BY SIZE. .............................................................. 10

FIGURE 2.9.—CONFIGURATION OF LASER OPTICAL GROUND FACILITIES. ........................... 11

FIGURE 2.10.—CONCEPT FOR BEAM-COMBINING SYSTEM WITH CENTRAL 

ROTATING OUTPUT MIRROR. ..................................................................................................... 12

FIGURE 2.11.—TYPICAL LAUNCH CONCEPT OF OPERATIONS WITH L1 (BOOST) 

LASER ARRAY. MARKERS: (0) LAUNCH. (1) MAIN ARRAY “SEES” VEHICLE AND 

BEGINS TRACKING. (2) MAIN ARRAY TURNS ON; L1 ARRAY AT MAXIMUM 

FULL-POWER RANGE. (3) MAIN ARRAY AT FULL POWER; L1 ARRAY. ............................ 12

FIGURE 2.12.—CONFIGURATION OF LASER THERMAL GROUND FACILITIES. ....................... 13

FIGURE 2.13.—CONFIGURATION OF MILLIMETER-WAVE THERMAL GROUND 

FACILITIES. ...................................................................................................................................... 13

FIGURE 2.14.—MILLIMETER-WAVE THERMAL LAUNCH SYSTEM AND ASCENT 

TRAJECTORY. ................................................................................................................................. 14

FIGURE 2.15.—LASER POWER LEVEL VERSUS COST. KARE TECHNICAL 

CONSULTING, USED WITH PERMISSION. ................................................................................. 16

FIGURE 2.16.—LASER OPTICAL ISSUES. ........................................................................................... 17

FIGURE 2.17.—LEO–GEO TUG DIAGRAM. RCS, REACTION CONTROL SYSTEM. ..................... 20

FIGURE 2.18.—LAUNCH ELEMENTS AND SEQUENCE. .................................................................. 21

FIGURE 2.19.—STOWED DRM 3 PROPULSION MODULE IN FALCON 9 FAIRING. ..................... 23

FIGURE 2.20.—SEQUENCE OF OPERATIONS FOR DRM 3: BEP-ENHANCED GALILEO 

MISSION. .......................................................................................................................................... 24

FIGURE 2.21.—BEP CONCEPT—INTEGRATED PROPULSION, POWER, 

COMMUNICATIONS, AND NAVIGATIONAL/ATTITUDE CONTROL ON THE SAME 

LASER BEAM (PV, PHOTOVOLTAICS; RCS, REACTION CONTROL SYSTEM). .................. 26

FIGURE 2.22.—BEP OPTICAL PLASMA ENGINE. .............................................................................. 27

FIGURE 2.23.—TIMELINE OF BEAMED-ENERGY DEVELOPMENT. ............................................. 27

FIGURE 2.24.—BEP DEVELOPMENT APPLICATIONS FROM LOW TO HIGH POWER. .............. 28

FIGURE 2.25.—DEMONSTRATIONS TO DEVELOP NEAR-TERM REMOTE POSITIONAL 

CONTROL TECHNOLOGY. ............................................................................................................ 28

FIGURE 2.26.—IN-SPACE FLIGHT DEMONSTRATION FROM THE INTERNATIONAL 

SPACE STATION TO FREE FLYER. .............................................................................................. 29

FIGURE 2.27.—DEMONSTRATIONS TO DEVELOP NEAR-TERM OPTICAL PLASMA 

ENGINE TECHNOLOGY. ................................................................................................................ 30

FIGURE 3.1.—12-MW LASER LAUNCH FACILITY AT WSMR, NEW MEXICO (ARTIST 

CONCEPT BY P. RAWLINGS). ..................................................................................................... 112

FIGURE 3.2.—LIGHTCRAFT PLASMA ENGINE PULSING IN 1-BAR LABORATORY AIR 

(PHOTOGRAPHS BY J.E. SHRYNE, III). LEFT: 14.7-CM-DIAMETER #200 VEHICLE. 

RIGHT: 23-CM #100 VEHICLE. .................................................................................................... 112

FIGURE 3.3.—ORBCOMM MESSAGING SATELLITE IN 800-KM CIRCULAR ORBIT 

(FROM REF. 4). COURTESY OF ORBITAL SCIENCES CORPORATION; USED WITH 

PERMISSION. ................................................................................................................................. 113

FIGURE 3.4.—PROGRESSIVE TECHNOLOGY ROADMAP AND POSSIBLE 

DEVELOPMENT PROGRAM: GROUND AND FLIGHT DEMONSTRATIONS 

LEADING TO 35-CM NANOSAT LAUNCH SYSTEM. .............................................................. 114

FIGURE 3.5.—FINAL LAUNCH CONFIGURATION OF 120-CM-DIAMETER DRM 1–A 

SPACECRAFT. ................................................................................................................................ 115
NASA/TM—2012-217014 vii 

FIGURE 3.6.—PRINCIPAL SPACECRAFT SYSTEMS. SA, SOLAR ARRAY; GN&C, 

GUIDANCE, NAVIGATION, AND CONTROL; C&DH, COMMAND AND DATA 

HANDLING; PMAD, POWER MANAGEMENT AND DISTRIBUTION; HE, HELIUM. ......... 116

FIGURE 3.7.—SEMITRANSPARENT VIEW OF 120-CM SPACECRAFT AND SATELLITE 

ELECTRONICS. IMU, INERTIAL MEASUREMENT UNIT. ...................................................... 119

FIGURE 3.8.—FUNCTION OF EXTERNAL SURFACES IN AIR-BREATHING 

PROPULSION MODE..................................................................................................................... 120

FIGURE 3.9.—PULSEJET ENGINE CYCLE BASED ON LSD WAVES. (DATA FOR 11- TO 

14-CM LIGHTCRAFT ENGINE TESTS ON PULSED LASER VULNERABILITY TEST 

SYSTEM (PLVTS) LASER AT WSMR.) ....................................................................................... 120

FIGURE 3.10.—PROPELLANT INJECTOR SCHEMES FOR ROCKET PULSEJET MODE 

(REF. 16). ......................................................................................................................................... 121

FIGURE 3.11.—CUTAWAY OF ENGINE “HOT SECTION,” SHOWING COOLING AND 

PROPELLANT INJECTOR SYSTEMS (REF. 16). ........................................................................ 122

FIGURE 3.12.—SIDE AND TOP VIEWS OF SHROUD AND SUPPORT STRUTS, 

EXPLAINING COOLANT SCHEME (REF. 16). ........................................................................... 122

FIGURE 3.13.—THRUST VECTORING, RCS, AND PROPELLANT INJECTOR DESIGN. ............. 122

FIGURE 3.14.—LEFT: CENTERED VERSUS OFFSET BEAMS. RIGHT: THRUST 

IMPULSE, SIDE IMPULSE, AND PITCHING ANGULAR IMPULSE ARE FUNCTIONS 

OF LATERAL BEAM OFFSET DISTANCE (REFS. 13 TO 15). .................................................. 123

FIGURE 3.15.—PERFORMANCE VARIATIONS IN AXIAL IMPULSE AND 

RESTORATIVE (I.E., BEAM-RIDING) IMPULSE VERSUS LATERAL OFFSET ARE 

ACTIVELY EXPLOITED IN THE AIM-POINT STRATEGY (REFS. 13 TO 15). ...................... 124

FIGURE 3.16.—AIM-POINT STRATEGY IN LAUNCHING A BEAM-RIDING 

LIGHTCRAFT (REFS. 13 TO 15). .................................................................................................. 124

FIGURE 3.17.—LIGHTCRAFT DRM 1–A CONFIGURATION—OBLIQUE, TOP, AND SIDE 

VIEWS. ............................................................................................................................................ 125

FIGURE 3.18.—AVIONICS FUNCTIONALITY (120-CM LAUNCHER; I/O, 

INPUT/OUTPUT). ........................................................................................................................... 128

FIGURE 3.19.—COMMUNICATION OPERATIONS DIAGRAM. ..................................................... 129

FIGURE 3.20.—LAUNCH GATEWAY SITES. ..................................................................................... 130

FIGURE 3.21.—LEFT: FLEXIBLE PV ARRAY. RIGHT: POWER CIRCUIT SCHEMATIC. ........... 132

FIGURE 3.22.—LIGHTCRAFT THERMAL SUBSYSTEMS. LEFT: DURING LAUNCH. 

RIGHT: ON-ORBIT. ....................................................................................................................... 135

FIGURE 3.23.—SURFACE MOUNT RADIATOR SIZING—THERMAL ANALYSIS. ..................... 137

FIGURE 3.24.—LEFT: SAMPLE COLD PLATE FOR ELECTRONICS COOLING. RIGHT: 

THERMAL ANALYSIS ASSUMPTIONS. .................................................................................... 137

FIGURE 3.25.—MLI FOR SPACECRAFT ON-ORBIT. ........................................................................ 138

FIGURE 3.26.—REFLECTANCE OF MULTILAYER DIELECTRIC FILM COATINGS 

VERSUS NORMALIZED FREQUENCY. ..................................................................................... 139

FIGURE 3.27.—RADIATION HEAT TRANSFER FROM ELECTRIC ARC IN 1-BAR 

ATMOSPHERE. .............................................................................................................................. 140

FIGURE 3.28.—DIFFUSE REFLECTING SURFACE FOR SCATTERING LASER LIGHT.............. 141

FIGURE 3.29.—ALTERNATIVE PAYLOAD FOR 120-CM SPACECRAFT—DUAL P-POD 

LAUNCHER. ................................................................................................................................... 142

FIGURE 3.30.—RELATIVE SCALE OF LIGHTCRAFT LAUNCHERS. LEFT: 

MICROSATELLITES. RIGHT: NANOSATELLITES. ................................................................. 143

FIGURE 3.31.—EXPONENTIAL GROWTH OF CONTINUOUS WORKING SSDP LASERS 

(REF. 20). ......................................................................................................................................... 145

FIGURE 3.32.—LLNL’S 67-KW SSHCL (REF. 23). ............................................................................. 146

FIGURE 3.33.—MERCURY HAPL AT LLNL (50- TO 100-J PULSES AT 10 HZ) (REF. 25). .......... 146
NASA/TM—2012-217014 viii 

FIGURE 3.34.—LEFT: IPG’S 10-KW SINGLE-MODE FIBER LASER WELDER. RIGHT: 

MASTER OSCILLATOR POWER AMPLIFIER (MOPA) SCHEMATIC (REF. 26). 

COPYRIGHT IPG PHOTONICS CORPORATION (OXFORD, MA) 10,000 W SINGLE 

MODE LASER; USED WITH PERMISSION. ............................................................................... 147

FIGURE 3.35.—SCATTER PLOT OF SSDP LASER SPECIFICATIONS RELATIVE TO THE 

DESIRED BEAMLINE GOAL. ...................................................................................................... 147

FIGURE 3.36.—11.5-MW RP SSDP LASER SYSTEM (24 BEAMLINES PER LEVEL, 

STACKED FOUR LEVELS DEEP FOR 96 TOTAL) (ARTIST CONCEPT BY P. 

RAWLINGS). .................................................................................................................................. 148

FIGURE 3.37.—WSMR FACILITY FOR HOUSING 12-MW LAUNCH LASER SYSTEM 

(ARTIST CONCEPT BY P. RAWLINGS). .................................................................................... 148

FIGURE 3.38.—SCHEMATIC FOR SINGLE SSDP BEAMLINE DESIGNED FOR ORBIT 

DEBRIS MITIGATION (REFS. 28 AND 29). ................................................................................ 149

FIGURE 3.39.—LEFT: DIODE BARS. RIGHT: HE GAS COOLING OF AMPLIFIER 

SLABLETS IN HAPL PROTOTYPE (REFS. 28 AND 29). ........................................................... 150

FIGURE 3.40.—HE GAS-COOLED AMPLIFIER SLABLETS AND OPTICAL SWITCHES IN 

HAPL PROTOTYPE (REF. 30). ..................................................................................................... 150

FIGURE 3.41.—ALRU DESIGN FOR FLOWING-GAS HE COOLING (REF. 30). ............................ 151

FIGURE 3.42.—HIGH-POWER VCSEL DIODE TECHNOLOGY (REF. 31). ..................................... 151

FIGURE 3.43.—ROTATING MIRROR BEAM COMBINING SYSTEM (ARTIST CONCEPT 

BY P. RAWLINGS). LEFT: PLANFORM VIEW. RIGHT: OBLIQUE VIEW. ............................ 152

FIGURE 3.44.—BEAM-COMBINING SYSTEM WITH CENTRAL ROTATING OUTPUT 

MIRROR (CENTER) AND TILT-TIP CORRECTION MIRRORS AND AO (RIGHT) 

(ARTIST CONCEPT BY P. RAWLINGS). .................................................................................... 152

FIGURE 3.45.—ROTATING MIRROR BEAM-COMBINER UNIT STABILIZED BY 

VERTICAL 12- BY 12-M OPTICAL BENCH/WALL (ARTIST CONCEPT BY P. 

RAWLINGS). .................................................................................................................................. 152

FIGURE 3.46.—BEAM DELIVERY TUBE (LEFT), 3-M TELESCOPE (CENTER), AND 

RETRACTABLE LAUNCH RIG WITH 1-M 45 TURNING FLAT (RIGHT) (ARTIST 

CONCEPT BY P. RAWLINGS). ..................................................................................................... 153

FIGURE 3.47—SEMI-TRANSPARENT VIEW OF 3-M TELESCOPE AND DOME (ARTIST 

CONCEPT BY P. RAWLINGS). ..................................................................................................... 153

FIGURE 3.48.—VIEW INSIDE BEAM DELIVERY TUNNEL, SHOWING RETRACTABLE 

45° FLAT MIRROR THAT DIVERTS THE UNEXPANDED 25- BY 25-CM LASER 

BEAM VERTICALLY UPWARD INTO A 3-M-DIAMETER TELESCOPE (ARTIST 

CONCEPT BY P. RAWLINGS). ..................................................................................................... 154

FIGURE 3.49.—AERIAL VIEW OF LASER HIGHBAY (LEFT), BEAM TUBE (CENTER), 

AND VAB WITH LAUNCH PAD (RIGHT) (ARTIST CONCEPT BY P. RAWLINGS). ........... 156

FIGURE 3.50.—LAYOUT OF VAB (ARTIST CONCEPT BY P. RAWLINGS). ................................. 156

FIGURE 3.51.—ROBOTIC HANDLING OF 120-CM LIGHTCRAFT (318-KG WET MASS) 

(ARTIST CONCEPT BY P. RAWLINGS). .................................................................................... 156

FIGURE 3.52.—POI–1000 SPIN BALANCE MACHINE (COURTESY OF SPACE 

ELECTRONICS, LLC; USED WITH PERMISSION), WITH SCALE 120-CM 

LIGHTCRAFT (REF. 32). ............................................................................................................... 157

FIGURE 3.53.—ROBOT LOADS 120-CM LIGHTCRAFT ONTO MOBILE LAUNCHER RIG 

INSIDE VAB (ARTIST CONCEPT BY P. RAWLINGS ). ............................................................ 157

FIGURE 3.54.—MOBILE LAUNCHER TRANSPORTS 120-CM LIGHTCRAFT OUT TO 

LAUNCH PAD AT END OF TRACK (ARTIST CONCEPT BY P. RAWLINGS). ...................... 158

FIGURE 3.55.—AERIAL VIEWS (ARTIST CONCEPT BY P. RAWLINGS). (A) 

LIGHTCRAFT ON LAUNCH PAD. (B) UNDER LASER BOOST. ............................................. 158
NASA/TM—2012-217014 ix 

FIGURE 3.56.—DAWN LAUNCH OF LIGHTCRAFT AT WHITE SANDS MISSILE RANGE, 

NEW MEXICO (BY L. MYRABO AND P. MOLONEY). ............................................................ 159

FIGURE 3.57.—MISSION TIMELINE ALONG TRAJECTORY FOR DRM 1–A. LTD, 

LIGHTCRAFT TECHNOLOGY DEMONSTRATOR. .................................................................. 160

FIGURE 3.58.—ON-ORBIT TIMELINE GIVES SPECIFICATIONS FOR DRM 1–A. ....................... 161

FIGURE 3.59.—RP SOLID-STATE LASER COSTS VERSUS PULSE ENERGY (FROM 

REFS. 2 AND 33). ........................................................................................................................... 162

FIGURE 3.60.—RP SOLID-STATE LASER COSTS VERSUS PRF (FROM REF. 2 AND 33). ......... 162

FIGURE 3.61.—LASER BEAM DIRECTOR COST VERSUS APERTURE DIAMETER 

(FROM REFS. 2 AND 33). .............................................................................................................. 163

FIGURE 3.62.—NASA (N) VERSUS COMMERCIAL (C) PRODUCTION COSTS FOR 35-

CM LIGHTCRAFT. ......................................................................................................................... 164

FIGURE 3.63.—LASER LAUNCH GEOMETRY FOR LIGHTCRAFT TRAJECTORY 

SIMULATIONS. .............................................................................................................................. 166

FIGURE 3.64.—NET CM VERSUS FLIGHT MACH NUMBER FOR LTD AIR-BREATHING 

MODE (REFS. 16 AND 17). (”BASELINE” 3-CM INLET GAP, 500-MW/CM2

 AT 

FOCAL RING.) ................................................................................................................................ 167

FIGURE 3.65.—THREE-DIMENSIONAL (3D) PLOT OF CM VERSUS FLIGHT MACH 

NUMBER FOR LTD AIR-BREATHING MODE (REFS. 16 AND 17). (”BASELINE” 

3-CM INLET GAP, 500-MW/CM2

 AT FOCAL RING.) ................................................................ 168

FIGURE 3.66.—EP VERSUS FLIGHT MACH NUMBER FOR LTD AIR-BREATHING 

ENGINE MODE (REFS. 16 AND 17). (“BASELINE” 3-CM INLET GAP, 500-MW/CM2

AT FOCAL RING.) ......................................................................................................................... 168

FIGURE 3.67.—PRF VERSUS FLIGHT MACH NUMBER FOR LTD AIR-BREATHING 

ENGINE MODE (REFS. 16 AND 17). (“BASELINE” 3-CM INLET GAP, 500-MW/CM2

AT FOCAL RING.) ......................................................................................................................... 168

FIGURE 3.68.—AERODYNAMIC DRAG MODEL FOR LTD VEHICLE (REFS. 16 AND 17). ....... 169

FIGURE 3.69.—BURN PHASES, RANGES, AND ANGLES FROM THE SATELLITE ORBIT 

ANALYSIS PROGRAM (SOAP) SIMULATION FOR 1.2-M LTD (DRM 1–A). ........................ 170

FIGURE 3.70.—OTIS SIMULATED LAUNCH TRAJECTORY FOR 1.2-M LTD (DRM 1–A). 

(A) ALTITUDE VERSUS TIME. (B) VELOCITY VERSUS TIME. (C) ALTITUDE 

VERSUS RANGE FROM LAUNCH PAD. .................................................................................... 170

FIGURE 3.71.—LASER POWER AND ENGINE THRUST VERSUS TIME FROM OTIS 

TRAJECTORY FOR 1.2-M LTD (DRM 1–A). .............................................................................. 171

FIGURE 3.72.—ALTITUDE VERSUS TIME FOR SIMULATED TRAJECTORIES OF 35-CM 

LCC (REF. 5). .................................................................................................................................. 174

FIGURE 3.73.—CAPTURED 1.62-ΜM LASER POWER VERSUS SLANT RANGE FROM 

FOR 35-CM LIGHTCRAFT (REF. 5). ............................................................................................ 175

FIGURE 3.74.—MACH NUMBER VERSUS TIME FOR 7-DOF SIMULATED 

TRAJECTORIES. ............................................................................................................................ 176

FIGURE 3.75.—ALTITUDE VERSUS TIME FOR 7-DOF SIMULATED TRAJECTORIES. ............. 177

FIGURE 3.76.—VELOCITY VERSUS TIME FOR 7-DOF SIMULATED TRAJECTORIES. ............. 177

FIGURE 3.77.—BEAM POWER VERSUS TIME FOR 7-DOF SIMULATED TRAJECTORIES 

FOR 120-CM LTD CRAFT. ............................................................................................................ 177

FIGURE 3.78.—BEAM POWER VERSUS TIME FOR 7-DOF SIMULATED TRAJECTORIES 

FOR 35-CM CRAFT. ....................................................................................................................... 178

FIGURE 3.79.—MACH NUMBER VERSUS ALTITUDE FOR 7-DOF SIMULATED 

TRAJECTORIES. ............................................................................................................................ 178

FIGURE 3.80.—ZENITH ANGLE VERSUS TIME FOR 7-DOF SIMULATED 

TRAJECTORIES. ............................................................................................................................ 178
NASA/TM—2012-217014 x 

FIGURE 3.81.—KEY ISSUES AFFECTING THE FEASIBILITY OF A LIGHTCRAFT 

VEHICLE AND A COMBINED-CYCLE ENGINE. ...................................................................... 180

FIGURE 3.82.—PROGRAM 1—LASER LAUNCH DEMONSTRATIONS TO 1 KM AT 

HELSTF USING 10-KW PLVTS CO2 LASER............................................................................... 182

FIGURE 3.83.—PROGRAM 2—SUBORBITAL LAUNCH DEMONSTRATIONS TO 10 TO 

30 KM AT HELSTF USING THE 105-KW JHPSSL. .................................................................... 182

FIGURE 3.84.—PROGRAM 3—SUBORBITAL LAUNCH DEMONSTRATIONS USING 

ALTB’S MULTIMEGAWATT COIL. ............................................................................................ 183

FIGURE 3.85.—PROGRAM 4—ON-ORBIT LAUNCH DEMONSTRATIONS FROM THE ISS 

USING A SPACE-RATED VERSION OF THE HAPL 1- KW LASER. ....................................... 184

FIGURE 3.86.—TYPICAL LAUNCH CONCEPT OF OPERATIONS WITH L1 (BOOST) 

LASER ARRAY. MARKERS: (0) LAUNCH. (1) MAIN ARRAY “SEES” VEHICLE AND 

BEGINS TRACKING. (2) MAIN ARRAY TURNS ON; L1 ARRAY AT MAXIMUM 

FULL-POWER RANGE. (3) MAIN ARRAY AT FULL POWER; L1 ARRAY TURNS 

OFF. (4) VEHICLE PITCHES DOWN TO CIRCULARIZE ORBIT. (5) ORBIT 

INSERTION; MAIN LASER TURNS OFF. COPYRIGHT KARE TECHNICAL 

CONSULTING; USED WITH PERMISSION. ............................................................................... 228

FIGURE 3.87.—HX THRUSTER CONCEPT; LH2, LIQUID HYDROGEN. COPYRIGHT 

KARE TECHNICAL CONSULTING; USED WITH PERMISSION. ........................................... 229

FIGURE 3.88.—CONVENTIONAL HX VERSUS LAMINAR-FLOW MICROCHANNEL HX. 

COPYRIGHT KARE TECHNICAL CONSULTING; USED WITH PERMISSION..................... 230

FIGURE 3.89.—CONCEPTUAL DESIGN OF 100-MW HX LAUNCH VEHICLE. 

COPYRIGHT KARE TECHNICAL CONSULTING; USED WITH PERMISSION..................... 235

FIGURE 3.90.—GROUND FACILITIES FOR LASER HX LAUNCH SYSTEM. MIDDLE 

IMAGE: COPYRIGHT LASERMOTIVE, LLC; USED WITH PERMISSION. RIGHT 

IMAGE: COPYRIGHT KARE TECHNICAL CONSULTING; USED WITH 

PERMISSION. ................................................................................................................................. 240

FIGURE 3.91.—BLOCK DIAGRAM OF 60-KW BEAM MODULE. COPYRIGHT KARE 

TECHNICAL CONSULTING; USED WITH PERMISSION. ....................................................... 242

FIGURE 3.92.—IPG YLS–10000–SM 10-KW SINGLE-MODE LASER. COPYRIGHT IPG 

PHOTONICS CORPORATION (OXFORD, MA) 10,000 W SINGLE MODE LASER; 

USED WITH PERMISSION. .......................................................................................................... 243

FIGURE 3.93.—ADAPTIVE OPTICS BEACON OFFSET VERSUS TIME FOR A TYPICAL 

LEO TRAJECTORY. THE NOTCH AT 420 S CATES A POINT IN THE TRAJECTORY 

WHERE THE VEHICLE PITCHES NOSE DOWN SLIGHTLY (~10) TO 

CIRCULARIZE ITS ORBIT AND IMPROVE THE LASER HX VIEWING ANGLE. ................ 245

FIGURE 3.94.—BEAM MODULES IN THE FIELD. COPYRIGHT KARE TECHNICAL 

CONSULTING; USED WITH PERMISSION. ............................................................................... 247

FIGURE 3.95.—MODULAR DATA CENTER, AS ENVISIONED IN IEEE SPECTRUM (REF. 

25). COPYRIGHT KARE TECHNICAL CONSULTING; USED WITH PERMISSION. ............ 247

FIGURE 3.96.—A 2000-MODULE ARRAY, WITH 40 BY 50 M SPACING, IS 2  1.6 KM, 

OR ABOUT THE SIZE OF A 18-HOLE GOLF COURSE. COPYRIGHT KARE 

TECHNICAL CONSULTING; USED WITH PERMISSION. ....................................................... 248

FIGURE 3.97.—BOOST LASER BEAM MODULE CONCEPT. .......................................................... 251

FIGURE 3.98.—ATMOSPHERIC TRANSMISSION VERSUS WAVELENGTH FOR A 20-KM 

HORIZONTAL PATH AT MIDLATITUDE DURING WINTER (MOLECULAR 

ABSORPTION/SCATTERING ONLY; NO PARTICULATES). .................................................. 253

FIGURE 3.99.—VARIABLE LAUNCH AZIMUTH AND DOGLEG TRAJECTORY OFFSET 

CAPABILITY WITH A SINGLE LAUNCH SITE AND LASER SITE. ....................................... 258

FIGURE 3.100.—LAWRENCE LIVERMORE NATIONAL LABORATORY LASER DIODE 

COST ESTIMATES. ........................................................................................................................ 262
NASA/TM—2012-217014 xi 

FIGURE 3.101.—HAITZ’S LAW FOR LIGHT-EMITTING DIODE (LED) PRICE AND 

PERFORMANCE VERSUS TIME. ................................................................................................ 263

FIGURE 3.102.—EXAMPLE OF SCALING ULTRAVIOLET LIGHT-EMITTING DIODE 

(LED) PRICES OVER 6 DECADES OF QUANTITY. .................................................................. 264

FIGURE 3.103.—ROADMAP FOR LAUNCH SYSTEM DEVELOPMENT AND 

DEPLOYMENT. .............................................................................................................................. 272

FIGURE 3.104.—POSSIBLE GROUND (BLUE) AND FLIGHT (GREEN) 

DEMONSTRATION PROGRAM FOR HX VEHICLE. ................................................................ 273

FIGURE 3.105.—LEFT: THE ADVENT OF GYROTRON TECHNOLOGY HAS RAISED 

THE POWER OUTPUT OF SOURCES IN THE KEY MILLIMETER-WAVELENGTH 

RANGE BY 3 ORDERS OF MAGNITUDE SINCE 1970. RIGHT: VGB–8125 

GYROTRON MANUFACTURED IN PALO ALTO BY CPI INC. THIS TUBE 

PRODUCES A 2.5-MW BEAM AT 95 GHZ. (IMAGE COURTESY OF 

COMMUNICATIONS & POWER INDUSTRIES (CPI); USED WITH PERMISSION). ............. 324

FIGURE 3.106—MICROWAVE THERMAL LAUNCH SYSTEM AND ASCENT 

TRAJECTORY ARISING FROM DRM–1C. ................................................................................. 325

FIGURE 3.107.—PYROLYTIC GRAPHITE WAS fiRST PRODUCED IN THE LATE 1800S 

FOR LAMP fiLAMENTS. LEFT: PYROLYTIC GRAPHITE SHEETING. RIGHT: 

THERMAL CONDUCTIVITY IN WATTS PER METERS KELVIN OF PYROLYTIC 

GRAPHITE SHEET VERSUS OTHER METALS (PANASONIC PGS DATA SHEET). 

COPYRIGHT ©2011 PANASONIC CORPORATION OF NORTH AMERICA; USED 

WITH PERMISSION. ...................................................................................................................... 327

FIGURE 3.108.—(A) PLAN AND SECTION VIEWS OF AN ANGLE-INTERLOCK 

ARCHITECTURE. IN THE LOOM, THE WARP YARNS ARE MOVED UP AND 

DOWN SELECTIVELY BY HARNESSES, WHEREAS THE WEFT YARNS ARE 

INSERTED BETWEEN THEM BY A FLYING SHUTTLE OR RAPIER. THE SECTION 

VIEWS SHOW HOW THE LOOM MOTIONS CAUSE WARP YARNS TO PASS RIGHT 

THROUGH THE THICKNESS OF THE RESULTING FABRIC, TYING TOGETHER 

LAYERS OF WEFT YARNS. THUS, THE ANGLE-INTERLOCK ARCHITECTURE 

SUPPRESSES DELAMINATION. (B) SCHEMATIC OF AN ADAPTATION OF AN 

ANGLE-INTERLOCK ARCHITECTURE TO FORM AN INTEGRAL SET OF ROCKET 

NOZZLE TUBES. THE TUBE WALLS HAVE AN INTERNAL ARCHITECTURE 

SIMILAR TO THAT SHOWN IN PANEL (A), WITH THE TOPS AND BOTTOMS OF 

THE TUBES JOINED AT NEXUS REGIONS FORMED IN THE LOOM. (C) FORMED 

C-SIC COMPOSITE STRUCTURE, SHOWING INTEGRALLY WOVEN STRUTS FOR 

CONNECTING TO THE ROCKET BODY AND DETAILS OF THE HIGH-DENSITY 

MATRIX FORMED WITHIN THE TUBE WALLS (REF. 1). ...................................................... 328

FIGURE 3.109.—LEFT: CONFORMAL SECTIONS OF WOVEN H2 HX FOR ROCKET 

NOZZLE COOLING. RIGHT: LINEAR SECTION OF WOVEN HX UNDERGOING 

TESTING AT GLENN. THE TESTS ENTAILED CONNECTING THE COMPOSITE 

TUBES TO METAL MANIFOLDS FOR THE SUPPLY AND RETURN OF 300-TO 400-

BAR H2 COOLANT. ONE SIDE OF THE PANEL WAS EXPOSED TO A 

HYDROGEN/OXYGEN COMBUSTION PLUME FROM A SMALL ROCKET 

COMBUSTION CHAMBER DELIVERING A FREE-STREAM TEMPERATURE 

APPROACHING 3,300 K. PANELS SURVIVED MULTIPLE CYCLES SIMULATING 

OPERATING CONDITIONS OF A BOOST ROCKET ENGINE (COLD-WALL HEAT 

FLUX OF 20 MW M−2

), WITH INDIVIDUAL CYCLES AS LONG AS 8 MIN AND 

CUMULATIVE TIMES OF 40 MIN. .............................................................................................. 329

FIGURE 3.110.—REACTION CONTROL SYSTEM (RCS) PODS. ..................................................... 329

FIGURE 3.111.—GRAPHITE ANNULAR AEROSPIKE/PLUG NOZZLE DEVELOPED AND 

TESTED BY CALSTATE LONG BEACH/GARVEY SPACECRAFT CORPORATION IN 
NASA/TM—2012-217014 xii 

2003. COPYRIGHT GARVEY SPACECRAFT CORPORATION; USED WITH 

PERMISSION. ................................................................................................................................. 330

FIGURE 3.112.—LEFT: SCANNING ELECTRON MICROSCOPE (SEM) IMAGE (CROSSSECTION) OF ZRO2 INTERFACE COATING ON CARBON FIBERS (WHITE RINGS 

AROUND EACH FIBER) SHOWING UNIFORMITY OF COVERAGE ACHIEVED BY 

THE ULTRAVIOLET CHEMICAL VAPOR DEPOSITION (UVCVD) TECHNIQUE 

(ULTRAMET.COM). RIGHT: SEM IMAGE SHOWING FIBER PULLOUT IN 

CERAMIC MATRIX COMPOSITE WITH ZRO2 INTERFACE COATING 

(ULTRAMET.COM). ...................................................................................................................... 334

FIGURE 3.113.—SATELLITE DATA INDICATING THE INTERNATIONAL LOCATIONS 

WHERE ATMOSPHERIC WATER VAPOR CONTENT IS LOWEST, AND 

THEREFORE MILLIMETER-WAVE PROPAGATION IS MOST EFFICIENT. 

COPYRIGHT EUROPEAN SPACE AGENCY (ESA); USED WITH PERMISSION. ................. 337

FIGURE 3.114.—EXAMPLES OF CONUS SITES WITH LOW LOSSES DUE TO WATER 

VAPOR ABSORPTION. SUCH SITES ARE PREDOMINANTLY IN THE 

SOUTHWEST. FOR MT. EVANS AND PIKES PEAK, A MILLIMETER-WAVE BEAM 

VERTICALLY TRANSMITTED TO SPACE WOULD LOSE 5% TO 10% OF ITS 

ENERGY TO ATMOSPHERIC ABSORPTION AT 170 GHZ. ..................................................... 338

FIGURE 3.115.—BEAM FACILITY ARRANGEMENT ALONG AN ASCENT 

TRAJECTORY. THE RED LINE IS THE ASCENT TRAJECTORY. DURING INITIAL 

ASCENT, A BOOST BEAM FACILITY WITH A MAXIMUM RANGE OF 60 KM 

PROPELS THE VEHICLE, WHICH IS THEN HANDED OVER TO A LARGER 

SUSTAIN FACILITY WITH A 125-KM RANGE AND HIGHER POWER. 

PROGRAMMATICALLY, A SUBSCALE DEMONSTRATION BEAM FACILITY 

WOULD PROBABLY BECOME THE BOOST FACILITY, AND THE DECISION TO 

PROCEED ON THE SUSTAIN FACILITY WOULD BE BASED ON ITS SUCCESSFUL 

OPERATION. THE BLUE LINE IS AN ALTERNATE ASCENT TRAJECTORY 

CONSIDERED BY THE COMPASS TEAM. BECAUSE TWICE THE RANGE WOULD 

REQUIRE TWICE THE SOURCE OR TARGET DIAMETER, IT WAS NOT FURTHER 

CONSIDERED FOR THESE POINT DESIGNS. ........................................................................... 339

FIGURE 3.116.—THREE GENERATIONS OF MILLIMETER-WAVE TELESCOPES. TOP: 

1990S-ERA GBT (LEFT) AND SECONDARY (RIGHT). MIDDLE: 2000S-ERA 

ATACAMA PATHFINDER EXPERIMENT (APEX) AT LLANO CHAJNANTOR 

(COURTESY OF APEX; USED WITH PERMISSION). VIEWED HERE AT DUSK, THE 

PANELS AND RETROREFLECTORS OF ITS ACTIVE SURFACE ARE CLEARLY 

VISIBLE. THE TELESCOPE WAS MANUFACTURED BY VERTEX 

ANTENNENTECHNIK IN DUISBURG, GERMANY. BOTTOM: 2010S-ERA 

CORNELL-CALTECH ATACAMA TELESCOPE (CCAT) (COURTESY TMT 

OBSERVATORY CORPORATION; USED WITH PERMISSION). ............................................ 341

FIGURE 3.117.—LEFT: 12-CM-DIAMETER FAST-STEERING MIRROR FOR HIGHPOWER LASER APPLICATIONS (IMAGE COURTESY OF ATA; USED WITH 

PERMISSION). A 30-CM-DIAMETER MODEL FEATURES >1-KHZ OPEN-LOOP 

BANDWIDTH, >300 RAD/S2

 ANGULAR ACCELERATION, <0.2-μM SURFACE 

ACCURACY, AND 8-MRAD MECHANICAL TRAVEL WITH <10-μRAD (2-

ARCSEC) POINTING ACCURACY. RIGHT: YAL–1A ABL PROTOTYPE IN FLIGHT, 

1.5-M TELESCOPE, MEGAWATT-CLASS 1.3-M LASER WITH UNDISCLOSED 

RANGE . .......................................................................................................................................... 342

FIGURE 3.118.—MILLIMETER-WAVE BEAM FACILITY OPTICAL ARRANGEMENT. ............. 343

FIGURE 3.119.—TOP: BEAM FACILITY COST BASED ON ASSUMPTIONS IN TABLE 

3.37. BOTTOM: EFFECT OF CHINES AND GREATER SPILLAGE. ........................................ 346
NASA/TM—2012-217014 xiii 

FIGURE 3.120.—TRL, RISK, AND UNCERTAINTY SUMMARY FOR THE MICROWAVE 

THERMAL ROCKET. FSM, FAST STEERING MIRROR; FADIS, FAST 

DIRECTIONAL SWITCH; KKV, KINETIC KILL VEHICLE; CVD, CHEMICAL VAPOR 

DEPOSITION; PSI, PRESSURE SYSTEMS, INC.; AND TPS, THERMAL PROTECTION 

SYSTEM. ......................................................................................................................................... 350

FIGURE 3.121.—MICROWAVE THERMAL LAUNCH ROADMAP. ................................................ 351

FIGURE 3.122.—LABORATORY DEMONSTRATION OF MICROWAVE THERMAL 

HEAT-EXCHANGE PRINCIPLE IN A 2.45-GHZ RESONANT CAVITY (REF. 12). 

SINGLE MULLITE THERMOCOUPLE TUBE WITH H2 PROPELLANT. DH, 1.19 MM; 

LENGTH OF TUBE, L, 36.6 CM; INLET POWER, PINLET, 1.1 BARS; INLET 

REYNOLDS NUMBER, REINLET, 280; AND PEAK WALL TEMPERATURE, TWALL, 

(1800 ± 100) K. THE END OF THE TUBE WAS DIPPED IN BRINE TO MAKE THE 

HYDROGEN PLUME VISIBLE. .................................................................................................... 352

FIGURE 3.123.—LEFT: THE 10-KW TEST STAND USES A LARGE PARABOLIC MIRROR 

TO CONCENTRATE THE ENERGY FROM A 20-KW TUNGSTEN LIGHT BULB 

ONTO A CREDIT-CARD-SIZED HX (YELLOW BOX). RIGHT: ALUMINUM PROOFOF-CONCEPT HX, SOON TO BE ATTEMPTED IN GRAPHITE. SUCH SMALL 

CHANNEL SIZES AND A CREDIT-CARD SIZE ARE NECESSARY AT THIS POWER 

LEVEL. ............................................................................................................................................ 353

FIGURE 3.124.—SCENE FROM THE 2009 SPACE ELEVATOR GAMES, WITH PRIZE 

MONEY PROVIDED BY THE NASA CENTENNIAL CHALLENGES PROGRAM. THE 

CONTAINERIZED, MOBILE, AND FLEXIBLE CONOPS WAS A GREAT SUCCESS. .......... 355

FIGURE 3.125.—GODDARD’S PROGRESS IN ROCKETS AND ALTITUDE FROM 1925 TO 

1936 SUGGESTS THE SCALE, TIMEFRAME, NUMBER OF TRIALS AND SUCCESS 

RATE AT WHICH A MICROWAVE THERMAL ROCKET MIGHT BE EXPECTED TO 

DEVELOP. EACH DATA POINT IS A LAUNCH ATTEMPT; SOME MORE 

SUCCESSFUL THAN OTHERS. ................................................................................................... 356

FIGURE 3.126.—THERMAL IMAGE OF THE ABL (RIGHT) ENGAGING A DUMMY 

TARGET (LEFT) GRAPHICALLY DEMONSTRATES THAT A SHAKY, SLEWING 

PLATFORM CAN LOCK ONTO AND TRACK A DISTANT TARGET ASCENDING TO 

ORBIT. WHAT IF THE UNCOOPERATIVE DUMMY WERE REPLACED BY A 

COOPERATIVE THERMAL ROCKET? HEL, HIGH ENERGY LASER; . ................................. 357

FIGURE 3.127.—A SINGLE BEAM FACILITY SIZED FOR A 130-KM RANGE COULD, IN 

PRINCIPLE, ILLUMINATE A ROCKET THROUGHOUT THE ENTIRE ASCENT 

TRAJECTORY. THE AREAS WITHIN THE BEAM FACILITY RANGE ARE SHADED 

PURPLE, AND THE BASELINE ASCENT TRAJECTORY IS THE RED (LOWER) 

LINE. A KEY UNKNOWN IS THE MINIMUM PRACTICAL ELEVATION ANGLE AT 

WHICH THE BEAM COULD ACQUIRE THE ROCKET ON LIFTOFF (ALT, 

ALTITUDE). .................................................................................................................................... 359

FIGURE 3.128.—COMPARISON OF TAYLOR THEORETICAL EQUATION WITH 

EXPERIMENTAL DATA (REF. 13). ............................................................................................. 360

FIGURE 3.129.—PRESSURE VERSUS RF BREAKDOWN FIELD STRENGTH AT 2.4 GHZ 

(REF. 14). SOLID SYMBOLS DENOTE BULK AIR, AND OPEN SYMBOLS DENOTE 

WITH SURFACE. ........................................................................................................................... 360

FIGURE 3.130.—RESULTS OF BREAKDOWN MEASUREMENTS WITH A HOT WIRE IN 

THE PRESENCE OF GAS FLOW (REF. 15). ................................................................................ 361

FIGURE 3.131.—LAUNCH TIME VERSUS RANGE OF INTERSECTING SATELLITE. 

EACH COLOR REPRESENTS A SINGLE DAY OF LAUNCHES (I.E., THE 8 DAYS 

ARE OVERLAID). THE NORTH AMERICAN AEROSPACE DEFENSE COMMAND 

(NORAD) SATELLITE ID IS LISTED ALONG WITH THE INTERSECTION 

DURATION AND RANGE. ............................................................................................................ 362
NASA/TM—2012-217014 xiv 

FIGURE 3.132.—LAUNCH TIME VERSUS INVERSE RANGE SQUARED (1/R2) OF 

INTERSECTING SATELLITE. EACH COLOR REPRESENTS A SINGLE DAY OF 

LAUNCHES (I.E., THE 8 DAYS ARE OVERLAID). ................................................................... 363

FIGURE 3.133.—PROBABILITY OF AN INTERSECTION HAPPENING IN EACH HALF 

HOUR PERIOD, OVER THE 8 DAYS. .......................................................................................... 363

FIGURE 3.134.—HISTOGRAM OF CLEAR LAUNCH WINDOWS—PERIODS THAT HAVE 

NO INTERSECTIONS. OVER THE 8 DAYS THERE WERE 29 PERIODS WITH MORE 

THAN 1 HR OF CLEARANCE. ..................................................................................................... 364

FIGURE 3.135.—HISTOGRAM SHOWING THE DISTRIBUTION OF INTERSECTION 

MAXIMUM INTENSITY, IMAX, DURING THE 8 DAYS. NOTE THAT MOST DAYS 

HAVE IMAX LESS THAN 100 W/M2

 (I.E., 8% OF SOLAR CONSTANT). .................................. 364

FIGURE 3.136.—HISTOGRAM OF THE MAXIMUM ENERGY INCIDENT ON A 

SATELLITE PER SQUARE METER DURING AN INTERSECTION. THIS WAS 

CALCULATED USING IMAX  INTERSECTION DURATION. .................................................. 365

FIGURE 3.137.—LAUNCH TIME VERSUS RANGE OF INTERSECTING SATELLITE. 

EACH COLOR REPRESENTS A SINGLE DAY OF LAUNCHES (I.E., THE 8 DAYS 

ARE OVERLAID). THE NORAD SATELLITE ID IS LISTED ALONG WITH THE 

INTERSECTION DURATION AND RANGE. .............................................................................. 365

FIGURE 3.138.—LAUNCH TIME VERSUS RANGE OF INTERSECTING SATELLITE. 

EACH COLOR REPRESENTS A SINGLE DAY OF LAUNCHES (I.E., THE 8 DAYS 

ARE OVERLAID). THE NORAD SATELLITE ID IS LISTED ALONG WITH 

INTERSECTION DURATION AND RANGE. .............................................................................. 366

FIGURE 3.139.—PROBABILITY OF AN INTERSECTION HAPPENING IN EACH HALFHOUR PERIOD, OVER THE 8 DAYS. .......................................................................................... 366

FIGURE 3.140.—HISTOGRAM OF CLEAR LAUNCH WINDOWS—PERIODS THAT HAVE 

NO INTERSECTIONS. OVER THE 8 DAYS, THERE WERE 41 PERIODS WITH MORE 

THAN 1 HR OF CLEARANCE. ..................................................................................................... 367

FIGURE 3.141.—HISTOGRAM SHOWING THE DISTRIBUTION OF INTERSECTION 

MAXIMUM INTENSITY DURING THE 8 DAYS. NOTE THAT MOST HAVE IMAX

LESS THAN 200 W/M2

 (I.E., 15% OF SOLAR CONSTANT). ..................................................... 367

FIGURE 3.142.—HISTOGRAM OF THE MAXIMUM ENERGY INCIDENT ON A 

SATELLITE PER SQUARE METER DURING AN INTERSECTION. THIS WAS 

CALCULATED USING IMAX  T, WHERE T IS THE INTERSECTION DURATION. ............... 368

FIGURE 3.143.—CONCEPTUAL TUG AND PROPELLANT STAGE. RCS, REACTION 

CONTROL SYSTEM. ..................................................................................................................... 434

FIGURE 3.144.—CONCEPTUAL TUG, PROPELLANT STAGE, AND GEO SATELLITES. ........... 434

FIGURE 3.145.—HALF-SYMMETRIC (HS) AXICON MIRROR ANNULAR RESONATOR 

FOR POSSIBLE DPAL EMITTER. ................................................................................................ 435

FIGURE 3.146.—STARFIRE OPTICAL RANGE (SOR). ..................................................................... 436

FIGURE 3.147.—30-METER-TELESCOPE (TMT). COURTESY TMT OBSERVATORY 

CORPORATION; USED WITH PERMISSION. ............................................................................ 437

FIGURE 3.148.—TUG RENDEZVOUS WITH PROPELLANT AND GEO SATELLITES. ............... 437

FIGURE 3.149.—LEO–GEO LASER TRANSPORT ORBITAL TRANSFER. ..................................... 438

FIGURE 3.150.—LEO–GEO LASER TRANSPORT STOWED FOR LAUNCH. ................................ 439

FIGURE 3.151.—PROPOSED TECHNOLOGY ROADMAP. ............................................................... 441

FIGURE 3.152.—GRAPHIC REPRESENTATION OF BEPS DRM 3 TRANSFER 

TRAJECTORY FROM EARTH ORBIT TO JUPITER DIRECT TRANSFER ORBIT 

SHOWING PATH OF BEP ACCELERATION. ............................................................................. 505

FIGURE 3.153.—DRAWING OF STOWED DRM 3 PROPULSION MODULE IN FALCON 9 

FAIRING. ......................................................................................................................................... 506
NASA/TM—2012-217014 xv 

FIGURE 3.154.—DRAWING OF THE DRM 2 ORBIT RAISING-TUG SHOWING THE 

MAJOR VEHICLE SYSTEMS. THE PRIMARY REFLECTOR IS NOT SHOWN FOR 

CLARITY. ........................................................................................................................................ 508

FIGURE 3.155.—DRAWING OF THE DRM 3 EXPENDABLE PROPULSION BOOSTER 

SHOWING THE MAJOR VEHICLE SYSTEMS. .......................................................................... 508

FIGURE 3.156.—SCHEMATIC OF LASER TRACKING AND GUIDANCE SYSTEM. .................... 510

FIGURE 3.157.—DRAWING OF CRITICAL LASER ELEVATION, FLIGHT PATH ANGLE 

AND REFLECTOR ANGLE AT BEGINNING AND END OF JUPITER TRANSFER 

TRAJECTORY INSERTION BURN. ............................................................................................. 511

FIGURE 3.158.—SEQUENCE OF OPERATIONS FOR DRM 3: BEP-ENHANCED GALILEO 

MISSION. ........................................................................................................................................ 512

FIGURE 4.1.—BEP EVALUATION SUMMARY. LEO TO GEO, LOW EARTH ORBIT TO 

GEOSYNCHRONOUS EARTH ORBIT; TRL, TECHNOLOGY READINESS LEVEL; 

OCT, OFFICE OF THE CHIEF TECHNOLOGIST; BAA, BROAD AGENCY 

ANNOUNCEMENT. ....................................................................................................................... 515

FIGURE D.1.—BASELINE POWER BEAMING CONFIGURATION AND MTR DETAILS. ........... 575

FIGURE D.2.—MTR ALTITUDE, RANGE, AND ELEVATION ANGLE WITH RESPECT TO 

THE TRANSMITTER. .................................................................................................................... 576

FIGURE D.3.—CLEAR SKY ATMOSPHERIC SPECIFIC ATTENUATION AT CHINA 

LAKE, CA AND THE LLANO DE CHAJNANTOR OBSERVATORY IN THE 

ATACAMA DESERT, CHILE. ....................................................................................................... 577

FIGURE D.4.—CLEAR SKY ATMOSPHERIC ATTENUATION AT CHINA LAKE AND 

ATACAMA. ..................................................................................................................................... 577

FIGURE D.5.—POWER DENSITIES AT THE MTR FROM POWER BEAMING TURN-ON 

TO TURN-OFF AT (A) 95 GHZ, (B) 140 GHZ, (C) 170 GHZ, AND (D) 220 GHZ. .................... 578

FIGURE D.6.—JPL’S 1993 UPLINK PHASED ARRAY DEMONSTRATION (REF. 2). ................... 580

FIGURE D.7.—UPLINK ARRAY PHASED LOCKED LOOP SCHEMATIC AND RECEIVE 

SIGNAL STRUCTURE (REF. 2). ................................................................................................... 580

FIGURE F.1.—PROPOSED EUROPEAN SPACE AGENCY LASER LAUNCH FACILITY 

(REPRODUCED FROM SOURCE 1). ............................................................................................ 597

FIGURE F.2.—EXISTING LAUNCH SITES AROUND THE WORLD. .............................................. 598

FIGURE F.3.—AVERAGE ANNUAL NUMBER OF CLOUD OBSCURATION DAYS 

ACROSS THE U.S. (REPRODUCED FROM 2). ........................................................................... 599

FIGURE F.4.—MEAN MAXIMUM DAILY PRECIPITATION ACROSS THE U.S. 

(REPRODUCED FROM 2). ............................................................................................................ 599

FIGURE F.5.—ANNUAL MEAN DAILY AVERAGE TEMPERATURE (REPRODUCED 

FROM 2). ......................................................................................................................................... 600

FIGURE F.6.—POTENTIAL BEP GROUND COMPLEX LOCATIONS ............................................. 600

FIGURE F.7.—ILLUSTRATION OF SPECULAR REFLECTION (REPRODUCED FROM 

SOURCE 4). ..................................................................................................................................... 603

FIGURE F.8.—ILLUSTRATION OF DIFFUSE REFLECTION (REPRODUCED FROM 

SOURCE 4). ..................................................................................................................................... 603

FIGURE F.9.—DIFFUSE REFLECTION NOMINAL HAZARD ZONE .............................................. 604

FIGURE F.10(A).—MPE COMPARED WITH 3 M BEP EXPOSURE AS A FUNCTION OF 

DISTANCE. ..................................................................................................................................... 606

FIGURE F.10(B).—MPE COMPARED WITH 1 M BEP EXPOSURE AS A FUNCTION OF 

DISTANCE ...................................................................................................................................... 607

FIGURE F.11.—INITIAL SAFETY ZONE COMPARISON WITH CONVENTIONAL 

ESTABLISHED FACILITIES. ........................................................................................................ 609

FIGURE F.12.—GEOMETRIC SAFETY ZONE COMPARISON WITH CONVENTIONAL 

ESTABLISHED LAUNCH FACILITIES. ...................................................................................... 610
NASA/TM—2012-217014 xvi 

FIGURE F.13.—PLOT OF LAUNCH TIME AGAINST RANGE OF INTERSECTING 

SATELLITE. EACH COLOR REPRESENTS A SINGLE DAY OF LAUNCHES (I.E., 

THE 8 DAYS ARE OVERLAID). THE NORAD SAT ID IS LISTED ALONG WITH 

INTERSECTION DURATION AND RANGE. .............................................................................. 611

FIGURE F.14.—PLOT OF LAUNCH TIME AGAINST 1/R2

 OF INTERSECTING 

SATELLITE. EACH COLOR REPRESENTS A SINGLE DAY OF LAUNCHES (I.E., 

THE 8 DAYS ARE OVERLAID). ................................................................................................... 611

FIGURE F.15.—PROBABILITY OF AN INTERSECTION HAPPENING IN EACH HALF 

HOUR PERIOD, OVER THE 8 DAYS. .......................................................................................... 612

FIGURE F.16.—HISTOGRAM OF CLEAR LAUNCH WINDOWS—PERIODS THAT HAVE 

NO INTERSECTIONS. OVER THE 8 DAYS THERE ARE 29 PERIODS WITH MORE 

THAN 1 HOUR OF CLEARANCE. ............................................................................................... 612

FIGURE F.17.—HISTOGRAM SHOWING THE DISTRIBUTION OF INTERSECTION 

MAXIMUM INTENSITY DURING THE 8 DAYS—NOTE THAT MOST HAVE IMAX

LESS THAN 100 W/M2

 (I.E., 8% OF SOLAR CONST.). .............................................................. 613

FIGURE F.18.—HISTOGRAM OF THE MAX ENERGY INCIDENT ON A SATELLITE PER 

SQUARE METER DURING AN INTERSECTION—THIS IS CALCULATED USING 

IMAX X INTERSECTION DURATION. .......................................................................................... 613

FIGURE F.19.—PLOT OF LAUNCH TIME AGAINST RANGE OF INTERSECTING 

SATELLITE. EACH COLOR REPRESENTS A SINGLE DAY OF LAUNCHES (I.E., 

THE 8 DAYS ARE OVERLAID). THE NORAD SAT ID IS LISTED ALONG WITH 

INTERSECTION DURATION AND RANGE. .............................................................................. 614

FIGURE F.20.—PLOT OF LAUNCH TIME AGAINST RANGE OF INTERSECTING 

SATELLITE. EACH COLOR REPRESENTS A SINGLE DAY OF LAUNCHES (I.E., 

THE 8 DAYS ARE OVERLAID). THE NORAD SAT ID IS LISTED ALONG WITH 

INTERSECTION DURATION AND RANGE. .............................................................................. 614

FIGURE F.21.—PROBABILITY OF AN INTERSECTION HAPPENING IN EACH HALF 

HOUR PERIOD, OVER THE 8 DAYS. .......................................................................................... 615

FIGURE F.22.—HISTOGRAM OF CLEAR LAUNCH WINDOWS—PERIODS THAT HAVE 

NO INTERSECTIONS. OVER THE 8 DAYS THERE ARE 41 PERIODS WITH MORE 

THAN 1 HOUR OF CLEARANCE. ............................................................................................... 615

FIGURE F.23.—HISTOGRAM SHOWING THE DISTRIBUTION OF INTERSECTION 

MAXIMUM INTENSITY DURING THE 8 DAYS—NOTE THAT MOST HAVE IMAX

LESS THAN 200 W/M2

 (I.E., 15% OF SOLAR CONST.). ............................................................ 616

FIGURE F.24.—HISTOGRAM OF THE MAX ENERGY INCIDENT ON A SATELLITE PER 

SQUARE METER DURING AN INTERSECTION—THIS IS CALCULATED USING 

IMAX X T, WHERE T IS THE INTERSECTION DURATION. ........................................................ 616

FIGURE F.25.—MIGRATORY BIRD PATTERNS (REPRODUCED FROM SOURCE 6). ................ 618
NASA/TM—2012-217014 1 

Beamed-Energy Propulsion (BEP) Study 

Patrick George and Raymond Beach 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

1.0 Executive Summary 

The official scope of this study was (1) to review and analyze the state-of-art in beamed-energy 

propulsion (BEP) by identifying potential game-changing applications, (2) to formulate a roadmap of 

technology development, and (3) to identify key near-term technology demonstrations to rapidly advance 

elements of BEP technology to Technology Readiness Level (TRL) 6. The two major areas of interest 

were launching payloads and space propulsion. More generally, the study was requested and structured to 

address basic mission feasibility. 

The attraction of BEP is the potential for high specific impulse (Isp) while removing the powergeneration mass. The rapid advancements in high-energy beamed-power systems and optics over the past 

20 years warranted a fresh look at the technology. BEP could help meet the known needs of NASA and/or 

the Department of Defense (DOD), providing low-cost, rapid access to space for operationally responsive 

military systems, scientific payloads, and the commercialization of space. 

For launching payloads, the study concluded that using BEP to propel vehicles into space is 

technically feasible if a commitment to develop new technologies and large investments can be made over 

long periods of time. Such a commitment would include specific technologies like multimegawatt power 

lasers and microwave sources as well as building new launch facility infrastructure. The costs of the 

infrastructure are high. From a commercial competitive standpoint, if an advantage of beamed energy for 

Earth-to-orbit (ETO) is to be found, it will rest with smaller, frequently launched payloads. 

For space propulsion, the study concluded that using beamed energy to propel vehicles from low 

Earth orbit to geosynchronous Earth orbit (LEO–GEO) and into deep space is definitely feasible and 

showed distinct advantages and greater potential over current propulsion technologies. However, this 

conclusion also assumes that upfront infrastructure investments and commitments to critical technologies 

will be made over long periods of time. BEP energy source requirements are much less than for the 

launch applications. Lower propulsion costs with shorter transit times for LEO–GEO servicing missions 

and faster science missions to the outer planets are the major benefits. The chief issue, similar to that for 

payloads, is high infrastructure costs. 

Since cost (nonrecurring and recurring) emerged as a critical element of this study for both launch 

and space missions, the following should be noted. The cost estimate method used in this study was the 

same as is used by the NASA Glenn Research Center on all of its spaceflight projects, and the expert 

personnel involved in such estimates were employed for this study. Launch facility requirements were 

obtained by working with the NASA Kennedy Space Center, which has extensive experience with launch 

facilities. Per specific direction received at the final outbrief to the study’s sponsors, independent 

estimates for the laser portions of the infrastructure were gathered from knowledgeable DOD personnel; 

their estimates of cost were considerably higher than the NASA Glenn estimate, that is, the Glenn cost 

estimate was nearly an order of magnitude too low. On the other hand, outside expert consultants to the 

study have stated their belief that the Glenn estimates were somewhat too high; the study’s authors and 

the expert cost estimators believe that the consultants’ estimates fail to include several elements of the 

necessary infrastructure. 

It was hoped that beamed-energy technology would be a “silver bullet” solution to low-cost access to 

space and less-expensive deep space missions, or that at least it would find a niche where the benefits 

could be great. The authors of this report would like to make it clear that there are factors, as with all 
NASA/TM—2012-217014 2 

studies, that, if the ground rules of the study change or if costs for critical elements (e.g., lasers, ground 

facilities) come down significantly, it follows that the conclusions could change. 

As can happen in detailed examinations of innovative approaches, the study uncovered two 

unanticipated areas that appear to have great potential and therefore are worthy of consideration for new 

investment, at least at the “detailed study” level. BEP could provide the key to combining propulsion, 

power, and communications systems for in-space applications, thus creating a new class of small 

spacecraft with minimal systems and reducing their complexity, cost, and mass. Beaming energy to a 

thermal optical plasma engine could provide a highly efficient technology for high-thrust and high-Isp

space propulsion. The specifics of these concepts are provided in the main body of the report. 

Finally, the intention of this study was to work within the time and resources allotted to determine the 

feasibility of BEP. This involved achieving a single closed-design solution for each mission analyzed 

which is, in all probability, not the optimum solution. As such, proponents of the investigated concepts 

may not agree with the study’s findings and final design solutions, citing that better performance could be 

achieved with additional design effort. The authors do not disagree with that perspective. Nonetheless, the 

study management team was satisfied that the results were sufficient to meet the primary objective, which 

was to prove or disprove the feasibility of the concept and missions. It is also noted that at this time all 

possible future demonstrations of beamed energy and power transfer will be pursued solely by NASA. 

2.0 Study Synopsis 

The following subsections summarize the detailed reports from the design analysis team, consultant 

team, and cost team, which appear in full in Section 3.0. 

2.1 Study Background 

In June 2010, NASA and the Defense Advanced Research Projects Agency (DARPA) agreed to 

co-fund this study to determine the feasibility of using beamed energy for propulsion. The reason for the 

study stemmed from a joint interest in the technology, but without a definitive in-depth review clearly 

stating the case. NASA’s main area of interest is in-space applications, and DARPA’s is for launch to 

LEO, although each organization has at least some interest in both areas. The study concluded in January 

2011, with the Summary Presentation containing the basic results given on April 14, 2011. 

2.1.1 Rationale 

Since the early conceptualization/realization of the laser in the late 1950s, one of its theoretical 

applications has been to utilize photonic energy for wireless power transmission and propulsion. By the 

early 1970s, BEP was envisioned as a launch system using conventional gas-dynamic expansion of laserheated propellant to accelerate payloads up to a ton into orbit, with the hot gas using a de Laval nozzle to 

produce thrust. Another early concept was the use of laser energy that was subsequently converted into 

electric energy for propulsion. This type of power transmission has been constrained primarily by the 

electrical-to-photonic conversion efficiencies at the laser source and the ability for a photovoltaic receiver 

device to efficiently reconvert the beam back to electrical energy at high levels of irradiance. Other 

limitations are described later in this document. 

Since the 1970s, the continual development of low-cost, compact, and efficient high-energy sources 

and supporting optics has enabled a limited number of directed energy demonstrations. For example, in 

the late 1980s, under the Strategic Defense Initiative Organization (SDIO)-sponsorship, Prof. Leik 

Myrabo developed a laser-propelled vehicle concept envisioned to propel a 120-kg sensor payload using a 

100-MW-class laser. At that time, a goal of $100/kg was set. By the late 1990s, ambitious estimates of 

$1000/kg for a 1- to 10-kg lightsat were being reported. Myrabo’s concept used a reflector on the 

underside of the vehicle to increase the intensity of the beam and produce a region of extremely high 

temperature. The result was a pulsed-detonation concept that propels the vehicle. This concept was 

validated with the Air Force Research Laboratory’s (AFRL’s) 10-cm Lightcraft flight test demonstrations 
NASA/TM—2012-217014 3 

at White Sands in the late 1990s. In 2000, these flight demonstrations reached 233 ft and lasted 12.7 s. 

Succeeding this work was a 9-year effort including a scaled-up X–25LR bench model demonstration, and 

the X–50LR program, in which propellant and vehicle fabrication techniques were advanced through a 

series of successful test flights of a 50-cm laser ramjet vehicle. 

In 2003 NASA demonstrated a sustained flight of an indoor aircraft by providing 7 W of power to the 

electrical system through 24 triple-junction photovoltaic cells with a near-infrared laser. 

A more recent system demonstration concentrated on utilizing optical directed energy as a solution to 

provide power to space elevators ascending a composite cable, with the eventual desire to carry payloads 

into space. This was and is the fundamental premise of today’s Spaceward Foundation’s Power Beaming 

(Climber) Competition. The competition’s recent success, LaserMotive’s competitive system, was able to 

achieve a successful power transfer of several hundred watts to the climber over a distance of 1 km, to 

accomplish a climbing rate of 2 m/s. 

Microwave technology has been used to demonstrate wireless power transmission. In 1975, the joint 

Jet Propulsion Laboratory (JPL)/Raytheon program at the Goldstone complex was able to achieve 

ground-to-ground system conversion efficiencies up to 84 percent at 30 kW at a range of 1 mile. Since 

then, flight demonstrations such as the Stationary High Altitude Relay Platform (SHARP) in Canada and 

the Phased Array Model Airplane Experiment in Japan have successfully demonstrated the highefficiency benefit of microwave-power-beaming utilizing rectennas. This work has extended into endoatmospheric applications, such as the successful firing of an “Ion Breeze” engine from a 6-kV rectenna, 

and orbital concepts, such as a space-based microwave power station to provide a boost to lightcraft 

ascending from LEO–GEO. 

2.1.2 Previous Component Demonstrations 

At the component level, recent advancements in both diode laser technology and high-energy vertical 

multi-junction (VMJ) photovoltaic cells allowed for significantly higher power optical systems to be demonstrated, such as the AFRL’s high-intensity laser-power beaming program, which has demonstrated up to 

20-W/cm2

 output at the receiver and 44-percent optical-to-electrical conversion efficiencies for recharging 

electric aircraft. This performance was achieved with nonoptimized commercially available components. 

Multikilowatt systems may be assembled by fiber-combining several diode sources. At this point in time, 

the end-to-end system efficiencies are around 10 to 15 percent, with the primary limitations residing in the 

diode laser efficiency (30 percent typical). Future advances in laser technology—fueled by an emerging 

market demand for high-power, highly efficient sources—and a refinement in chemical composition and 

antireflective coatings in the photovoltaic cells will continue to improve overall system performance. 

In 2007, the NASA Goddard Space Flight Center demonstrated creating, modulating, and detecting 

x rays for communication applications, and this has implications for both power beaming and propulsion. 

The ability for an x ray to penetrate through hypersonic plasmas and exhibit narrow beamwidths because 

of their short wavelengths makes this an attractive region in which to operate in a point-to-point directed 

energy system. The potential applications for x-ray-directed energy may be limited to the space 

environment, because of high ionospheric attenuation. The concepts for x-ray power have grown from 

driving nanorobots to large 1-TW sources with 1-km apertures propelling a craft to circumnavigate the 

universe. Although large controllable x-ray sources do not exist today, the growing need for this 

technology beyond biomedical imaging will facilitate its development into higher output energies, which 

may eventually be used for power transmission and propulsion. 

The attraction of BEP is the potential for high-Isp while removing the power-generation mass. The rapid 

advancements in high-energy beamed-power systems and optics over the past 20 years warrant a fresh look 

at the technology and potential game-changing applications. For example, BEP could help meet the 

following known needs of NASA or DOD, providing low-cost, rapid access to space for operationally 

responsive military systems, scientific payloads, and the commercialization of space through 

 Launch to Earth orbit of small selected payloads 

 Stationkeeping at very low Earth orbit (DOD use—reconnaissance; NASA use—Earth observing) 
NASA/TM—2012-217014 4 

 LEO–GEO transportation system 

 Deep space exploration 

2.1.3 Study Scope 

The study scope was to review and analyze the state-of-art in BEP and to identify potential gamechanging applications, formulate a roadmap of technology development, and identify key near-term 

technology demonstrations to rapidly advance elements of BEP technology to TRL 6. The study includes 

the following specific areas: 

 ETO and space-to-space BEP 

 Thermal propulsion using beamed power 

 Electric (plasma) propulsion using beamed power 

 Synergies of integrated systems (i.e., beamed power and the propulsion implications) 

 Candidate technologies and solutions to achieve the reference missions (drawn from the 

categories of the electromagnetic spectrum: microwave, optical, and x ray) 

 Trades to evaluate not only the directed-energy trades but also the various energy-conversion 

mechanisms from the beam into thrust 

 Concepts defined to a level that will allow effective comparison of performance and operation 

costs with those of conventional combustion-based, self-contained fuel/oxidizer systems 

2.1.4 End Goals of Study 

Initial goals were established and later refined at the mid-term review to provide a more precise focus 

of the study. The following list provides the combined DARPA/NASA direction. 

 Via analysis, prove or disprove the feasibility of ETO via BEP, or at a minimum provide 

significant new insights into the limiting performance factors of BEP technology, quantify 

efficiency losses, and analyze atmospheric and environmental considerations 

 Identify synergies of BEP, if any, with crew servicing at GEO, horizontal launch access to space, 

and deep space exploration 

 Identify a list of proposed ground-based and space-flight demonstrations 

 Develop rough order-of-magnitude (ROM) nonrecurring costs for vehicles and ground systems 

 Develop ROM vehicle and ground facility costs for recurring launches 

 Investigate and analyze the effects of thermal blooming on laser and millimeter-wave propagation 

through the atmosphere 

 Emphasize concepts utilizing and extending from high-TRL commercial lasers like diodepumped lasers; descope concepts that require low-TRL lasers 

 The primary product of this study needs to be a determination of TRLs for near-term concepts of 

beamed-energy technology and rocket vehicle technology in BEP 

2.2 Study Approach 

The study incorporated the use of Design Reference Missions (DRMs) to determine BEP feasibility. 

Having specific missions on which to focus allowed preliminary design efforts to be conducted at a high 

level, although not with a complete systems analysis, with the intention of including all aspects of 

aerospace design practices, including applying standard design margins. The approach was successful in 

exposing major and minor issues and allowed exploring potential solutions to arrive at closed designs for 

feasibility evaluation. 

The study team consisted primarily of external experts selected for their depth of knowledge in the 

following fields: advanced propulsion technology, high-power microwave and lasers, continuous- and 
NASA/TM—2012-217014 5 

pulsed-laser satellite pointing and tracking, and aerospace mission design. Experts from industry, the U.S. 

Air Force, and academia were included. Resumes of team members are shown in Appendix B. The full 

team worked together to create a set of DRMs that covered a wide envelope of applications. 

The intention of this study was to work within the time and resources allotted to determine the 

feasibility of BEP. This involved achieving a single closed-design solution for each DRM, which is, in all 

probability, is not the optimum solution. As such, proponents of the investigated concepts may not agree 

with the findings and the study final design solutions, citing that better performance could be achieved 

with additional design effort. However, the study management team was satisfied that the results were 

sufficient to prove the feasibility of the technology. 

Existing mission integrated-design-analysis teams were utilized to speed the process of mission 

design. Glenn’s COllaborative Modeling for Parametric Assessment of Space Systems (COMPASS) team 

was assigned to the launch DRMs and the LEO–GEO DRM. Goddard’s Integrated Design Center Mission 

Design laboratory team was assigned to the deep space DRM. In both cases, the assignments were based 

on previous experience with the type of mission, team availability, and cost. Both are multidisciplinary 

collaborative engineering teams whose primary purpose is to perform integrated vehicle systems analyses. 

Through these analyses, the teams conduct and provide system designs based on trades studies. 

In order to narrow the trade space for the launching payload (laser) and space propulsion (space 

DRMs), an initial comparison was made between electric and thermal engines to determine which offered 

the best efficiency. Our results indicated that thermal engines would be the better choice. The details are 

shown in Figure 2.1 and were based on technology previously demonstrated at system and component 

levels in projects like Integrated Solar Upper Stage (ISUS) and High Delta-V (HiDVE). 

Figure 2.1.—Beamed thermal system efficiency—launching payloads and space propulsion. DDCU, DC–DC 

conversion unit; PPU, power processing unit. 

The DRMs were selected and created during the first meeting with the consultants. The collaborative 

effort signified concurrence within the group that the technologies selected were indeed the major ones to 

consider and spanned the envelope of those available. 

The cost estimates shown in this report represent the estimated “prime contractor” cost for each 

spacecraft based on Glenn’s COMPASS mission design team analysis. These estimates were generated at 

the subsystem and component levels using mostly mass-based, parametric relationships developed with 

historical cost data. Commercial costs also were estimated on selected DRMs as part of the trade space. 

These commercial estimates also were based on the COMPASS spacecraft design but were generated 

using PRICE System’s costing model. These estimates followed a methodology delineated by PRICE 

Systems for estimating commercial satellites with their software. A more detailed description and a list of 

cost assumptions for each of the various estimates are included in the individual reports. In addition, 

several lifecycle cost estimates were generated to evaluate the total cost of the technology over an 

assumed time horizon. The ground facilities necessary to beam energy proved to be the major cost driver 

in each case. The launch facility cost estimates are based mainly on data provided by the NASA Kennedy 
NASA/TM—2012-217014 6 

Space Center and were adjusted to address each case. A detailed description of the lifecycle and facility 

costs is included with each DRM report. 

2.3 Study Findings 

This section brings together the information generated by the mission design teams, consultant teams, 

and cost teams in each of the major study areas. 

2.3.1 Launch DRMs 

Three launch concepts were selected for evaluation, each having unique characteristics that span the 

envelope of relevant current technologies. This was to ensure that the study touched potential areas 

offering opportunities to uncover promising technologies. The decision to limit the study to laser and 

microwave energy sources for launch was made by the management team during preliminary discussions 

with NASA experts at GRC and an extensive literature search. The energy from x-rays could be used for 

transmitting; however, the technology for developing high-power transmitters and receivers was 

considered to not be available for a long time in comparison to the other technologies. 

2.3.1.1 DRM Envelope Selection and Definition 

The first DRM created utilized laser optical technology. This technology was selected because it 

involved the options of a continuous or pulsed-laser source and the optical plasma engine utilizing air and 

water as propellants. Some of the related technology challenges (TRL 2 to 4) and resultant launch profiles 

are very tough to meet; however, the potential of very low vehicle mass made this concept a good 

candidate for investigation. The estimated payload range is 1 to 100 kg which, for example, could consist 

of six CubeSats plus support hardware totaling 40 kg. Ultimately, two vehicle sizes were studied. The 

first, a 120-cm-diameter unit, was based on previous studies. The second, a 35-cm-diameter unit, was 

studied to determine the characteristics and performance of a CubeSat-sized payload option (see 

Figure 2.2). 

Figure 2.2.—Two sizes of laser optical vehicles studied. 
NASA/TM—2012-217014 7 

The engine operates in air-breathing laser ramjet mode up to Mach 6 to 7 and an altitude of 35 km. 

Then it switches into rocket mode, whereupon H2O is pressure fed to the 120-cm propellant injector ring 

for the remainder of the flight to LEO. As indicated in Figure 2.3, the entire vehicle aeroshell is designed 

to serve multiple structural- and propulsion-related functions. The axisymmetric nose/forebody serves as 

an external compression inlet in supersonic flight, precompressing air through a short annular slit into the 

engine “hot section.” 

Figure 2.3.—Function of external surfaces in air-breathing propulsion mode. 

A wealth of material exists on this technology because of the previous development and tests of 

Dr. Leik Myrabo, who was a part of the BEP study team. Dr. Myrabo was relied upon to provide details 

of the technology as it applied to this DRM. His previous analysis and tests on the lightcraft design 

provided the basis for use as a launch vehicle for 1- to 100-kg payloads. For this reason, this design was 

chosen as the first DRM and a good starting point for developing our missions. A full report is in Section 

3.2.1 and covers two sizes of vehicles. 

The second DRM was selected because it utilizes laser thermal technology. Dr. Jordin Kare agreed to 

be a consultant and team lead because of his years of experience and successful practical testing that used 

lasers to power a multitude of experiments. This application included a single-stage vehicle with a heat 

exchanger (HX) using liquid hydrogen (LH2) and water powered by a continuous-wave (CW) laser. 

Payloads in the 10- to 100-kg range were estimated to be launched. A full report is shown in Section 

3.2.2.1, Section 3.2.2.2, and Section 3.2.2.3. 

The propulsion system operates by using a liquid propellant that is pressure-fed or pumped through an 

HX, which absorbs the laser energy and transfers it to the propellant, vaporizing it and heating it to high 

temperature. The hot propellant is exhausted from a conventional nozzle (Figure 2.4). 
NASA/TM—2012-217014 8 

Figure 2.4.—Laser thermal vehicle. Laser thermal rocket configuration showing the 

laser receiver HX mounted on its side. 

The third DRM featured a millimeter-wave source and an LH2 thermal engine on a two-stage vehicle 

with a solid booster to enable orbit. The target payload range was 10 to 100 kg. Dr. Kevin Parkin headed 

this team because of his previous work on high-power transfer using millimeter waves and his recent 

work on building a 1-MW system for launching small payloads. A full report is shown in Section 3.2.3.1, 

Section 3.2.3.2, and Section 3.2.2.3. The operation of the propulsion system is similar to that of the laser 

thermal DRM, but the energy source is a millimeter wave and the HX material and configuration matches 

the source (Figure 2.5). Over the past 60 years, the time-average power output of high-power microwave 

sources in the key millimeter-wavelength range has increased by over 6 orders-of-magnitude, for the first 

time putting ETO launch within economic reach by using gyrotron technology. Today, a 1-MW gyrotron 

oscillator can be purchased for about $2M. By combining the output of many such oscillators, the power 

and spot size needed to propel a vehicle to orbit are possible. 
NASA/TM—2012-217014 9 

Figure 2.5.—Millimeter-wave thermal DRM. Millimeter-wave thermal rocket configuration showing a 

circumferential receiver HX. 

Central to the microwave thermal rocket is a microwave-absorbent HX covering the underside of the 

vehicle. The refractory HX bypasses the specific enthalpy limits of conventional combustion, imparting 

30 to 40 MJ/kg of H2 propellant as opposed to 16 MJ/kg for conventional hydrogen peroxide (H2/O2) 

combustion. This higher specific energy halves the propellant flow rate needed to produce a given thrust 

and reduces the total propellant needed to put a given dry mass into orbit by a factor of 3. 

The downside to an LH2 thermal rocket is the relatively low density of hydrogen, making the tank and 

turbopump larger and heavier than for dense propellants. The larger tank adds to the drag losses of the 

vehicle, requiring extra propellant to compensate. 

2.3.1.2 DRM Evaluation Method and Comparisons 

Each DRM was designed by defining the necessary subsystems, conducting trades, examining various 

options to achieve a launch profile to the required orbit altitude, defining an operational procedure, and 

analyzing costs. The subsystems included thermal, power, communications and data handling (C&DH), 

navigation and control, structure, propulsion, mechanical systems, and vehicle configuration. 

For a reference point of feasibility, the Pegasus XL/ORBCOMM (Figure 2.6) was selected as an 

existing launch/payload operation. It was the closest comparable system found, but is not directly 

applicable, so allowances must be made when making comparisons. The Pegasus XL (Figure 2.7) can 

carry a payload up to 420 kg, is air launched from 40,000 ft, and uses three stages to orbit. The 

ORBCOMM satellite has been optimized to fit eight units into the Pegasus XL available mass and 

volume limitations. Each satellite costs $1.2M and has a mass of 40 kg. 
NASA/TM—2012-217014 10 

Figure 2.6.—ORBCOMM 

spacecraft. Courtesy of Orbital 

Sciences Corporation; used 

with permission. 

 

Figure 2.7.—Pegasus XL. 

Figure 2.8 compares the sizes of the vehicles. It clearly shows that the laser optical vehicle is much 

smaller. Other figures of merit (FOMs) are shown in Table 2.1. All DRMs require two beaming sites to 

provide sufficient energy to the vehicle to achieve orbit altitude. 

Figure 2.8.—Launch vehicles compared by size. 

TABLE 2.1.—OTHER FOMS FOR THE DRMs 
NASA/TM—2012-217014 11 

2.3.1.3 Energy Source Facilities 

All three missions utilize three facilities; a control center where the launch vehicle and payload are 

assembled and readied for launch; a “boost” beaming station for the primary source that takes the launch 

vehicle to high altitude; and a “main” beaming station that provides the energy to take the vehicle to orbit. 

Some variations exist among the missions because of their different technologies. The following 

subsections briefly outline their differences. For a detailed description, see Section 3.0 for the individual 

COMPASS reports and consultant reports. 

2.3.1.3.1 Laser Optical Ground Facilities 

The laser optical ground facilities for the 120-cm vehicle have a combined boost station/vehicle 

assembly building since the laser beam has to be placed under the vehicle. The boost laser consists of a 

continuous or a pulsed 140-MW (Figure 2.9) output laser constructed of up to 2900 (continuous or 

pulsed), 120-kW commercial-off-the-shelf (COTS) modules. Two options exist for the “main” laser: (1) a 

ground system of up to 7250 COTS lasers (20 kW each) or (2) a space-based relay station that utilizes the 

boost laser facility. 

Figure 2.9.—Configuration of laser optical ground facilities. 

Only one ground facility is required for the 35-cm craft with a CubeSat payload. The beam is 

redirected from a vertical arrangement for the boost phase to a tiltable mirror used for the main phase. 

The source requires only 96 beamlets (12 kW each) for total of ~12-MW of power output. Laser pulsing 

is in the kilohertz range for efficient air-breathing engine performance: for example, 96 beamlines at 

60 Hz each = 5760 Hz, which falls in the viable range. Note that only with pulsed lasers can a “temporal” 

approach to beam-combining be realized, wherein all beamlines are fired sequentially like a “Gatling” 

gun onto a rotating mirror that efficiently shepherds pulses into a single 3-m telescope. With 60-Hz 

beamlines, the mirror must rotate at 3600 rpm (Figure 2.10). The ring-mirror diameter is set by the 96 

adjacent turning flats positioned around the ring, each sized at 25 by 25 cm. These flat mirrors would be 

remotely adjusted with servodriven, three-axis mirror mounts to precisely target the on-axis rotating 

mirror. 
NASA/TM—2012-217014 12 

Figure 2.10.—Concept for beam-combining system with central rotating output mirror. 

2.3.1.3.2 Laser Thermal Ground Facilities 

The laser thermal boost station is located 20 km down range from the launch site to allow energy to 

be transmitted to the vehicle in the initial part of the ascent. The main station is located 400 km down 

range from the boost site, as shown in Figure 2.11. 

Figure 2.11.—Typical launch concept of operations with L1 (boost) laser array. Markers: (0) Launch. 

(1) Main array “sees” vehicle and begins tracking. (2) Main array turns on; L1 array at maximum 

full-power range. (3) Main array at full power; L1 array. 

The vehicle launch site is where the vehicles actually leave the ground. The launch site will include at 

least one launch tower along with auxiliary equipment for handling vehicles, payloads, and propellant. 

The boost, or L1, laser site will provide laser power to the vehicle for its initial launch and climb from the 

launch site through the atmosphere, until the vehicle is high enough to receive power from the main laser. 
NASA/TM—2012-217014 13 

Currently available laser sources have been selected to eliminate development of single high-powered 

lasers (Figure 2.12). However, the arrangement and control of the many low-powered laser modules is 

only at a concept stage, although this is seemingly fairly straightforward. 

Figure 2.12.—Configuration of laser thermal ground facilities. 

The boost laser site will have a layout generally similar to the main laser site, with hundreds to a few 

thousand modules installed over an area on the order of 1 km2

 (possibly less, if land area is limited). Each 

module will be largely independent, with its own power supplies, tracking, and so forth. The laser power 

and aperture area per module are not yet defined, but a plausible configuration might use 20-kW diode 

array assemblies, each with its own ~1-m mirror, mounted in groups of 10 on a common elevation mount. 

Because the boost laser needs to point at, or very close to, the horizon in one direction, but only slightly 

past the zenith in the opposite direction, the boost laser array should be built on a slope facing the launch 

site. The altitude of the boost laser array is somewhat less important than that of the main laser array 

(because when the vehicle is at maximum range for the boost laser, it is also near zenith and atmospheric 

effects are minimized). It may be desirable to put the boost laser array slightly below the launch-site 

altitude (if the intervening terrain is lower still) so that even at launch, the beam would never be aimed to 

intersect the Earth’s surface (i.e., it would always point at least slightly upward). 

The main laser site is the heart of the launch system and consists of an array of beam modules—

independent lasers with small beam directors and tracking systems—with a total output power (for the 

DRM baseline system) of ~140 MW. It is typically several hundred kilometers downrange (east) of the 

launch site, so that the vehicle trajectory passes over the main laser near the midpoint of its trajectory. 

2.3.1.3.3 Millimeter-Wave Thermal Ground Facilities 

The millimeter-wave thermal ascent trajectory consists of two segments: In the first (Figure 2.13), the 

vehicle is powered by a short-range “boost” beam facility, maintaining vertical ascent of the rocket at 

40% throttle to minimize drag losses. The first segment transitions to the second at an altitude of 30 to 

60 km. In this time, the “boost” beam facility hands the rocket over to the “main” or “sustain” beam 

facility as the rocket levels off and begins to accelerate horizontally at 100% throttle. The acceleration of 

9g to 19g during this segment of the ascent raises the velocity from 1.5 to 8 km/s in only 60 s and enables 

the rocket to achieve orbital velocity within the 125-km range of the beam source (Figure 2.14). This 

minimizes the initial cost of the beam facility and could eventually be expanded to allow a gentler ascent 

trajectory suitable for human launch (Figure 2.13). 

Figure 2.13.—Configuration of millimeter-wave thermal ground facilities. 
NASA/TM—2012-217014 14 

Although the 1990s-era Green Bank Telescope aperture is about the right size for application to a 

millimeter-wave thermal rocket, it is the current era of closed-loop active surfaces that provide the 

necessary surface accuracy. Surface accuracies now exceed those needed for the microwave thermal 

rocket by nearly an order of magnitude, albeit for smaller apertures. The key question to answer in a beam 

facility conceptual design will be what combination of structural rigidity and active surfaces can correct 

for changing gravitational distortions and mechanical bending modes as the structure slews; sudden 

thermal deformations can be negated using heaters. If greater mechanical stiffness is required, there are 

Fresnel reflector and even holographic techniques that would be unsuitable for astronomical work but 

might be advantageous for beam direction, provided they turn out to be compatible with the spectral 

beam-combining approach used in the optical back end. 

Figure 2.14.—Millimeter-wave thermal launch system and ascent trajectory. 

2.3.1.4 Costs 

Table 2.2 shows the nonrecurring and recurring costs for the launch vehicles. The costs were 

determined during the 2- to 3-week mission design process—a very intense exercise that made a fair set 

of assumptions. As a result, the cost estimates are at the top level but are sufficient to expose any major 

issues. A list of what is and is not included in the estimate is in the full reports in Section 3.2.1.3, Section 

3.2.2.3, and Section 3.2.3.3. The most noteworthy assumption is that the estimates do not include the 

costs for technology development. 

The nonrecurring costs are for design, development, testing, and engineering (DDT&E). The 

recurring cost is for the first unit produced using government-based cost models. Realizing that DDT&E 

government costs would be high in comparison to a commercial venture, we processed the estimates 

through a commercial model. As expected, the costs were lower and give some idea of what the 

differences could be between government and commercial approaches. A proportionate drop in cost was 

seen using both models after a number of vehicles were produced. 
NASA/TM—2012-217014 15 

Table 2.2 shows the estimated ground facility nonrecurring and operating costs for each mission. The 

nonrecurring costs were based on a costing model used by Kennedy to estimate the construction of their 

facilities. The beaming-source costs were provided by the respective consultants using estimates based on 

current technologies. Ground operations costs were based on 360 launches per year. This number was 

chosen to represent the existence of a market for the launch capabilities so that an operations cost per 

launch could be determined. 

TABLE 2.2.—NONRECURRING AND RECURRING 

COSTS FOR LAUNCH VEHICLES 

2.3.1.4.1 Cost Variation 

In order to obtain a sanity check on the cost estimates for the launch facilities, two efforts were 

undertaken. First, knowledgeable personnel from the Directed Energy Directorate at Kirtland Air Force 

Base in New Mexico, where the Starfire Optical Range (SOR) is based, were contacted to obtain 

information on the cost to build the facility. Data on the cost of the laser and the optics used for imaging 

and identification of space objects is limited. However, the authors were directed to two reports that 

contain estimates of megawatt-class laser systems. Both “New Concepts for Space-Based and GroundBased Laser Weapons” and “A Preliminary Assessment of the Potential Cost and Cost-Effectiveness of 

Space Based Weapons” were used as a reference for facility costing. The cost estimates in the reports 

were determined to have low fidelity such as no factor for scaling or taking into consideration the cost 

reduction as the number of lasers increased. The facility estimates range from $165/W to $1070/W. 

The second effort to get a better estimate on facility costs involved discussions with a major defense 

contractor with many years of expertise in building lasers. Although this input was voluntary, per study 

guidelines, it was felt to be more accurate than other sources. Their estimate, which was based on a 

description of the application and the energy required, was $250/W. 

This study’s facility cost estimate was $28/W, which is a dramatic difference from both of the 

preceding estimates. The difference is explainable by the assumed cost of the laser source used in each 

case. This report optimistically took into account a number of what the authors believe to be realizable 

factors, such as the likely increases in laser power over the next 10 years as well as the associated cost 

reduction due to a gradually increasing commercial demand. This approach takes advantage of using 

hundreds of smaller powered lasers versus single large units, drastically reducing the nonrecurring costs 

and individual laser costs. Figure 2.15 shows an anticipated laser power level development versus cost. 
NASA/TM—2012-217014 16 

Figure 2.15.—Laser power level versus cost. Kare 

Technical Consulting, used with permission. 

Using a large number of lasers creates the additional problem of combining a large number of beams, 

but this is taken to be an achievable engineering challenge via fiber optics rather than a technical barrier. 

Details of this methodology and the cost for a launch system can be found in the 2006 paper by Jordin 

Kare and Kevin Parkin, both consultants on this study, “A Comparison of Laser and Microwave 

Approach to CW Beamed Energy Launch.” They reported that a 100-MW laser system would cost $24/W 

and that a 250-MW microwave system would cost $22/W. Adding in the cost for land, buildings, and 

other infrastructure, the study cost of $28/W falls in line. 

The two referenced “external” estimates indicate that the study cost data could be off by at least a 

factor of 10 on the ground facility costs. If our assumptions are incorrect and the reference estimates are 

correct, take our laser costs (68% of total) and multiply them by 10. This yields $207/W versus our 

original estimate of $28/W. It is probable that the correct value lies somewhere between the two and that 

it would depend on the rate of higher power laser development and commercial demand. However, for the 

purposes of this study, whose goal was to prove feasibility, the conclusion that infrastructure costs would 

be too high to make this economically feasible at this time still holds. The higher costs from the other 

sources cited here would only exacerbate the problem and makes the conclusion stronger. 

2.3.1.5 Risks and Issues 

The following subsections discuss a set of risks and issues with the proposed launch methods. These 

are intended to express the challenge to be overcome for success. 

2.3.1.5.1 Laser Optical 

The main risk with the thermal system components pertains to the laser reflectors and the plasma 

degradation of the mirror and heat transfer during ascent (Figure 2.16). 

Any damage or degradation to the mirrored surface experienced during launch can cause excessive 

local heating because of the intensity of the laser beam (approximately 3000 W/cm2

). Debris impact, or 

erosion from the nearby plasma can potentially degrade the mirror reflectivity, which would cause 

catastrophic heating and loss of the vehicle. The very high reflectivity of the mirror (99.99%) must be 

maintained throughout the launch phase. 
NASA/TM—2012-217014 17 

Figure 2.16.—Laser optical issues. 

If the plasma gets within a few centimeters of the mirror surface, there could be excessive heating. 

The mirror surface is not as reflective in the frequency in which the plasma will be radiating. Therefore, it 

can absorb significant amounts of heat if the plasma moves too close to the mirror surface. Depending on 

the distance of the plasma from the wall, heating can range from 20 to 80 W/mm2

. 

Laser light scattered off of the reflective shield can potentially be reflected back to Earth. This could 

be a health hazard to someone who inadvertently looks into the reflected beam. The power of the reflected 

beam that reaches the Earth’s surface will depend on the curvature of the shield mirror’s surface and the 

distance from the Earth. The higher the shield curvature, the greater the laser light dispersion and, 

therefore, the lower the intensity of the reflected beam. To reduce this risk, developers would need a 

diffuse reflecting surface. A mirrored surface that has conical reflecting surfaces etched onto it may be 

able to disperse the incoming laser perpendicular to the incoming beam, limiting the reflection back 

toward the surface. 

2.3.1.5.2 Laser Thermal 

The main risk with the thermal system concerns the operation and structural integrity of the HX, 

which is to achieve the required heat transfer capabilities, designers made the HX walls very thin. This 

thin-wall construction could cause significant structural problems because of the various loadings that the 

HX will experience. These include 

 Aerodynamic loading during ascent 

 Thermal stresses due to the large temperature variations encountered during the launch 

 The temperature gradient experienced across the HX during operation 

 The high internal pressure due to the heating and expanding hydrogen gas 

The operational feasibility of the HX concerns the ability to transfer the required amount of laser 

energy to the hydrogen gas and the ability to construct the HX to provide the flow characteristics needed 

to achieve this heat transfer. 

A simplifying assumption made during the analysis was that the dividing walls between the fluid 

passages of the HX were operating at the same temperature as the HX surface. This assumption is very 

optimistic. Under actual conditions, this inner wall will be at a temperature gradient less than that of the 

surface. Further, more detailed analysis should be performed to determine what the temperature gradient 

is and what effect it will have on the HX operation. 

The HX design utilized a complex manifold that brought in the cool hydrogen gas and collected the 

hot hydrogen gas at spacing of 15 cm over the total area of the HX. This complex manifold arrangement 

and the subsequent heat losses associated with it need to analyzed in much greater detail to assess the 

feasibility of the design. 
NASA/TM—2012-217014 18 

2.3.1.5.3 Millimeter-Wave Thermal System 

The main risks with the thermal system are similar to those described in Section 2.3.1.5.2. The 

operational feasibility of the HX concerns the ability to transfer the required amount of microwave energy 

to the hydrogen gas and the ability to construct the HX to provide the flow characteristics needed to 

achieve this heat transfer. Convective heat transfer coefficients of greater than 14,000 were required. 

These were calculated to be theoretically possible on the basis of the flow conditions. However, a more 

detailed analysis and experimental work would be required to determine if this level of heat transfer is 

truly possible. 

2.3.1.6 Summary of Launch Mission Findings 

The technical feasibility evaluation factors selected for all missions included a complete closeddesign solution, a defined basis for technology and technology improvements, designs that include 

allowable margins for mass, engine performance, and thermal and structural subsystems. All three 

missions were able to achieve feasibility, although some particular tough challenges lie ahead for all 

because of high infrastructure costs (Table 2.3) and because some key technologies are at low TRLs, 

especially with the optical laser option. 

TABLE 2.3.—LAUNCH VEHICLE AND INFRASTRUCTURE COSTSa

[MM, millimeter.] 

a

Launch vehicle cost assumptions: (1) Prime contractor costs (no fee), (2) No technology development costs, (3) Protoflight 

development approach, and (4) Before reserves, insight/oversight, launch services. 

 

A game-changing technology evaluation criteria defined by the NASA Office of the Chief 

Technologist (OCT) is “Technology that is innovative, unique and promises to enable revolutionary 

improvements to the efficiency and effectiveness of our country’s space capability.” The study interpreted 

this for launch missions as reusability, improved payload mass fraction, improved mission capability (like 

an increase in available operations time), and providing a new method for access to space. As can be seen 

in the comparison table (Table 2.3), the air-launched Pegasus XL/ORBCOMM combination is difficult to 

beat in most of the areas, such as the cost for getting payload to orbit. However, the launch missions 
NASA/TM—2012-217014 19 

studied excel in the future capability to enable revolutionary access to space. Using high-powered lasers 

or millimeter waves to send payloads into orbit is definitely a new way of launching. This technology is 

in its early phase of development, and investments could lead to increases in performance and reduced 

costs. These benefits, with the added possibility for other uses of the beaming station like orbital debris 

removal, makes each of the technologies game-changing. 

Although economic viability could be considered to be a game-changing factor, the study anticipated 

that it was so important that it was deemed to be worthy of its own category. Its evaluation factors were 

lower vehicle costs ($/kg); lower launch facility operation costs like range safety, vehicle preparation and 

handling, and infrastructure costs. This area was the biggest challenge for all of the launch missions. 

Launch vehicle recurring costs in comparison with the selected point of reference Pegasus XL/ 

ORBCOMM could be debated. The costs are in the same range, but this is not a good comparable mission 

because the designs have been optimized and the vehicles can launch to a variety of trajectories. Facility 

operations costs were based on a standardized proportion of the facility costs for a potential market for 

launching 360 flights per year, so determining the facility costs will play a large role in economic 

viability. Because there are common needs for all three missions—land needed for buildings, range 

safety, security, transportation accessibility, and electrical power lines—the major differencing factor was 

the beaming-source costs. However, the total costs for all missions were considered to be nonfeasible 

because of the extremely high infrastructure costs. 

It was determined that all three missions have technical feasibility. Achieving game-changing 

viability would rely on large investments of funds to develop the technologies required. None of the 

missions were able to meet the economic requirements for feasibility. Therefore, the overall evaluation of 

this study was that utilizing beamed-energy for launching small payloads into orbit is not feasible. 

2.3.2 Space Design Reference Missions (DRMs) 

The mission analyses for the space DRMs were scheduled and configured to take advantage of the 

analysis and lessons learned from the launch DRMs. An ETO launch power requirement, estimated in 

previous studies at 1 MW/kg, was used as an early gauge to decide on the maximum satellite size to be 

addressed in this study. In addition to the power requirement, the potential atmospheric affects at high power 

levels led to the decision to limit the ETO launch to microsatellites or smaller CubeSats. Thus, the initial 

maximum power expected for the ETO application was estimated at 100 MW or less. To take full advantage 

of the launch facilities’ potential for the space applications, the BEP team chose a LEO–GEO reusable tug 

as the first space DRM. This decision led to a synergistic relationship for a deep space DRM that would 

utilize the ground facilities, as well as the LEO–GEO tug design and analysis. The following subsections 

provide a synopsis of the space DRMs, with pointers and references to the mission designs and analyses. 

2.3.2.1 LEO–GEO DRMs 

2.3.2.1.1 Vehicle Description 

The beamed-energy tug propulsion stage shown in Figure 2.17 is 5.78 m long and cylindrical, 

consisting of a 4.0- by 5.6-m elliptical primary reflector cantilevered off the side of the vehicle on a 

±60 gimbal. Two degrees of freedom will be required to track the ground-based emitter, and either the 

entire vehicle will be rolled during a burn or another degree of freedom will be added to the gimbal 

(although a two-degree-of-freedom gimbal creates complications because the rotation does not occur 

about the focal point). The secondary reflector will be fixed. Both reflectors and shielding to protect the 

entire vehicle from the laser flux will be 97%-efficient, 0.9999 reflective carbon-fiber composite. A 

1.5- by 3.3-m solar array will supply 560 W of nominal operating power, and a 2000-W-hr lithium-ion 
NASA/TM—2012-217014 20 

Figure 2.17.—LEO–GEO tug diagram. RCS, Reaction Control System.

(Li-ion) battery pack will supply an additional 1625 W during 5- to 15-min burns to run pumps, tracking, 

and so forth. The main engine will provide 5000 N of thrust with 50-psia pressure in a 30-cm spherical 

combustion chamber. The engine will use water heated to 6000 K to provide 813-s Isp. The plasma ball 

will be electromagnetically suspended in the chamber, and the nozzle will be regeneratively cooled or 

film cooled. The beamed energy will pass through a 10-cm-diameter sapphire refractory window into the 

combustion chamber. The nozzle will have an area ratio of 50, will be regeneratively cooled, and will be 

gimbaled for communications pointing on a screw or ball mechanical motor. 

2.3.2.1.2 Concept of Operations (CONOPS) 

A standard CONOPS involves the launch of two 1970-kg GEO satellites and 4895 kg of water 

propellant on a Falcon 9 Block 2 from Kwajalein into a 400-km circular orbit with an inclination of 9°. 

The tug will rendezvous with the stack of two GEO satellites and propellant and begin orbit-raising burns 

within 3 days of launch. Six 5- to 8-min burns are required to raise the initial orbit to a 3000-km, 9° 

inclination circular orbit. Three 15-min burns at 3000 km are required to raise apogee to GEO: one 

15-min apogee burn is required to change planes to 0° inclination, and two more 15-min apogee burns are 

GEO

Satellite

Primary Reflector

4.0 x 5.6 meter elliptical

Propulsion Stage

(Tug)

Single Launch

5.78 m 2.06 m

Propellant

Stage

GEO

Satellite

Single Launch

1.60 m

2.46 m

RCS Pod

Solar Array

Secondary

Reflector

Patch Antennas 

(every 90 deg around 

bus)

Propellant 

(Water) Tank

Engine

Shielding for Engine

RCS Tank

Pressurant

Tank

Primary Reflector
NASA/TM—2012-217014 21 

required to circularize. The total transit time to GEO assuming burns every orbit is 32 hr, requiring 

4.2 km/s Δv. 

Following undocking from the GEO spacecraft, the tug returns to a 400-km, 9° inclination, using 

residual water propellant that was launched with the two GEO spacecraft, and is ready for the next transit 

in 24 hr (Figure 2.18). The tug itself is assumed to have been launched separately. The overall concept is 

summarized in Table 2.4. 

 

Figure 2.18.—Launch elements and sequence. 

TABLE 2.4.—LEO–GEO TUG CONOPS 

2.3.2.1.3 Ground Facilities 

The DRM 2 mission as defined, requires two types of laser ground facilities, a 2- to 3-m-diameteraperture Perigee Facility for burns for altitudes between 400 and 3000 km, and a much larger diameter 

Apogee Facility for burns for GEO. Both types of laser facilities will be required to deliver about 50 MW 

of laser energy to the spacecraft. On the basis of atmospheric loses, each of the ground facilities will need 

an 80-MW laser to deliver the required power to the tug. The apertures for these ground facilities will 

resemble large astronomical telescopes such as SOR and the proposed 30-Meter Telescope (TMT). 

All of the facilities will have two main components, a Beam Generation system and a Beam Control 

system. In addition, each facility will require sufficient power and heat-dissipation capacity for the 

operation of the laser. 
NASA/TM—2012-217014 22 

Perigee facilities.—Three Perigee Facilities spaced evenly around the globe on or near the equator 

should provide the best viewing opportunities to enable perigee burns on nearly every orbit. DRM 2–A 

uses the Kwajalein Atoll as the launch site, which is located at a latitude of 9o. The Perigee Facilities will 

require only 2-m apertures to compensate for beam diffraction, but the ability of the optical coatings to 

handle the 80-MW power levels may drive larger optics to lower the flux. The Perigee Facilities will most 

likely resemble the 3-m telescope at SOR at Kirtland Air Force Base (AFB) (see Figure 2.17) or the 

Advanced Electro-Optical System (AEOS) 3.67-m telescope on Maui. The near-equator locations of the 

Perigee Facilities implies that they could be ship based. 

Apogee facility.—The Apogee Facility has more flexibility on location, but it will need to be at least 

30 m in diameter to focus the spot on the spacecraft’s reflector at GEO. The facility could be located 

below a latitude of 45° and still provide good lasing opportunities for both the 3000-km altitude (and so 

could be used for perigee burns) and GEO. The TMT being constructed at Mauna Kea, Hawaii, is a good 

representation of a baseline Apogee Facility. The laser requires the same 80-MW power as the Perigee 

Facility and for practical reasons could be identical. 

2.3.2.1.4 Costs 

The costs associated with the LEO–GEO tug are shown in Table 2.5. These costs were determined 

based on a reuse of the tug for 10 round trip transfers from LEO, with fuel transported with the satellite 

payload for each mission. For comparison purposes recurring costs for a LEO–GEO transfer, using a 

centaur stage, were developed. Reuse of the tug for 10 transfers provides a cost saving of $15M per 

launch, with further cost saving potential if additional reuse can be provided. 

The nonrecurring costs are for DDT&E. The recurring cost is for the first unit produced, and uses a 

government based cost model for the estimate. Table 2.6 shows the estimated ground facility nonrecurring 

and operating costs for each mission. The non-recurring costs were based on a costing model used by the 

NASA Kennedy Space Center to estimate construction of their facilities. The beaming source costs were 

provided by the respective consultants using estimates based on current technologies. Ground operations 

costs were based on a dual use scenario for launch and space operations. 

As an alternative to the ground-based infrastructure, a space-based system was analyzed and is shown 

in the second row of Table 2.5. For the space-based system a lower power level of 500 kW was analyzed, 

with the corresponding increase in costs and transfer time given in the table. 

TABLE 2.5.—RECURRING AND NONRECURRING COSTS FOR LEO–GEO DRM 

. 
NASA/TM—2012-217014 23 

2.3.2.2 Deep Space Mission 

Several possible applications of BEP were explored for deep space missions. The primary goal of this 

portion of the study was to find an application of BEP that would either enable a mission not otherwise 

achievable or measurably enhance myriad possible missions. 

Early work on the direct propulsion of gossamer bodies such as the Forward-Landis Star Wisp 

microwave-propelled interstellar craft and the success of the Japanese Interplanetary Kite-craft 

Accelerated by Radiation Of the Sun (IKAROS) solar sail mission to Venus led to the consideration of 

laser-powered sails to propel payloads to the outer planets. 

The concept finally adopted for DRM 3 was to show how BEP as used in DRM 2 for an orbit-raising 

tug could improve the performance of a planetary or deep space mission. The study was performed in 

collaboration with the Mission Design Laboratory at Goddard. The approach was to use the results of the 

Glenn’s COMPASS team design for the orbit-raising tug with the minimal changes necessary to adapt the 

design to provide acceleration for the one-way trip to Jupiter: the object being to investigate performance 

enhancements possible with BEP. The Galileo mission to Jupiter (1989 to 2003) was chosen for the 

comparison. 

2.3.2.1.5 Vehicle Description 

The vehicle design for DRM 3 was adapted directly from the 

orbit-raising tug designed by Glenn’s COMPASS design team 

for DRM 2. Like DRM 2, the payload and propulsion vehicles 

are launched separately and joined on orbit. However, whereas 

the orbit-raising tug was designed to be reusable, to deliver selfcontained payloads to geostationary orbit, the propulsion vehicle 

in DRM 3 was designed to be an expendable booster. For the 

orbit-raising tug to be reusable under the DRM 2 CONOPS, the 

propellant would have to be replaced for each mission and be 

manifested as part of the satellite payload launch. In the DRM 3 

configuration, the propulsion module includes the water 

propellant for the laser plasma engine for near-Earth orbit-raising 

and Jupiter Transfer Trajectory Insertion burns. The propulsion 

module is by far the heavier payload for the two Falcon-9-class 

launches used in DRM 3. 

The Galileo payload analog for the study was updated from 

the original Galileo mission by replacing the original 

Radioisotope Thermoelectric Generators with Advanced Stirling 

Radioisotope Generators (ASRGs), otherwise all power, mass, 

and size specifications from the original Galileo vehicle were 

used to define the payload. Three of the four ASRGs needed to 

power the Galileo-like science payload will be launched with the 

propulsion module. Including these ASRGs with the propulsion 

module will eliminate the need for the solar panel used for power 

on the DRM 2 orbit-raising tug and will have the advantage of 

additional mass fraction being available on the payload launch 

vehicle, should it be needed. 

The propulsion module will easily fit within the 5.2-m 

Falcon 9 fairing (see Figure 2.19). The stowed module will have 

an overall length of about 8 m. 

The major components of the propulsion module follow: 

(1) the BEP system, consisting of the beam collecting and 

focusing optics, the thrust chamber and nozzle, and propellant 

Figure 2.19.—Stowed DRM 3 propulsion module in Falcon 9 fairing. 
NASA/TM—2012-217014 24 

storage and delivery; (2) the electrical energy system consisting of the ASRGs, a battery, and power 

management and distribution; (3) the attitude and control system consisting of reaction wheels for fine 

pointing, an autonomous docking system, and monopropellant thrusters with fuel for maneuvering; (4) the 

thermal management system; and (5) the communications system. 

Electric power for the propulsion module will be provided by three ASRG units and a rechargeable 

Li-ion battery launched with the propulsion module. A fourth ASRG will be launched with the Galileolike payload and added when the payload and propulsion modules are docked. The ASRGs (with 

beginning-of-life power of 143 W each) will be used for primary power for all functions except 

propulsion, when the 100-Ah Li-ion battery will provide supplemental power for the propellant pumps 

during laser engine operation. For a typical 10-min orbit-raising propulsion firing, the battery depth of 

discharge will be 8.3%; and for the Jupiter Transfer Trajectory Insertion firing of 100 min, the depth of 

discharge will be 83%. The power system will be a 28-V-battery-dominated bus, with separate 28-V 

buses on the payload and propulsion modules. The ASRGs will be retained by the payload during 

separation of the payload from the expended booster and used for power during the remainder of the 

mission. 

2.3.2.1.6 CONOPS 

Differences between the LEO–GEO tug (DRM 2–A) and the deep space mission led to design 

changes in the propulsion module design and spacecraft systems allocations (e.g., photovoltaic panel 

replaced by ASRGs, greater battery capacity, and water propellant launched as part of the propulsion 

module). The notional operational timeline (from launch) is shown in Figure 2.20. 

Figure 2.20.—Sequence of operations for DRM 3: BEP-enhanced Galileo mission. 

The initial launch and orbit-raising sequence closely mirror that of DRM 2 and will employ two 

Falcon 9 launch vehicles (one for the propulsion module and one for the Galileo payload). The propulsion 

module and Galileo mission spacecraft will be launched separately from Kwajelein Atoll into an 800-km, 

9 inclined circular LEO. 
NASA/TM—2012-217014 25 

The propulsion module and Galileo mission spacecraft will be joined on orbit using Next Generation 

Advanced Video Guidance Sensors and automated rendezvous and docking hardware, firmware, and 

software. The time allocated for the launch, docking, and spacecraft diagnostics is 1 month. 

Orbit-raising protocols from DRM 2 will be followed to first transfer the combined spacecraft from 

an 800-km, 9 inclined circular orbit to a 3000-km, 9 inclined circular orbit and then to a 3000- by 

35,000-km, 9 inclined Jupiter Departure Orbit. The time allocated for the orbit change maneuvers is 

15 days. 

The final phase of near-Earth operation will be the 100-min laser-beam-powered Jupiter Transfer 

Trajectory Insertion acceleration from Jupiter Departure Orbit to Jupiter Transfer Orbit, followed by 

separation of the propulsion module from the Galileo mission craft with a collision avoidance maneuver. 

Here, the BEP phase of the mission will end with the propulsion module flying away and the Galileo 

craft continuing to Jupiter, where it will use onboard hydrazine monopropellant to perform a Jupiter Orbit 

Insertion burn of 70 min and enter the highly elliptical Galileo Jupiter Orbit of 200,000 by 12,000,000 km 

with a 98-day period. 

2.3.2.1.7 Cost 

The design of the propulsion module and the CONOPS for DRM 3 were adapted with minimal 

changes from DRM 2, and cost estimates for the BEP-enhanced Galileo mission can be derived from 

DRM 2 estimates. 

BEP can open new opportunities for deep space exploration and science missions. The cost and time 

to develop the 80-MW ground station will be the most significant obstacles to the realization of the 

potential of BEP for this deep space application. 

2.3.2.2 Deep Space Costs 

The costs associated with the deep space DRM are shown in Table 2.6. Since the deep space DRM 

utilizes the LEO–GEO tug, the costs are similar with addition of the advanced Stirling radioisotope 

generator needed for power operation at the outer planets. Also, the recurring cost of $85M is based on 

the ten times reuse for the LEO–GEO tug. 

As with the LEO–GEO DRM, an alternative to the use of a ground-based system was analyzed. The 

high power requirement for the deep space mission (50 MW) necessitates the need for a low-cost, highpower, long-life system. Both nuclear and solar options were considered for this application, but the costs 

based on the Jupiter Icy Moon Orbiter and the ISS were considered impractical for this alternative. 

 TABLE 2.6.—RECURRING AND NONRECURRING COSTS FOR DEEP SPACE DRM 
NASA/TM—2012-217014 26 

2.4 Unexpected Findings 

The following subsections describe two technologies that evolved out of discussions during the 

process of the study. They represent concepts that have high potential and should be considered for 

further investigation. 

2.4.1 Spacecraft Remote Operation 

The first concept is an integrated power, communications, propulsion, and navigation and attitude 

control system. This concept, shown schematically in Figure 2.21, could control the operation of a 

cooperating spacecraft. By utilizing the communications aspect of the power beam through optical 

Figure 2.21.—BEP concept—integrated propulsion, power, communications, 

and navigational/attitude control on the same laser beam (PV, photovoltaics; 

RCS, reaction control system). 

retrodirective beam control, it is possible to achieve a closed-loop control capability between the beamed 

source and the receiving spacecraft. Thus, the spacecrafts’ propulsion system could be controlled by the 

amount of energy transmitted. This, in combination with the communications link, could also be used to 

control the spacecraft’s attitude, move it into close proximity, and berth it, if required, without having to 

have the dynamics of the two spacecraft physically coupled during the rendezvous and berthing operation. 

Navigation data could be transmitted on the laser beam, relieving the receiving craft of the need for that 

systems equipment. 

If successfully developed, this technology could lead to a new class of spacecraft with enhanced 

operational capability, lower complexity, higher reliability, and lower mass. 

2.4.2 Propulsion Engine 

The second concept is not new but one that has an application with energy beaming to provide 

propulsion with both high thrust and high Isp (concept shown in Figure 2.22). The optical plasma engine 

requires temperatures above 6000 K to heat propellant within a cavity, expanding the material that is 

creating the propulsive force. This system is able to utilize diverse new propellant types such as water and 

ammonia, thus eliminating toxic chemicals and associated safety issues both in space and on the ground. 

The environmentally friendly water propellant is potentially available on the Moon and Mars. The energy 

source can be solar, laser, or millimeter wave, making it flexible to match the technologies as they 

develop. Since this system is an optical-to-optical conversion, it naturally has high efficiency with the 

potential for the highest propulsion engine performance because the oxidizer/energy source is eliminated 

on the spacecraft. 

Development work could begin on this concept with very little investment. 
NASA/TM—2012-217014 27 

Figure 2.22.—BEP optical plasma engine. 

2.5 BEP Development Direction 

It is shown in Section 2.3 that the high infrastructure cost of ground- and space-based beamed-energy 

infrastructure makes BEP unfeasible at this time. But this does not mean that BEP should be dropped as a 

potential game-changing technology. There are near-term applications, as explained in Section 2.4, that 

have the potential for high investment return. The launch and in-space applications, deemed unfeasible 

now, should be pursued by attacking the key issue of high energy-source costs. However, the cost of 

developing megawatt energy sources should be borne by those in need of them at this time. NASA and 

DARPA should take on role of developing the key technologies that will be needed later to implement 

their specific applications. Figure 2.23 provides a proposed timeline of how the various applications of 

beamed energy should align to maximize the total overall investment. 

Figure 2.23.—Timeline of beamed-energy development. 

Development of beamed-energy transmission protocols and standards (P&S), and along with the 

previously mentioned applications, is worth pursuing since these P&S and applications will provide the 

foundation and early applications on which to build the more complex and technologically challenging 

components of the program. 

Development of P&S should be addressed initially to lay the groundwork and direction for BEP. The 

P&S should define the most appropriate organizations to baseline and maintain the regulations governing 

such areas as the characterization of transmitted and received beam quality. The scope could include 

describing the limitations on ground, aviation, and space applications along with equipment performance. 
NASA/TM—2012-217014 28 

It is recommended that a set of low-cost ground development demonstrations related to the near-term 

applications should be funded to begin the development process. These demonstrations could be 

combined with industry and academia competitions similar to NASA’s Centennial Challenge. Other 

longer term developments could include laser and millimeter-wave beam-combining techniques to allow 

the use of numerous low-cost commercial energy sources, long-distance pointing and tracking techniques, 

and optical systems to direct high-power energy sources. Figure 2.24 shows how the gradual buildup of 

knowledge through testing at low flux and low power levels to high flux and high power levels could be 

infused into applications. 

Figure 2.24.—BEP development applications from low to high power. 

As reported in Section 2.4, the integration of power, propulsion, control/communications, and 

navigation is possible via a single laser beam to enable remote positional control. Near-term ground 

applications include unmanned aerial vehicles (UAVs) and low-flying aircraft. Space applications, such 

as solar electric propulsion and DARPA’s Exo-SPHERES, could utilize this technology. In the long term, 

these developments could enable lower cost access to space by minimizing the complexity of systems 

needed for smaller craft operations. Suggested development tests and demonstration are shown in 

Figure 2.25. 

Figure 2.25.—Demonstrations to develop near-term remote positional control technology. 
NASA/TM—2012-217014 29 

One of the key demonstrations that would provide a critical stepping stone for evolving wireless 

power beaming in space would be to place a laser on the International Space Station (ISS) and beam 

energy for propulsion to a free-flyer. The free-flyer would reposition by changing orbit as a result of the 

energy. Performing a test in the space environment would anchor the design models, confirm the 

environmental assumptions, and establish energy beaming as a new way of thinking about the design of 

future spacecraft (Figure 2.26). It could also demonstrate safe beaming control for space and terrestrial 

energy transfer, along with determining beam locator sensor efficiencies and accuracies in a controlled 

application. It would support growth and future applications for beaming energy in space to the Earth, to 

the Moon, and to other satellites for both propulsion and power. 

To show the potential of BEP and beamed energy in general, Figure 2.26 describes a similar 

demonstration, also designated for the ISS, that was detailed in 2008 for another study. 

Figure 2.26.—In-space flight demonstration from the International Space Station to Free Flyer. 

The other early technology development reported in Section 2.4 that would have a far-reaching effect 

in the development of BEP is the optical plasma engine. The most promising factor in this case is the 

potential of high systems efficiency in comparison to existing or near future propulsion technologies. As 

with most high-potential undeveloped technologies, the rewards could be great but come with the need for 

investment and major challenges. Maintaining control of the plasma ball within the chamber and away 

from the walls is considered to be the major hurdle to be proven. The work could begin by using solar 

energy as the source rather than laser energy and, therefore, is not dependent on the development of 

lasers. Some early work (10 or more years ago) has been performed in this area with a first prototype 

model. A low funding level could go a long way to determining the feasibility of this technology. Figure 

2.27 lists a set of tests/demonstrations that could be performed, gradually building to a space test. 
NASA/TM—2012-217014 30 

 

Figure 2.27.—Demonstrations to develop near-term optical plasma engine technology. 

Development demonstrations for each of the DRMs appear in the specific consultant reports in 

Sections 3.1 and 3.2. 

2.6 Future Potential of BEP 

As mentioned in this report, and as shown in the investigations into the DRMs, BEP has the potential 

to realize low-cost access to space and the transportation of vehicles from LEO to GEO and deep space. If 

successfully developed, these applications alone would be game changing. It should be kept in mind that 

there are applications of beamed energy in general that also have game changing potential. 

Beamed-energy technology, when developed and appropriately demonstrated, has the potential to 

significantly impact future NASA Exploration and Science Missions. The flexible path strategy for space 

exploration envisions a hybrid approach involving robots and men to explore the Moon and near-Earth 

asteroids. Early missions may not land men on the surface but would utilize robots controlled by 

astronauts on an orbiting spacecraft to explore the surface and return samples to the orbiting spacecraft. 

Expectations of water ice, located in craters or buried in crevices, will require power to be delivered to the 

robot to enable exploration of these areas. Beamed energy has the potential to deliver power and enable 

control of the mobile robot, thus providing a virtual wire allowing the astronauts the freedom to explore 

the surface from the orbiting spacecraft. Knowledge obtained from the remotely controlled robots will 

then be utilized to prepare for human operations on the surface. 

Lunar studies have identified the potential for laser power beaming that utilizes orbital platforms to 

capture and convert solar energy for transmission to the surface. Lunar in situ resource utilization (ISRU) 

work, done in anticipation of water ice in the Shackleton crater, has demonstrated the value of beamed 

energy in processing ice to obtain hydrogen and oxygen. Thus, the power-beaming technology and assets 

developed and utilized during the early hybrid phase of exploration will form the basis for the 

infrastructure for future human campaigns. 

The flexible-path strategy, by design, couples the Exploration and Science Missions. Thus, robotic 

technology developed for science missions could form the technology base for robots utilized in the 

controlled robot phase of crewed missions, or vice versa. The nature of beamed energy allows for the 

integration of power, propulsion, communications, and navigation/attitude control. This integrated 

capability could be utilized to enable satellite constellations for planet finder, other formation-flying 

missions, and future fractionated satellite applications. When properly scripted, the power-beaming 
NASA/TM—2012-217014 31 

technology utilized for human-controlled robotic exploration of craters and crevices could be utilized for 

autonomous robotic science applications. 

Development of BEP could provide the key to a new method of launch and in-space transportation as 

well as aid in exploration and science missions. 

3.0 Study Reports 

The approach to this study was to engage consultants, external to NASA, to obtain the widest possible 

viewpoints on launch and in-space beamed-energy propulsion (BEP). Their input was incorporated into 

the formulation of Design Reference Missions (DRMs) and then processed through NASA-based mission 

design teams to bring a sense of practicality to the actual vehicle configuration and mission operations 

utilizing NASA standard design practices. NASA Glenn Research Center project costing techniques and 

methods were used to apply a known basis of cost estimating for costing the missions. The complete set 

of reports from the mission design teams, the consultants, and the cost-estimating teams are included here 

without editing by the authors to provide the reader with original data. This approach was an attempt to 

ensure that the conclusions of the report were not biased and to provide the reader with information by 

which their own conclusions may be drawn. 

3.1 Launch DRMs 

The rationale, description, and subsequent evaluation of each DRM are given in the following 

subsections. It is important to note that it was not the intention of this study to make a comparison and 

select a “best choice.” Each technology has distinct features and was evaluated to answer the feasibility 

question. Side-by-side evaluations were made only to point out the distinguishing characteristics of each 

technology. The details must be read for a full understanding of the potential and challenges of each 

technology. 

Three specific DRMs were deemed necessary to determine feasibility for the launch case because of 

the wide range of technologies that have been applied to this specific area and the high level of 

development activity that has occurred over the last few years. 

Each DRM was put through the same mission analysis process with the same team to maintain 

consistency between developments of the technology applications. This approach removed the relative 

effect of any bias within the team’s process.
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3.1.1 Laser Optical Collaborative Modeling for Parametric Assessment of Space Systems 

(COMPASS) Report 
NASA/TM—2012-217014 33 

National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov

Lightcraft Summary 

• 1.2 m top shaped launcher/spacecraft: no 

separate launch vehicle required

• Sample vehicle emulates ORBCOMM 

messaging spacecraft

– laser can be reused for orbit maintenance, 

repositioning & deorbit

• Launched 30° elevation, requires two Laser 

stations (up to 350 MW) for complete launch

• Propulsion and thermal technologies still in 

TRL 2-4 range

• Propulsion: 

– Parabolic mirror doubles as plug nozzle

– Pulsed Laser focused onto ring where temps 

> 6000 ° K theoretically provide I Isp > 900 

seconds

– Laser power up to 350 MW required

– Water propellant best for density, cooling, ground 

handling

• Guidance, Navigation, Control

– Vehicle spun up to 120 rpm for stabilization at 

launch, differential water inlet injection for pitching

• Some Launch power, C&DH, Comm systems 

reused for spacecraft functions

3

On-orbit

configuration

Launch 

configuration

Cost bottom line: Eliminating launch vehicle 

reduces launch vehicle costs by more than half!

(does not include differences in launch support, 

range, and laser facilities)
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National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov

Lightcraft Reuse as a spacecraft

• Constellations

• Self-assembled bigger spacecraft

• Communications

– Messaging

– Cell

– Internet

• Reuse Laser engine

– Laser communications (in-space)

– Optical imaging

• Earth resource: multi-spectral, almost continuous imaging with 4000 S/C – new 

paradigm

• Astronomy: customers want 6-15 m apertures (planet finders)

• SETI search for alien laser communications

– Repositioning, orbit maintenance, deorbit

• LEO science spacecraft

• Quick way to demonstrate/qualify new technologies

– Electronics, solar cells, lubricants

• Propellant for Depot

• Servicing spacecraft (viewing, deliver ORUs)

• Assist in deorbit of spacecraft 10
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Bottom image courtesy of Orbital Sciences Corporation; used with permission. 
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Image copyright Space Electronics LLC; used with permission. 

National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov

L - 10 days:

• The vehicle, including payload is assembled in a clean room

• The vehicle is placed on a spin balance machine and undergoes static and

dynamic spin balancing, and mass properties (c.g., moments and cross products

of inertia) are determined

• Any guidance updates to the vehicle’s flight computer are done

• Vehicle is fueled with water propellant, cleaned and loaded into its transport 

container for roll-out to the launch 

pad.

• the container supports the vehicle at its ring fins during transport to the pad 

• the container is open at the bottom

• the container has horizontal and vertical split-line so that its top half can be

removed to mount the vehicle to the pad, and so its bottom half can be

opened for removal from around the vehicle

L – 2 hours:

• The vehicle in its transport container is rolled out to the launch pad

• A crane lifts the transport container and lowers it onto the pad spin bearing

• viewing ports in transport container allow for visually aligning vehicle on pad

• Latches holding the transport container’s top cover in place are released and

the top cover is removed by the crane

• The pad service tower extends a balance arm with spin bearing, ground power

umbilical and water fueling line. The power umbilical and fueling line are

automatically mated when the spin cap in placed on the vehicle

Processing and Launch
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National Aeronautics and Space Administration

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Transport Container

Service Tower

Pad Spin bearing

National Aeronautics and Space Administration

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Transport Container

Service Tower

Pad Spin bearing
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National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov

Service Tower

Transport Container

Pad Spin bearing

National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov

Service Tower

Transport Container

Pad Spin bearing
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National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov 47

Balance arm with spin bearing, ground power umbilical

Service Tower

Pad Spin bearing

National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov 48

Service Tower

Pad Spin bearing

Balance arm with spin bearing, ground power umbilical
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National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov 49

Service Tower

Pad Spin bearing

Balance arm with spin bearing, ground power umbilical

National Aeronautics and Space Administration

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Service Tower

Pneumatic spin-up arm

Pad Spin bearing

Balance arm with spin bearing, ground power umbilical

Bottom retracts just before laser firing
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National Aeronautics and Space Administration

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Service Tower

Pneumatic spin-up arm

Pad Spin bearing

Balance arm with spin bearing, ground power umbilical
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National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov 53

Service Tower

Pad Spin bearing

Pneumatic spin-up arm

National Aeronautics and Space Administration

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Service Tower

Pad Spin bearing
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National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov 55

Service Tower

Pad Spin bearing

National Aeronautics and Space Administration

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Service Tower

Pad Spin bearing

T= 0 sec

At this point differential 

thrusting to pitch over to the 

30° elevation is required
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National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov

BEP GNC

• Requirements

– Inject the Lightcraft into a 400 km circular orbit with a 28.5 degree 

inclination

• Assumptions

– BEP Lightcraft has zero products of inertia

– Pitch control on ascent 

• Provided by differential thrust while in sight of the laser 

• Provided by cold gas thrusting when not in sight of the laser

– Needed for orienting the vehicle to the correct attitude for the 2nd laser 

burn (LEO insertion burn)

• Navigation Design

– One LN-200s IMU

• IMU is low mass (0.75 kg) with spaceflight proven design on 

Clementine, Deep Space I and the MER rovers

• Gyros – 3 solid state fiber optic with a bias stability of 0.1 deg/hr

• Accelerometers – 3 solid state silicon with a bias of 0.3 mg

– One Ithaco IM-103 3-axis magnetometer to measure the strength and 

direction of Earth’s magnetic field

– 4 patch antennas used for GPS orbit and attitude determination

89
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Thermal Analysis: Radiator Sizing

• A radiator panel was sized for use on the vehicle to reject 

the excess heat from the electronics once in orbit. 

•A surface mount radiator was utilized.

•The radiator model was based on a first principles analysis of 

the area needed to reject the identified heat load to space. 

From the area a series of scaling equations were used to 

determine the mass of the radiator. Worst case thermal 

environment of LEO was used to size the radiator. 

•No louvers were utilized

•Heat Pipes were used to conducted heat to the radiator 

surface. 

Assumptions used

Variable Value

Radiator Solar Absorptivity 0.14

Radiator Emissivity 0.84

Radiator Sun Angle 70°

Radiator Operating 

Temperature

320 K

Total Radiator Dissipation 

Power

90 W

View Factor to Earth 0.5

View Factor to the Solar 

Array

0.15
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National Aeronautics and Space Administration

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Individual Metalized Layers

Mesh Spacer

MLI Blanket

MLI can be conformed to fit 

over various shapes. 

It can be held in place with 

Velcro or glue.

MLI covering the outside of a 

spacecraft
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National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov

Risks

The main risk with the thermal system components 

pertains to the laser reflectors and the plasma 

degradation of the mirror and heat transfer during ascent. 

Multi-index laser reflectivity Plasma Frequency

Laser Frequency

Plasma Heating

Localized heating due to the plasma. If the plasma gets 

within a few centimeters of the mirror surface there could be 

excessive heating. The mirror surface is not as reflective in 

the frequency the plasma will be radiating in. Therefore it 

can absorb significant amounts of heat if the plasma moves 

too close to the mirror surface. Depending on the distance of 

the plasma from the wall heating can be in the range of 20 to 

80 W/mm2.

Mirror Reflectivity

Any damage or degradation to the mirrored surface experienced 

during launch can cause excessive local heating due to the 

intensity of the laser beam (approximately 3000 W/cm2). Debris 

impact, or erosion from the near by plasma can potentially 

degrade the mirror reflectivity which would cause catastrophic 

heating and loss of the vehicle. The very high reflectivity of the 

mirror (99.99%) must be maintained throughout the launch 

phase. 
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3.1.2 Laser Optical Consultant Report—Design Reference Mission 1–A 

This section covers the design and feasibility of pulsed-laser-launching microsatellites (10 to 100 kg) 

and nanosatellites (1 to 10 kg) into low Earth orbit (LEO) (see Figure 3.1). Note, however, that 

picosatellites (0.1 to 1 kg) may be the nearest “target of opportunity” for emergent ~1-MW-class, pulsed,

1-μm solid-state diode-pumped (SSDP) laser technology. Provided herein are descriptions of the requisite 

lightcraft engine/vehicle/payload, ground-based laser (GBL) facility, concept of operations (CONOPS), 

costs, and overall feasibility assessment. 

The Beamed Energy Propulsion Study (BEPS) examined the state-of-the-art (SOA) in BEP (both 

lightcraft and lasers) applicable to (1) identifying potential game-changing applications (e.g., CubeSats, 

fractionated satellites, and Responsive Space), (2) formulating a roadmap of technology development, and 

(3) pinpointing key near-term technology demonstrations to rapidly advance elements of BEP technology 

to Technology Readiness Level 6 (TRL 6). This section explores the application of existing lightcraft 

pulsed-plasma engines (Figure 3.2) that have seen 150 successful flights at White Sands Missile Range 

(WSMR, Refs. 1 and 2), and set the current 71-m world altitude record on October 2, 2000 (Ref. 3). 

Figure 3.1.—12-MW laser launch facility at WSMR, New Mexico (artist concept by 

P. Rawlings). 

Figure 3.2.—Lightcraft plasma engine pulsing in 1-bar laboratory air 

(photographs by J.E. Shryne, III). Left: 14.7-cm-diameter #200 

vehicle. Right: 23-cm #100 vehicle. 
NASA/TM—2012-217014 113 

Figure 3.3.—ORBCOMM messaging satellite in 800-km 

circular orbit (from Ref. 4). Courtesy of Orbital 

Sciences Corporation; used with permission. 

One key feature of lightcraft technology is its scalability which, in fact, helps to mitigate or manage 

technological and financial uncertainties. The specific DRMs considered in this section are for singlestage-to-orbit (SSTO) 120-cm lightcraft (DRM 1–A) and 35-cm lightcraft (DRM 1–A) powered by 

repped-pulse 1 μm SSDP lasers. 

For the DRM 1–A mission (120-cm lightcraft launcher), the COMPASS team produced a detailed 

conceptual design configured as an ORBCOMM (Ref. 4) satellite replacement with full functionality (see 

Figure 3.3). These design results were then scaled down to shed insight into a 35-cm-diameter nanosat 

launcher, which was subsequently compared and contrasted with the Air Force Research Laboratory’s 

(AFRL’s) 2007 lightcraft concept (Ref. 5), including engine and vehicle design, mass breakdown, 

payload specifications, and boost performance. 

3.1.2.1 Lightcraft Technology Development 

The lightcraft development process is an interdisciplinary endeavor involving engine/optics/airframe 

integration, mostly through applying traditional engineering design practices. The process begins with 

first creating an efficient engine concept, engineered for maximum/optimum efficiency with a given 

photon source: that is, source specifications must closely match engine requirements (and vice versa) to 

avoid a performance-degrading “impedance mismatch.” Once a workable engine concept is demonstrated, 

it can be integrated into the whole lightcraft vehicle concept. 

Both airbreathing and rocket engine options, as well as combined cycles that transition between these 

propulsion modes, have to be investigated. Each engine design will involve the construction of one or 

several prototypes, exploiting variations on a theme (i.e., on one or multiple design features). Initial rough 

prototype engines for laboratory testing can be heavy and employ heat-sink cooling (i.e., lots of thermal 

mass). Later, increasingly refined flight-weight engines will exploit more exotic high-temperature 

structural materials, refractory metal coatings, and so forth. 

The most efficient and most cost-effective approach to lightcraft engine/vehicle development is to 

thoroughly exploit existing laser sources located at Government, Department of Defense (DOD), and 

industry laboratories, and some of these are already set up as photon users facilities. Such facilities can be 

used to quickly extend engine performance data bases, at reasonable cost. 

The lightcraft concept has inherent scalability characteristics that can facilitate a progressive 

technology development roadmap (see Figure 3.4) that minimizes risks and upfront investment for the 

infrastructure. A principal measure of technological progress in this roadmap is the setting of increasingly 

higher altitude records (e.g., as was done by Goddard and Von Braun). 
NASA/TM—2012-217014 114 

Figure 3.4.—Progressive technology roadmap and possible development program: ground and 

flight demonstrations leading to 35-cm nanosat launch system. 

The development process will emphasize hardware creation (i.e., build, test, break, redesign, and 

retest), while judiciously exploiting available engineering design software to expedite the development 

cycle. (Extensive testing on repetitive-pulsed (RP) plasma engines has established the proof of concept, 

but a similar endeavor is still needed for continuous-wave (CW) plasma engines.) This approach enables 

rapid optimization, refinement, and evolution of any engine/vehicle geometry, especially in areas of 

aerodynamics, structural and heat-transfer analysis, flight dynamics and stability/control, and so forth. 

The essential development steps will be 

 Bench-top static tests: Thrust stand and ballistic pendulum tests with pulsed and CW photon 

sources; beam-riding performance measurements with additive-increase multiplicative-decrease 

(AIMD) (Refs. 6 to 9)) apparatus 

 Wind tunnel spinning lightcraft tests: Measuring aerodynamics of unpowered spinning lightcraft 

aeroshells in subsonic and/or transonic wind tunnels versus angle of attack and revolutions per 

minute (rpm) (Ref. 10) 

 Direct-connect engine tests: Measuring laser propulsion performance in subsonic and transonic 

wind tunnels (blow-down and/or Ludwig tube) with two-dimensional lightcraft plasma engine 

geometries 

 Supersonic and hypersonic tests: Measuring laser engine performance in blow-down and/or 

impulsive supersonic- and hypersonic shock tunnels with two- and three-dimensional engine 

and/or vehicle geometries; hypersonic tests at Mach 6 to 10 are ongoing in Brazil (Refs. 11 

and 12) 

 Indoor vertical beam-riding tests: Short 5- to 10-m free flights up to the lab ceiling with flightweight lightcraft models (likely spin-stabilized, Ref. 2)) 

 Outdoor vertical beam-riding tests: Flights to altitudes of 10, 30, 100 m for proof of concept; no 

flight control system needed (Refs. 1 to 3) 

 Numerical flight dynamics studies: Incorporate above engine performance data, vehicle 

aerodynamics data, trajectory data, and so forth into a comprehensive numerical simulation model 

for optimization and control systems studies (Refs. 13 to 15)

 Flights to extreme altitudes: Boosts to 1, 10, 30, and 100 km requiring full flight control systems; 

altitude records demonstrably track technology progress 

At the present point in time, RP plasma engine development is substantially more mature than its CW 

counterpart and was, thus, the focus of the BEPS. However, CW plasma engines have great potential to 
NASA/TM—2012-217014 115 

leverage the recent rapid evolution of high-power CW beam sources by extending into entirely new 

propulsion territory. In addition, the engineering development of such CW engines is expected to benefit 

from hard lessons learned in RP engine research and testing.

3.1.2.2 Description of 120-cm Vehicle 

The BEPS objective was to create an independent concept design for a laser-powered launch vehicle 

utilizing the combined-cycle, pulsed detonation engine (PDE) in Figure 3.3 that was first postulated in 

References 16 and 17. The basic operating principles and performance parameters for that engine are well 

established and were characterized in early theoretical studies (Ref. 16 and 17), before hardware 

development campaigns (lab and outdoor free-flight tests) began at WSMR in 1996 (Refs. 1 and 3). 

Hence, the BEPS 120-cm spacecraft in Figure 3.5 was conceived as a point of reference for assessing the 

feasibility of the concept and for identifying what technologies and demonstrations needed to be carried 

forward to expedite development. 

Figure 3.5.—Final launch configuration of 120-cm-diameter DRM 1–A spacecraft. 

The toy-top-shaped, nonreusable spacecraft requires no separate launch vehicle: that is, the satellite 

and vehicle hardware are assumed to be inseparable, and in this 120-cm BEPS design option, both fly into 

orbit. (An alternative option could be a separable satellite payload with an integral chemical kick rocket 

for orbit circularization. This was studied by the COMPASS team, but the launch trajectory “closed” on 

180 km instead of the desired 400 km altitude.) 

The COMPASS team began this design and feasibility study by examining a wide variety of missions 

and applications for the reuse of a single-stage lightcraft: that is, the vehicles become spacecraft or 

satellites after attaining orbit: 

 Satellite constellations 

 Self-assembled bigger spacecraft 

 Communications (messaging, cell, and Internet) 

 Reuse of laser propulsive engine parts and/or functionality 

– Laser communications (in-space) 

– Optical imaging 

o Earth resource: multispectral, almost continuous imaging with 4000 spacecraft—

new paradigm 

o Astronomy: customers want 6- to 15-m apertures (planet finders) 

o SETI Institute search for alien laser communications 

– Repositioning, orbit maintenance, and deorbit 
NASA/TM—2012-217014 116 

 LEO science spacecraft 

 Quick way to demonstrate and/or qualify new technologies (electronics, solar cells, and 

lubricants) 

 Propellant for orbiting depots 

 Servicing spacecraft (viewing and delivering orbital replacement units (ORUs); low-Earth-orbit 

to geosynchronous-Earth-orbit (LEO–GEO) shuttles 

 Assist in deorbit of spacecraft and/or orbital debris 

After reviewing the options, the COMPASS team selected a spacecraft that emulates the ORBCOMM 

satellite (as mentioned previously), except that its 40-kg payload would be delivered to a 400-km circular 

orbit. This slightly less energetic launch trajectory may demand a slightly larger number of satellites in 

the constellation for the same coverage. ORBCOMM is a commercial venture that provides global 

messaging services (handling 50,000/hr), using a constellation of 26 LEO satellites in 785-km circular 

(two polar), three planes, eight spacecraft per plane, 45 ° inclined. Each spacecraft would weigh 40 kg, 

cost $1.2M, and consume 160 W of power over its 4-year lifetime. Very-high-frequency (VHF) 

communication (138 and 400 MHz) will be at 57.6 kbps. There will be 17 data processors and 7 antennas. 

The satellite will be gravity-gradient stabilized, with magnetic torquers and cold gas nitrogen (N2) 

Reaction Control System (RCS) thrusters. 

3.1.2.3 Spacecraft Key Features 

Figure 3.6 highlights the essential DRM 1–A spacecraft systems for which key features are itemized 

in the following list: 

 The GBL laser could be reused for orbit maintenance and reposition of the spacecraft, and 

could finally deorbit when the satellite electronics fail or become obsolete—that is, using 

“planned retirement” rather than aggravating the orbital debris problem.

Figure 3.6.—Principal spacecraft systems. SA, solar array; GN&C, guidance, navigation, and control; C&DH, 

command and data handling; PMAD, power management and distribution; He, helium. 
NASA/TM—2012-217014 117 

 The 120-cm COMPASS vehicle will be launched at a 30° elevation off the horizon and will 

require just one GBL station (350–MW maximum in air-breathing mode; 250 MW in rocket 

mode) at the launch point, and a laser relay mirror in LEO for orbit circulation; 

 Combined-cycle pulsed-detonation propulsion system and thermal control technologies (which 

are still in the TRL 2 to 4 range) include 

o A rear parabolic mirror will serve the double function of a focusing optic and a plug 

nozzle. 

o Pulsed laser power will be concentrated into a 1.2-m-diameter annular ring, where 

working fluid temperatures may exceed 6000 K and theoretically will provide specific 

impulse Isp in excess of 900 s. 

o Water propellant is best for density, cooling, and ground handling, but liquid nitrogen 

(LN2) and hydrogen peroxide (H2O2) are alternative options. 

 Guidance, navigation, and control 

o The vehicle is assumed to be spun at 120 rpm for stabilization during launch. Higher 

vehicle rotation may be required for desired stability levels in flight, but the payload bay 

could be despun to avoid acceleration constraints on payload. 

o Differential water propellant injection into the laser focal ring (60-cm radius) will be used 

for pitch and yaw control of the vehicle under laser boost, in both air-breathing and 

rocket modes. 

 Some launch power, C&DH, and communications systems will be reused for satellite functions 

once in orbit. 

 Eliminating the separate (first-stage) launch vehicle should reduce the lightcraft costs by more 

than half. (This does not include differences in launch support, range, and laser facilities.) 

3.1.2.4 Spacecraft Requirements and Assumptions 

Table 3.1 summarizes the principal requirements and assumptions made by the COMPASS team in the 

DRM 1–A design and feasibility study, which considered various important tradeoffs. 

TABLE 3.1.—DRM 1–A SUMMARY OF REQUIREMENTS AND ASSUMPTIONS 

 Item Requirements and assumptions 

1 Top-level/ 

science 

Lightsat shall perform both the launch and satellite functions. Representative satellite 

function to emulate ORBCOMM. Figures of merit (FOM): Cost per n units, payload mass, 

and safety. 

2 System Single fault safe only; otherwise zero fault tolerant; 4-year lifetime; launch year open 

ended—current TRL assessment is an output. 

Mass growth per AIAA 120–2006 (additional growth to make system level 30%). 

3 Mission, 

operations, and 

GN&C 

Lightcraft engine attached to spacecraft performs launch from ground to suborbit, 

circularization with relay mirror satellite or onboard chemical (circular, 400-km altitude, 

45o inclined (±0.2o). Desire reuse of lightcraft engine for repositioning, reboost, and deorbit; 

high rotation rate for spacecraft stability. 

4 Launch vehicle Not applicable (N/A) (integral SSTO vehicle reused as satellite/spacecraft in orbit) 

5 Propulsion Lightcraft engine primary propulsion: transitions through pulse detonation, laser ramjet, and 

laser rocket modes at appropriate altitudes. Water chosen as rocket propellant because of 

better density and cooling, and easier handling on pad. 

6 Power Solar photovoltaic (PV) arrays and secondary batteries 

7 C&DH and 

communications 

Preuse of flight computer to handle launch; uprange telemetry communications; equivalent 

ORBCOMM messaging payload 

8 Thermal/ 

environment 

Remove laser and plasma heat from body by vaporizing propellant. 

(Trades: propellant type for cooling: N2, water (H2O), or ammonia (NH3)). Maintain high 

reflectivity (>0.9999) of rear parabolic optic during laser boost. 

9 Mechanisms Air-breathing inlet covers (open for laser ramjet mode, closed for rocket modes); nosecone 

ejection; solar PV array and tether deployment 
NASA/TM—2012-217014 118 

TABLE 3.1.—DRM 1–A SUMMARY OF REQUIREMENTS AND ASSUMPTIONS 

 Item Requirements and assumptions 

10 Structures High-temperature-resistant inlet ring/annular shroud; maintain 0.9999 reflective parabolic 

mirror/plug nozzle under laser boost 

11 Cost Assume that lightcraft engine/vehicle replaces need for separate launch vehicle. Assume that 

commercial spacecraft building cost models are used. 

12 Risks Temperature leaks from laser and plasma; fouling mirror HX, laser pointing, thrust direction, 

and angle of attack, high spin rate 

At the top-level, the major tradeoff was the choice for an integrated vehicle/satellite with shared 

satellite functions (selected for DRM 1–A with a LEO laser relay mirror for orbit circularization) versus a 

smaller separable satellite payload (favored for DRM 1–A' with an onboard chemical kick rocket to 

circularize orbit). 

In Item 3, a vehicle rotation of 120 rpm was assumed, but higher spin rates may be required for 

adequate stability under laser boost. Under Item 5, two main propellants (H2O and LN2) were initially 

considered for the pulsed-detonation (PD) rocket mode, but H2O was selected because of its 20% higher 

density and better cooling capacity. N2 has better potential than H2O for dual use on orbit in the RCS. 

Onboard RCS propellant trades were cold gas N2 (favored) versus hydrazine. The trade in Item 6 was 

additional primary batteries for launch needs. In Item 7, the obvious trades were computer type, data 

storage, data transfer rates, and communication frequencies. Mechanism trades in Item 9 faced materials, 

power, and operations issues. Finally in Item 10, the overriding concern was for minimum structural mass 

to sustain expected thermomechanical loads. Microchannel water (H2O) -cooled titanium and hightemperature ceramic matrix composites (carbon/silicon carbide (C/SiC) and SiC/SiC)) were both 

considered. The rear parabolic mirror/plug nozzle must maintain 0.9999 reflectivity under laser boost. 

3.1.2.5 Satellite Function and On-Orbit Configuration 

Figure 3.7 gives a semitransparent view of the satellite electronics packages required for the 

ORBCOMM-like messaging function. The COMPASS team’s objective was to conceptually design and 

assemble all essential electronics into the conical-shaped payload bay (i.e., nose volume) of the 120-cm 

lightcraft launcher—for example, the avionics, communications, PV/battery power system, gravitygradient stabilization system, RCS, magnetometer, magnetic torquers, and inertial measurement unit—

using components largely scaled from existing technology and commercially available units. 
NASA/TM—2012-217014 119 

Figure 3.7.—Semitransparent view of 120-cm spacecraft and satellite electronics. IMU, inertial measurement unit. 

A more integrated, compact, and mass-efficient spacecraft/satellite electronics package could be 

created for the 120-cm lightcraft by using an integrated systems engineering approach, rather than the 

“first-order” subsystem-by-subsystem aggregate, as the COMPASS team has spearheaded. With a critical 

eye toward small payload and microsatellite engineering practices, the lightcraft electronics package 

could be designed for minimum volume, mass, and power needs. For example, by tightly integrating 

separate electronics enclosures into larger boxes (rather than “piece-meal” assemblies), a significant 

reduction in mass budget can be obtained, which in turn will reduce costs through mass production. Engineers already experienced in microsatellite and launch systems engineering and design, coming 

from major microsatellite companies, would welcome this challenge and make quick work of this. 

3.1.2.6 Combined-Cycle Propulsion System 

This section provides a brief overview of the lightcraft’s combined-cycle propulsion system, for 

which the energy is provided via a GBL. 

The engine operates in air-breathing laser ramjet mode up to Mach 6 to 7 and an altitude of 35 km 

and then switches into rocket mode (whereupon H2O is pressure fed to the 120-cm propellant injector 

ring, with no need for a boost pump) for the remainder of the flight to LEO (Refs. 16 and 17). As 

indicated in Figure 3.8, the entire vehicle aeroshell is designed to serve multiple structural- and 

propulsion-related functions. The axisymmetric nose/forebody serves as an external compression inlet in 

supersonic flight (Ref. 18), precompressing air through a short annular slit into the engine “hot section.” 

(Incidentally, the nose also supports four patch antennas—90 apart—for pre-deployment 

communications.) 
NASA/TM—2012-217014 120 

Figure 3.8.—Function of external surfaces in air-breathing propulsion mode. 

The entire vehicle aft surface is a parabolic mirror that concentrates incident laser power into the 

annular shroud region to heat the engine working propellant, which is either (1) atmospheric air or 

(2) onboard liquid propellant (H2O; liquid nitrogen, LN2; hydrogen peroxide, H2O2, etc.). Laser heating of 

the working fluid causes ionization, dissociation, and the creation of a high-pressure, high-temperature 

plasma toroid inside the shroud and laser-supported detonation (LSD) waves generate pulse-periodic 

high-pressure blast waves (see Figure 3.9). Hence, a quasi-continuous propulsive force is produced by the 

rapidly pulsing laser. The rear parabolic optic also serves as a plug nozzle surface for the expanding blast 

waves to efficiently impart impulse. Flight control (pitch and yaw) in the air-breathing mode is achieved 

by differential injection of H2O sacrificial coolant (very low flow rates). 

Figure 3.9.—Pulsejet engine cycle based on LSD waves. (Data for 11- to 14-cm 

lightcraft engine tests on Pulsed Laser Vulnerability Test System (PLVTS) laser at 

WSMR.) 
NASA/TM—2012-217014 121 

The annular shroud functions to enclose the engine “hot section” (laser energy absorption chamber) 

and to inject onboard liquid monopropellant (H2O) for subsequent laser heating in the rocket mode. 

Again, this monopropellant is still energized by the GBL, using repetitive LSD waves to process the 

propellant into a high-temperature plasma that generates successive high-pressure blast waves 

(Figure 3.9). The injectors (see Figure 3.10) employ self-acoustic valving to meter the propellant flow, 

tuned to operate at the laser pulse repetition frequency. This injector system is designed to prevent 

“chug.” 

Figure 3.10.—Propellant injector schemes for rocket pulsejet mode (Ref. 16). 

Note in Figure 3.5 and Figure 3.7 the inverted “Hershey’s kiss”-shaped (or pear-shaped) H2O main 

propellant tank and its 15.2-cm-diameter helium pressurant tank that is bonded inside the bottom tip. The 

helium tank is constructed of titanium alloy and reinforces the lower tip of the main tank. The H2O tank, 

also made from titanium, is bonded to the regeneratively cooled, 100-cm-diameter parabolic optic so that 

the resultant sandwich structure reinforces the mirror’s precise surface contour. 

The propellant flow provides regenerative cooling of the primary optic surface and the 24 cowl 

support struts (titanium), as well as film/transpiration cooling on the inner cowl surface (and possibly, the 

cowl and strut leading edges), as shown in Figure 3.11 and Figure 3.12. An integral RCS and main 

propellant storage system employ four additional equal-spaced, pulsed injectors on the rotating shroud, 

specifically for the RCS function (see Figure 3.13). The pear-shaped titanium H2O propellant tank with 

integral He spherical pressurant tank is a blow-down design—that is, pressure fed (no pump). 

The preliminary propellant injection and distribution (PI&D) in Figure 3.13 (left diagram) reveals the 

essential elements for this blow-down system equipped with a single propellant tank and single He 

pressurant tank. The pear-shaped, titanium alloy propellant tank is equipped with slosh baffles and has a 

channel propellant management device (PMD) around the tank periphery because of centrifugal forces on 

the propellant. The heat leak into the main and He tanks aids in the pressurization of the blow-down 

system. 
NASA/TM—2012-217014 122 

Figure 3.11.—Cutaway of engine “hot section,” showing 

cooling and propellant injector systems (Ref. 16). 

Figure 3.12.—Side and top views of shroud and support struts, explaining coolant scheme (Ref. 16). 

Figure 3.13.—Thrust vectoring, RCS, and propellant injector design.
NASA/TM—2012-217014 123 

This zero-fault-tolerant feed system has three branches: (1) differential injectors/RCS (four); 

(2) uniform rocket-mode injectors (evenly distributed around the 120-cm shroud); and (3) film/ 

transpiration cooling flow. As mentioned earlier, the H2O propellant flow also regeneratively cools the 

vehicle aftbody (parabolic mirror). The pulsed injectors provide vehicle attitude control through 

differential propellant injection into the toroidal beam focus area (i.e., the engine’s annular “absorption 

chamber” or “hot section”) during both air-breathing and rocket modes. Relying on commercial-off-theshelf (COTS) high-speed RCS valves, the system would also provide on-orbit attitude control of the 

spacecraft. 

3.1.2.6.1 Propulsion System Requirements and Assumptions 

Briefly summarized, the principal requirements and assumptions for this DRM 1–A propulsion 

system are 

 Requirements 

o Zero-fault-tolerant system 

o Tanks remain internal to vehicle 

o No staging 

o Laser-based air-breathing mode to gain altitude and velocity (Mach 6 to 7 and 35 km) to 

minimize the added ΔV of the rocket mode to reach orbit 

o Laser-based rocket mode to build up velocity using onboard propellant 

 Assumptions 

o Isp of 950 s with water monopropellant in rocket mode 

o Laser system can continuously track and beam energy to engine/vehicle. 

o He pressurization of main propellant tank (H2O) 

o Effective regenerative and film/transpiration cooling of optic and other surfaces 

3.1.2.6.2 Beam-Riding Feature and Aim-Point Strategy 

A unique and special feature of this lightcraft engine bears further elaboration; that is, the propulsion 

system generates a lateral restorative force that allows the vehicle to be a beam rider in both atmospheric 

and vacuum space environments. Figure 3.14 shows the impulsive forces and moments applied to a 

lightcraft engine when the laser beam is offset from the vehicle’s axis of symmetry. 

Figure 3.14.—Left: Centered versus offset beams. Right: Thrust 

impulse, side impulse, and pitching angular impulse are 

functions of lateral beam offset distance (Refs. 13 to 15). 
NASA/TM—2012-217014 124 

Figure 3.15 explains how axial impulse and lateral impulse vary as a function of laser beam offset and 

angular offset from the engine’s longitudinal axis. (This engine performance data has been extensively 

measured for a family of lightcraft engine geometries under stationary conditions in Refs. 6 to 8). 

Figure 3.15.—Performance variations in axial impulse and restorative (i.e., beam-riding) impulse 

versus lateral offset are actively exploited in the aim-point strategy (Refs. 13 to 15). 

The autonomous beam-riding feature significantly reduces the total RCS propellant load consumed in 

the laser boost phase, because only the vehicular pitch and yaw must be controlled. Furthermore, by 

directing the laser beam with a precise lateral offset on the lightcraft’s rear optic, the engine thrust may be 

vectored at will. By targeting a given “aim point,” the laser beam can effectively lead the vehicle (with a 

prescribed offset), coaxing it to fly along the desired boost trajectory. 

Figure 3.16.—Aim-point strategy in launching a beam-riding lightcraft (Refs. 13 to 15). 
NASA/TM—2012-217014 125 

3.1.2.6.3 Propulsion System Risks 

The following is an itemized list of propulsion system risks identified in the DRM 1–A study: 

 Confirm that rocket-mode time-averaged Isp value of 950 s can be achieved with H2O propellant 

(estimated at TRL 2) 

 Maintenance of appropriate beam angle and offset relative to vehicle spin axis 

 Adequate control authority of beam to enable engine thrust vectoring via beam lateral (and 

angular) offset and differential propellant injection 

 Excessive propellant pressure drop along cooling channels resulting in inadequate propellant flow 

rates 

 Acoustic valve failure resulting in incorrect propellant injection timing and/or inadequate 

propellant flow rates 

 Inadequate regenerative cooling of optic surfaces and exposed hot surfaces 

 Unknown film cooling requirements for exposed hot surfaces 

 Unproven performance of air-breathing ramjet/scramjet mode 

 Proper air inlet annular door/gap closing operation in air-breathing to rocket mode transition 

 Robust design: allow for margin in component design to support zero fault tolerant requirement 

3.1.2.6.4 Recommended Future Trades 

Future trades considered in the combined-cycle laser propulsion system design include (but not 

limited to): 

 Dense noncryogenic propellants: for example, ammonia, hydrazine, and hydrogen peroxide 

 Moving the rear optic focus from the inner shroud surface, into the flowfield (i.e., smaller focal 

radius than 60 cm) in order to reduce cooling requirements 

 Varying the number and configuration of pulsed propellant injectors 

 Pumped propellant system versus He pressurized (blow down) 

3.1.2.7 Structures and Mechanisms Subsystems—Description 

The high-level design requirements for structures and mechanisms are 

 Contain the necessary hardware for research instrumentation, communications, avionics, 

propulsion, and power (see Figure 3.17) 

 Withstand applied mechanical and thermal loads from launching 

 Provide minimum deflections, sufficient stiffness, and vibration damping 

 Minimize mass 

Figure 3.17.—Lightcraft DRM 1–A configuration—oblique, top, and side views. 
NASA/TM—2012-217014 126 

The lightcraft structures assumptions are 

 Provide main structures backbone for BEP vehicle 

 Materials include graphite/polyimide, C/SiC or SiC/SiC, titanium, and aluminum 

 Shells have beam and/or stringer reinforcement 

 Fastener assembly is bonded, welded, and threaded 

The mechanisms assumptions are 

 Provide support for (1) forebody clam shells and nose cone, (2) separation and tether (for 

forebody shells), and (3) shroud air inlet flaps 

 Mass of mechanisms for pyrotechnic bolts and springs— 

o Forebody shells and nose cone, 1.82-kg basic mass total 

o Forebody shells tethers, 0.4 kg each; quantity, 2 

o Gravity gradient tether, 0.11 kg 

 Mass of mechanisms for shroud air inlet gap/flaps linear actuators—0.05 kg each; quantity, 24 

(i.e., one flap sandwiched between each pair of struts). 

The structural design was ultimately to exploit advanced composites and titanium components, with 

flanges for joining sections and mounting internal hardware made from aluminum. The structure 

subsystems will employ shells to bear all operational loads. Aluminum composite sandwich structures 

will see application in deck and beams for mounting hardware (science probes, payload support, etc.). 

The primary mass components of the structure will be (1) the forebody of carbon/carbon and 

graphite/polyimide with aluminum flanges and stringers, (2) the annular shroud thrust structure (with 24 

support struts) of titanium, (3) the main propellant tank of titanium, and (4) the afterbody parabolic mirror 

of C/SiC. The carbon/carbon nose cone is at TRL 6, as is a graphite/polyimide shell forebody, with 

aluminum flanges and stringers in the primary structure. The annular shroud and inlet gap flaps will be of 

titanium alloy (TRL 6). 

The secondary mass components of the structure will be (1) the aluminum composite sandwich 

structure instrumentation deck and installations, (2) the spacecraft adaptor and separation mechanism 

details, and (3) the pyrotechnic bolts that clamp the flanges of the two-part forebody. 

The favored design choice for the main structure material would ultimately be refractory metal-coated 

C/SiC, especially for the annular shroud and parabolic mirror afterbody. This choice would likely lead to 

the minimum mass and maximum high-temperature capability (i.e., the primary drivers); the NASA 

Technology Readiness Level of such ceramic matrix composites is at TRL 4. Analytical methods were 

applied for stress analysis scoping studies. For example, an advanced C/SiC mirror, supported on the 

launch pad with a point load of ~300 kg (at 1g) applied to the mirror’s lower tip, must sustain ~43 MPa 

(including a 1.5 safety factor), which is well below the 240-MPa failure stress of C/SiC. In the early 

stages of development, the forebody sections could be made of aluminum, but final construction would 

likely use graphite/polyimide. Structural mass could be reduced (or margin increased) with a more 

detailed stress analysis using FEA.

Table 3.2 gives the mass breakdown for the 120-cm DRM 1–A spacecraft. The spacecraft dry mass 

will be 141 kg, so with 188 kg of H2O propellant, the estimated total (wet) launch mass will be 318 kg 

(based on the Optimal Trajectories by Impact Simulation (OTIS)). 
NASA/TM—2012-217014 127 

TABLE 3.2.—MASS BREAKDOWN FOR 120-cm DRM 1–A SPACECRAFT 

3.1.2.8 GN&C Overview 

The objective of the GN&C system will be to facilitate lightcraft injection into a 400-km circular 

orbit with a 28.5 inclination. The pitch control on ascent was assumed to be provided by (1) differential 

thrust while in sight of the laser and (2) cold gas thrusting when not in sight of the laser (i.e., needed for 

orienting the vehicle to the correct attitude for the second laser burn (LEO insertion) using a relay mirror 

satellite). For this first-order analysis, the BEP lightcraft was assumed to have zero products of inertia. 

The navigation electronics included the following: 

 One LN–200S IMU 

 An IMU with low mass (0.75 kg) and a design proven in space flight on Clementine, Deep 

Space I, and the Mars Exploration Rovers (MERs) 

 Three solid-state fiber-optic gyros with a bias stability of 0.1/hr 

 Three solid-state silicon accelerometers with a bias of 0.3 mg 

 One Ithaco IM–103 three-axis magnetometer to measure the strength and direction of Earth’s 

magnetic field 

 Four patch antennas to be used for Global Positioning System (GPS) orbit and attitude 

determination 

The control system electronics will manage both launch and on-orbit functions. The launch trajectory 

control system will encompass 

 Passive stability will be provided by spinning the vehicle at 120 rpm. 

 The vehicle will use the inherent beam-riding feature while in sight of the laser. 

 The main pitchover maneuver will be provided by differential thrusting. 

 Prior to orbit insertion, the nosecone will separate and act as yo-yo weights to despin the vehicle 

from 120 to 60 rpm. 

 Cold gas thrusting will be used to pitch the vehicle to the correct attitude for the LEO insertion 

burn. 

The on-orbit system will utilize 

 Three Ithaco TR10CFR magnetic torquers to provide up to a 15-Am2

 dipole moment magnetic 

torquers to despin the vehicle after injection into the circular orbit as well as to provide three-axis 

control for on-orbit operations 

 Gravity gradient stabilization will be activated after the vehicle has been despun in orbit. A tether 

will be deployed with a 2.6-kg mass attached. 

 A nutation damper will be used to provide additional energy dissipation for increased stability. 
NASA/TM—2012-217014 128 

3.1.2.9 Avionics Overview 

The avionics package design requirements follow: (1) avionics for systems command, control, and 

health management; (2) single-string processor architecture; (c) low-mass optimization; and 

(4) operational power modes that vary according to requirements, quiescent until needed. The space 

allocation for the avionics enclosure will specify a semicircular package (see Figure 3.18). 

Figure 3.18.—Avionics functionality (120-cm launcher; I/O, input/output). 

The principal assumptions include (1) 50-kRad avionics will be used for stable operation for the 

duration of the mission, (b) cabling mass will be estimated with a Monte Carlo simulation (mass and 

length), and (3) all avionics are at TRL 5—needing only to fly because of electrical, electronic, and 

electromechanical (EEE) and Class-S parts usage. 

The following design features will be included in the avionics subsystem: 

 Radiation-tolerant field-programmable gate arrays (FPGAs) packaged with Internet Protocol (IP) 

for processor capability, embedded software kernels without major operating system (O/S) 

overhead (optimized GN&C) 

 Memory of 5 GB with error correction 

 Power supply stabilized with switching regulators, filter, and electromagnetic interference (EMI) 

shielding to accept wide variations in supply voltage 

 Discrete devices and power driver cards for general vehicle functions 

 I/O cards for sensors and drivers 

 Housekeeping data and health and safety (H&S) data to ground via a communication channel 

at ~ to be determined (TBD) bits per second 

 High-TRL EEE parts in custom card designs, –40 C ratings 

The general avionics processor could be a Power Personal Computer (PC), chip, or intellectual 

property in an FPGA. The semicircular avionics card cage will use stacked custom boards in an open 

frame design. Other processor functionalities include 

 Gimbals and thrust vector control (TVC) cards will be mass reduced and will control any 

necessary positioning (e.g., solar array deployment and gimbaling on orbit; communication 

antenna gimbaling to be resolved (TBR) on orbit). 

 GN&C navigation—data from IMUs will be used to derive and control TVC parameters 

(0.16 MIPS nominally). 

 Systems health and status reporting 
NASA/TM—2012-217014 129 

 Power will be managed for various power modes (power trimmed with specialized sentinel 

switching) and will be full on for launch. 

 System management will include control of valves and heaters. 

 There will be a TBD bps serial interface between C&DH and the communications channel. 

The sensors and drivers list includes 24 valves, TBD igniters, pyros, RCS, and pressure and 

temperature sensors. The software rough-order-of-magnitude (ROM) source lines of code (SLOC) 

estimate of ~140,000 to 180,000 includes launch. 

Avionics risks include 

 RAD 750 FX obsolescence likely, replace with FPGA (very-high-speed integrated circuits 

(VHSIC) hardware description language (VHDL) intellectual property) 

 Particle radiation tolerance (shielding, high-radiation EEE parts; FPGAs exist with high radiation 

tolerance) 

 Elevated vibration levels from the pulsejet engine during launch—may require “tuning” of the 

structure and board layout for stiffness 

The geometry of stiffening members for enclosure and printed circuit boards will be determined 

through vibration analyses (experimental and numerical). 

3.1.2.10 Communications—Description 

The communications system will provide communication to the vehicle during launch and will 

demonstrate capabilities similar to that of ORBCOMM in space, using a single-string system. The S-band 

at NASA frequencies was chosen to reduce the mass and size of the communication system, but other 

frequency bands could be visited and analyzed later. The system will communicate from individual 

transmitters at 4.8 kbps or greater, and from space gateways at 57.6 kbps or greater with a nongimbaled 

antenna system (Figure 3.19). 

Figure 3.19.—Communication operations diagram. 
NASA/TM—2012-217014 130 

At liftoff, the spacecraft will communicate initially to a balloon or airplane near the launch site before 

switching to down-range gateways on the ground as it overflies these terminals along the trajectory (see 

Figure 3.20). The number of gateways will be determined by the trajectory and antenna gain chosen. A 

four-patch antenna (see Figure 3.7 and Figure 3.16) on the spacecraft nose will enable S-band 

communications with the airplane or balloon, as well as Earth stations down range of the launch site. In 

orbit, the satellite will communicate to the ground with a helix antenna. Other types, like patch antennas, 

should be considered in the future. The helix antenna, which is entirely enclosed by the nosecone until it 

is jettisoned, will be pointed along the spacecraft axis toward the ground, and it will not be gimbaled. The 

satellite will communicate to individual users and to gateway users and will transfer the information 

between users and gateways. A store-and-forward function will facilitate the relay of information 

between sites. Gateway users will be connected to the Internet. The only identifiable risk is the need for 

flipping the spacecraft over for communications with the ground. 

Figure 3.20.—Launch gateway sites. 

A trade study was conducted on the communication link budget versus the antenna gain and Earth 

coverage. The largest feasible helix antenna was assumed in the master equipment list (MEL) for 

estimating communications system mass. The specific trades performed were 

 Communication spot size versus antenna gain 

 Maximum communication range from the user versus antenna gain 

 Data rate versus antenna gain 

 Different antenna gain for an individual user versus a gateway user 

 Data rate versus low-noise amplifier (LNA) noise temperature 

From Table 3.3 on the gateway to the satellite/spacecraft, one can see that almost all beam angles will 

meet minimum data rate requirements from the satellite to the gateway station, when the gateway station 

is using a 1-m-diameter dish. Only the largest beam angle is not achievable for the helix antenna design, 

because the antenna will have less than 1 coil. The helix antenna will become too long for the spacecraft 

for beam angles of less than 30o. 

Table 3.4 displays data rates for two margins (3 and 20 dB) for three different user transmit powers: 

0.25, 1, and 2 W. The green boxes fulfill the minimum requirements of the study. For low-data-rate users 

requiring 4.8 kbps, most antenna beam angles will work even with only 0.25 W of transmit power, if only 

3 dB of margin is needed. However for many users, especially within cities, a higher margin will be 

required. 
NASA/TM—2012-217014 131 

TABLE 3.3.—TRADE STUDY—DATA RATE AND COVERAGE 

VERSUS BEAM WIDTH TO GATEWAY

Many communication applications exist for a constellation of laser-launched satellites, ranging from 

Earth-observing to mining, trucking, and just about any low-data-rate system that is power sensitive. 

Examples of Earth-observing applications include the monitoring of 

 Sea (and land) temperatures, salinity, carbon dioxide (CO2)/pollutant concentrations, wind speed, 

wave actions (height, frequency, etc.), and solar flux 

 Seismic activities in remote locations (e.g., earthquakes) 

 Vegetation coverage and assessments 

 Forest fires 

 Tagged-animal movements 

 Data from remote mining, oil pumping stations, and so forth. 

TABLE 3.4.—TRADE STUDY—DATA RATE AND COVERAGE 

VERSUS BEAM WIDTH TO USER
NASA/TM—2012-217014 132 

3.1.2.11 Power System—PV Array and Battery Description 

Itemized in the following subsections are the requirements and assumptions used in the design of the 

DRM 1–A spacecraft’s PV and battery power system (see Figure 3.21), followed by the power system 

risks. Table 3.5 shows the power equipment list (PEL) for the 120-cm spacecraft. 

3.1.2.11.1 Power System Requirements 

 No redundancy, single string (i.e., no redundant battery cells, battery packs, or solar array strings) 

 Orbital user power (day and night) = 105 W (includes 30% growth) 

 Orbit = 400 km (93-min orbit period/36-min eclipse); no orbital altitude degradation, eclipse 

period is worst case 

 Life = 4 years 

TABLE 3.5.—PEL FOR 120-cm LIGHTCRAFT LAUNCHER 

Figure 3.21.—Left: Flexible PV array. Right: Power circuit schematic. 

3.1.2.11.2 Power System Assumptions 

 Full Sun-tracking will be used to minimize the solar array area and mass. 

 Two solar array wings will be used with only one axis gimbal per array and yaw steering (full 

Sun-tracking to minimize solar array area); each wing gimbal axis must be parallel, thus one or 

two will be possible. 

 10% of solar array power will be degraded over the mission because of radiation, atomic oxygen, 

ultraviolet radiation, thermal cycling, contamination, and orbital debris and micrometeorites. 
NASA/TM—2012-217014 133 

 Array thermal properties: emissivity top = 0.85, emissivity bottom = 0.8, absorptivity top = 0.92. 

and absorptivity bottom = 0.6. 

 Triple Junction Solar Cell efficiency: at 28C (no degradations) = 36%; 70C (400 km, no 

degradations) = 32%; 70C (400 km, with degradations) = 29%. 

 10% losses in power system (due to electronics, harness, and connector efficiencies). 

 Areal mass= 3.66 kg/m2

 (if 3-mil coverglass assumed on each side of solar cell). 

 Cell packing factor into solar array = 0.85. 

 Lithium batteries = 200 W-hr/kg, 50% depth of discharge (limited because of number of cycles). 

 28 Vdc 

 Technology status: solar arrays = TRL 4, gimbals = TRL 8, and batteries = TRL 4. 

3.1.2.11.3 Power System Risks 

 The complex launch environment (i.e., vibrations induced by the pulsed detonation engine) will 

impact the power system’s hardware and design.

 High-spin acceleration combined with pulsed accelerations along the vehicle launch direction and 

spin axis, as well as high thermal loads from both the laser and aero-heating, will add uncertainty 

to the power system design. 

 Although stowed flexible arrays can be designed for high uniform load accelerations (20g to 

50g), this concept requires more analysis than a conceptual design to confirm, especially for a 

low to moderate TRL array (not off the shelf). 

 Even lithium batteries cannot be assumed to be immune from spin accelerations due to internal 

electrolytes that may require the battery package to have a specific orientation to function 

properly. 

 Thermal loads require more analysis. The present stowed solar array design was focused on being 

able to separate/partition the power system from this problem as much as possible, but the need 

for thermal blankets or isolation has not been quantified. 

 Consider reducing the vehicle spin rate (or despin payload), orienting stowed solar arrays parallel 

with the spin axis to minimize differential loading, and structural analysis to determine required 

supports for the stowed solar array during launch. 

3.1.2.11.4 Power System Trades 

Fixed versus tracking solar arrays.—Although fixed arrays could be assumed, this would increase 

the size of the solar array (50% to 100%) because of a decreased effective Sun period and cosine losses; 

the energy storage mass would also increase because of an increased effective eclipse period. Therefore, 

tracking arrays were selected. 

Tracking method of solar arrays.—There are various ways to Sun-track the solar arrays: 

(1) Inertial flight (i.e., pointing the entire spacecraft at the Sun) 

(2) Two-axis gimbals for each solar array 

(3) One-axis gimbal for each solar array wing with yaw steering (i.e., this gimbal tracks about the 

lengthwise axis of the solar array; the other axis is handled using yaw steering of the spacecraft) 

For small spacecraft, yaw steering is the most economical method for Sun tracking since inertial 

flight modes are somewhat costly, as are too many gimbals. 

Energy storage type.—The options for projected available technology for energy storage include 

(1) Electrodynamic tether: Extracts energy by passing through the Earth’s magnetic field; it is likely 

heavier than other options, has a tether dynamics risk, and causes orbit altitude reduction 

(2) Lithium sulfur battery: Has very high specific energy but presently has too few demonstrated 

charge/discharge cycles to meet the LEO orbit life of >1 year; also has a low operating depth of discharge 

to limit degradation (likely not ready by 2020) 
NASA/TM—2012-217014 134 

(3) Lithium ion battery: Has extensive GEO and some LEO testing and the potential is high for LEO 

cycling rates at a moderate depth of discharge 

Although the lithium ion battery type assumed has a higher specific energy than for SOA LEO 

lithium ion batteries, it is anticipated that the technology will advance rapidly. If not, then a moderate 

mass penalty due to a lower specific energy SOA lithium ion battery can be assumed. 

Power generation type.—The options for projected power generation technologies include 

(1) Electrodynamic tether: Can provide power during the entire orbit but will really require some 

method of orbit-raising, which is likely not a minimum mass solution 

(2) Solar PV array painted or grown on the outer structure: Low efficiency (~10%), high 

development risk because of stress, vibrations, and thermal problems; need conductive surface to keep 

solar cells cool to operate efficiently, but the outer surface must be an insulator; unknown radiation, 

ultraviolet light (UV), and atomic oxygen effects on solar cells; and low TRL—likely not ready by 2020 

(3) Concentrator PV cells using BEP mirror pointing at Sun: Low TRL; unknown degradation of 

reflective surface over time; requires high Sun pointing accuracy or larger area of solar cells; unknown 

integrated impacts of this on the vehicle; requires significant radiator area to keep the cells cool to have 

reasonable efficiency and not melt, with likely some sort of pumped loop cooling; and requires much 

more analysis to determine if it can be practical 

(4) Thin-film solar cells: Low efficiency, thus too large of an area if deployed or used on limited 

surface area of the vehicle, but they are at a high TRL 

(5) Deployable rigid panel wings: High TRL, fairly heavy—2 to 3 times heavier than flexible panel 

wings; hinges are simple and reliable, but fitting into the limited volume vehicle “bay” may present a 

problem because the stowed volume is large; risk due to centripetal and launch pulse accelerations and 

vibrations and due to integration into nonoptimal vehicle structure; if integrated into the “nose” of the 

spacecraft inner surface because of other components, adequate surface area would be unlikely for the 

power levels required; and, because articulation is required to minimize the solar array surface area, it is 

harder to articulate the heavier surfaces and extend them) 

 (6) Inflatable PV wings: Low TRL, need significant development, could integrate various cell types 

into the inflatable surface, and could rigidize the wings by UV 

(7) Deployable flexible PV wings: Moderate TRL, Roll-Out Solar Array (ROSA) or similar simple 

unrolling or unfolding types, no thick substrate or mass to attach cells to, highly compact during launch, 

very light, thermally favorable to keep solar cells at low operating temperature and at high efficiency, low 

mass, and minimizes gimbal requirements 

3.1.2.11.5 Power System Recommendations and Work to be Done 

General 

 Factor the range of orbit altitudes (i.e., what is the minimum operational altitude?) into the 

design. 

 Reduce payload compartment spin rate and increase payload volume—could despin the entire 

payload and/or nose components. 

 Consider jettisoning the annular shroud of the spacecraft. 

 Batteries/solar arrays need spin-testing for this unique acceleration/impulsive force launch 

environment. 

 Since this is a commercial vehicle and COMPASS does space-rated/government vehicle designs, 

cost is not minimized by using lower quality solar cells, non-space-rated batteries or cells, or 

gimbals. This needs to be considered in some manner. 

Solar Arrays 

 Perform more thermal load and stress analysis on the solar array, and perform more detailed 

layout of the array. 
NASA/TM—2012-217014 135 

 If lightcraft is reusable, obtain reentry requirements to determine the need for retraction or 

jettison of solar array. 

 Consider a modified design of the solar array deployment boom and articulation by increasing the 

length of the solar array and permitting the bottom quarter or third to be unpopulated by solar 

cells (possibly by a radiative surface) to simplify the deployment booms and mechanisms. 

 Consider using only one solar array wing. This will require stowed packaging and orbital 

dynamics analyses. It could help to reduce complexity and part count, increase reliability (one 

wing deploy rather than two), and reduce shadowing impacts. However, the balance of the 

spacecraft through the orbit would be impacted. 

 Consider one gimbal for both solar arrays. They would both rotate in the same direction, so some 

cost could be saved by connecting them to eliminate one gimbal. 

Battery 

 Obtain and analyze the time-phased load profile for the launch phase (and reentry phase, if 

applicable, since solar power is not available at this time) to better size the battery and determine 

nominal and peak current impacts. 

 Consider the use of a higher TRL, heavier battery (somewhat heavier) to reduce design risk. 

 Confirm thermal analysis of the battery to ensure that it will survive the launch and orbit 

environment. 

3.1.2.12 Thermal Subsystem—Description 

Lightcraft thermal subsystems must address two entirely different flight environments: (1) during 

launch and (2) on-orbit (see Figure 3.22). The total thermal subsystem mass is estimated at 15.23 kg. 

Figure 3.22.—Lightcraft thermal subsystems. Left: During launch. Right: On-orbit. 

The ascent thermal subsystem will cool the aft parabolic mirror and the engine “hot section” parts 

(i.e., annular shroud and 24 support struts), largely through regenerative, film, and transpiration cooling. 

The principal assumptions are 

 The plasma toroid (created by pulsed laser heating) will be located sufficiently far from the aft 

parabolic mirror and shroud lower surface such that the needed cooling will be reduced to that 

available by the propellant flow. 

 The aft mirror reflectivity will be 0.9999 through the use of multilayer high-reflectivity coatings 

tuned for the laser wavelength (~1 mm). 
NASA/TM—2012-217014 136 

The on-orbit thermal subsystem must cool the electronics payload, including avionics, 

communications, and GN&C components. The assumptions include 

 Heat dissipation for 100 W while in orbit 

 An Earth-view factor of 0.5 

 A solar-array-view factor of 0.15 

 A Sun angle of 30° 

LEO environmental conditions were used to determine the heat transfer while in orbit. 

The thermal subsystem components include the following: 

 Microchannel heat exchanger (HX) passages and pressure-fed H2O to regeneratively cool the aft 

parabolic mirror and the shroud undersurface where the laser plasma is formed during ascent 

 Additional transpiration and/or film cooling of the engine “hot section” parts as needed during 

ascent 

 Cold plates and heat pipes to transfer heat from the electronics to the radiator section 

(Figure 3.22) in space 

 Possible use of the aft parabolic mirror as a radiator when in orbit 

 Multi-Layer Insulation (MLI) used to insulate the electronics boxes 

3.1.2.12.1 Trades Considered in Thermal Analysis 

 High reflectivity of the aft parabolic mirror will be critical to the successful operation of the 

engine/vehicle. If reflectivity degrades badly during launch, insufficient H2O propellant flow will 

be available to cool this surface and complete the boost (could lead to catastrophic failure: i.e., 

loss of integrity of mirror reflectivity leads to increased thermal load on the mirror). 

 Hence, placement of the high-temperature, pulsed plasma toroid at a safe distance from the optic 

(and shroud) walls is critical to engine functionality. If the plasma is positioned too close (i.e., 

within a few centimeters) to these walls, excessive heating will occur—potentially leading to 

structural failure (melting and disintegration). This could be resolved by varying engine design 

geometry. 

3.1.2.12.2 Radiator Sizing 

 A radiator panel (Figure 3.23) was sized to reject excess heat from the electronics packages once 

in orbit. 

 A surface mount radiator was utilized. 

 The radiator model was based on a first-principles analysis of the area needed to reject the 

identified heat load to space. The assumptions are given in Figure 3.23. From this radiator area, a 

series of scaling equations were used to determine radiator mass. The worst case thermal 

environment of LEO was used to size the radiator. 

 No louvers were utilized. 

 Heat pipes were used to conduct excess heat to the radiator surface. 
NASA/TM—2012-217014 137 

Figure 3.23.—Surface mount radiator sizing—thermal analysis. 

3.1.2.12.3 Cold Plates 

 Electronics cooling will be accomplished by cold plates (see Figure 3.24) and heat pipes. 

 Electronic components will be mounted onto cold plates that have integral heat pipes to transport 

excess heat to the radiator. 

 Cold plates also will have heaters integrated onto them for maintaining electronics at desired 

temperatures throughout the mission. 

 Flat plate heaters on these cold plates can provide heat to electronics if necessary. Additional 

heaters will be placed throughout the spacecraft interior to facilitate adequate temperature control 

of the vehicle. 

 Thermal control will be accomplished through a network of thermocouples, whose data will be 

used to control the thermal power input to various heaters. A data acquisition and control 

computer will operate this thermal management system. 

Figure 3.24.—Left: Sample cold plate for electronics cooling. Right: Thermal analysis assumptions. 

3.1.2.12.4 Spacecraft Insulation 

A first-order thermal analysis of the spacecraft insulation needs was carried out for which the 

principal assumptions are given in Table 3.6. 

 MLI will insulate the spacecraft for operation in orbit. 

 MLI provides its function by limiting radiation heat transfer to and from the spacecraft. 

 MLI will be constructed of numerous layers of metalized material with a nonconductive spacer 

between the layers (see Figure 3.25). 
NASA/TM—2012-217014 138 

 Each layer of insulating material will have a low emissivity, which will limit radiation heat 

transfer between the layers. 

 The metalized material will have a low absorptivity, which will resist radiative heat transfer 

between the layers. 

TABLE 3.6.—THERMAL ANALYSIS ASSUMPTIONS—SPACECRAFT INSULATION 

Figure 3.25.—MLI for spacecraft on-orbit.

3.1.2.12.5 Heaters and Thermal Control 

Flat-plate heaters on the cold plates will provide heat to the electronics when necessary. Additional 

heaters will be placed throughout the spacecraft interior to facilitate adequate temperature control of the 

vehicle. Thermal management will be accomplished through a network of thermocouples whose data 

output will be used to control the power of the various heaters. A data acquisition and control computer 

will operate the thermal management system. 

3.1.2.12.6 Thermal Risks 

The main risk with the ascent thermal subsystem components pertains to the rear laser reflector, 

plasma degradation of that mirrored surface, and high potential heat transfer rates during ascent: 

Mirror reflectivity.—Any damage or degradation to the mirrored surface experienced during launch 

could cause excessive local heating because of the intensity of the laser beam (approximately 3-kW/cm2

incident). Debris impact or erosion from the nearby hot plasma could potentially degrade mirror 

reflectivity and cause catastrophic heating and loss of the vehicle. The very high reflectivity of the mirror 
NASA/TM—2012-217014 139 

(99.99%—using multilayer coatings tuned to the laser wavelength) will have to be maintained throughout 

the launch phase. This is indeed possible with high multilayer dielectric film coatings today 

(see Figure 3.26). 

Figure 3.26.—Reflectance of multilayer dielectric film coatings versus normalized frequency. 

Plasma heating.—Localized heating is due to the laser-created plasma “working fluid.” If this pulsed 

plasma is deliberately positioned within a few centimeters of the mirror surface, excessive heating of 

nearby walls will result. The mirrored surfaces cannot be highly reflective over the broad range of 

frequencies radiated by this hot plasma. Therefore, these walls will absorb significant amounts of heat if 

the plasma moves too close to the mirrored surface—likely in the range of 20 to 80W/mm2

 depending on 

the plasma distance (see Figure 3.27). In comparison, the throat section of the Space Shuttle Main Engine 

(SSME) nozzle saw ~10 W/mm2

. Such thermal loadings can be estimated from electric arcs (e.g., in 1-bar 

atmosphere), wherein plasma temperatures easily reach 30,000 K—quite typical of peak temperatures in 

pulsed laser propulsion. 
NASA/TM—2012-217014 140 

Figure 3.27.—Radiation heat transfer from electric arc in 1-bar atmosphere. 

Reflected laser light.—Laser light scattered off of the annular shroud’s rear-facing (convex) surfaces 

and aft mirror (concave) surface will be reflected back to Earth. In essence, this could be a health hazard 

to someone inadvertently looking into the reflected beam, because ~1-μm laser light is not eye-safe (i.e., 

the beam will propagate through the cornea and lens and onto the retina). The incident intensity in watts 

per centimeter squared of the reflected beam that reaches the Earth’s surface will depend on the curvature 

of the mirrored surface and the line-of-sight range to that spot on the Earth. And, of course, the smallest 

radius of mirror curvature yields the greatest dispersion and therefore lowest intensity of the reflected 

beam. For this reason, all external rear-facing lightcraft surfaces were designed as convex or concave 

optics to quickly dissipate reflected intensities to nonhazardous levels. In contrast, for planar surfaces, 

such as the flat sides of the 24 struts that support the annular shroud, another approach will have to be 

taken. Diffuse reflecting surfaces (Figure 3.28) will be applied to prevent planar reflections of incident 

laser light back toward the Earth. However, these struts would only be exposed to the incident laser power 

beam if the lightcraft were tumbling out of control, whereupon the beam would have normally been shut 

off anyway. 
NASA/TM—2012-217014 141 

Figure 3.28.—Diffuse reflecting surface 

for scattering laser light. 

3.1.2.12.7 Options and Recommendations (Thermal Subsystems) 

 A more detailed analysis of plasma thermal interaction with the rear optics and other engine “hot 

section” components must be performed to accurately assess its impact and identify options for 

further reductions in thermal subsystem mass. 

 Further analysis and experimentation should be performed to determine the amount of mirror 

degradation that will occur (if any) during launch and ascent. 

3.1.2.12.8 CONOPS—Description 

See section 3.1.2.18, CONOPS—Frequency of Launch and Repeat Time. 

3.1.2.12.9 OTIS Trajectory Simulation—Description 

See section 3.1.2.20, Launch Trajectory Simulations. 

3.1.2.13 Multi-CubeSat Launcher Design for 120-cm Lightcraft 

After completion of the ORBCOMM emulation study, the COMPASS team briefly explored the 

option to launch six CubeSats (~1 kg, 10 by 10 by 10 cm each) with the 120-cm lightcraft concept—see 

Figure 3.29. The basic ground rules were as follows: 

 Retain the same launch trajectory, beam power levels, and infrastructure/stations. 

 Remove only the ORBCOMM equivalent and satellite equipment, and keep everything needed 

for launch and insertion. 

 Add a dual Poly-Pico Satellite Orbital Deployer (P-POD) CubeSat dispenser for deploying six 

CubeSats (1 kg) once on orbit (Figure 3.27). 

The P-POD dispenser fits nicely in the conical payload bay, but the 120-rpm spin rate issue will likely 

dictate a redesign of the P-POD and CubeSats unless the payload section (spacecraft nose) is despun from 

the moment of launch. However, since the 120-cm launcher may still cost ~$2M ROM, would it make 

economic sense to launch six $30,000 CubeSats with it? 
NASA/TM—2012-217014 142 

Figure 3.29.—Alternative payload for 120-cm spacecraft—dual P-POD launcher. 

The MEL for this modified 120-cm CubeSat launcher is given in Table 3.7. The P-POD deployment 

device and CubeSat mass is carried (i.e., accounted for) in the structures and mechanisms allotment. An 

additional 3 kg is unallocated but is held for redesigning the P-POD and CubeSats to handle the high-g 

spin rate. 

TABLE 3.7.—MEL FOR 120-cm LIGHTCRAFT LAUNCHER FOR 

SIX CUBESATS, DEPLOYED FROM DUAL P–POD ON ORBIT 

However, considering the embryonic state of diode-pumped, solid-state lasers (e.g., now at the 100- 

to 120-kW level), it seems more prudent to first examine prospects for launching just a single CubeSat, 

with a much smaller and lighter launcher—for example, a 35-cm craft like that proposed in AFRL’s 

recent study (Refs. 5 and 19) invokes a 12-MW GBL. That far smaller launch infrastructure would be 20 

to 30 times less powerful than what the 120-cm lightcraft requires (i.e., 250 to 350 MW). Clearly, a single 

nanosat launcher is the more prudent and attainable objective in the near term. 

3.1.2.14 Description of 35-cm Vehicle 

Figure 3.30 shows the relative scales of the 120-cm (i.e., DRM 1–A) and 35-cm lightcraft (designated 

DRM 1–A') next to AFRL’s 35-cm nanosat launcher concept from References 5 and 19. The objective of 

launching 1.33-kg CubeSats (DRM 1–A') with a 12-MW laser is consistent with the progressive 

technology roadmap (see sections 3.1.2.1 Lightcraft Technology Development and section 4.0 Summary 

and Conclusions), a growth strategy that may begin with 0.1- to 1-kg picosatellites (picosats)—the first 

target of opportunity for a ~1-MW laser. The natural progression would be to demonstrate boosting 
NASA/TM—2012-217014 143 

increasingly larger payloads to ever higher altitudes over time: that is, starting with picosat (0.1- to 1-kg), 

then nanosat (1- to 10-kg), then microsat (10- to 100-kg) payloads. 

Figure 3.30.—Relative scale of lightcraft launchers. Left: Microsatellites. Right: Nanosatellites. 

Table 3.8 compares the MEL component mass breakdown of the two BEPS lightcraft launchers. Note 

the assumed CubeSat payload of 1.33 kg. To the first order, lightcraft mass can be assumed to scale with 

the cube of its diameter, so the 35-cm craft will be ~40 times lighter. Table 3.9 gives the mass breakdown 

for AFRL’s 35-cm lightcraft, designed for a 12-MW GBL and showing a liftoff mass of 8.0 kg. 

TABLE 3.8.—MEL MASS BREAKDOWN FOR 35-cm 

VERSUS 120-cm BEPS LIGHTCRAFT 
NASA/TM—2012-217014 144 

TABLE 3.9.—MEL BREAKDOWN FOR AFRL 35-CM 

NANOSAT LAUNCHER (REF. 5) 

3.1.2.15 BEPS Versus AFRL Vehicle/Engine Comparison 

 Both 35-cm lightcraft will use combined-cycle air-breathing rocket engines to minimize onboard 

liquid propellant mass (consumed in the rocket mode). 

 The BEPS craft will have a 30 nose half angle versus AFRL’s more pointed 10 to 20. Both 

vehicle forebodies will function as external compression air inlets. 

 The BEPS craft will transition to rocket mode at Mach 6 to 7 and 30 km (Refs. 16 and 17), 

whereas the AFRL craft will transition at Mach ~9+ and 68 km. 

 The AFRL air-breathing engine mode will transition from ramjet to scramjet mode beyond Mach 

5 or 6 (in contrast, at this Mach number, the BEPS ramjet will switch into rocket mode). 

 The AFRL craft will accelerate in scramjet mode to Mach ~9+ to reduce liquid propellant load 

down to ~4 kg (the BEPS craft has a larger propellant load of 4.66 kg). 

 The physics of pulsed laser scramjet propulsion is presently under investigation in Brazil 

(Refs. 11 and 12). The Mach 6 to 10 impulsive flow will be produced by a hypersonic shock 

tunnel; twin gigawatt pulsed CO2 lasers will be able to supply two sequential pulses to simulate 

high pulse repetition frequency (PRF) engine operation. 

 Both 35-cm vehicles will invoke onboard chemical kick rockets for orbit circularization, without 

the need for a laser relay mirror in LEO. 

3.1.2.16 Technical Challenges 

 Can pulsed laser scramjets generate positive thrust in the Mach 6 to 10 range? 

 AFRL’s concept has a larger internal nose volume and external surface area, indicating the 

potential for higher structural mass fractions and heat transfer loads in the hypersonic regime. 

 Can the AFRL craft be spin-stabilized, for autonomous beam riding? Will its center of mass, 

aerodynamic center, and moment of inertia ratios allow stable flight, or will the AFRL vehicle 

design tumble? 

 These technical issues should be investigated in the BEPS follow-on hardware/flight 

demonstration program for lightcraft technology. 

3.1.2.17 Description of Ground Facility 

Arguably, one viable technological approach to constructing a 12-MW repped-pulse SSDP laser can 

be described as follows. Table 3.10 lists specifications for existing SOA, high-power 1-μm SSDP lasers, a 

technology that has experienced exponential growth (Figure 3.31) since 1995. 
NASA/TM—2012-217014 145 

TABLE 3.10.—EXISTING SOLID STATE, DIODE-PUMPED LASERS

Figure 3.31.—Exponential growth of continuous working SSDP 

Lasers (Ref. 20). 

For example, in January 2009, Northrop Grumman’s seven-chain Joint High Power Solid-State Laser 

(JHPPSL) CW laser demonstrated 105 kW with a beam quality better than 3 times the diffraction limit 

(D.L.), and wall-plug efficiency of 19.3%. Adding an eighth chain to JHPPSL would increase its output 

power to 120 kW (from Ref. 21). 

Other strong contenders are Textron’s 60-kW Thin-Zag CW laser, and Lawrence Livermore National 

Laboratory’s (LLNL’s) 67-kW solid-state heat capacity laser (SSHCL), which is repetitively pulsed

(335 J pulses at 200 Hz; 500-μs pulse duration)—see Figure 3.32. SSHCL, which has no active cooling 

system, emits 335 J/pulse at 200 Hz, with a 500-μs pulse duration; diode bar pumping is 900 μs at 

872 nm. 
NASA/TM—2012-217014 146 

Figure 3.32.—LLNL’s 67-kW SSHCL (Ref. 23). 

LLNL’s 1-kW solid-state, diode-pumped Mercury High Average Power Laser (HAPL) (Figure 3.33) 

puts out 50- to 100-J pulses at 10 Hz, with a selectable duration of 5 to 20 ns. Finally, note that IPG 

Photonics manufactures extremely reliable, 10- to 50-kW fiber laser welders (Figure 3.34) that run 

continuously. Their 10-kW single-mode unit (YLS–10000–SM) has an outstanding 1.3 times D.L. beam 

quality and a wall-plug efficiency exceeding 23%. One must conclude that, yes indeed, high-average 

power, RP SSDP lasers are already here, but they are generally more expensive than their CW 

counterparts. 

Figure 3.33.—Mercury HAPL at LLNL (50- to 100-J pulses at 10 Hz) (Ref. 25). 
NASA/TM—2012-217014 147 

Figure 3.34.—Left: IPG’s 10-kW single-mode fiber laser welder. Right: Master Oscillator Power Amplifier 

(MOPA) schematic (Ref. 26). Copyright IPG Photonics Corporation (Oxford, MA) 10,000 W Single Mode 

Laser; used with permission. 

Figure 3.35 is a scatter plot of existing SSDP laser specifications relative to the desired 120-kW 

“beamline” module for our 35-cm lightcraft launcher (i.e., a short pulse version of JHPPSL and/or 

SSHCL technology, e.g., 2 kJ at 60 Hz). Demonstrated short-pulse (50- to 100-ns) technology would be 

favored to minimize risk, development time, and cost. 

Scaling up from the 10- by 10-cm output beam of SSHCL (at 335 J/pulse), our aperture must increase 

to 25 by 25 cm for emitting 2 kJ pulses. Note that the desired 60-Hz PRF is 6 times higher than HAPL, 

but only 30% of SSHCL’s 200 Hz. HAPL’s pulse duration of 5 to 20ns is 5 to 10 times too short for our 

specifications, and SSHCL’s 500-μs pulses are 500 to 1000 times too long. One is tempted to conclude 

that SSDP laser technology is now sufficiently mature to spawn a 120-kW short-pulse, 2 kJ at 60 Hz, 

beamline with 10% to 20% wall-plug efficiency, but no government or military agency has asked for it! 

Figure 3.35.—Scatter plot of SSDP laser specifications 

relative to the desired beamline goal. 

The 35-cm lightcraft requires laser PRF’s in the kilohertz range for efficient air-breathing engine 

performance: for example, 96 beamlines at 60 Hz/each = 5760 Hz, which falls in the viable range. Note 

that only with pulsed lasers can a “temporal” approach to beam-combining be realized, wherein all 

beamlines are fired sequentially like a “Gatling” gun onto a rotating mirror that efficiently shepherds 

pulses into a single 3-m telescope; with 60-Hz beamlines, the mirror must rotate at 3600 rpm. 
NASA/TM—2012-217014 148 

Figure 3.36 shows an artist concept for a 12-MW repped-pulsed SSDP launch laser system designed 

to fit within a 60- by 12- by 12-m highbay, a positive-pressure cleanroom facility (Figure 3.37), to prevent 

contamination of exposed beam-combining and adaptive optics (AO). Although it could take many forms, 

the artist concept is laid out similar to one-half of the National Ignition Facility (NIF, LLNL, Livermore, 

CA): that is, 24 beamlines per level, stacked four levels deep. 

Figure 3.36.—11.5-MW RP SSDP laser system (24 beamlines per level, 

stacked four levels deep for 96 total) (artist concept by P. Rawlings). 

Our launch laser beamline concept could resemble a recent LLNL design proposed for orbit debris 

removal (see Figure 3.38 from References 3 and 4). That 1-μm ground-based, RP SSDP laser design 

would have a 25- by 25-cm2

 output aperture emitting 8-kJ pulses of 5-ns duration at 15-Hz PRF for a 

time-average power of 120 kW. 

Figure 3.37.—WSMR facility for housing 12-MW launch laser system (artist concept by P. Rawlings). 
NASA/TM—2012-217014 149 

Figure 3.38.—Schematic for single SSDP beamline designed for orbit debris mitigation 

(Refs. 28 and 29). 

Omitting the need for 3ω harmonic conversion parts would result in a simplified and lower cost 

beamline design. LLNL’s intent was to deorbit a representative ~10-cm debris chunk of aluminum with a 

weight of 70 g and an orbital altitude of 500 km. If we assume a ~1-μm beam, a range of 800 km, a target 

area of ~100 cm2

, a transmitter diameter of 3 m, a pulse duration of 5 ns, a pulse energy of 6 kJ, and a 

momentum coupling coefficient (CM) of 6 dynes/W, then about 240 pulses would be required to change 

the orbit of that debris chunk into an elliptical one with perigee of 100 km (for subsequent atmospheric 

entry). Therefore, a ΔV of 115 m/s would be needed (Refs. 28 and 29). A 12-MW pulsed laser would take 

only 0.167 s to accomplish this task. The cost of launch laser facilities for a 12-MW system could 

approach $900M (in U.S. dollars, USD) for diode pumping. Fortunately, this nanosat launcher investment 

could potentially be shared for dual use in space debris removal, possibly with international collaboration. 

A multipurpose laser with a selectable pulse duration of 5 ns for orbit debris removal or of 50 to 100 ns

for laser LEO launch propulsion, has been judged to be technically feasible. LLNL cost estimates for its 

RP, solid-state, 1-μm laser range from $2500/J to $5000/J (Ref. 29), depending on the choices for laser 

gain medium, amplifier pump source, and thermal management method. The latter cost is representative 

of diode-pumped systems, which would be required for high wall-plug efficiency and reliable RP 

operation. 

Our 35-cm lightcraft engine will require a beamline with identical average power of 120 kW to that 

portrayed in Figure 3.44 but with 4 times smaller pulse energies, 4 times higher PRF, and a much longer 

pulse duration (e.g., 10 to 20 times longer). Hence, our desired launch laser specifications are 

 Output wavelength = ~1000 nm (maybe 1053 or 1047 nm) 

 Time-average power = 2 kJ  60 Hz  96 beamlines yields 11.52 MW 

 PRF = 60  96 = 5760 Hz with rotating mirror beam combining 

 Square output beam = 25 by 25 cm 

 Laser output pulse duration = 50 to 100 ns (for good pulsejet engine efficiency) 

 Diode pulse length = 900 μs; pumping wavelength = 808, 872, or 900 nm 

 Wall-plug efficiency = 10% or better 

Fortunately, several kilowatt-level RP SSDP prototype lasers (with short pulses) are in the hardware 

stage today, including the SILL (Ref. 24) and Mercury HAPL (Refs. 25 and 29), as mentioned earlier. 

The intent here is to highlight examples of critical technology that will enable upgrades to ~100+ kW RP

lasers in short order. Take for example, the Mercury HAPL at LLNL, which features the two novel 

elements in Figure 3.39 and Figure 3.40: (1) pulsed high-power pump diodes (in eight 100-kW, 900-nm 

arrays) and (2) high-speed He gas cooling of amplifier slablets (3 W/cm2

 average) and optical switches. 
NASA/TM—2012-217014 150 

Figure 3.39.—Left: Diode bars. Right: He gas cooling of amplifier slablets in HAPL prototype (Refs. 28 and 29).

Figure 3.40.—He gas-cooled amplifier slablets and optical switches in HAPL prototype (Ref. 30). 

Figure 3.41 shows the Amplifier Line Replaceable Unit (ALRU) concept proposed by LLNL for 

flowing-gas He cooling of four vertically stacked beamlines (e.g., the laser geometry in Figure 3.36). 

Another innovation with potential to cut diode pumping costs by more than threefold in the near future is 

high-power Vertical Cavity Surface-Emitting Laser (VCSEL) diode technology (Figure 3.42). VCSEL 

has already enabled reliable high-temperature operation for optical pumping, industrial, and medical 

applications (Ref. 31). 
NASA/TM—2012-217014 151 

Figure 3.41.—ALRU design for flowing-gas He cooling (Ref. 30).

Figure 3.42.—High-power VCSEL diode technology (Ref. 31). 

Figure 3.43 is the concept for a 96-element, 850-cm-diameter beam-combining system for the 

envisioned ~12-MW launch laser. The central mirror would rotate at 3600 rpm. The ring-mirror diameter 

would be set by the 96 adjacent turning flats positioned around the ring, each sized at 25 by 25 cm. These 

flat mirrors would be remotely adjusted with servodriven, three-axis mirror mounts to precisely target the 

on-axis rotating mirror. The 850-cm ring trusswork would be rigidly attached to a vertical 30-cm-thick, 

steel-bar-reinforced, poured 12- by 12-m concrete wall (see Figure 3.44 and Figure 3.45) that will serve 

as its optical bench. A similar 60- by 12-m reinforced-concrete optical bench would support and stabilize 

the 96 beamlines in Figure 3.36. The entire laser highbay will be temperature controlled to reduce 

thermal-induced movements of optical elements to negligible levels. Note in Figure 3.44 and Figure 3.45 

that tip and tilt errors in the 25- by 25-cm “on-axis” output beam are removed by the first pair of turning 

flats, and the second pair serves as AO to counter atmospheric distortion effects (turbulence, thermal 

blooming, etc.), before the beam enters the 3-m telescope. 
NASA/TM—2012-217014 152 

Figure 3.43.—Rotating mirror beam combining system (artist concept by P. Rawlings). Left: Planform 

view. Right: Oblique view. 

Figure 3.44.—Beam-combining system with central rotating output mirror (center) and tilt-tip correction 

mirrors and AO (right) (artist concept by P. Rawlings). 

Figure 3.45.—Rotating mirror beam-combiner unit stabilized by vertical 12- by 12-m optical 

bench/wall (artist concept by P. Rawlings). 
NASA/TM—2012-217014 153 

Note that in Figure 3.44 and Figure 3.45 the output laser beam will vacate the highbay toward the 

right, then enter a “light-tight” beam-tube (tunnel) stretching out to the launch pad. This 25- by 25-cm 

square beam will be able to be delivered selectively onto three different “targets”: (1) a large calorimeter 

to measure the total output beam power, (2) a 3-m telescope, or (3) a 1-m flat turning mirror, positioned at 

45 just below the launch pad to direct the beam vertically onto the lightcraft’s rear parabolic optic.

Handoff of the unexpanded beam from the 1-m flat mirror to the 3-m telescope takes place when the 

lightcraft reaches an altitude of 1 km or more. Redirection (i.e., “handoff”) of this high power beam will 

be done by rapidly inserting (i.e., in 1/10 s or less) track-mounted mirror flats, which will be driven by 

linear actuators like that used in the 150-kW Laser Hardened Materials Evaluation Laboratory (LHMEL) 

II CO2 laser at Wright Patterson Air Force Base, Ohio. 

Figure 3.46.—Beam delivery tube (left), 3-m telescope (center), and 

retractable launch rig with 1-m 45 turning flat (right) (artist concept by 

P. Rawlings). 

Figure 3.47—Semi-transparent view of 3-m telescope 

and dome (artist concept by P. Rawlings). 
NASA/TM—2012-217014 154 

Figure 3.48.—View inside beam delivery tunnel, showing retractable 45° flat 

mirror that diverts the unexpanded 25- by 25-cm laser beam vertically 

upward into a 3-m-diameter telescope (artist concept by P. Rawlings). 

3.1.2.18 CONOPS—Frequency of Launch and Repeat Time 

This section provides (1) the CONOPS Description for both 120- and 35-cm lightcraft (with mission 

timelines), employing a LEO laser relay satellite or onboard chemical kick rocket options, respectively, 

and (2) the Mission Description resulting from the OTIS Trajectory Simulation, probably necessitating a 

laser relay satellite for the orbit circularization burn. The fully-mature commercial laser launch site will 

be built at an altitude ≥3 km (mountain top) to minimize atmospheric attenuation, although the prototype 

infrastructure is likely to be first deployed on the High Energy Laser Systems Test Facility (HELSTF) at 

WSMR. This laser launch facility will include the following 

 Vehicle Assembly-processing Building (VAB) 

o Vehicle and payload assembly in Class 1 (ISO3) clean room environment (see Table 11) 

o Vehicle balancing (static and dynamic) and mass properties determination 

o Cleaning before placement on launcher, active queue rack, or installation in storage 

container 

o Robotic placement of vehicle on launch stabilization system 

 Vehicle transport out to the launch pad 

 Launcher pad functions 

o Umbilical cord for ground power 

o Propellant (water) loading 

o Both terminated before vehicle spin-up 

 Laser launch systems 

o Laser facility building 

o Electrical power supply including energy storage system (flywheels, SMES, etc.) 

o Telescope/transmitter with AO 
NASA/TM—2012-217014 155 

 TABLE 3.11.—ISO 14644-1 CLEANROOM STANDARDS; 

(CLASS 1 IS PROPOSED FOR THE VAB)

The Launch Commit criteria will include 

 Surface winds survey 

 Winds aloft survey 

 Clear skies over launch site and downrange trajectory 

 Launch clearance (LEO spacecraft)—from Laser Beam Clearing House.

 Local air traffic clearance 

Figure 3.49 shows an aerial perspective of the VAB with its track-mounted launch pad; the 3-m 

telescope and dome (Figure 3.46 and Figure 3.47) are omitted for clarity. This mobile, rail-guided launch 

rig will transport the lightcraft out from the protective shed, through a dust curtain (to mitigate dust 

infiltration), and to the launch point. The VAB three-view drawing (Figure 3.50) highlights the use of 

robotic handling of lightcraft (see Figure 3.51) to facilitate (1) autonomous retrieval of the correct vehicle 

from the warehouse, (2) spin balancing (Figure 3.52) and measurement of three-axis moments of inertia 

(IXX, IYY, IZZ)—e.g., POI–1000 from Space Electronics, LLC (Ref. 32), and IZZ), and (3) loading the vehicle 

onto the mobile launcher (Figure 3.53). 
NASA/TM—2012-217014 156 

 Figure 3.49.—Aerial view of laser highbay (left), beam tube (center), and 

VAB with launch pad (right) (artist concept by P. Rawlings). 

Figure 3.50.—Layout of VAB (artist concept by P. Rawlings). 

Figure 3.51.—Robotic handling of 120-cm lightcraft (318-kg wet mass) (artist concept by P. Rawlings).
NASA/TM—2012-217014 157 

Figure 3.52.—POI–1000 spin balance machine (courtesy of Space Electronics, LLC; used with permission), with 

scale 120-cm lightcraft (Ref. 32). 

Figure 3.53.—Robot loads 120-cm lightcraft onto mobile launcher rig inside VAB (artist concept by P. Rawlings ). 

Accurate spin-balancing and measurement of vehicle inertial properties is essential. Such data will be 

required input for the flight simulation and control software that will ultimately dictate the launch 

trajectory and vehicle dynamics. With its all-up wet mass of 318 kg, robotic handling of the 120-cm 

lightcraft will be essential, but at 8 kg, the nearer-term 35-cm vehicle will be clearly light enough for 
NASA/TM—2012-217014 158 

manual processing. (Dust covers on the 3-m beam expander and 45 turning flat below the launcher will 

be retracted just prior to beam delivery and liftoff.)

3.1.2.18.1 Processing and Launch 

Launch minus 10 days (or perhaps only 1 day) 

 The vehicle, including payload is assembled in a clean room. 

 The vehicle is placed on a spin balance machine and undergoes static and dynamic spin 

balancing, and mass properties (center of mass (C.M.), moments, and cross products of inertia) 

are determined. 

 Any guidance updates to the vehicle’s flight computer are done. 

 Vehicle is fueled with water propellant inside the VAB, cleaned, and robotically placed on the 

mobile launcher for roll-out to the launch pad. 

o The mobile launcher supports the lightcraft from a fixed lower stabilizing bar (bearing 

mates to rear optic tip) during transport to the pad. 

o The upper stabilizer arm, which is retractable, secures the lightcraft nose with another 

bearing (see Figure 3.54d and Figure 3.55a). 

Figure 3.54.—Mobile launcher transports 120-cm lightcraft out to launch pad at end of track (artist concept 

by P. Rawlings). 

Figure 3.55.—Aerial views (artist concept by P. Rawlings). (a) Lightcraft on launch pad. (b) Under laser boost. 
NASA/TM—2012-217014 159 

Launch minus 2 hr (or as little as 1⁄2 hr) 

 The vehicle on its mobile launcher is rolled out to the launch pad. 

 The upper balance arm has a spin bearing, ground power umbilical, and water fueling line. The 

power umbilical and fueling line are automatically mated when the spin cap is placed on the 

vehicle nose. 

 After vehicle spin-up using compressed air jets impinging on the lightcraft rim, the upper 

stabilizer bar is lifted (Figure 3.54d), the laser is fired, and the vehicle lifts off. 

 Once the vehicle clears the launcher (T = ~ 1 s), the launcher is retracted out of the beam so that 

the lower stabilizer is no longer illuminated (i.e., robbing power). 

Figure 62 is an artist concept drawing of a lightcraft launch over the White Sands National 

Monument, showing a visible H2O vapor trail and low power red tracking laser beam; the 1-μm 

propulsive laser beam would of course not be visible with the naked eye. 

Figure 3.56.—Dawn launch of lightcraft at White Sands Missile Range, New Mexico (by L. Myrabo and P. Moloney). 

3.1.2.18.2 Mission Timeline 

Figure 3.57 describes the Mission Timeline following liftoff. The COMPASS simulation assumes 

liftoff at a 30° flight path angle (FPA), from a 3-km mountain site. But in practice, the vehicle would 

launch vertically and then immediately pitch over to the desired FPA before climbing beyond 1 km above 

the pad. During this climbout, the 1-m launch mirror (flat) would simply be rotated from vertical through 

60°. At T = 50 s (altitude = 18.5 km), the 30° FPA hold will be released and the FPA will begin to 

decline. 

At 80.4 s and 35 km (optimally determined), the combined-cycle engine will switch from airbreathing mode into a laser-heated rocket mode with Isp = 950 s using onboard H20 propellant. At 418 s 

(altitude = 296 km), the lightcraft will begin an unpowered coast, since it will no longer be visible from 

the launch laser. Shortly thereafter, the lightcraft’s nosecone will split in two halves and separate for use 

as a yo-yo to despin the vehicle from 120 to 60 rpm. The vehicle will then reorient for the circularization 
NASA/TM—2012-217014 160 

burn, which will occur at T = 660 s (down range = 1330 km) using a 3.5-m laser relay mirror in 400-km 

circular orbit. The length of the laser-heated rocket burn is estimated to be between 30 and 50 s, and the 

firing will also despin the vehicle. 

Figure 3.57.—Mission timeline along trajectory for DRM 1–A. LTD, Lightcraft Technology Demonstrator. 

Note: the original COMPASS study imagined a second laser—placed downrange—for orbit 

circularization. However, a “dual-use” elliptical 3.5 by 5-m “Monocle” laser relay mirror in ~400 km 

orbit might be substantially less costly. A fully “commercial” laser launch system may require at least six 

such relay stations in LEO; that could also enable orbital debris mitigation. 

3.1.2.18.3 On-Orbit Timeline 

Figure 3.58 describes the on-orbit timeline for a microsatellite deployed in LEO. First, the vehicle is 

despun using magnetic torquers and then oriented vertically. Next a countermass is deployed on a 10-mlong tether to begin the gravity gradient stabilization mode. Finally, the solar photovoltaic arrays are 

deployed to begin activation and checkout of the payload. 

3.1.2.19 Costs—Comments 

The total cost of this advanced launch-on-demand system were derived from (1) the ground-based 

power-beaming infrastructure cost and service life plus (2) the lightcraft vehicle cost and (3) the launch 

rate. A representative payload cost estimate for 35-cm lightcraft (based on of the amortization/ payback 

costs, the laser station operating costs, etc.) is provided at the end of this section. 
NASA/TM—2012-217014 161 

Figure 3.58.—On-orbit timeline gives specifications for DRM 1–A. 

3.1.2.19.1 Laser Power-beaming Infrastructure Cost Estimate 

At the present state of RP SSDP laser technology, the cost of laser launch infrastructure for nanosat 

and picosat payloads is the main cost driver—even more so with larger microsat launchers. However, 

SSDP laser and related photonic technologies are now rapidly evolving in a climate where innovation is 

predominantly driven by DOD, commercial, and industrial applications (e.g., laser welders, optical 

pumping, medical usages, etc.), and these markets have had little to do with BEP. The cost of such 

sources can be leveraged by mass production with attractive learning curves in the long run. Note that, for 

whatever reasons, CW sources clearly dominate the marketplace today. 

The cost-scaling algorithms plotted in Figure 3.59 and Figure 3.61 were generated by J.P. Reilly of 

Northeast Science & Technology (Ref. 33), specifically for a recent feasibility study on the LTD. 

Sponsored by the AFRL and NASA (Ref. 2), the LTD Program’s technical objective was to launch a 35-

cm lightcraft into LEO using a ~10-MW ground-based laser. Figure 3.59 shows the estimated cost for 

0.01-, 0.1-, 1-, and 10-MW RP 1-μm solid-state lasers as a function of output energy in joules/pulse. Note 

the asterisks placed at 2 kJ at 5-kHz PRF (launch mission) and 8 kJ at 1.25-kHz PRF (debris-removal 

mission)—both “point designs” on the 10-MW curve. Reilly’s costing applies to flashlamp-driven Ndglass lasers with only 2% wallplug efficiency, whereas SSDP lasers demonstrate 10% to 20% today. 

Hence, for our pulsed SSDP laser system, the reduced cost of 5 to 10 times smaller electric power 

supplies will be offset by the increased cost of diode-bars (vs. the much cheaper flashlamp pumping 

technology). 

Using the Consumer Price Index (CPI) inflation index to bring Reilly’s 1995 estimates up to 2010, we 

find the 2 kJ at 5-kHz laser would cost $787M today, or $79/W. For our 11.5-MW laser (2 kHz), the price 

would be $904M for the 96 beam-lines running at 60 Hz each. Both Figure 3.59 and Figure 3.60 indicate 

that for a fixed time-average laser output power, cost falls with increasing pulse-repetition frequency (and 

lower output pulse energies). 
NASA/TM—2012-217014 162 

Figure 3.59.—RP solid-state laser costs versus pulse energy (from Refs. 2 and 33). 

Figure 3.60.—RP solid-state laser costs versus PRF (from Ref. 2 and 33). 

Table 3.12 summarizes the laser power-station infrastructure for a 35-cm nanosat and 17.5-cm picosat 

launchers. In the nanosat case, a ballpark cost figure for the 11.5-MW SSDP laser is $904M in 2010 

USD, and a 3-m telescope would add another $17 to $30M. Hence, the total infrastructure cost would be 

$920M to $934M. These cost algorithms, originally generated by J.P. Reilly of Northeast Science & 

Technology, were taken from Reference 2 and the CPI inflation index was used to adjust for 2010 USD to 

account for intervening years. Note that the laser price point for the 11.5-MW system comes to $4710/J—

nearly identical to LLNL’s $5000/J estimate for RP SSDP lasers (Ref. 29). 
NASA/TM—2012-217014 163 

TABLE 3.12.—SSDP LASER LAUNCH SYSTEM 

SPECIFICATIONS AND COST 

Parameter Nanosat 

launcher 

Picosat 

launcher 

Lightcraft diameter, m 0.35 0.175 

Launch mass—wet, kg 8.0 1.0 

Payload mass, kg 1.33 0.125 

Laser power, MW 11.5 1.0 

Laser wavelength, μm ~1.06 ~1.06 

Pulse energy, kJ 2.0 0.33 

Repetition frequency, kHz 5.76 3.5 

Telescope diameter, m 3 6 

Telescope cost,a $M 17 to 30 40 to 70 

Laser cost,a $M 904 110 

Infrastructure cost,a $M 921 to 934 150 to 180 a

2010 USD, based on cost algorithms from Reference 2. 

Figure 3.61.—Laser beam director cost versus aperture diameter (from Refs. 2 and 33). 

Since lightcraft and plasma engine technology is eminently scalable, even a picosat-class (i.e., 100 to 

1000 g) launcher could be viable. Imagine a 17.5-cm-diameter lightcraft with wet launch mass of 1 kg, 

designed for boosting 125 g of electronics payload to LEO. (Note that a 16.1-cm aluminum lightcraft 

model was flown on the 10-kW PLVTS pulsed CO2 laser at WSMR on Feb. 6, 1998, in the air-breathing 

pulsejet mode with 54.4-g launch mass.) This mission would likely require a ~1-MW pulsed SSDP laser 

(e.g., 330 J/pulse at 3.5 kHz; ~$110 M cost) with a 6-m telescope (about $40M to $73M in Figure 3.61)—

for a total infrastructure investment of $150 to $180 M. Incidentally, this is roughly the size of NASA’s 

6.5-m Next Generation Telescope (NGT), which will comprise 18 hexagonal-shaped mirror 

subassemblies. 

But what cost reductions can we anticipate in the next 3 to 5 years of high-power laser development? 

J.T. Kare (Ref. 34) suggests that the lower efficiency and more difficult cooling of pulsed slab lasers 

versus fiber lasers probably puts the cost of a high-power RP system at 3 to 5 times the cost per watt of a 

CW system. Presently, the 10- to 50-kW single-mode and multimode fiber lasers from IPG Photonics run 

$135/W and $50/W to $70/W, respectively. However, multimode fiber lasers could fall to $35/W in 1-

MW quantities within 3 years and could fall as low as $20/W with 10- to 100-MW production rates 

within 5 years. (Note that the present kilowatt-level commercial fiber-laser systems include the power 
NASA/TM—2012-217014 164 

supply, cooling, delivery fiber, control system, and packaging, and are actually overdesigned in terms of 

operating life—i.e., multiyear operation, 24/7.)

Furthermore, if LLNL’s suggestion in Reference 30 that high-power VCSEL arrays (Ref. 31) will 

become available for about one-third of the cost of edge-emitting laser diode arrays proves to be correct, 

then CW SSDP lasers could drop to $10/W for systems of 100 MW or more, and pulsed SSDP lasers 

might become as low as $35/W for systems of 10 MW or more. This would halve the price of our 

11.5-MW laser for 35-cm nanosats to ~$400M—and maybe cut a 1-MW picosat laser cost to just $35M. 

3.1.2.19.2 Lightcraft Vehicle Cost Estimate 

Figure 3.62 gives the COMPASS team’s production cost estimates for 35-cm lightcraft (DRM 1–A) 

generated by two different routines that were applied to confirm differences between the curve sets. The 

differences reveal an opportunity (with learning curves) to reduce per unit cost by going from NASA (N) 

government to commercial (C) requirements and production methods. 

Figure 3.62.—NASA (N) versus commercial (C) production costs for 35-cm lightcraft. 

However, this first-order lightcraft vehicle cost analysis (Figure 3.59) faces several unresolved issues 

that, taken altogether, may present an erroneous picture in the near term. One critical issue relates to the 

type of costing models presently in use by NASA, the Aerospace Corp, and others, which have been 

shown, repeatedly, to break down in the lower mass regime. These vehicle cost models perform well for 

systems of around 1000 kg and up. However, especially below 500 kg and critically below 200 kg, they 

break down completely and tend to show much higher than actual costs in comparison to space systems 

already deployed and in use. Such broad parametric models were applied for cost analyses of DRM 1–A 

and DRM 1–A' vehicles by the COMPASS team—talented aerospace engineers who are not necessarily 

microsatellite and launcher specialists. 

The lifetime issues present another problem. Although an ORBCOMM mission would have a 

multiyear lifetime, the launcher components would not. Similarly, the 30th unit would not have a full set 

of ground spares (e.g., no launch operator would build 60 launch systems when 30 will actually be 

launched). Operationally, there would be some optimum ratio of launches to spares when the five- to 

seven-launch milestone has past. 

In addition, the COMPASS learning curve analysis may not have been on a level playing field. From 

the first unit to the 30th, the analysis assumed a 75% reduction in costs for conventional ORBCOMM 

satellites and only a 66.67% reduction in costs for the BEPS vehicle. This alone biases the outcome 

because a similar 75% reduction in costs would put BEPS several million dollars lower than the 

traditional ORBCOMM. 

Furthermore, a meticulous systems engineering approach must be taken, rather than just a subsystemby-subsystem “piece-meal” approach, with an eye toward small payload and microsatellite engineering 

practice in order to properly assess the costs. For example, by combining electronics enclosures into 
NASA/TM—2012-217014 165 

larger boxes rather than subsystem by subsystem, a significant reduction in mass budget could be 

obtained, which in turn, would reduce overall costs. 

 These and other considerations must be taken into account to secure a true assessment of BEP 

launcher costs. To resolve this conundrum, analysts should solicit input from those engineers and 

designers presently working on microsatellite and launch systems (i.e., in major microsatellite 

companies). We can now build upon what the COMPASS team has started by enlisting help from 

personnel who already know this satellite and launch system design regime very well. Economic viability 

of revolutionary lightcraft launchers is the critical issue here, and to move forward without properly 

assessing vehicle costs seems ill advised.

3.1.2.19.3 AFRL’s Life Cycle Cost (LCC) Study for 35-cm Lightcraft Launcher 

Table 12 from Reference 19 reveals some details on the AFRL Life Cycle Cost (LLC) Summary for a 

35-cm lightcraft capable of delivering 2 kg of payload to 185-km circular orbit, using the 10-MW class 

SSDP ground-based laser. 

Judging from the $635M acquisition cost listed for the 10-MW laser, the table likely represents 

FY2000 USD since the laser would be $550M (Figure 3.58) in FY1995 USD, and the CPI from 1995 to 

2000 was ~ 1.13. Note that the 3-m telescope cost ($12M to $21M in 1995 USD) was not broken out 

separately under DDT&E/Acquisition Costs. 

AFRL’s lightcraft launcher LCC model specifies a mission length of 10 years, with a laser service life 

equaling this term. The “Average Cost per Flight” listed in Table 3.13 ($74,141), then, absorbs the 

aggregate of DDT&E/Acquisition Costs and Operations Costs over that period (i.e., $743 M) into 10,000 

launches, which might turn out to be conservative; a launch rate of 1000 per year represents only 20 per 

week or 3 per day, on average. Also stated in Reference 19: “Laser acquisition and operation costs were 

assumed to be shared with another user, and all operations costs are reduced to 1⁄2 of those values 

estimated from historical data.”

The AFRL study apparently assumes a single production run of 10,000 vehicles at the stated price of 

$3760 per unit. For a vehicle burnout mass of 3.52 kg, this comes to just $1062/kg which seems highly 

optimistic and not likely representative of the specialized 1.33- to 2-kg electronics payloads that must be 

tailored for each customer’s mission. 

TABLE 3.13.—LIFE-CYCLE COST MODEL FOR 35-CM LIGHTCRAFT 

LAUNCHED WITH 10-MW SOLID-STATE LASER (REF. 19) 

Lightcraft model cost summary Share SSDP laser to 

cut ops costs by 50%

Percentage 

Mission model length (yr) 10 

Launch rate per year 1,000 

Payload per launch (kg) 2.0 

Mission flight time (s) 221.52 

Total program cost ($M) 741.41 

DDT&E/acquisition costs ($M) 680.358 91.77 

Operations costs ($M) 61.053 8.23 

Average cost per flight ($) 74,141 

Average cost per kg (based on operations costs) ($) 3,052 

DDT&E / acquisition costs ($M) 680.358 

Laser lightcraft (LLC) development cost ($M) 18.000 2.65 

Laser lightcraft (LLC) acquisition ($M) 37.596 5.53 

10-MW ground-based laser acquisition ($M) 624.762 91.83 

Launch site facility costs (construction) ($M) 5.000 0.73 

Operations costs, annual ($M) 6.105 

Laser annual operations cost ($M) 3.750 61.42 

Laser refurbishment, annual ($M) 0.750 12.28 

Laser consumables, annual ($M) -------- ------ 

Energy cost, annual ($M) 0.101 1.65 
NASA/TM—2012-217014 166 

TABLE 3.13.—LIFE-CYCLE COST MODEL FOR 35-CM LIGHTCRAFT 

LAUNCHED WITH 10-MW SOLID-STATE LASER (REF. 19) 

Lightcraft model cost summary Share SSDP laser to 

cut ops costs by 50%

Percentage 

Launch site facility cost, annual ($M) 0.250 4.09 

USAFa systems program office (SPO) cost, annual ($M) 0.250 4.09 

NORADb coordination cost, annual ($M) 0.500 8.19 

FAAc coordination cost, annual ($M) 0.250 4.09 

Range (safety, tracking, telemetry), annual ($M) 0.255 4.17 a

U.S. Air Force. 

b

North American Aerospace Defense Command. 

c

Federal Aviation Administration. 

For the sake of argument, though, let us assume that the basic 35-cm lightcraft vehicle/engine 

technology is updated yearly in annual production runs of 1000 units at the same costs displayed in Table 

3.12 (i.e., spacecraft development cost of $18M plus acquisition cost of $37.6M = $55.6M total). Then, 

for a 1000-unit production run, each vehicle comes to $55.6K—which is somewhat more in line with the 

infrastructure’s “Average Cost per Flight” in Table 12 of $71K. Note that this is still 33 times less 

expensive than the COMPASS estimate of $1.85M each, but the correct/realistic vehicle pricing may fall 

between them. 

3.1.2.20 Launch Trajectory Simulations 

3.1.2.20.1 Compass Team’s Simulation Details for 1.2-m LTD (DRM 1–A) 

The laser launch geometry for lightcraft trajectory simulations is shown in Figure 3.63; the relay 

mirror satellite in LEO (e.g., 400-km altitude) is required under some conditions. Compass Team’s 

DRM 1–A simulation overview: 

(1) Point-mass is assumed for the vehicle 

(2) 3-DOF simulation using OTIS code, with overall objective to maximize payload mass for a 

minimum gross liftoff mass (GLOM) to a 400-km suborbital target 

(3) Launch from a 3-km altitude mountain-top laser station with a 30° fixed laser pitch angle and 30° 

vehicle flight path angle for 50 s 

(4) Climb out in air-breather/rocket mode until switch point into laser rocket mode (switch point 

optimally determined) 

 (5) Climb in laser rocket mode to 400-km suborbital target 

 (6) Coast until laser-heated circularization burn into final orbit 

Figure 3.63.—Laser launch geometry for lightcraft trajectory simulations. 
NASA/TM—2012-217014 167 

Air-breather mode assumptions and simulation details 

 Air-breathing engine performance data consists of Mach and altitude table lookup of CM (N/MW), 

EP (J), and PRF (Hz)—from References 16 and 17 (see Figure 3.64 and Figure 3.68). Note in 

Figure 3.65 that CM is modeled as a smoothly varying function of both altitude and Mach number. 

Data for 35-km altitude had to be estimated; unfortunately, the original code burned up in a fire. 

 Reference data for a 100-MW-class pulsed laser. Simulation includes a throttle (to reduce thrust 

when or if necessary) and a scale factor (range 1 to 3.5) to simulate a more powerful laser (with 

higher PRF) when required. 

 In the air-breathing mode, the LTD’s pulsed laser-ramjet engine heats surrounding air with an 

infinite Isp (no on-board mass flow is expended), and thrust is developed by momentum exchange 

with the atmosphere. 

 The vehicle drag model in Figure 3.68 is taken from References 16 and 17 with a reference 

frontal area of 0.7854 m2

—representing the 1-m-diameter vehicle forebody, which is modeled as 

a 30° half angle cone. The annular shroud’s frontal area is accounted for within the engine model.

Rocket mode assumption and simulation details 

 Thrust is a function of laser power (PRF  EP). Thrust is then found by multiplying CM  PRF  EP. 

 Laser atmospheric propagation loss was modeled after a formulation in Reference 38 (pp. 82 and 

85). 

 Reference laser power is 100 MW, scaled as necessary to be consistent in power level with the 

air-breathing engine model’s needs. 

 Laser rocket mode “burns” LH20 with an Isp of 950 s. 

 Vehicle drag model is taken from References 16 and 17 for a reference frontal area of 0.7854 m2

(see Figure 3.68). 

Note that key 3-DOF (point mass) simplifying assumptions are used for entire trajectory (i.e., optimistic 

trajectory). More rigor would require additional coordinate system in OTIS and/or full 6-DOF simulation 

to circumvent the present code’s limitations 

 No loss term for off-angle laser pointing. 

 No steering (restoring forces or moments) from off-centerline beam targeting upon the LTD 

engine. 

Figure 3.64.—Net CM versus flight Mach number for LTD airbreathing mode (Refs. 16 and 17). (”Baseline” 3-cm inlet gap, 

500-MW/cm2

 at focal ring.) 
NASA/TM—2012-217014 168 

Figure 3.65.—Three-dimensional (3D) plot of CM versus flight Mach 

number for LTD air-breathing mode (Refs. 16 and 17). (”Baseline” 

3-cm inlet gap, 500-MW/cm2

 at focal ring.) 

Figure 3.66.—EP versus flight Mach number for LTD air-breathing 

engine mode (Refs. 16 and 17). (“Baseline” 3-cm inlet gap, 

500-MW/cm2 at focal ring.) 

Figure 3.67.—PRF versus flight Mach number for LTD airbreathing engine mode (Refs. 16 and 17). (“Baseline” 3-cm 

inlet gap, 500-MW/cm2

 at focal ring.) 
NASA/TM—2012-217014 169 

Figure 3.68.—Aerodynamic drag model for 

LTD vehicle (Refs. 16 and 17). 

OTIS results summary (see Figure 3.69 to Figure 3.71, and Table 3.14 to Table 3.15) 

 The OTIS optimal trajectory for a specified flight profile to a 400-km suborbital target with final 

velocity of 5.5 km/s is shown in Figure 3.69 to Figure 3.71. 

 The closed-form rocket-mode calculation to circularize LTD orbit using a relay mirror satellite 

(or second laser source) is in the results summary in Table 3.14 and Table 3.15. 

 Key trajectory parameters and timing of events are summarized in the CONOPS section. 

 Future work should include 

o An investigation of different trajectories for optimality. The work should also include the 

implementation of a full air-breathing propulsion model: at minimum, a complete 

propulsion table, but preferably direct formulations of propulsion performance as a 

function of altitude, Mach number, and required/received laser power. 

o A full assessment of the relay mirror satellite (or second laser station) and location of the 

LTD vehicle pointing, should full inclusion of an air-breathing propulsion model fail to 

eliminate the necessity of performing the rocket-mode circularization burn with the laser 

(i.e., to capitalize on the high Isp associated with laser-heated LH20 propellant). 

o Secondary effects of off-centerine angle thrust terms and aim-pointing steering (i.e., a 

more complete beam-riding engine model). These effects would require the addition of a 

coordinate system into OTIS or a full 6-DOF representation of the vehicle (outside the 

scope of the COMPASS/OTIS study). 
NASA/TM—2012-217014 170 

Figure 3.69.—Burn phases, ranges, and angles from the Satellite Orbit 

Analysis Program (SOAP) simulation for 1.2-m LTD (DRM 1–A). 

Requirements for insertion into circular 400-km orbit: 

 

 OTIS trajectory ends in elliptical orbit; the following are the requirements necessary to circularize 

from the OTIS end point (see Table 3.14 and Table 3.15) 

o Required ΔV to circularize: 1772 m/s 

o Too far downrange (~1335 km) from laser/launch site to use launch laser directly 

 Options to circularize 

o Onboard chemical “kick” rocket gives insufficient performance to achieve orbit unless 

electronics payload (with smaller rocket) is separated from vehicle 

o Orbital-based reflector: unknown cost and phasing problems 

o A second ground-based laser site: high cost, nonoptimal thrust pointing 

 To first order, a 400-km relay mirror satellite (Figure 3.63) or second GBL station is needed 

o Angle from the station’s power beam to the vehicle centerline axis needs to be assessed 

(Figure 3.69) 

o For the laser rocket circularization “burn,” the angular offset of the vehicle centerline axis 

to the laser beam appears to be extreme for the GBL option, and this is probably best 

accommodated with a 400-km relay satellite. 

Figure 3.70.—OTIS simulated launch trajectory for 1.2-m LTD (DRM 1–A). (a) Altitude versus time. (b) Velocity 

versus time. (c) Altitude versus range from launch pad. 
NASA/TM—2012-217014 171 

Figure 3.71.—Laser power and engine thrust versus time from OTIS trajectory for 1.2-m LTD (DRM 1–A). 

TABLE 3.14.—MISSION ∆V SUMMARY FOR 

1.2-M LTD (DRM 1–A) USING OTIS 

 TABLE 3.15.—MASS, ∆V, AND PROPELLANT 

SUMMARY FOR 1.2-M LTD 

Variable Value 

Gross lift-off mass (kg) 334.99 

Nose cone mass (kg) 4.73 

Total propellant (H20) consumed (kg) 171.20 

Air-breather mode 

Mass at end of air-breather mode (kg) 334.99 

Propellant (air) consumed (kg) N/A 

ΔV (m/s) N/A 

Rocket mode (propellant = H20) 

Mass at end of rocket mode (pre circ) (kg) 192.20 

Mass at end of mode less nose cone (pre circ) (kg) 187.47 

Propellant (H20) consumed (kg) 137.86 

ΔV (m/s) 4935.30 

Propellant (H20) to circularize (kg)a 33.34 

Final estimate mass to circular 400 km orbit (kg)a 159.15 

ΔV to circularize (m/s)a 1771.90 a

Indicates that estimates are from closed form calculations. 
NASA/TM—2012-217014 172 

3.1.2.20.2 AFRL 3-DOF Simulation Results for 35-cm Lightcraft 

The laser launch geometry for AFRL’s 35-cm lightcraft launch simulation (Refs. 5 and 19) is as 

indicated in Figure 3.63, requiring no LEO relay mirror satellite for orbit circularization. The combinedcycle propulsion system encompasses two air-breathing modes (laser ramjet and laser scramjet) and a 

laser rocket mode. 

OTIS simulation overview.— 

(1) Point-mass assumed for 35-cm vehicles (Figure 3.30) with a GLOM of 8 kg (Table 3.9) 

(2) 3-DOF simulation using the OTIS code, with overall objective to transport 2-kg payload mass 

into a 185-km circular orbit 

(3) Launch vertically from mountain-top laser power station, then pitch over to 30° fixed laser pitch 

angle and 30° vehicle flight path angle 

(4) Climb out in laser ramjet mode (air-breather) until the switch point into laser scramjet mode 

(switch point optimally determined) 

(5) Climb in laser scramjet mode (air-breather) until switch point into laser rocket mode (switch point 

optimally determined) 

(6) Climb in laser rocket mode to 185-km suborbital target 

(7) Coast until chemical “kick” circularization burn into final orbit 

Air-breather mode assumptions and simulation details 

 The air-breathing engine performance model used thrust coefficients and delivered Isp (as a 

function of Mach number and altitude) from an Aerojet Techsystems fixed geometry scramjet. 

 Reference data were for a 10-MW-class pulsed laser with optimum wavelengths of 1.06 and 

1.62 μm. As with the COMPASS BEPS effort, the simulation likely includes a throttle to reduce 

thrust when or if necessary or a scale factor to model a more powerful laser when required. 

 In air-breathing mode, the laser scramjet engine exploits the surrounding air for an infinite Isp (no 

onboard mass flow expended), developing thrust by momentum exchange with the atmosphere. 

 The 35-cm vehicle drag model is based on vehicle geometries in Figure 3.30. Transonic drag 

coefficients of 0.2167 and 0.45 are suggested for 10° and 20° conical forebodies (frontal area = 

0.096 m2

)—for much higher fineness ratios than the LTD’s 30° half-angle cone with CD = 0.85. 

Such 2 to 4 times lower drag coefficients were mandatory to extend the air-breathing envelope to 

the Mach 9 to 10 regime. 

Rocket mode assumption and simulation details 

 Thrust is a function of available laser power versus range collected by 30-cm-diameter lightcraft 

rear parabolic optic. Laser rocket performance (CM, Isp) was defined as a function of laser pulsewidth (tP) and pulse energy (EP). 

 Laser atmospheric propagation losses were modeled with code, include thermal blooming, 

atmospheric turbulence, and extinction (molecular and particulate absorption and scattering; 

telescope jitter was not mentioned in Refs. 5 and 19) 

 Reference laser power is 10 MW, scaled versus range to account for atmospheric propagation 

losses. 

 Laser rocket mode can “burn” 4 kg of propellant. Apparently Isp varies from 100 s at beginning of 

firing 1000 s at end. 

 Vehicle drag model is likely based on 10° and 20° half-angle cones. 

 
NASA/TM—2012-217014 173 

The 3-DOF (point mass) simplifying assumptions are likely applied for the entire trajectory (i.e., 

optimistic trajectory). The OTIS-based trajectory simulation describes all laser beam and lightcraft 

kinematics and associated combined-cycle air-breathing/rocket propulsion performance. As with the 

COMPASS simulations, more precision would require an additional coordinate system in OTIS and/or a 

full 6-DOF simulation to eliminate the present code’s limitations 

 No loss term for off-angle laser pointing (i.e., incidence relative to centerline axis of rear optic) 

 No steering (restoring forces or moments) from off-centerline beam targeting on the lightcraft engine 

 No beam-riding engine/vehicle performance modeling 

OTIS results summary 

 The OTIS optimal trajectory for the specified flight profile to the 185-km suborbital target uses 

laser rocket mode burnout at 221 s into the final transfer orbit of 185 by 92.5 km, coast to 

185 km, and a final injection burn with a chemical rocket (Figure 3.72 (a)). 

 Near final trajectory plots in Figure 3.72 (b) to (e) are for an 8-kg GLOM vehicle with 3 kg 

expended of 4-kg total propellant allotment (see Table 3.9) for laser rocket burn. Logically, 1 kg 

more boosts altitude from 82 to 92.5 km. 

 The OTIS trajectory ends in a 92.5- by 185-km elliptical orbit. This used a closed-form rocketmode calculation to circularize the orbit at 1953 s with a very low thrust bipropellant rocket. 

o Chemical “kick” rocket system mass = 0.46 kg; this exploits microelectromechanical 

systems (MEMS) technology 

o Motor mass = 0.408 kg; bipropellant tank = 0.01 kg (19.2% mass penalty) 

o Monomethyl hydrazine (MMH)/mixed oxides of nitrogen (MON) propellant mass = 

0.042 kg to circularize orbit (Table 3.9) 

 Key trajectory parameters and timing of events are 

o Laser ramjet to scramjet transition near Mach 5 and 45 km at 105 s 

o Laser scramjet mode ends at 2.6 km/s (Mach 9), 70 km, and 155 s 

o Laser rocket burnout at >7.5 km/s, 92.5 km, 221 s 

o Coast to 185-km altitude from 221 to 1953 s 

o Very low thrust circularization burn of 25 s duration begins at 1953 s 

 
NASA/TM—2012-217014 174 

Figure 3.72.—Altitude versus time for simulated trajectories of 35-cm LCC (Ref. 5). 

According to Reference 5, this chemical “kick” rocket would exploit “emerging MEMS technologies 

(being developed under the national Nanotechnology Initiative), to secure a 7 to 12 times mass advantage 

over existing small off-the-shelf chemical rockets”— for example, the S-400 control motor, a 400-N 

(thrust) bipropellant rocket designed for MMH/MON and built by MBB, now called Astrium (an EADS 

company). It had an Isp performance of 3000 N-s/kg and mass of ~2.8 kg—and hence the MEMS mass 

reduction objective of 2.8/0.408 = 6.9. 
NASA/TM—2012-217014 175 

Further Observations 

 The OTIS trajectory optimizer indicates that it is best to launch a lightcraft at a very near vertical 

attitude and then rotate its flight path angle downward as it accelerates upward. Lifting off at 30° 

off-vertical (as suggested in Ref. 16 for the LTD), may be less than optimum. 

 Lightcraft drag and light capture ratio (see Figure 3.73) are critical issues affecting mission 

performance. The LTD’s higher drag coefficient reduces laser ramjet air-breathing performance 

in comparison with AFRL’s 35 cm, lower CD vehicles with 10° to 20° conical forebodies. 

However, the 1.2-m LTD’s larger laser energy capture efficiency gives it greater laser rocket 

performance at high altitude and in space. 

 Although the LTD reaches Mach 7 at 35 km with the laser ramjet mode, it is not likely to reach 

Mach 8 to 10 (in air-breathing mode) because of its fairly low forebody fineness ratio. However, 

the LTD’s heavier payload configuration and higher laser power capture efficiency (in terms of 

vehicle mass/light capture area) reaching Mach 8 to 10 might not be so important to the mission. 

 AFRL’s 35-cm craft accelerates in laser scramjet mode to Mach ~9+ to reduce the liquid 

propellant load (for laser rocket mode) down to ~ 4 kg. With its laser ramjet engine, the BEPS 

35=cm craft requires a larger propellant load of 4.66 kg. 

 Candidate liquid propellants: Water or ammonia are likely more viable liquid propellants for the 

laser rocket mode than liquefied natural gas (LNG), whose low density could make tank volume 

and structural mass penalties excessive. 

Figure 3.73.—Captured 1.62-μm laser power versus slant range from for 35-cm 

lightcraft (Ref. 5). 

3.1.2.20.3 DOF Simulation Results for 35-cm and 120-cm Lightcraft 

The study objectives were to (1) attempt to replicate the LEO launch trajectory performance for the 

120-cm (focal-ring diameter) LTD vehicle (120-kg dry mass, plus 140 kg LN2 and 2-kg He = 262 kg 

GLOM; described in Ref. 16), (2) simulate flights to LEO of a BEPS 35-cm, 8-kg GLOM lightcraft (see 

Figure 3.30, and see Table 3.8 for mass breakdown), and (3) investigate beam-riding potential (Figure 

3.15 and Figure 3.16) of both lightcraft to orbital velocity. To enable such simulations, it was first 

necessary to upgrade the existing 7-DOF code (Refs. 13 to 15) with a suitable combined-cycle, airbreathing/rocket pulsejet engine model by incorporating the pulsed laser ramjet performance given in 

Figure 3.64 to Figure 3.67, along with a pulsed laser rocket mode with specified performance and 

propellant mass inventory, using the vehicle drag model shown in Figure 3.68. No previous 7-DOF code 

runs had modeled the pulsed laser ramjet mode. 
NASA/TM—2012-217014 176 

3.1.2.20.3.1 7-DOF Study Tasks 

The 7-DOF study tasks are (1) upgrade the existing 7-DOF operational code for simulating 120- and 

35-cm lightcraft with combined-cycle engines in trajectories to orbit: for example, transitioning from airbreathing to rocket mode at about Mach 7 and 35 km (see Ref. 16); (2) perform flight dynamics and 

launch simulations with refined 7-DOF code, simulating trajectories to ~8 km/s; identify optimum pitch 

angle for launch and boost to orbital velocity; and (3) plot out associated output variables for comparison 

with the 120-cm BEPS vehicle trajectory (COMPASS team) and AFRL’s 35-cm vehicle trajectory—both 

of which employed OTIS. 

3.1.2.20.3.2 7-DOF Code Overview 

The 7-DOF code utilized for this endeavor has been developed, validated, and calibrated against 

existing experimental lightcraft flight trajectory data to improve the understanding of beam-riding flight 

dynamics in order to accurately predict the flight performance of future lightcraft engine/vehicle designs 

(Refs. 13 to 15, 39, and 40). Both the axial thrust characteristics and beam-riding characteristics of 

theType-200 lightcraft engine have been modeled and verified against extensive laboratory experimental 

data. The flight dynamics system model was originally a 6-DOF code, and was later extended to include a 

seventh DOF for a despun forebody—a feature not invoked in the present study. 

3.1.2.20.3.3 7-DOF Simulation Results 

Using the 7-DOF code, the flight dynamics, stability, and control of 35- and 120-cm diameter 

lightcraft were successfully modeled in autonomous beam riding under laser boost to extreme altitudes 

and orbital velocities. The interim results are presented in Figure 3.74 to Figure 3.80. The 35-cm craft 

used the same CM and PRF schedules as the 120-cm craft (see Figure 3.64 and Figure 3.67), but the EP

schedule in Figure 3.66 was reduced by 11.75 times (EP scales with engine cross-sectional area) to 

account for the smaller scale craft. As shown in Figure 3.74 to Figure 3.80, this goal was achieved by 

both 35- and 120-cm vehicles. Note that the Mach number “bumps” appearing in Figure 3.74 at 80 and 

130 s (35 and 120 cm, respectively) are due to the atmospheric model, not actual changes in velocity—as 

verified in Figure 3.76. 

Figure 3.74.—Mach number versus time for 7-DOF simulated trajectories. 
NASA/TM—2012-217014 177 

Figure 3.75.—Altitude versus time for 7-DOF simulated trajectories. 

Figure 3.76.—Velocity versus time for 7-DOF simulated trajectories. 

Figure 3.77.—Beam power versus time for 7-DOF simulated trajectories for 120-cm LTD craft. 
NASA/TM—2012-217014 178 

Figure 3.78.—Beam power versus time for 7-DOF simulated trajectories for 35-cm craft. 

Figure 3.79.—Mach number versus altitude for 7-DOF simulated trajectories. 

Figure 3.80.—Zenith angle versus time for 7-DOF simulated trajectories. 

The 35-cm vehicle exhibited much larger accelerations and a shorter “burn” time in laser rocket mode 

than its 120-cm counterpart. The resultant laser rocket mode ΔVs calculated from Equation (1) are nearly 

identical for both craft, but the propellant burn rates m and total propellant masses from Equations (2) 

and (3), respectively, are different. The same propellant ablation rate (kg/J) is assumed to be constant for 

both engines (throughout the burn), whereas the beam power and, consequently, the burn rate (kg/s) 

scales roughly with the square of the engine diameter. The propellant mass scales with the tank volume or 

the cube of the craft diameter. Because of these scaling effects, the smaller craft’s laser rocket propellant 

60

62

64

66

68

70

72

0 50 100 150

Zenith Angle (degrees)

Time (s)

35 cm

120 cm
NASA/TM—2012-217014 179 

“burns out” more quickly than that of the larger craft while providing roughly the same ΔV. Hence, the 

35-cm craft sees larger accelerations over the burn duration, and the higher “burnout” velocity achieved 

benefits from smaller losses to atmospheric drag along the trajectory. 

1

0

0 ln

m

m v  Isp g (1) 

m  Beam Power (W ) Ablation Rate (kg/J) (2) 

 

 propellant density  volume mPropellant (3) 

3.1.2.20.3.4 Summary and Future Work 

A simple model for the air-breathing laser ramjet engine has been added to the 7-DOF code, giving 

CM, PRF, and EP as functions of altitude and Mach number. The resulting lightcraft “burnout” velocities 

and attainable altitudes are well within the targeted range, indicating that the launch concept is feasible 

under autonomous beam-riding flight. Further work is required to optimize such trajectories to reduce the 

required laser power and increase flight performance under boost. 

3.1.2.20.4 Feasibility Assessment 

The present feasibility study began with an “end-game” vision of a laser launcher for 130-kg payloads 

(DRM 1–A)—an ideal final result (IFR) requiring a monolithic, 100-MW-class infrastructure. Although 

this IFR was clearly useful in structuring technology roadmaps for lightcraft R&D and for assessing 

feasibility, the optimal laser launch concept likely resides at a smaller scale, and perhaps that decision 

must await the completion of subsequent trade-off studies. Hence, the strategy for a progressive

technology roadmap has been presented herein, calling for a smaller 12-MW laser launcher for 1.33-kg 

CubeSat-class payloads (DRM 1–A’)—an initially more attainable system and closest “target of 

opportunity,” that is, 1- to 10-kg nanosats (and maybe even 0.1- to 1-kg picosats).

Baring a rapid evolution of free-electron lasers from the present 14-kW level into long-planned 100-

-kW and 1-MW devices (Refs. 35 to 37), the RP SSDP laser technology is our source of choice for 

lightcraft propulsion of nanosats and picosats to LEO in the near future—simply because it enables the 

highest engine performance and most energetic spacecraft for the DRM 1–A mission. However, because 

of the sizeable cost of GBL infrastructure, the feasibility/affordability of this launch system may hinge on 

its potential for shared use with another crucial role such as orbital debris mitigation, anti-satellite 

(ASAT), anti-missile, ballistic missile defense (BMD), or other use—a conclusion shared with AFRL’s 

nanosat launcher study (Refs. 5 and 19). 

Moreover, note that the SSDP laser and other competitive source technologies are rapidly evolving, 

and innovation appears to be driven both by commercial and DOD weapons applications (e.g., laser 

welders, fusion plasma heaters, etc.). In addition, the cost of such photon sources can leverage mass 

production with attractive learning curves in the long term. 

To ensure the feasibility of the spacecraft/lightcraft design, we recommend further work on the 

following fronts: 

 Propulsion/Vehicle Dynamics: (1) Heat leaks, injector options, propellant selection, off-nominal 

performance (thrust vector/Isp/thrust as a function of off-pointing, angle of attack, and/or offcenter beam displacement), and (2) atmospheric compensation for the laser power delivery 

(determine if beacon is required); see Figure 3.81 for key issues affecting engine performance and 

efficiency 

 Thermal: (1) Modeling of reflector and injector-ring cooling (heating from laser, plasma—

conductive, convection, and radiation—in all modes), in addition to pressure loss in cooling 
NASA/TM—2012-217014 180 

channels (pressure design); (2) estimate thermal damage of mirror reflectivity due to thermal 

cycling; and (3) final velocity profile (with Mach < 0.6 “pop-up” hold removed) was not 

evaluated for its impact on thermal and structures (max-Q considerations) 

 Mission: Improved fidelity of trajectory design to include (1) beam pointing loss terms and offcenter beam pointing for steering (current trajectory is optimistic) and (2) examination of optimal 

trajectories including a study of a vertical launch followed by an immediate pitch-over to remove 

the current 30° launch conditions and to simplify the launch apparatus 

 GN&C: (1) Final launch profile change to 30° elevation—pitch requirements not yet evaluated 

for this trajectory, (2) differential injection/thrust for pitching, (3) error margin analysis on the 

laser pointing/position on the spacecraft body, (4) required spin-rate for adequate stability, and 

(5)thermal control in various propulsion modes

 Microelectronics: A critical issue for 35-cm nanosat, and for the smaller picosat launchers is the 

microminiaturization of electronics packages (both size and mass).

Figure 3.81.—Key issues affecting the feasibility of a lightcraft vehicle and a combined-cycle engine. 

We must also maintain focus on (1) increasing design margins (e.g., materials and costs), (2) reliance 

on multiple technological choices for each element, and (3) flexible integration (e.g., Responsive Space 

goals). Again, the real game changer here is a very simple and CHEAP vehicle that relies on mass 

production to offset operational costs (associated with the laser launch facility), in addition to inherent 

fail-safe features, the use of environmentally friendly H2O propellant, and so forth. 

Succinctly stated, engineering research and development (R&D) is needed across the two main BEP 

subsystems: (1) the lightcraft engine/vehicle and (2) ground-based power-beaming infrastructure. These 

subsystems are intimately and inseparably linked, and as mentioned earlier, cost may likely be the major 

driver in the selection of ground-based power-beaming infrastructure (RP laser versus CW laser versus 

millimeter wave). 

3.1.2.21 Ground and Flight Demonstrations 

In the meantime, until such multimegawatt SSDP laser launch infrastructures become available and 

operational, the “go-forward” plan for advancing embryonic BEP technology in the next half-decade 

should include the following agenda, leading swiftly to high-visibility flight demonstrations to extreme 

altitudes: 

 Develop a variety of lightcraft plasma engine/vehicle designs, at various scales, by exploiting 

existing photon sources (RP-lasers, CW-lasers, CW, and RP millimeter wave) and diverse engine 

modes (e.g., pulsejet, ramjet, scramjet, rocket, combined-cycle, etc.) 
NASA/TM—2012-217014 181 

 Systematically assess, through joint experimental/theoretical/numerical studies, the relative merits 

and performance of CW versus RP plasma engines (laser vs. millimeter wave), prior to downselecting a specific power-beaming infrastructure for initial deployment 

 Continue to expand the BEP database by validating plasma engine models and vehicle flight 

simulation results; then define the feasible and optimal range of beam source functional 

requirements, margins, and tolerances 

 Closely monitor the development of high-power CW and RP photon sources (laser and millimeter 

wave), and influence/encourage trends that enhance BEP capability 

Furthermore, in a broader context, we must 

 Create a progressive technology roadmap with a focus on managing risks, both technological and 

financial 

 Design and prototype a modular launch infrastructure that can grow over time, starting with a 

small-scale system, then gradually expanding it as technology evolves and uncertainties are 

resolved 

 Maintain focus on increased design margins (e.g., materials and costs), reliance on multiple 

technological choices for each element, and flexible integration (Responsive Space goals) 

 Identify technology gaps and leverage the flexibility of the private sector—both small and large 

enterprises—specialized in addressing exacting technological challenges, within efficient and 

cost-effective working cultures 

Figure 3.64 to Figure 3.67 lay out four candidate programs (each progressively more ambitious) for 

high-visibility flight demonstrations to extreme altitudes using existing photon sources with powers 

ranging from 10 kW upwards into the multimegawatt range. (As mentioned earlier, an attractive feature 

of lightcraft technology is its scalability which, in fact, helps to mitigate or manage such technological 

and financial uncertainties.)

The least expensive, Program 1 (see Figure 3.82), would exploit the Pulsed Laser Vulnerability Test 

System (PLVTS) 10-kW CO2 pulsed laser presently operational at HELSTF (WSMR) to accomplish the 

following objectives: (1) demonstrate the “handoff maneuver” in flight, by passing the 10- by 10-cm 

PLVTS unexpanded beam off to a larger telescope (e.g., 25 to 30 cm or more), enabling higher flights 

beyond the current 71-m record; (2) establish the 1-km world altitude record with this new “hand-off” 

beam expander; and (3) exhibit controlled flight during “pitch-over” of the telescope wherein the vehicle 

must ride a nonvertical, slewing beam and accelerate horizontally for cross-range capability. The 

estimated cost is ~$2.5M for 18 months. 

Figure 3.83 gives the objectives for Program 2, which would utilize the Army’s JHPSSL 105-kW CW 

laser linked to the 70-cm diameter Tactical High Energy Laser (THEL) beam expander, already recoated 

for ~1.06 m. The laser facility is being set up at HELSTF now that its new building has been completed. 

The $6M, 2- to 3-yr program would require the development of a new CW plasma engine with 

autonomous beam-riding abilities, and an attitude control system for lightcraft flights to altitudes of 10 to 

30 km. The program objectives include (1) accelerating the vehicle to transonic speeds while climbing 

vertically to an altitude of 10 km, (2) demonstrating a location communications link to the JHPSSL 

facility, (3) exhibiting successful attitude control, and (4) proving safe laser beam control, pointing, and 

tracking capability. 
NASA/TM—2012-217014 182 

Figure 3.82.—Program 1—Laser launch demonstrations to 1 km at HELSTF 

using 10-kW PLVTS CO2 laser. 

Figure 3.83.—Program 2—Suborbital launch demonstrations to 10 to 30 km at 

HELSTF using the 105-kW JHPSSL. 

Program 3 in Figure 3.84 seeks several ambitious objectives using the multimegawatt chemical 

oxygen iodine laser (COIL) installed in the Airborne Laser Test Bed (ALTB), and boosting lightcraft with 

a new CW plasma engine (mentioned above). Several air-launched lightcraft vehicles would be fired in 

succession (two to six per sortie) by a fighter jet cruising at an altitude of 35 to 40 km in formation with 

the ALTB mothership. Each air-to-air missile would eject its lightcraft at a range suitable for direct 

engagement by COIL’s 1.5-m telescope—aided by ultrabright light-emitting diodes (LEDs) and/or 

corner-cubes installed around the lightcraft’s shroud. 
NASA/TM—2012-217014 183 

Figure 3.84.—Program 3—Suborbital launch demonstrations using ALTB’s multimegawatt COIL. 

The technical objectives would be to (1) exhibit beamed energy transfer through the upper 

atmosphere to a beam-riding lightcraft, (2) accelerate the vehicle through the transonic and into the 

supersonic regime, (3) validate safe ALTB beam control for 15-to 30-s lightcraft flights, and (4) prove 

ALTB pointing and tracking capability with beam-riding lightcraft. The estimated cost of $12 M (2- to 

3-yr duration) includes ~$2.5 M for ALTB operations. 

Program 4 in Figure 3.85, clearly the most expensive at $50M to $150M (3 to 4 yr), takes place on 

orbit, utilizing the ISS as the platform for a 1-kW laser based on the Mercury HAPL prototype at LLNL. 

This prototype RP, 1-μm laser emits 55 to 100 J at a pulse repetition frequency of up to 100 Hz, and must 

be space-hardened for a high-vibration launch environment in order to fulfill its role as a multipurpose 

laser power-beaming station. The “wall-plug” efficiency should be in the 10% to 20% range. 

The objectives would be to (1) establish a multipurpose 1-kW space-power-beaming infrastructure, 

(2) develop a laser-propelled orbit transfer vehicle around a simple laser solid ablative rocket engine, 

(3) launch repeatedly from a magazine loaded with numerous vehicles (6 to 12 or more), (4) explore 

fractionated spacecraft architectures, with maneuvering using BEP, and (5) impart a 100-m/s velocity 

change to deorbit vehicles. Lightcraft vehicles would be 10 to 12 cm in diameter with a mass of 25 to 

50 g, with each spin stabilized prior to launch. Furthermore, the 1-kW power-beaming station might also 

demonstrate recharging of satellites (batteries) by illuminating their PV arrays, and enable experiments on 

long-range laser communications. 
NASA/TM—2012-217014 184 

Figure 3.85.—Program 4—On-orbit launch demonstrations from the ISS using 

a space-rated version of the HAPL 1- kW laser. 

As a final note, a high-visibility BEP Launch Demonstration program funded at the $3M to $10M 

level by a government or private entity would be exceptionally effective in promoting the development of 

low-cost space access. The program could be structured for record-breaking flights to extreme altitudes 

(e.g., 30 to 100 km) in a timeframe of 3 to 5 yr or less, using existing 10- to 100-kW photon sources. For 

example, organized as a $10M BEP X Prize competition, it could easily attract a dozen or more entries 

from multinational participants, spawning a cornucopia of novel approaches. Once these demonstration 

flights have proven successful, then attracting private capital for the commercialization of BEP 

technology should be at hand. At this point, even raising $75M to $180M for launching nanosats into 

LEO should be feasible. 

3.1.2.22 Responsive Space and Fractionated Architectures

It is common knowledge that today’s space launch industry is quite unresponsive to changes and 

uncertainties. Making progress in the traditional launch infrastructure—based on chemical rocket 

boosters—has relied on small optimizations and innovative business strategies, but it has finally “hit the 

wall.” A change in Vision is badly needed, and responsive space architecture appears to be the key. Its 

merit has already been demonstrated in industry segments. However, flexible space systems must rely on 

fractionated architectures, and the extreme rigidity of traditional chemical-fueled launch systems 

prevents progress in this direction. 

Fundamental change will not be achieved unless the same level of flexibility is extended to the launch 

apparatus—which is exactly what BEP brings to the table. In addition, because of BEP’s ultra-energetic 

performance, it is uniquely able to embrace the subtleties of (1) increased design margins, (2) reliance on 

multiple technological choices for each element, and (3) flexible integration—that is, the Responsive 

Space goals. 

Exploitation of BEP technology can achieve this vision by offering 

 A multipurpose laser power-beaming infrastructure on the ground (where BEP is only one of its 

functions) 

 A scalable laser launch system that responds rapidly to all sources of uncertainty, rather than 

adhering to rigid missions 

 A game-changer technology that will create new markets and new opportunities unforeseen today 
NASA/TM—2012-217014 185 

Note that dramatically reduced launch costs would in itself qualify as game-changing, but if the BEP 

launcher design could also support “responsive and fractionated space architectures,” the fusion may be 

unstoppable. Indeed, this “scaled versatility” will soon enable lightcraft to open a new paradigm for space 

access. 

3.1.2.23 References for DRM 1–A 

1. L.N. Myrabo, “Brief History of the Lightcraft Technology Demonstrator (LTD) Project,” in 1st

International Symposium on Beamed Energy Propulsion, Huntsville, Alabama, November 2002,

edited by A.V. Pakhomov, AIP Conference Proceedings 664, American Institute of Physics, 

Melville, NY, 2003, pp. 49-59. 

2. F.B. Mead, Jr., “Part 1 – The Lightcraft Technology Demonstration Program,” Final Report, 

AFRL-RZ-ED-TR-2007-0078, Nov. 2007. 

3. L.N. Myrabo, “World Record Flights of Beam–Riding Rocket Lightcraft: Demonstration of 

‘Disruptive’ Propulsion Technology,” AIAA Paper 2001-3798, 37th AIAA/ASME/SAE/ASEE 

Joint Propulsion Conference, Salt Lake City, Utah, July 2001. 

4. http://www.sncorp.com/news/press/snc_2008_snc_orbcomm.shtml 

5. D. Froning, L. McKinney, F. Mead, W. Larson, and A. Pike, “Some Results of a Study of the 

effectiveness and cost of a laser-powered “lightcraft” vehicle system,” in Proceedings of HighPower Laser Ablation (HPLA) 2004, April 2004, pp. 25-30; see also: D. Froning, L. McKinney, 

F. Mead, W. Larson, and A. Pike, “Some Results of a Study of the effectiveness and cost of a 

laser-powered ‘lightcraft’ vehicle system,” Technical Report AFRL-PR-ED-TP-2004-094, 

AFRL/PRS, Edwards AFB, CA, 26 March 2004; see also: D. Froning, L. McKinney, F. Mead, 

W. Larson, and A. Pike, “Some Results of a Study of the effectiveness and cost of a laserpowered “lightcraft” vehicle system,” in 2nd International Symposium on Beamed Energy 

Propulsion, edited by K. Komurasaki, AIP Conference Proceedings 702 American Institute of 

Physics, Melville, NY, 2003, pp. 242-250. 

6. D.A. Kenoyer, I.I. Salvador, and L.N. Myrabo, “Axial Impulse Generation of lightcraft Engines 

with ~ 1 μs Pulsed TEA CO2 Laser,” in 7th International Symposium on Beamed Energy 

Propulsion, edited by H.-A. Eckel and S. Scharring, AIP Conference Proceedings, American 

Institute of Physics, Melville, NY, 2011, pp. 98 - 108. 

7. D.A. Kenoyer, I.I. Salvador, and L.N. Myrabo, “Beam-Riding Behavior of Lightcraft Engines 

with ~ 1 μs Pulsed TEA CO2 Laser,” in 7th International Symposium on Beamed Energy 

Propulsion, edited by H.-A. Eckel and S. Scharring, AIP Conference Proceedings, American 

Institute of Physics, Melville, NY, 2011, pp. 109 - 121. 

8. D.A. Kenoyer, I.I. Salvador, S.N. Notaro, and L.N. Myrabo, “Flow Visualization of ThrustVectoring Lightcraft Engines with ~1μs Pulsed TEA CO2 Laser,” in 7th International Symposium 

on Beamed Energy Propulsion, edited by H.-A. Eckel and S. Scharring, AIP Conference 

Proceedings, American Institute of Physics, Melville, NY, 2011, pp. 122 - 130. 

9. L.N. Myrabo, P.W. Lyons, R.A. Jones, S. Liu, and C. Manka, “Airbreathing Laser Propulsion 

Experiments with 1 μm Terawatt Pharos III Laser: Parts 1 & 2,” in 7th International Symposium 

on Beamed Energy Propulsion, edited by H.-A. Eckel and S. Scharring, AIP Conference 

Proceedings, American Institute of Physics, Melville, NY, 2011, pp. 203-239. 

10. D.A. Kenoyer, L.N. Myrabo, “Subsonic Aerodynamics of Spinning and Non-Spinning Type 200 

Lightcraft: Progress Report,” in 6th International Symposium on Beamed Energy Propulsion, 

edited by C. Phipps, K. Komurasaki, and J. Sinko, AIP Proceedings 1230, American Institute of 

Physics, Melville, NY, 2010, pp. 30-40. 
NASA/TM—2012-217014 186 

11. I.I. Salvador, L.N. Myrabo, M.A.S. Minucci, A.C. de Oliveira, P.G.P. Toro, J.B. Channes Jr. and 

I.S. Rego, “2-D Airbreathing Lightcraft Engine Experiments in Quiescent Conditions,” in 7th

International Symposium on Beamed Energy Propulsion, edited by H.-A. Eckel and S. Scharring, 

AIP Conference Proceedings, American Institute of Physics, Melville, NY, 2011, pp. 174 – 189. 

See also: I.S. Salvador, “Static and Hypersonic Experimental Analysis of Impulse Generation in 

AirBreathing Laser Thermal Propulsion,” Ph.D. Thesis, Rensselaer Polytechnic Institute, Troy, NY, 2010. 

12. I.I. Salvador, L.N. Myrabo, M.A.S. Minucci, A.C. de Oliveira, P.G.P. Toro, J.B. Channes Jr. and 

I.S. Rego, “2-D Airbreathing Lightcraft Engine Experiments in Hypersonic Flow Conditions,” in 

7th International Symposium on Beamed Energy Propulsion, edited by H.-A. Eckel and S. 

Scharring, AIP Conference Proceedings, American Institute of Physics, Melville, NY, 2011, 

pp. 190 – 202. See also: I.S. Salvador, “Static and Hypersonic Experimental Analysis of Impulse 

Generation in AirBreathing Laser Thermal Propulsion,” Ph.D. Thesis, Rensselaer Polytechnic Institute, 

Troy, NY, 2010. 

13. C.G. Ballard, K.S. Anderson, and L.N. Myrabo, "Flight Dynamics and Simulation of Laser 

Propelled Lightcraft" ASME Journal of Computational and Nonlinear Dynamics, Vol. 4, Issue 4, 

Oct. 2009. 

14. D.A. Kenoyer, K.S. Anderson, L.N. Myrabo, “Trajectory Simulations For Laser-Launched 

Microsatellites Using a 7-DOF Flight Dynamics Model,” Paper 86664, Proc. of 

IDETC/MSNDC7, 2009. 

15. D.A. Kenoyer, K.S. Anderson, and L.N. Myrabo, “Calibration and Validation of a 6-DOF Laser 

Propelled Lightcraft Flight Dynamics Model Upon Experimental Data, in 5th International 

Symposium on Beamed Energy Propulsion, edited by A.V. Pakhomov, AIP Proceedings 997, 

American Institute of Physics, Melville, NY, 2008, pp. 325-337.

16. L.N. Myrabo (ed.), et.al., “Lightcraft Technology Demonstrator,” Final Technical Report, SDIO 

Laser Propulsion Program, prepared under contract to Lawrence Livermore National Laboratory, 

No. 2073803, 1989. 

17. Richard, J.C., and Myrabo, L.N., “Analysis of Laser-Generated Impulse in an Airbreathing 

Pulsed Detonation Engine - Parts 1 & 2,” in 3rd International Symposium on Beamed Energy 

Propulsion, edited by A.V. Pakhomov and L.N. Myrabo, AIP Conference Proceedings 766, 

American Institute of Physics, Melville, NY, 2005, pp. 265-278; see also: J.C. Richard, “Analysis 

of Pulsed Laser-Generated Impulse in an Advanced Airbreathing Thruster,” PhD Thesis, 

Rensselaer Polytechnic Institute, Troy, NY, 1989. 

18. T. Langener, L.N. Myrabo, and Z. Rusak, “Inlet Aerodynamics and Ram Drag of Laser Propelled 

Lightcraft Vehicles,” in 6th International Symposium on Beamed Energy Propulsion, edited by C. 

Phipps, K. Komurasaki, and J. Sinko, AIP Proceedings 1230, American Institute of Physics, 

Melville, NY, 2010, pp. 41-60. 

19. E.W. Davis and F.B. Mead, Jr., “Review of Laser Lightcraft Propulsion System,” in 5th

International Symposium on Beamed Energy Propulsion, edited by A.V. Pakhomov, AIP 

Proceedings 997, American Institute of Physics, Melville, NY, 2008, pp. 283-294. 

20. A. Krepinevich, T. Ehrhard, and B. Watts, “Solid-State Laser Weapon Systems, Bridging the 

Gap—or Bridge Too Far?,” Center for Strategic & Budgetary Assessments (CSBA), May 20, 

2009. 

21. S. Weinberger, “Defense Technology International: Solid-State Lasers; 100 Kilowatts or Bust; 

Rivals Vie To Develop Next Generation of Lasers,” Aviation Week & Space Technology, 22 May 

2006. 

22. http://www.textrondefense.com/news/2010/02_18_10.html 

23. B. Yamamoto, et.al., “Evolution of a solid state laser,” SPIE 6551: 55205 (2007); see also: 

https://lasers.llnl.gov/programs/psa/directed_energy/sshcl.php 

24. http://www.sflorg.com/technews/tn101206_02.html 

25. A. Bayramian, et.al., “The Mercury project: A high average power, gas-cooled laser for inertial 

fusion energy development,” Fus. Sci and Tech., 52, pp. 383-387 (2007). 
NASA/TM—2012-217014 187 

26. www.ipgphotonics.com; see also: http://www.optoiq.com/index/photonics-technologiesapplications/lfw-display/lfw-article-display/371319/articles/laser-focus-world/volume-45/issue12/features/photonic-frontiers-fiber-lasers-fiber-lasers-ramp-up-the-power.html 

27. www.trassa.org. 

28. A.M. Rubenchik, C.P.J. Barty, R.J. Beach, A.C. Erlandson, J.A. Caird, “Laser Systems for Orbit 

Debris Removal,” LLNL PROC-423323, Lawrence Livermore National Laboratory, February 8, 

2010; also given at the High Power Laser Ablation ’11 conf. (HPLA 2010), April 18-22, 2010. 

29. C.P.J. Barty, J.A. Caird, A.E. Erlandson, R. Beach, A.M. Rubenchik, “High Energy Laser for 

Space Debris Removal,” LLNL-TR-419114, Lawrence Livermore National Laboratory, October 

31, 2009. 

30. J.A. Caird, et.al., “Nd:glass laser design for laser ICF fission energy (LIFE),” Fusion Science and 

Technology, 56, 607 (2009); also, see ppt with same title presented by J.A. Caird to the American 

Nuclear Society 18th Meeting, 29 Sept. – 2 Oct. 2008. 

31. http://www.optoiq.com/index/photonics-technologies-applications/lfw-display/lfw-articledisplay/311588/articles/laser-focus-world/volume-43/issue-11/features/vcsels-powerful-vcselarrays-beat-the-heat.html 

32. http://www.space-electronics.com/Literature/POISeriesMetric.pdf; http://www.spaceelectronics.com/Literature/POI%20Specs.PDF 

33. J.P. Reilly, “NASA Design Study on Laser Propulsion Concepts – Requirements and Options,” 

Northeast Science and Technology, Inc., East Sandwich, MA, for Anteon Corporation, Dayton, 

OH, under Contract No. F33615-97-D-5403, dated 31 Dec. 1999, 62 pages. 

34. J.T. Kare (personal communication), Kare Technical Consulting, Seattle, WA. 

35. http://www.engadget.com/2011/02/21/us-navys-free-electron-laser-breaks-another-record-takesaim-a/; and http://www.geek.com/articles/geek-cetera/navys-record-breaking-laser-burnsthrough-20-feet-of-steel-per-second-20110221/ 

36. S.E. Sampayan, et.al., “Performance Characteristics of an Induction Linac Magnetic Pulse 

Compression Modular at Multi-Kilohertz Pulse Repetition Frequencies,” Lawrence Livermore 

National Laboratory, UCRL-JC-105442 Preprint, prepared for 1991 IEEE Particle Accelerator 

Conference 6-9 May 1991. 

37. R. Whitney, D. Douglas, and G.Neil, “Airborne megawatt class free-electron laser for defense 

and security,” Jefferson Lab, 12000 Jefferson Avenue, Newport News, VA USA 23606; 

whitney@jlab.org; phone 757.269.7536; www.jlab.org 

38. D.P. Resendes, S. Mota, G. Sorasio, J.T. Mendonca, B. Sanders, and J. Encarnacão, “Laser 

Propulsion for ESA Missions: Ground to Orbit Launch,” Final Report, ESA CR 

17048/03/NL/PA, Dec. 2004, pp. 82 and 85; see also: D.P. Resendes, S. Mota, J.T. Mendonca, B. 

Sanders, J. Encarnacão, J. Gonzalez del Amo, and L.N. Myrabo, “Laser Propulsion for ESA 

Missions: Ground to Orbit Launch Project Overview - Part I,” in 4th International Symposium on 

Beamed Energy Propulsion, edited by K Komurasaki, T. Yabe, S. Uchida, and A. Sasoh, AIP 

Conference Proceedings, American Institute of Physics, Melville, NY, 2006, pp. 576-587. 

39. D.A. Kenoyer, "Validation and Calibration of a 6-DOF Laser Propelled Lightcraft Flight 

Dynamics Model vs. Experimental Data," M.S. Thesis, Rensselaer Polytechnic Institute, Troy, 

N.Y., Dec. 2007. 

40. D.A. Kenoyer, “Combine Experimental and Numerical Investigations into Laser Propulsion 

Engineering Physics,” Ph.D. dissertation, Rensselaer Polytechnic Institute, Troy, NY. June, 2010. 
NASA/TM—2012-217014 188 

3.1.3 Laser Optical Cost Analysis 

National Aeronautics and Space Administration

Beamed Energy Propulsion Facility 

ROM Cost Estimate 

March 28, 2011

National Aeronautics and Space Administration

General BEPS Launch Facility Assumptions

• All cost are rough order of magnitude (ROM) estimates in 

FY11M$.

• Estimates assume all facilities are built new on government 

land, no existing infrastructure is used. Available infrastructure 

will reduce costs.

• Costs do not include remediation of potential significant 

environmental impact issues that may be identified.

• Range estimates are presented to highlight the uncertainty 

associated with the lack of definition and supporting cost 

information available at this time. 

• Range estimates for the launch site also includes the 

uncertainty associated with the selected location and the cost 

impact of materials that may be required to protect against 

environmental conditions.
NASA/TM—2012-217014 189 

National Aeronautics and Space Administration

General BEPS Launch Facility Assumptions

• Launch site ROM estimates are based on a NASA KSC launch 

site estimating approach. Microwave and laser sites use the 

same approach for ‘like function’ facilities which are scaled for 

expected differences in size or function.

• All sites include site design and testing and activation costs.

• Launch facility design is estimated at 12% of estimated total site 

costs less testing and activation.

• Launch facility testing and activation is estimated at 30% of the 

capital facilities cost.

• Microwave and laser site design is estimated at 8% of estimated 

total site costs less testing and activation.

• Microwave and laser site testing and activation is estimated at 

15% of the capital facilities cost.

National Aeronautics and Space Administration

Operations Cost Assumptions

• All costs are estimated for one year of operation in FY11$. 

• Yearly operations cost assume 360 launches per year 

(one/day). Additional study is required to determine the feasible 

number of launches per year from a single launch pad. The 

number of launches per pad is affected by the CONOPS, pad 

turnaround time, launch facility maintenance, weather, and 

launch window requirements.

• Total estimated facility operations costs are assumed to be 

allocated to the beamed launch operations. Facility sharing can 

reduce costs. 

• Facility equipment and maintenance is estimated at 5% of the 

total value of capital assets per year resulting in an estimated 

70-80% of total operations costs.
NASA/TM—2012-217014 190 

National Aeronautics and Space Administration

Laser Optical Launch Site Assumptions

• Launch Site facility requirements

– General – site work, roads

– Support Facilities/Launch Pad – see details below

– Site Infrastructure – water, fire, sewer, power, communications, HVAC, environmental 

control system

– Ground support equipment

• Key cost estimate assumptions

– New facility on government land in secluded area due to danger associated with 

beamed energy. Launch rate 1-4 vehicles a day, if CONOPS allows.

– Launch pad includes all related hardware/services (scaled from microwave estimate). 

Rocket will be small, approximately 1.5 meters in length, and less than 500 kg total wet 

mass with payload.

– Command center includes all related services.

– Payload integration facility (assuming an integrated payload) can process/integrate up 

to four payloads per day (scaled from microwave estimate).

– Water storage facility sized for ~6,000 kg with no on-site processing.

– Rocket storage facility size for 8 rockets (scaled from microwave estimate).

– Admin facility sized for launch site staffing needs (estimated at 100 people).

– Security facilities/equipment includes fencing, cameras, etc.

National Aeronautics and Space Administration

Laser Optical Launch Site Assumptions

• Launch site boost laser facility requirements

– Laser/Optics – diode lasers(140 MW), optics (Qty.1150, 120 kWe modules)

– Power Storage/Distribution – energy storage, power supply, power 

distribution

• Key cost estimate assumptions

– Diode Laser – Projected multi-MW price $2/W (ref: J. Kare) used as low 

(ML +50%, high +100%)

– Optics – Projected $200K/module (ref: J. Kare) use as low (ML +25%, high 

+50%)

– Energy Storage - $50K/kWh estimated as most likely based on GRC 

experience with wind tunnels (low -40%, high +20%)
NASA/TM—2012-217014 191 

National Aeronautics and Space Administration

Laser Optical Main Laser Site Assumptions

• Main Laser Site facility requirements

– General – site work, roads

– Support Facilities – admin facility (50 people), command center (75% of 

launch site center), guard gate facility

– Site Infrastructure – water, fire, sewer, power, communications, HVAC, 

environmental control system

– Ground support equipment

– Laser/Optics – diode lasers(350 MW), optics (Qty. 3000, 120 kWe modules)

– Power Storage/Distribution – energy storage, power supply, power 

distribution

• Key cost estimate assumptions

– Diode Laser – Projected multi-MW price $2/W (ref: J. Kare) used as low 

(ML +50%, high +100%)

– Optics – Projected $200K/module (ref: J. Kare) use as low (ML +25%, high 

+50%)

– Energy Storage - $50K/kWh estimated as most likely based on GRC 

experience with wind tunnels (low -40%, high +20%)

National Aeronautics and Space Administration

Laser Optical Launch Facility Development ROM

Low ML High Mean Est Low ML High Mean Est Low ML High Mean Est

Project Total ROM ($M) $1,116.8 $1,713.4 $2,741.1 $1,857.1 $2,103.3 $3,119.0 $4,460.7 $3,227.6 $3,220.1 $4,832.4 $7,201.7 $5,084.7

Design $95.1 $145.9 $233.3 $158.1 $137.1 $203.2 $290.7 $210.3 $232.2 $349.1 $524.0 $368.4

Testing & Activation $228.9 $351.7 $563.7 $381.5 $253.0 $375.2 $536.0 $388.1 $481.9 $726.9 $1,099.8 $769.5

Laser Optical Facilities ROM Subtotal $792.7 $1,215.8 $1,944.1 $1,317.5 $1,713.3 $2,540.6 $3,633.9 $2,629.2 $2,506.0 $3,756.3 $5,578.0 $3,946.8

WBS Item

Site Work $11.2 $15.4 $18.8 $15.1 $8.4 $11.6 $14.1 $11.3 $19.6 $27.0 $32.9 $26.5

Roads $18.4 $27.9 $46.2 $30.8 $18.4 $27.9 $46.2 $30.8 $36.9 $55.8 $92.4 $61.7

Support Facilities $22.0 $45.7 $95.7 $54.5 $7.1 $14.9 $36.1 $19.4 $29.2 $60.6 $131.7 $73.8

Concrete Administration Facilty $7.4 $10.8 $28.5 $3.7 $5.4 $14.2 $11.0 $16.1 $42.7

8 Rocket Storage Facility $2.8 $3.8 $6.1 $0.0 $0.0 $0.0 $2.8 $3.8 $6.1

Concrete Command Center $3.3 $10.4 $25.3 $2.5 $7.8 $19.0 $5.8 $18.2 $44.4

Payload Integration Facilty $7.6 $19.1 $32.9 $0.0 $0.0 $0.0 $7.6 $19.1 $32.9

Guard Gate Facility $1.0 $1.8 $2.8 $1.0 $1.8 $2.8 $1.9 $3.5 $5.6

Pad Structure $6.6 $9.8 $12.2 $9.5 $0.0 $0.0 $0.0 $0.0 $6.6 $9.8 $12.2 $9.5

Pad Structure $2.2 $3.1 $4.2 $3.2 $0.0 $0.0 $0.0 $0.0 $2.2 $3.1 $4.2 $3.2

Emergency Egress System $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0

Flame Trench $3.4 $4.0 $4.5 $3.9 $0.0 $0.0 $0.0 $0.0 $3.4 $4.0 $4.5 $3.9

Ligtning Protection Towers $1.1 $2.7 $3.5 $2.4 $0.0 $0.0 $0.0 $0.0 $1.1 $2.7 $3.5 $2.4

Infrastructure $99.2 $174.5 $435.5 $236.4 $91.3 $142.7 $260.5 $164.8 $190.5 $317.2 $696.0 $401.3

GSE $13.5 $23.4 $53.3 $30.0 $8.4 $14.6 $33.3 $18.8 $21.9 $38.0 $86.5 $48.8

Laser/Optics Equipment $510.0 $707.5 $905.0 $707.5 $1,300.0 $1,800.0 $2,300.0 $1,800.0 $1,810.0 $2,507.5 $3,205.0 $2,507.5

Power Storage/Processing Facilities $111.8 $211.5 $377.5 $233.6 $279.6 $528.9 $943.8 $584.1 $391.4 $740.4 $1,321.3 $817.7

Launch Site Complex ‐ ROM Subtotal ($M) Main Laser Site ‐  ROM Subtotal ($M) Laser Optical Facilities ‐ Total ROM ($M)

Laser Optical Launch Facilities ROM (FY11$M)

(Excludes FTE/WYE personnel and O&Mcosts)
NASA/TM—2012-217014 192 

National Aeronautics and Space Administration

Laser Optical Launch Facility Operations ROM

Low ML High Mean Est

Laser Optical Operations Total ROM ($M/Year) $152.2 $228.1 $331.7 $237.3

Facility/Equipment Maintenance $122.5 $183.7 $272.6 $192.9

Operations Personnel $22.5 $30.0 $37.5 $30.0

Propellant/Energy/Consumables $7.2 $14.4 $21.6 $14.4

Other $0.0 $0.0 $0.0 $0.0

Launch Site Complex ‐ ROM Subtotal ($M/Year) $56.8 $85.8 $129.8 $90.8

Facility/Equipment Maintenance $38.2 $58.6 $94.0 $63.6

Operations Personnel $15.0 $20.0 $25.0 $20.0

Propellant/Energy/Consumables $3.6 $7.2 $10.8 $7.2

Other $0.0 $0.0 $0.0 $0.0

Main Laser Site ‐  ROM Subtotal ($M/Year) $95.4 $142.3 $202.0 $146.6

Facility/Equipment Maintenance $84.3 $125.1 $178.7 $129.4

Operations Personnel $7.5 $10.0 $12.5 $10.0

Propellant/Energy/Consumables $3.6 $7.2 $10.8 $7.2

Other $0.0 $0.0 $0.0 $0.0

Laser Optical Operations ‐ Total ROM ($M/Yr)

Laser Optical Launch Facilities Operations ROM (FY11$M/Year)

(Does not include vehicle/payload)

National Aeronautics and Space Administration

Beamed Energy Propulsion Study

Laser Optical Vehicle Cost Estimate

as determined by the COMPASS Team

March 2011
NASA/TM—2012-217014 193 

National Aeronautics and Space Administration

COMPASS Cost Assumptions - Lightcraft

• DRAFT Cost Estimate based on COMPASS design

• All costs are in FY11$M

• This estimate assumes the following:

– Proto-flight development approach

– No ground spares included

– Model assumes TRL Level 6

• This estimate does not include any cost for technology development

– Represents the most likely estimate based on cost-risk simulation results

– Flight heritage is assumed as Off-The-Shelf (OTS) for most components

• Excluding new or advanced technology items

– Parametric estimate based on mostly mass based CERs from historical cost data

– Launch vehicle systems integration wraps

– Software not included

• Does not include:

– Contractor Fee (10-12% on top of the contractor estimates)

– Any insight/oversight costs (for NASA-managed estimates)

– Reserves (can be as high as 40-50%)

– Ground System Cost (ie. Laser)

– Launch Services Costs (ie. Special launch approval process)

– Technology costs for components lower than TRL-6 

National Aeronautics and Space Administration

Lightcraft

Preliminary COMPASS Cost ROM

(Represents estimated Prime Contractor cost) 

Costs in FY11$M

WBS Description

DDT&E 

Total 

BY $M

Flight 

HW 

BY $M

Total 

BY $M

06.1.1 Payload 0.0 0.0 0.0

06.1.2 Attitude Determination and Control 4.4 1.6 6.0

06.1.3 Avionics 3.8 1.7 5.5

06.1.4 Communications and Tracking 1.3 0.5 1.8

06.1.5 Electrical Power Subsystem 2.2 1.2 3.4

06.1.6 Thermal Control (Non-Propellant) 5.4 1.0 6.4

06.1.7 Propulsion (Hardware) 4.8 1.3 6.1

06.1.8 Propellant (Chemical) 0.0 0.0 0.0

06.1.9 Struc tures and Mechanisms 7.9 2.9 10.8

Subtotal 29.7 10.3 40.1

Systems Integration 33.1 4.4 37.5

Spacecraft Total 62.8 14.8 77.6
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3.1.5 Laser Thermal Consultant Report—Design Reference Mission 1–B 

Design Reference Mission 1–B (DRM 1–B) is based on the laser-heated heat exchanger (HX) vehicle 

originally proposed by Kare (Ref. 1) and since elaborated on (Refs. 2 and 3). The beam source is an array 

of “beam modules”— small, independent lasers and associated beam directors, not mutually coherent—

that individually track the vehicle (Ref. 4). This approach allows the laser array to be developed at the 

single-module scale for minimal cost. Current off-the-shelf industrial laser technology is sufficient to 

build a launch system, but more advanced lasers can be incorporated as they become available. 

The nominal reference mission was to launch 100-kg payloads (independent satellites or consumables 

such as tanks of propellant) to circular orbit (nominally 34 inclined, 300- to 400-km altitude), with a 

capacity of 1000 launches per year. The overall concept of operations (CONOPS) is shown in Figure 

3.86. 

3.1.5.1 Description of Vehicle 

3.1.5.1.1 Overview 

The general propulsion concept is shown in Figure 3.86. An inert liquid propellant is pressure-fed or 

pumped through an HX, which absorbs the laser energy and transfers it to the propellant, vaporizing it 

and heating it to high temperature. The hot propellant is exhausted from a more-or-less conventional 

nozzle. 

Figure 3.86.—Typical launch concept of operations with L1 (boost) laser array. Markers: (0) Launch. (1) Main array 

“sees” vehicle and begins tracking. (2) Main array turns on; L1 array at maximum full-power range. (3) Main array 

at full power; L1 array turns off. (4) Vehicle pitches down to circularize orbit. (5) Orbit insertion; main laser turns off. 

Copyright Kare Technical Consulting; used with permission. 
NASA/TM—2012-217014 229 

Figure 3.87.—HX thruster concept; LH2, liquid hydrogen. Copyright Kare Technical Consulting; used with permission. 

Relative to other beamed-energy propulsion (BEP) approaches, notably pulsed-detonation thrusters, 

the HX thruster has several important advantages: 

High efficiency.—Exhaust kinetic power/laser power up to ~90% 

Wide beam-acceptance angle. 

Thrust direction independent of receiver orientation.—Receiver can be oriented as needed to 

satisfy drag or beam-angle constraints 

Large receiver area.—Allows comparatively high beam divergence and small apertures 

No onboard focusing optics or windows. 

Easy to design/develop.—Remainder of propulsion system is essentially conventional 

Works with any laser.—Ground or suborbital flight testing can use any available laser or low-cost 

“bare” diode arrays 

Testable with non-laser heat sources.—Ground testing can use resistive or combustion heating 

However, the HX thruster does have two major (and related) disadvantages: It has an upper limit on 

the propellant temperature, set by the material properties of the HX, and it requires hydrogen propellant to 

achieve high (>500 s) specific impulse Isp. 

One feature of the HX thruster not shown in Figure 3.87 is that the propellant composition, and thus 

the exhaust velocity and Isp can be varied in flight. In particular, a denser propellant such as water or 

liquid nitrogen can be used instead of, or in addition to, liquid hydrogen for the early, low-velocity part of 

the launch trajectory. Increasing the mean molecular weight of the propellant (and the mass flow), while 

keeping the HX temperature constant, yields increased thrust for a given laser power. 

3.1.5.1.2 HX 

3.1.5.1.2.5 DRM vehicle HX baseline 

The baseline HX design approach is a laminar flow microchannel HX, based on the original work of 

Tuckerman (Ref. 5). This type of HX has heat transfer characteristics that are nearly independent of the 

gas flow velocity, and thus, it can be designed for low flow velocity with many parallel channels, as 

shown in Figure 3.88. This gives a very low pressure drop in the HX, which is enabling, or at least highly 

Propellant (LH 2) Tank

Pump

(optional)

Flat Plate

Heat Exchanger Nozzle(s)

Laser beam
NASA/TM—2012-217014 230 

desirable, for a pressure-fed system but is less critical if a propellant pump is used. It also keeps the gas 

flow far from sonic velocity. 

Figure 3.88.—Conventional HX versus laminar-flow microchannel HX. Copyright Kare Technical Consulting; 

used with permission. 

The microchannel HX inherently uses a short channel length (typically 0.1 to 0.5 m) which allows a 

large HX to be fabricated and tested in smaller segments (panels). A large HX also can be shaped to 

match the expected beam profile (e.g., approximating a circle, with panels near the edge having longer 

channels to match the lower flux). 

The microchannel HX has very low surface-to-gas thermal resistance because of the large thermal 

contact area (7 times the frontal area) and, thus, a very low surface-to-gas temperature drop. Analytical 

design calculations for the microchannel HX are given in (Ref. 6). Disadvantages of the microchannel HX 

include very small channel dimensions (nominally 0.2 by 1.2 mm) and relatively complex manifolds to 

interconnect panels. 

The microchannel HX was originally conceived of as being made of metal (nickel, Ni), with panels 

fabricated and joined by electroforming (Ref. 7), but Ni melts at 1726 K, limiting the maximum exit 

temperature to <1500 K. Higher temperature metals are very dense, and generally expensive and difficult 

to fabricate, but may be used for test articles or for supports, fasteners, coatings, and so forth. 

The current baseline HX material is silicon carbide (SiC), possibly with carbon fiber reinforcement 

(C-SiC). The SiC structure temperature required for 2000 K gas temperature is 2200 to 2300 K, well 

below the SiC vacuum melting point of ~2850 K. 

 
NASA/TM—2012-217014 231 

The current HX parameters, as used in the DRM–B design, are given in Table 3.16. 

TABLE 3.16.—HX PARAMETERS FOR CURRENT (DRM–1B) DESIGN 

Size 4 by 6 m 

(approx.); 

24 m2

Elongated to reduce drag and improve stiffness; foreshortened to ~1:1 

aspect ratio when beam is 50o off normal 

Type Laminar flow 

microchannel 

See discussion 

Material SiC or C/SiC Low-temperature manifolds/structure may be metallic or composite; 

high- temperature structure TBD 

Design flux 9 MW/m2 Absorbed flux, area average. Incident laser flux may be up to ~10% 

higher to allow for reflection and reradiation 

Inlet pressure 5 MPa 

Pressure drop <1 MPa 

Design flow 7 kg/s H2

Exit gas temperature 2000 K 

Maximum surface 

temperature 

2300 K Mean surface-gas T needs to be modeled; a simple approximation gives 

~300 K for 10-MW/m2

 absorbed flux. 

Panel size TBD Cross-channel width of SiC panels may be limited by fabrication process 

or thermal stresses 

Channel length 40 cm COMPASSa design used 15 cm because of a calculation error related to 

inlet flow velocity

Channel cross section 200 by 1200 m

Channel spacing 400 m 

Areal density 2.9 kg/m2 HX only, for a 100-m front and 200-m back facesheet (average) a

Collaborative Modeling for Parametric Assessment of Space Systems. 

In the current conceptual design, the HX would be fabricated as SiC channel sections that would be 

attached to high-temperature outlet manifolds to form panels, nominally 80 cm wide by the full HX 

length. These panels would be assembled onto the HX support structure and attached to the lowtemperature inlet manifolds. The individual panel outlets would be attached to a common ceramic or 

metallic outlet duct leading to the nozzle assembly. 

3.1.5.1.2.6 HX Options 

The major alternative would be a conventional turbulent-flow HX, consisting of parallel thin-walled 

tubes, typically ~5 mm in diameter and 2 to 3 m long, as used in the microwave-powered vehicle design 

(Ref. 8). A turbulent-flow HX has higher pressure drop, requiring more pump power, and higher surface 

temperature, but lower manifold complexity. Further study, and probably fabrication and testing of 

prototypes of both kinds of HX, are needed to make a final selection. 

The baseline HX is a single stage. This design requires that the hydrogen propellant be above critical 

pressure (so there will be no two-phase flow) and may require some tapering of the channels to maintain 

stable flow and prevent channel-to-channel flow instabilities. In addition, it is potentially fragile and may 

be expensive to fabricate. 

An alternative that should be considered is a two- or even three-stage HX, using the microchannel 

SiC HX only for the high-temperature stage. The low-temperature stage(s) could afford a much larger 

surface-to-gas thermal resistance, allowing a broader design space, including turbulent-flow designs. The 

low-temperature stage(s) could also use a wider range of materials, including metals, for lower cost or 

easier fabrication. Operating the SiC stage over a more limited temperature range would also decrease 

thermal stresses and potential flow instabilities. 

The current leading material alternative to SiC is pyrolytic graphite, which has an even higher 

limiting temperature and better thermal conductivity, but it may require greater protection from air at high 

temperatures, or be eroded by hot hydrogen. Other ceramics and metals capable of higher temperatures 

than with SiC seem to be either too expensive or too easily oxidized (or both) to be good candidates, but 

further research is needed. 
NASA/TM—2012-217014 232 

3.1.5.1.2.7 HX Fabrication 

The best way to fabricate SiC microchannel HXs is not known, but possibilities include 

 Fabricating graphite structures and converting them to SiC by reaction with liquid Si 

 Molding “green” SiC blanks of fins and one face sheet, and firing, with either prefiring or 

postfiring bonding of the other face sheet and manifolds 

 Preassembling fins and using vapor deposition to create one or both thin face sheets (similar 

to electrodeposition of metal face sheets.) 

Note that components other than the HX itself, especially the hot-gas manifolds, do not require high 

thermal conductivity, and can be comparatively thick (limited by mass budgets). In general, they will also 

be at least slightly cooler than the gas temperature because of conductive and/or radiative cooling. 

3.1.5.1.2.8 Laser HX Issues 

The HX is the lowest Technology Readiness Level (TRL) aspect of the HX laser launch system. 

Some of the issues that have been raised for laser HXs are discussed in Table 3.17. 

TABLE 3.17.—ISSUES FOR LASER HX 

Issue Description Potential solution/mitigation 

Absorption/ 

reflection loss 

Some laser light will reflect from 

the HX surface, wasting energy. 

(Reflected light is also a potential 

hazard; see Section 3.3.5.3—Issues 

and Resolutions) 

Reflection is assumed to be 

negligible in the baseline system 

design, since good-quality black 

surfaces are ~98% absorbing. 

 Add carbon to the SiC surface layer to make it more absorbing. 

 Deposit mechanically thin (<10 m) but optically thick absorbing 

layers on the HX surface. Note that , to the author’s knowledge, a 

highly absorbing surface that remains intact up to the full SiC 

temperature limits has not been demonstrated. 

 Structure the surface to be optically absorbing. Nanostructured 

black surfaces generated by ultrashort laser pulse irradiation have 

been demonstrated on refractory metals (Ref. 9). Narrow V-grooves 

parallel to the vehicle axis can act as cavity light traps. 

 Compensate for reflection loss by increasing laser power. 

Reradiation The hot HX surfaces will radiate 

energy, which is both a loss and a 

potential thermal problem for other 

vehicle components. 

At 2200 K, the black-body 

radiation flux is ~1.3 MW/m2

. For 

a linear temperature gradient (i.e., 

equal area for each temperature 

level) the mean radiated flux is 1/5 

of the peak, or 0.26 MW/m2

. 

This is 6.5 MW for a 25-m2

 HX, 

which is an acceptable loss for 

100- to 200-MW incident power. 

However, at lower power levels 

(near end-of-trajectory), the 

temperature gradient may be 

nonlinear, because the reradiation 

at the extreme hot end will 

approach the incident flux. 

 Make HX surfaces not intended to absorb the laser shiny or white, 

reducing emittance and thus radiated flux by 10 to 100 times. 

 Cover surfaces not facing the laser with opaque insulation to lower 

the radiated flux. 

 Minimize the surface-to-gas temperature drop by optimizing the HX 

design, for example, increasing the rib cross section and reducing 

the channel width for the highest-temperature part of the HX. Since 

emitted flux is proportional to T4

, even a slight reduction in surface 

temperature will significantly reduce reradiation. 

 Raise the laser flux on the HX, and reduce the HX area. Reduce the 

laser spot size or use reflective surfaces to keep the collecting area 

constant. 

 Use a surface or coating that absorbs efficiently (and thus has high 

emissivity) at the laser wavelength but has low emissivity at other 

wavelengths. The peak reradiation wavelength is unfortunately in 

the near-infrared (~1.5 m), close to the laser wavelength, so 

achieving a large emissivity ratio may be impractical. (In this 

respect, the COMPASS design is optimistic because it assumes an 

emissivity of 0.3 for the face of the HX, instead of ~1.0. However, 

the COMPASS estimate of reradiation is pessimistic by an even 

larger factor because it assumes that the entire HX surface is at the 

maximum temperature). 
NASA/TM—2012-217014 233 

Issue Description Potential solution/mitigation 

Convective loss In the atmosphere, some heat will 

be lost to the air passing over the 

HX. 

The COMPASS group estimated 

convective loss at 20.7 MW, 

although their assumptions 

concerning surface temperature 

distribution and flow conditions are 

not given. 

 Raise the boost laser maximum power. Since this is a surface loss, 

no change in HX design is required; the boost laser power can be 

adjusted to maintain constant absorbed power. 

 Maintain laminar flow over the HX (as over an airplane wing) to 

maximize boundary layer thickness 

 Insulate or enclose non-laser-absorbing surfaces (back of HX, 

manifolds) 

 Configure the HX surface (e.g. with grooves or flow fences) to 

create stagnant or low-velocity flow over at least part of the 

absorbing face. 

Oxidation In air, SiC begins to oxidize well 

below its vacuum melting 

temperature. The usual working 

temperature given for SiC in air is 

1600 to 1650 C (1873 to 1923 K), 

but at least one manufacturer offers 

SiC with a maximum service 

temperature in air of 1900 C 

(Ref. 10). However in a highvelocity airstream, the thin 

facesheet may fail quickly. 

 Flow or transpire an inert or reducing gas (nitrogen or hydrogen) 

over the HX face—possibly by deliberately making the HX 

facesheet slightly porous—to maintain a low oxygen concentration 

in the boundary layer. 

 Run the HX at lower temperature while in the atmosphere, by 

increasing the propellant flow or reducing the laser power, and 

either operating at lower Isp or generating more heat by combustion. 

 Use available multilayer coatings (HfC/SiC) to withstand oxidation 

for short periods at ~2400 K (Ref. 11). 

Stress and 

creep 

SiC does not melt until 2850 K, 

but it loses mechanical strength 

and begins to creep (plastically 

deform slowly under load) at 

lower temperatures. Both creep 

and stress failure can be caused by 

stress concentrations, hot spots, or 

weak points such as bond lines. 

The microchannel HX design is 

comparatively resistant to creep or 

pressure failure despite its thin 

structure because the unsupported 

span of material in the facesheets 

is very short (~200 m). 

 Make the HX channels with rounded corners or oval cross sections 

to reduce stress concentration, if the fabrication technique allows. 

 Add reinforcing material such as carbon fibers to the SiC matrix. 

 Reduce the number of hot seams by operating at higher gas pressure 

and flow velocity, allowing longer channels and fewer panel joints. 

 Eliminate hot spots on the HX face by applying reflective material 

(or removing absorber) wherever cooling is inadequate. This can be 

done in fabrication for known hotspots (such as along manifolds) or 

in testing (e.g., if there are blocked channels). 

Thermal stress Thermal gradients on the order of 

1000 K/cm normal to the surface 

and 100 K/cm along the channels 

will create large thermal stresses, 

potentially fracturing SiC. Surfacenormal gradients will stress the 

front faceplate in compression and 

the back face in tension. Alongchannel gradients will stress the 

cold edge of each HX panel in 

tension, in the direction 

perpendicular to the gas flow. 

 Put gaps in the ribs, or use a “diamond post” structure instead of 

continuous ribs, to relieve stress accumulation. 

 Incorporate expansion features into the back facesheet, or allow the 

entire HX to bow slightly to relieve front-to-back gradient stress. 

 Add tensile strength members or inclusions (e.g., carbon fiber) 

along the cold edge of the HX. 

 Limit the maximum width (in the perpendicular-to-flow direction) 

of individual heat-exchanger panels, and incorporate compliance in 

the input manifold-to-panel junctions, or “expansion joints” in the 

input manifolds. 
NASA/TM—2012-217014 234 

Issue Description Potential solution/mitigation 

Thermal shock 

and over/ 

underheating 

Sudden variations in propellant 

flow or laser intensity may cause 

large changes in exit temperature 

on millisecond timescales. 

Variations in the laser beam profile 

(either fast, because of atmospheric 

effects, or slow, as a function of 

range) will change the relative flux 

on different parts of the HX. 

 Limit gross thermal shock by controlling the ramping of the laser 

power and propellant flow at startup and shutdown. 

 Tailor the beam shape at shorter ranges (using pointing offsets 

between individual beam modules) to approximately match the 

beam profile at long range. 

 Sense the HX exit temperature at multiple points and adjust 

propellant flow via local valves. 

 Use a piston pump or pressure-fed system, rather than a turbopump, 

to allow fast changes in total propellant flow. 

 Use a two-stage HX that mixes the propellant between stages to 

smooth out heating variations from the first stage. 

Overall, the microchannel SiC HX will be challenging to develop and represents the largest technical 

risk for the HX launch system. Conversely, there are many options for optimizing the design that have not 

yet been explored even minimally: varying channel cross sections, surface coatings, curved or textured 

surfaces, integrated structures and manifolds, and other options. 

3.1.5.1.3 DRM Vehicle Design 

The starting point was the conceptual design shown in Figure 3.89. (See COMPASS report for the 

DRM–1B end-point design). 

 
NASA/TM—2012-217014 235 

Figure 3.89.—Conceptual design of 100-MW HX launch vehicle. 

Copyright Kare Technical Consulting; used with permission. 
NASA/TM—2012-217014 236 

The overall dimensions of this design were driven by the desire to enable full adaptive optics (AO) 

correction for atmospheric turbulence using only a beacon on the vehicle, which forces the vehicle to be 

~30 m long so the beacon can be far enough ahead of the HX (~25 m; see Section 3.1.5.2.2—Main Laser 

Site). The HX itself was sized conservatively, to minimize laser-pointing requirements (and to allow the 

system to work without AO, should they prove difficult to implement) and to minimize HX flux. 

Also, the vehicle was designed with the assumption of particular values of tank mass fraction (13% 

for hydrogen and 2% for liquid nitrogen, LN2) and structural mass fraction (10% of other dry mass) and 

the assumption of pressure-fed propellants (to avoid the cost of a turbopump) at a tank pressure of 

250 psi. With these values, and the assumption that the maximum beam power is approximately constant 

(i.e., the boost laser power is equal to the main laser power), the design optimizes with a fairly high 

proportion of the propellant as dense propellant, to give high thrust for liftoff. 

For the COMPASS team analysis, one change was made “going in”: 

 Piston-pumped propellant feed, with 600-psi “chamber” (nozzle entrance) pressure. This was 

enabled by the demonstration of an LH2 piston pump at roughly the correct scale by Xcor, 

allowing both a higher TRL and a basis for a mass estimate for the piston pump 

In preliminary discussions, the long-thin “pencil” hydrogen tank was deemed to have several 

problems: high tank mass that was due to minimum-gauge issues and the need to prevent buckling 

(particularly when supporting ~2000 kg of dense propellant), high insulation mass that was due to the 

high surface-to-volume ratio, and the risk of vibrational resonances and bending instabilities. A shorter 

tank with a nominal diameter of 2 m was assumed, which was based on a rough compromise between 

structural efficiency and drag. 

Initial estimates of the mass of a fixed tower/long shroud or a deployable boom to provide a pointahead beacon were relatively high, so for the purposes of this modeling effort, we assumed that we would 

not consider beacon location as a design requirement. 

The vehicle design as analyzed by NASA Glenn Research Center’s COMPASS team is discussed in 

Section 3.1.4. Key features of this design are given in Table 3.18: 

TABLE 3.18.—PROPERTIES AND MASS BREAKDOWN OF CONCEPTUAL 100-MW HX 

LAUNCH VEHICLE AND DRM–1B LASER HX LAUNCH VEHICLE 

 Concept COMPASS DRM–1B 

Design insertion altitude, km 300 400 

Boost laser power,a MW 100 220 

Main laser power, a MW 100 120 

GLOW,b kg 3250 5543 

LH2 propellant, kg 995 3166 

Dense propellant, kg 1750 (LN2) 1227 (H2O) 

Dense propellant tank Internal External drop tank 

Propellant feed system Pressure fed LH2: Piston pump 

H2O: Pressure fed 

Tank pressure, MPa 1.6 0.35 

“Chamber” pressure, MPa 1.0 4.0 

HX area, m2 25 24 

HX channel length, cm 10 40 

HX exit temperature, K 1700 2000 

Maximum HX power, MW 100 220 

Peak HX flux, MW/m2 5 10

Estimated HX mass/area,c including manifold and 

structure, kg/m2

~3 ~8 

a

Nominal maximum power absorbed at vehicle; actual laser output power would be higher. b

Gross liftoff weight 

c

Before margins. 
NASA/TM—2012-217014 237 

3.1.5.1.3.9 Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) 

Design Evolution 

 The COMPASS design was consistent with the author’s earlier estimates in many respects. The 

Optimal Trajectories by Implicit Simulation (OTIS) trajectory model eventually obtained results very 

close to those predicted using the custom LAUNCH code (a nonoptimizing code) written by the author. 

The propellant tank mass fractions from the COMPASS analysis were generally lighter than the earlier 

estimates (albeit with a 50-psi tank pressure rather than 250 psi). Several components and subsystems 

were heavier in the COMPASS design than in the initial design because of the use of heritage components 

and design rules or simply because of a more detailed analysis, but the increased masses were generally 

manageable. Most of the difficulty in closing the COMPASS design came from the unexpected (by this 

author) high structural masses, particularly in supporting and attaching the HX. 

An early (and possibly premature) decision was made to attempt to close the design without using 

dense propellants. When only hydrogen propellant was used, the initial vehicle mass was limited by the 

available thrust, and for constant laser power, the vehicle generally runs out of propellant well before it 

reaches the limits of the laser range, thus making inefficient use of the laser power. Increasing the boost 

laser power helps and is relatively inexpensive in terms of the ground system capital cost. However, even 

with a 200-MW boost laser, the design would not close for a true single-stage all-hydrogen vehicle. 

One option that did allow the design to close was the use of a vertical catapult or small “strap-on” 

chemical booster rockets. If the vehicle is given a brief boost in addition to the laser-powered thrust so 

that it reaches 150 to 250 m/s, it can have a GLOW that is 30% to 50% higher than if it needs to 

accelerate from rest. 

A catapult is a plausible solution as long as its cost is small compared with the rest of the system 

capital cost. The boost laser can acquire the vehicle, and the laser thruster can be running before the 

catapult fires, so no “in-flight engine start” is required. A subsonic catapult able to supply ~100-kN force 

for a few seconds is well within the state of the art (SOA) (comparable to an aircraft carrier catapult). 

However, a short vertical catapult (e.g., 200 m) necessarily applies substantial acceleration to the loaded 

vehicle, potentially increasing the vehicle structural mass considerably. A long vertical catapult (300+ m) 

would be an unprecedented structure, although large antenna towers are routinely built to heights >500 m. 

Solid booster rockets were suggested, but although they are a plausible solution for tests, they would 

tend to defeat both the cost and range-safety advantages of a beamed-energy launch system. 

The final design returns to the use of dense propellants to provide increased thrust, but with a much 

smaller ratio of dense propellant to hydrogen than in the earlier design, roughly 1:3 versus 1.75:1. 

Because of the time constraints of the study, the design uses the simplest possible form of dense 

propellant injection: pressure-fed injection after the HX (so that the HX does not need to handle varying 

propellant properties). The pressure-fed tanks and pressurizing system are too heavy to take to orbit, so 

they are configured as drop tanks. 

3.1.5.1.3.10 Drop Tanks Are Not Required 

The final COMPASS design can be made to close without drop tanks with only slight changes. The 

dense propellant tank is small (~1.2 m3

) and, if the propellant is pump-fed, very light (<20 kg). If we 

allow 28 kg for the tank, added plumbing and hardware, and extra pump mass (32 kg including systemlevel margin) and assume that the 27 kg allocated for structure to support the drop tanks is enough to 

support fixed tanks, we can maintain the same gross liftoff mass (GLOM) and flight profile, up to the 

point where the tanks were dropped, while carrying substantially extra hydrogen. By slightly increasing 

the main laser power (by 154 kg /2820 kg, ~5.5%) this extra hydrogen can be burned without lengthening 

the trajectory but giving substantially greater final mass injected into orbit. A self-consistent set of mass 

deltas (relative to the COMPASS final design) at various points in the trajectory, assuming fixed GLOM, 

are given in Table 3.19. 
NASA/TM—2012-217014 238 

TABLE 3.19.—DELTAS IN MASS FOR NON-DROP-TANK VERSION OF COMPASS DESIGN 

Change Mass delta, kg 

Liftoff Tank drop Burnout 

Remove drop tank assemblies –194 -- -- 

Add integral tanks +32 +32 +32 

Add extra LH2 (including 1% residual) +156 +156 +2 

Increase LH2 tank size +16 +16 +16 

1% allowance for residual dense propellant + pressurant -- +12 +12 

Decrease payload as needed to make design change self-consistent –10 –10 –10 

Net vehicle mass change 0 +206 +52 

Change in mass to orbit (mass ratio of 3.94 from tank drop) +52 

The mass deltas in Table 3.19 are estimates because there was no opportunity to rerun the COMPASS 

mass model or trajectory calculation. However, it is clear that a true single-stage vehicle, with no drop 

tanks or other dropped mass, will close even with the COMPASS mass margins, with a small decrease in 

payload. 

3.1.5.1.3.11 Potential Future Design Directions 

Optimized design.—Many aspects of the DRM–1B design are arbitrary, artifacts of earlier design 

choices, or chosen for ease of modeling, and are likely far from optimum. For example, the flat HX could 

be replaced by a sectioned or curved HX, allowing at least part of the HX to be supported directly by the 

tank wall, as in the DRM–1C design, and reducing the off-center aerodynamic loads on the vehicle. The 

shroud is considerably oversized, even for the large assumed payload volume, and could be replaced with 

a more tapered shroud with lower mass and drag. 

SOA electronics.—The largely heritage electronics design has a total mass for guidance and control, 

telemetry, attitude determination, and electrical power of ~37 kg. This is clearly excessive for a vehicle 

with a total dry mass of <1000 kg. A new design, developed specifically for production runs of thousands 

of units, would certainly be lighter, possibly by 20 kg or more. It is worth noting that highly sophisticated, 

space-qualified avionics packages weighing well under 10 kg were developed for the Strategic Defense 

Initiative Organization (SDIO) Brilliant Pebbles program over a decade ago. Some options for weight 

(and cost) reduction follow: 

 Microelectromechanical systems (MEMS) inertial sensors and gyroscopes, similar to those used 

in small unmanned aerial vehicles (UAVs)—The vehicle can also use optical beam sensing and 

differential Global Positioning System (GPS) to supply attitude data, and ground optical tracking 

data for guidance. 

 Lithium polymer or lithium-iron-polonium (LiFePo) batteries—Neither long-term storage (low 

self-discharge) nor large total energy capacity are required. 

 Single-board construction—A card cage is unnecessary. 

 Wireless or serial-bus interconnects 

 Separate batteries and controllers for the fore (payload) and aft (propulsion) bays, with wireless 

or optical-fiber connection between them; no multimeter electrical cable runs 

Thrust vectoring.—The expansion-deflection (E–D) nozzle is ideally suited for thrust vectoring by 

fluid injection, either before or after the radial throat. Injecting dense propellant differentially around the 

nozzle would potentially give much more attitude control authority at low altitude, where it is most 

needed, and reduce the size and propellant usage of the attitude control system (ACS) thrusters, which 

would provide only roll control for most of the mission. 

 
NASA/TM—2012-217014 239 

Dropped mass.—The DRM design assumes that, to minimize operational range-safety issues, no 

hardware at all is dropped from the vehicle other than, possibly, the dense propellant tanks, which would 

fall very close to the launch site. If the safety and operational impacts are acceptable, the vehicle could 

follow the common launch vehicle practice of dropping at least part of the payload shroud and any other 

hardware (such as tank insulation and parts of the HX structural support) needed only for prelaunch or for 

flight through the atmosphere. 

More radical options include 

Enhanced propellant cycles.—The DRM design feeds only hydrogen through the HX, thus 

requiring both hydrogen flow at all times and a small amount of combustion heating to maintain the 

propellant temperature when dense propellant is injected downstream of the HX. The HX may be able to 

accommodate both hydrogen and dense propellant (e.g., nitrogen, ideally with a fully variable mixture 

ratio). In that case, the vehicle could carry no oxidizing propellant at all and could turn the hydrogen flow 

down or completely off for at least the first few seconds of launch. An all-inert launch would protect both 

the launch site and the vehicle from any risk of a hydrogen-air fire or detonation. 

Alternatively, the oxidizer content of the dense propellant could be increased to provide more energy 

from combustion. This would reduce the power and cost of the boost laser array. In the limit, the vehicle 

would be primarily a chemical rocket for the first part of its trajectory. A low level of laser power could 

still be used to heat some or all of the hydrogen flow and power the propellant pump. This would ensure 

ignition and fast burning of the oxidizer and would also allow the system to fail safe by cutting off 

propellant flow if the laser beam was lost or turned off. 

Deployed reflectors.—Once above the atmosphere the vehicle could deploy lightweight reflectors to 

increase the effective beam-collecting area. These would typically be flat metal-foil reflectors, similar to 

the reflectors used on low-concentration solar panels. The HX area could be reduced or the useful laser 

range increased; either would increase payload or allow scaling to lower laser power. This is particularly 

beneficial for scaling the system to lower power and smaller vehicle size because the HX area and thus 

the mass of the HX and its support structure are set in part by the minimum practical laser spot size. 

Ducted rocket.—The temptation to use the atmosphere as “free” oxidizer and reaction mass is 

always present in designing launch vehicles. For the laser HX rocket it is particularly strong given that its 

exhaust is hot hydrogen, which will immediately burn in air. The COMPASS team considered a ducted 

rocket or rocket-assisted ramjet configuration, operating over a wide Mach range. This gave very 

promising estimated performance, but required a variable inlet and a potentially complex air-exhaust 

mixing arrangement. The COMPASS group lacked the time and expertise to complete a credible design. 

A high-performance air-breathing design, which can operate usefully from subsonic speeds and low 

altitudes up to, for example, Mach 5+ and 30+ km altitude, would require a major design effort and would 

add substantially to the cost and risk of development. An alternative that would be consistent with the 

general approach of minimizing vehicle complexity would be a fixed-inlet duct operating over a relatively 

narrow velocity range and discarded at an altitude of 30 to 40 km. 

Air-breathing modes should certainly be considered in future vehicle design studies. We note, 

however, that, despite decades of proposals and research and development (R&D) efforts, only a handful 

of test engines of any kind have actually flown above roughly Mach 3, and no chemical-fuel launch 

vehicles (and very few missiles) have used air-assisted stages, suggesting that it is very difficult to 

actually build an air-breathing “accelerator” (as opposed to a constant-speed-and-altitude “cruiser”) that 

outperforms a pure rocket thruster. It is usually easier and cheaper to simply carry more rocket propellant. 

Reusable vehicle.—We made no effort in this study to design the vehicle for intact reentry and reuse, 

but it is almost certainly possible to make a reusable version. The empty vehicle has low areal density, 

and a large part of its hot surface on reentry would be the already-refractory HX. Intact reentry would 

require some additional thermal protection, but possibly very little, plus adding a design constraint for 

stable reentry aerodynamics. The vehicle would need additional propellant for fully controlled reentry, as 
NASA/TM—2012-217014 240 

well as a landing mechanism—possibly wings for horizontal landing, but more likely a parachute/airbag 

or similar system, given the vehicle’s low empty weight and low terminal velocity. 

3.1.5.2 Description of Ground Facility 

3.1.5.2.1 Overview 

The overall ground system for DRM–1B consists of three facilities, as shown in Figure 3.90: 

 The vehicle launch site is where the vehicles actually leave the ground. The launch site includes 

at least one launch tower or rail (which may be a vertical catapult), along with auxiliary 

equipment for handling vehicles, payloads, and propellant. 

 The boost, or L1, laser site provides laser power to the vehicle for its initial launch and climb 

from the launch site through the atmosphere, until the vehicle is high enough to receive power 

from the main laser. 

 The main laser site is the heart of the launch system, and consists of an array of beam modules—

independent lasers with small beam directors and tracking systems—with a total output power 

(for the DRM baseline system) of ~120 MW. It is typically several hundred kilometers 

downrange (east) of the launch site, so that the vehicle trajectory passes over the main laser near 

the midpoint of its trajectory. 

Details of these facilities are discussed in the following subsections in the order of cost, which is the 

reverse of the preceding sequence. 

Figure 3.90.—Ground facilities for laser HX launch system. Middle image: copyright LaserMotive, LLC; used with 

permission. Right image: copyright Kare Technical Consulting; used with permission. 
NASA/TM—2012-217014 241 

3.1.5.2.2 Main Laser Site 

3.1.5.2.2.1 Beam Module Concept 

The overwhelming problem with laser-beamed-energy launch has generally been seen as the need for 

a very large laser. An approximate rule of thumb is that it takes at least 1 MW of average laser power per 

kilogram of payload launched to low Earth orbit (LEO),1

 so that even few-kilogram payloads require 

multiple megawatts of power. 

There is no requirement, however, that the laser energy come from a single coherent source. Many 

laser applications, including most directed energy weapon concepts, do require, or at least strongly prefer, 

a coherent source, since they are attempting to deliver as high a flux (power/area) as possible, or to 

deliver energy using as physically small an optical aperture as possible. In directed energy applications, 

pointing and tracking systems are often very expensive, since they need to track uncooperative targets, 

often at very high angular rates and under difficult conditions (e.g., from an airborne platform). 

 For laser launch, however, there is no inherent limit on the number of separate lasers and optical 

apertures used, as long as each individual laser and beam director can deliver its power efficiently, and 

the optics and tracking hardware do not dominate the system cost.2

There are multiple advantages to using many small lasers versus a single large laser: 

 Lower development cost and risk 

 Lower production costs through volume production and competitive production 

 Higher system reliability and availability (using a small number of “spare” lasers) 

 Continuous power expansion and incorporation of new technology 

 Reduced propagation issues: no thermal blooming 

3.1.5.2.2.2 DRM System Beam Module Design 

The conceptual beam module design is sketched in Figure 3.91. The major components follow: 

 Lasers: six by 10-kW single-mode fiber lasers 

 Beam projector (telescope and mount) 

 Tracking subsystem, including tip-tilt mirror 

 AO subsystem, including Rayleigh beacon (laser guidestar) if needed 

 Power supply, including local energy storage 

 

1

 This “rule” was originally derived following the first SDIO/Defense Advanced Research Projects Agency 

(DARPA) Laser Propulsion Workshop. It arises from the dual constraints of thrust/weight >1 at liftoff, and horizon 

distances at LEO altitudes, both of which, in a simple model, limit the time for horizontal acceleration to a few 

hundred seconds, and thus the total energy delivered to a few hundred megajoules per megawatt of source power. If 

100% of that energy ended up as payload kinetic energy (~32 MJ/kg for 8 km/s orbital velocity), we could put 10 to 

20 kg/MW in orbit, but various inefficiencies reduce the actual useful payload by a factor of at least 10 for every 

propulsion scheme and vehicle design proposed to date. 2

 There is a surprisingly common misconception that multiple lasers must be phase-locked together to deliver 

laser power efficiently. This is generally based on misinterpreting the fact that phase-locking increases the flux

delivered by an array of lasers; the peak flux possible using a coherent array of N lasers is N2

 times the peak flux 

from a single laser, versus N times for an incoherent array. However, the total power delivered is the same in both 

cases; the coherent array simply delivers the power into a smaller area. 
NASA/TM—2012-217014 242 

Figure 3.91.—Block diagram of 60-kW beam module. Copyright Kare Technical Consulting; used with permission. 

The overall design is virtually identical to the beam module concept described in (Refs. 2 and 4), 

except for the use of full AO, and a slightly larger mirror to take advantage of the improved beam quality 

possible with AO. 

As in the original design, the lasers are assumed to be completely independent, with each laser having 

its own “slice” of the telescope (i.e., the telescope acts as six separate off-axis telescopes). This is by no 

means required and other telescope and mount configurations are possible, including 

(1) Completely independent telescope (35 to 50 cm) and mount for each laser 

(2) Independent telescopes, but with some number of telescopes sharing part or all of a mount 

Independent on-axis telescopes would be more compact (shorter focal length) and probably less 

expensive than a large telescope, but they would have a lower transmission efficiency (typically by a few 

percent) because of the secondary mirror and its support. Independent off-axis telescopes would in the 

past have cost much more than a single larger telescope, but with developments in optical fabrication 

including diamond turning, magnetorheological polishing, and deterministic polishing, plus automated 

measurement, assembly, and alignment techniques, the cost may be similar or lower. 

The choice of a single 10-kW laser is based on the availability of such a laser as a commercial 

product, announced in July 2009 by IPG (Ref. 12) shown in Figure 3.92. Note that these are single-mode 

lasers with nearly diffraction-limited beam quality (M2 < 1.3). IPG fiber lasers are built expressly for 

extremely high reliability, long-life operation in industrial production environments, and in particular use 

telecommunications-grade single laser diodes (with an operating life >100 000 hr) to pump the fiber. The 

beam is delivered over a single high-power optical fiber that could be directly connected to the beam 

director or adaptive optics system. They are therefore, uniquely, a TRL 8 beam source for laser launch, 

requiring (to the author’s knowledge) no modifications for use in a launch facility. 
NASA/TM—2012-217014 243 

Figure 3.92.—IPG YLS–10000–SM 10-kW single-mode laser. Copyright IPG Photonics Corporation (Oxford, MA) 

10,000 W Single Mode Laser; used with permission. 

That said, the unmodified IPG 10-kW fiber laser is unlikely to be either the best or the least expensive 

laser for a launch system. It is an unusual design in that the actual 10-kW fiber is pumped by a set of 

lower-power fiber lasers rather than directly by diode lasers. This makes it more expensive and somewhat 

less efficient than IPG’s lower power fiber lasers (23% vs. >30% wall-plug efficiency). 

The main alternatives to the IPG 10 kW laser follow: 

Lower power (1- to 3-kW) single-mode fiber lasers.—These are available from IPG and potentially 

from other laser manufacturers (Refs. 13 and 14). Some form of beam combining, either coherent or, 

incoherent (i.e., wavelength combining), would be desirable to raise the unit beam power and thus keep 

the total area and cost of optics manageable, although even at 3 kW/beam it is unlikely that optics would 

dominate the main laser array cost. Incoherent combining of several continuous-wave (CW) fiber lasers 

using volume Bragg gratings has been demonstrated at power levels of a few hundred watts per beam 

(apparently limited by the available lasers) (Ref. 15), so this approach is at least at TRL 4. 

Thin-disk lasers.—These are manufactured by TRUMPF (which holds fundamental patents on the 

design) in power levels up to 16 kW, also for industrial cutting and welding. These are sold as multimode 

lasers, with nominal beam quality of 8 mm-mrad (about 4  diffraction limited) (Ref. 16), but Boeing has 

licensed the technology and has demonstrated weapons-quality beams (Ref. 17). Prices for TRUMPF’s 

welding lasers are comparable to or slightly higher than equivalent-power, equivalent-beam-quality fiber 

lasers from IPG because the two companies are direct competitors in this market. 

“Conventional” solid-state lasers.—Joint High Power Solid-State Laser (JHPSSL) has 

demonstrated 100-kW beams with high beam quality (exact value unknown, but almost certainly adequate 

for a launch system) and has run (cumulatively) for >6 hr (Ref. 18). Northrop Grumman announced the 

FireStrike 15-kW Diode-Pumped Solid-State Laser (DPSSL) module as an orderable product in 2008 

(Ref. 19). Textron Systems has also demonstrated 15 kW from its ThinZag diode-pumped ceramic slab 

laser, and other solid state lasers are under development (Ref. 20). This approach is thus at TRL 6, but 

costs and true operating lifetimes are unknown. 

Diode pumped alkali lasers (Refs. 21 to 23).—These have been demonstrated only at 100-W power 

levels, but which are comparatively easy to scale to higher power. There is a reasonable chance that 

>10 kW will be demonstrated in 2011, which would raise this approach to TRL 4 or higher. 

The “holy grail” would be a low-cost means of phase-locking large arrays of laser diodes at low cost, 

or, alternatively, of making high-power narrow-linewidth laser diodes that can be incoherently 

wavelength-combined to radiances of >1014 W/m-sr. There are tantalizing suggestions that one or both of 

these may be possible with external-cavity diode arrays, either edge-emitting or vertical-cavity, but so far 

these options remain at best at TRL 3. 
NASA/TM—2012-217014 244 

The overall system design is nearly independent of the choice of laser,3 and development and testing 

of beam modules can be done at the level of a single module, so it is not necessary to choose a winner in 

laser technology now. Indeed, new laser technology can be incorporated into beam modules at any point, 

even after the launch array is built and operating, if it provides significant benefits. 

Pointing and tracking.—The pointing and tracking subsystems in each beam module independently 

track the vehicle beacon through the launch. If the tracking subsystem for a particular laser does not have 

a positive signal from the beacon, it will shut the laser down for safety. 

The detailed design and operation of the tracking subsystem is beyond the scope of this report, but it 

is well within the current SOA. A typical system would use an imaging camera (charge-coupled device, 

CCD, or complementary metal-oxide-semiconductor, CMOS), viewing either through the main telescope 

or through a separate “finder scope,” to acquire the vehicle beacon, and a higher-bandwidth quadrant 

detector, viewing through the telescope and tip-tilt mirror, to keep the line of sight centered on the 

beacon. This detector can take advantage of modulation of the beacon light (typically at 0.1 to 1 MHz) to 

suppress interference from both background light and reflected laser light that leaks through filtering. 

A tip-tilt mirror is required to compensate for large-scale atmospheric turbulence, but it can also 

compensate for pointing errors and vibration of the telescope mount. (It is an open design question 

whether or not multiple beams sharing a telescope aperture should share a common tip-tilt mirror. Using 

separate mirrors and tracking sensors costs more, but it provides slightly better atmospheric correction if 

full adaptive optics are not used, and it reduces the adaptive optics dynamic range if they are.) 

Adaptive optics.—The 100-MW baseline laser launch system in prior work was deliberately sized to 

avoid any need for adaptive optics beyond simple tip-tilt-focus correction because full adaptive optics 

were assumed to be too expensive for a $1M-class beam module. However, progress in micro-optical 

devices, plus the predictable march of Moore’s Law in computing capacity, has driven the price of 

adaptive optics far below $1M. For example, at this writing, Thorlabs, Inc., is offering a complete 140-

element adaptive mirror with wavefront sensor and control electronics for roughly $23,000. Although this 

system uses a MEMS mirror with limited power-handling capability, it has enough sensor resolution and 

enough actuators (140) to correct a beam up to ~54 cm across (assuming that N = (d/r0)

2

 for aperture 

diameter d and Fried parameter r0 = 4.5 cm at a 60 zenith angle). 

A detailed reanalysis of the beam divergence and beam module optics remains to be done, but if we 

assume that atmospheric turbulence is largely corrected, a 54-cm aperture (combined with appropriate 

improvements in jitter and pointing errors) will reduce the expected beam divergence (1/e2 half-angle) 

from 8 to ~4 urad. 

This improvement would allow the laser range to be increased, the HX dimensions to be decreased, or 

both. Increasing the laser range would increase the maximum trajectory length and, therefore, the energy 

delivered to the vehicle. This would increase the payload size for a given laser power. This would require 

revising several aspects of the system design (e.g., the orbit insertion altitude would have to be higher to 

keep the vehicle from going over the laser’s horizon). 

Decreasing the HX area without changing the beam power would reduce the mass of the HX and 

(perhaps crucially) that of its support structure, but it would require the HX to accept higher flux. 

Alternatively, the entire vehicle could be scaled to lower mass (including smaller payload) and lower 

power, reducing the cost of an initial launch system. Decreasing the beam divergence by a factor f would 

allow a reduction in HX area and mass by factor of 2 (f2) (assuming constant HX thickness). Not all 

vehicle components would scale linearly with HX mass, nor would, for example, drag losses, so reducing 

the beam divergence to 4 μrad would not allow scaling a 100-MW system with a 100-kg unit payload to a 

25 MW/25 kg system, but preliminary modeling suggests that a 25-MW system would still have positive 

net payload. 

 

3

There would be differences in module design, safety requirements, and other parameters for different 

wavelengths, but all the proposed lasers are in the same general spectral range (0.7 to 1.1 m), so the differences 

among them would be small. 
NASA/TM—2012-217014 245 

Note that individual beam modules operate well below the power and beam diameter thresholds 

where significant thermal blooming occurs. If we assume that intermodule spacing would be reasonable, 

beam modules would not need to correct for thermal blooming. 

Figure 3.93.—Adaptive optics beacon offset versus time for a 

typical LEO trajectory. The notch at 420 s cates a point in 

the trajectory where the vehicle pitches nose down slightly 

(~10) to circularize its orbit and improve the laser HX 

viewing angle. 

Adaptive optics beacon requirement.—If full adaptive optics are used, a beacon located ahead of 

the HX by the point-ahead distance dpa = 2vR/c along the velocity vector is required, so that the 

downward-going beacon light samples the same part of the atmosphere as the outgoing power beam 

passes through. (Here v is the vehicle velocity, R the range to the laser, and c the speed of light; 2R/c is 

the round-trip time for a light signal from beacon to laser and back.) Note that if the vehicle has a nonzero 

angle of attack (i.e., velocity vector is not parallel to the vehicle axis) the physical offset between the 

beacon and HX may be different from the pointahead distance. 

A plot of the beacon offset as a function of time for the last 200 s of a typical launch trajectory is 

shown in Figure 3.93. The offset initially declines, even though the vehicle is accelerating, because it is 

getting closer to the laser (which is 400-km downrange from the launch site); once the vehicle passes over 

the laser, the offset increases rapidly. The maximum offset of 25 m is typical for launches to LEO. 

As noted in Section 3.1.5.1.3—DRM Vehicle Design, a vehicle long enough to support a beacon with 

a 25-m offset may be far from optimum structurally at the 100-MW scale. Further design work is needed 

to determine whether there is a practical solution using a fixed or deployable spike or boom on the nose of 

the vehicle. Note that the variation in beacon offset also means that either the beacon would have to be 

moveable or (more likely) that there would be beacon emitters (diode lasers or optical fiber tips) at 

several points. 

An alternative to having a long vehicle with multiple onboard beacons is to equip each beam module 

with a laser-generated beacon. Rayleigh beacons (e.g., Ref. 24) require moderate beacon laser power (a 

few watts time-average) and provide good performance over the aperture diameters of interest. They can 

be operated in a pulsed mode with gated detectors, with adjacent modules using different pulse timeslots 

to minimize mutual interference. This approach would raise the capital cost of the system (although 

probably only by on the order of 1%) but would minimize the vehicle complexity and cost. An onboard 

optical beacon would still be needed for coarse tracking and tip-tilt beam steering, but a single beacon at 

any convenient point on the vehicle would be sufficient. Rayleigh beacons for this class of aperture are at 

least TRL 6, having been implemented on various ground-based telescopes and on the Airborne Laser 

(ABL). 
NASA/TM—2012-217014 246 

The structure and scale length of atmospheric turbulence varies with time and location. Analytical 

modeling, simulations, and possibly onsite testing will be needed to predict the actual beam profiles to be 

expected for a particular launch site and a particular beacon configuration. The operational impact of a 

given degradation in beam quality also depends on the HX size and shape and the trajectory flown. We 

believe it is likely that an operational system will not require Rayleigh beacons for each beam module 

(i.e., a vehicle-mounted beacon will be sufficient even if it is not optimally placed) but that Rayleigh 

beacons may provide more than enough improvement in system performance to be worth their cost. 

3.1.5.2.2.3 Main Laser Site Design 

An artist’s conception of a beam module array is shown in Figure 3.94. This particular version has the 

beam module packaged in two standard 40-ft cargo containers, with the lasers in one container and the 

beam director and associated optics in the other (with a roll-away roof and walls); the two would be 

connected by optical fiber cables. 

 Using standard containers would allow the modules to be easily trucked from a factory to the site and 

returned to the factory for repairs if needed. The modules could be emplaced on simple concrete pads and 

connected to cooling water and power. This general configuration has been proposed for large data 

centers, as illustrated in Figure 3.95. A large data center would have roughly the same number of 

containers and much greater power and cooling requirements than the DRM launch system array, 

differing mainly in that server containers could be placed closer together, since they have no field-of-view 

requirements. Google and others have built such modular data centers, although generally on a smaller 

scale, and cargo-container server modules are available from Sun, IBM, and other manufacturers. It 

would, of course, be possible to install beam modules in more conventional buildings if cargo containers 

were too small or otherwise unsuitable. 

Figure 3.96 shows the scale of a 2000-module array, using modules on regular ~40- by 50-m centers. 

Many other configurations are possible, including regular or irregular patterns with closer or wider 

spacing, as long as the spacing is sufficient to keep one module from blocking another’s beam—typically 

<10 m—and to allow access for installation and maintenance. (Some spacing is also needed to avoid 

collective thermal-blooming effects.) If land area were at a premium (e.g., on a mountaintop) a 2000-

module array could fit into ~200 000 m2

, or 1/5 km2—slightly over 40 acres.
NASA/TM—2012-217014 247 

Figure 3.94.—Beam modules in the field. Copyright Kare Technical Consulting; used with permission. 

Figure 3.95.—Modular data center, as envisioned in IEEE Spectrum (Ref. 25). Copyright Kare Technical Consulting; 

used with permission. 
NASA/TM—2012-217014 248 

Figure 3.96.—A 2000-module array, with 40 by 50 m spacing, is 2  1.6 km, or about the size of a 18-hole golf 

course. Copyright Kare Technical Consulting; used with permission. 

Power and cooling.—If we assume a wall-plug efficiency of 30% (present efficiency of IPG singlemode fiber lasers) and a total main-site laser output power of 140 MW (corresponding to 120 MW at the 

vehicle as it passes over the site)4

, the main laser array would need 470-MW input electric power for 

~400 s. Some additional power would be needed for electronics, pumps, telescope mounts, and so forth.; 

we estimate this at 20 kW per beam module, or about 10% of the power needed by the lasers. As 

discussed by Kare and Parkin (Ref. 26), a heavily used system could simply install 520 MW of gasturbine generators, identical to those used for utility peaking power, at a cost of about $0.60 to $0.80/W. 

For the DRM launch rate of 1000/year (a few launches per day) it will probably be less expensive to 

use a lower-power generator, or a grid connection, plus an energy-storage buffer. Any of several types of 

batteries are capable of storing the necessary energy (~210 GJ or ~60 MWh per launch) and providing the 

required power for a fraction of the capital cost of generators. However, most existing battery 

technologies do not have sufficient cycle life to be economical. 

The availability of all-electric cars is encouraging, as they necessarily have battery packs designed for 

high cycle life. For example, the Chevrolet Volt battery pack characteristics from (Ref. 27) are given in 

Table 3.20. 

 

4

 This is the approximate power requirement for the COMPASS DRM vehicle design. Significantly lower 

power (100 to 150 MW at the vehicle) may be sufficient with an optimized design. 
NASA/TM—2012-217014 249 

TABLE 3.20.—CHARACTERISTICS OF CHEVROLET VOLT BATTERY PACK 

Energy storage 10.4 kWh (usable) 

Power output 111 kW peak 

Estimated usable power for 400 s 80 kW 

Estimated cycle life (based on 8 year warranty, 1 cycle/day) 3000 

Estimated cost today $7500 

Estimated cost in 5 to 10 yr $4000 

Nissan Leaf batteries are similar, with slightly higher cost estimated (<$750/kWh total capacity, about 

$18000 for 24 kWh (Ref. 28), presumably with similar power output. 

At 80 kW per pack, three packs, costing (at $4000 each) $12,000 total, would be needed to power a 

single beam module for one launch, or roughly 6500 packs costing $26 million to supply 520 MW 

($0.05/W). For operational flexibility, the launch system would probably have sufficient storage to allow 

at least two or three launches in quick succession, representing a capital cost of $52 to $78 million. With 

two launches worth of storage and 1000 launches/year, batteries would need to be replaced after 6 years 

(500 cycles/yr on each battery). With three launches worth of storage, the battery life would be 9 years, or 

very close to the nominal 10-yr system life. 

Note that electric car batteries are probably not the optimum energy storage solution, but they do 

represent the SOA in mass-produced high-discharge-rate batteries, they are TRL 8 (in volume production) 

and they happen to fit the cycle-life requirements for the DRM (if we assume that the claimed cycle lives 

are achieved in practice) and the power requirements of nominal beam modules. Other high-cycle-life 

batteries being developed for stationary applications (sodium sulfide, NaS; vanadium redox, etc.) or 

energy storage devices such as flywheels may be cheaper or more reliable, especially in another 10 years. 

For launch rates less than 100 per year, older battery technologies (lead-acid; nickel metal hydride, 

NiMH; or nickel cadmium, NiCd, if cadmium is acceptably recycled) would probably be less expensive, 

even with life of ~300 cycles. For launch rates higher than roughly 3000 per year, or bursts of >10 

launches, gas turbine generators would be competitive. 

Using batteries at each beam module (or perhaps a battery system for each small group of modules) 

would reduce power distribution costs within the launch array, and might reduce power supply cost by 

providing direct direct-current (DC) input to the laser drivers at an appropriate voltage. 

If we allow 8 hr to deliver energy for one launch, with 80% efficiency in battery charging, the input 

power needed would be only 75 MWh/8 to ~9.4 MW. Continuous housekeeping power of a few kilowatts 

per beam module would add another 5 to 10 MW, so the primary power source—generators or grid 

connection—for the main laser site should be rated at 20 MW or more. 

Cooling.—For the DRM launch rate, cooling to a thermal sump (water tank or pool) during launches 

is convenient and cost effective. The total energy dumped would be (470 – 140 + 50) MW  400 s = 

152 GJ. A water volume of approximately 10 000 m3

 (four Olympic swimming pools) would be heated by 

3.6 C by one launch. (A single 60-kW beam module, dissipating 160 kW  400 s = 64 MJ, would require 

3.5 m3

 of water, or about 900 gal (3407 liters), for a similar temperature rise). Depending on the local 

climate and water supply, the thermal sump temperature could be maintained by an evaporative cooling 

tower or a system of chillers removing 500 GJ of heat per day or ~6 MW continuously. Housekeeping 

cooling—removing constant heat loads and solar heat—may be substantially more. 

3.1.5.2.3 Boost Laser Site 

The boost, or L1, laser site is a separate laser array located relatively close to the vehicle launch site, 

nominally 20-km downrange, with line of sight to the vehicle launch stand(s). The primary purpose of 

having a boost laser is to supply power to the vehicle until it is high enough to be in view of the main 

laser array at a reasonable elevation. Because the HX vehicle has a side-mounted HX, it is difficult to get 

a workable trajectory if the main laser is close enough to the launch site to see the vehicle on the ground: 

the vehicle quickly becomes tail-on to the beam. (It is possible to design a HX-based vehicle that will 

work in this geometry, using, e.g., a pivoting HX or deployed reflector), but the resulting system is still 

less efficient than one with the main laser farther downrange.) 
NASA/TM—2012-217014 250 

A nonlaser “first stage” (in quotes because this option also includes non-rocket stages such as carrier 

aircraft or catapults) could be used to get the vehicle off the ground and up to a reasonable altitude, and 

this may be the preferred option for some testing, but it inherently makes the vehicle and operations 

concepts more complex. Also, many first-stage options, including air-dropping the vehicle from a carrier 

aircraft, cannot provide enough altitude to allow placing the main laser optimally far downrange.) 

If the cost per watt of the boost laser and main laser were similar, using a boost laser would nearly 

double the overall system capital cost, and would probably not be justified except for a very high-volume 

launch system. However, the boost laser needs a range of only 50 to 100 km and can, thus, have 

~10 times higher divergence and ~100 times lower brightness than the main laser. This makes lower-cost 

laser options usable. 

If the boost laser is substantially cheaper per watt than the main laser array, then the overall system 

design could be optimized by making the boost laser somewhat more powerful (typically 1.5 to 2 times) 

than the main laser array. This would provide extra power for liftoff and flight through the atmosphere, 

allowing the vehicle to be sized to use the full capacity of the main laser without being limited by liftoff 

weight or wasting propellant during a low initial acceleration. Similarly (although less significantly) if the 

boost laser is more efficient than the main laser, using it for a large fraction of the powered flight duration 

(even though only a small part of the trajectory length) would reduce the overall launch-system electric 

power consumption. 

3.1.5.2.3.4 Boost Laser Technology 

The preferred technology for the boost laser is direct diode arrays. These are threefold to tenfold less 

expensive than fiber lasers and substantially more efficient. The actual diodes are routinely 60% efficient 

(DC to light) today, and efficiencies of over 70% have been demonstrated (Refs. 29 and 30). 

The source radiance R (power/[area(solid angle)], in W/m2

-sr), which is a conserved quantity through 

optical systems, determines the total transmitter aperture area needed to deliver a specified combination 

of intensity and range. (Note that, for fixed radiance, the aperture area is independent of the source 

power.) 

Over the last few years, the radiance of laser diode array sources has increased significantly. As of 

2008, a typical 808-nm diode stack with two-axis microlenses to collimate the individual diodes could 

produce a radiance of 6 to 81010 W/m2

-sr (~500 W/cm2 with a divergence of ~20  4 mrad). Using 

polarization combining and optical interleaving would raise this by a factor between 3 and 4, to ~2 to 

31011 W/m2

-sr (Ref. 31). This is slightly too low to be practical for the DRM system boost laser: to 

deliver 10 MW/m2

 at 50 km (which would actually be about 8 MW/m2 at the vehicle, allowing for 

atmospheric absorption and scattering) would require ~100 000 m2

 of optical aperture. 

However, diode arrays themselves are improving in power and beam quality, so the raw array 

radiance available, particularly at a slightly longer wavelength (980 to 1060 nm) is about a factor of 2 

higher (and likely to increase). Demonstrated techniques exist for improving the optical collimation of the 

laser output, notably individual corrective micro-optics (Ref. 32), which can reduce the fast-axis 

divergence by several fold, essentially to the diffraction limit. Finally, dichroic combining of several 

(three to four) diode laser beams at kilowatt levels is done routinely, for example, by TRUMPF in their 

TruDiode industrial laser systems (Ref. 33). We thus assume that high-power diode sources at 

~41012 W/m2

-sr are at TRL 6, although this is slightly beyond what is currently sold commercially. 

With a source radiance of 41012 W/m2

-sr, the nominal DRM system boost laser array would require 

a total optical aperture of 6250 m2

—roughly 3 times the area for the baseline main laser array, but the 

required optical quality and pointing accuracy are very low, literally worse than modern hobby-grade 

telescopes from Celestron and similar sources. Pointing accuracy of several arcseconds (20 to 30 rad) is 

sufficient, and no adaptive optics are required except possibly tip-tilt steering. Such optics cost in the 

range of $10,000/m2

 to $30,000/m2, with the high end being conventional optics and the low end being 

alternatives such as replica optics, plate-glass honeycomb primaries (Ref. 34), or transmissive diffractive 

optics “printed” on thin glass substrates (Ref. 35). 
NASA/TM—2012-217014 251 

If the expected optics cost or diode-array performance is not achieved, there are multiple alternatives. 

Wavelength-stabilized high-power diode arrays have been demonstrated for alkali-vapor pumping 

(Ref. 36) and would allow much higher numbers of diode arrays to be combined with dichroics, 

diffraction gratings, or volume Bragg gratings. Single-mode high-power diodes are available with higher 

radiance (but also higher cost) than standard multimode arrays. Either of these would allow reducing the 

optics area by factors of 5 to 20. 

Finally, the worst-case fallback for high TRL would be to use multimode fiber, slab, or thin-disk 

lasers, which are currently available at power levels >10 kW and radiances of >1013 W/m2

-sr (e.g., the 

TRUMPF TruDisk 16002 laser is 16 kW with a beam parameter product of 8 mm-mrad; radiance R

= P/(

2

BPP2) = 2.51013 W/m2-sr) with higher radiance versions in development. This would make the 

boost array only 1.5 to 2 times cheaper than the main laser array, but very compact. 

3.1.5.2.3.5 Boost Laser Site Design 

The boost laser site will have a layout generally similar to the main laser site, with hundreds to a few 

thousand modules installed over on the order of 1 km2

 (possibly less, if land area is limited). Each module 

will be largely independent, with its own power supplies, tracking, and other capabilities. The laser power 

and aperture area per module are not yet defined, but a plausible configuration might use 20-kW diode 

array assemblies, each with its own ~1-m mirror. The arrays would be mounted in groups of 10 on a 

common elevation mount as sketched in Figure 3.97. Such a 200-kW beam module would, like the main 

laser array beam module, fit conveniently into two standard cargo container-sized packages, one for the 

actual laser/telescope system and one for power supplies and control electronics. 

Figure 3.97.—Boost laser beam module concept. 

Because the boost laser needs to point at, or very close to, the horizon in one direction, but only 

slightly past the zenith in the opposite direction, the boost laser array should be built on a slope facing the 

launch site. The altitude of the boost laser array is somewhat less important than that of the main laser 

array (because when the vehicle is at maximum range for the boost laser, it is also near zenith and 

atmospheric effects are minimized). It might be desirable to put the boost laser array slightly below the 
NASA/TM—2012-217014 252 

launch-site altitude (if it is assumed that the intervening terrain is lower still) so that even at launch, the 

beam would never be aimed to intersect the Earth’s surface (i.e., it would always point at least 

slightly upward). 

A 250-MW, 60% efficient boost laser array (delivering 220 MW to the vehicle) will require either a 

~420-MW prime power source (generators or grid connection) or, as with the main laser array, a 

distributed energy storage system and a much smaller prime power source. 

3.1.5.2.4 Launch Site 

The launch site provides facilities for 

 Storing vehicles before launch 

 Attaching payloads to vehicles (if they are not delivered with the vehicles) 

 Moving vehicles to one or more launch pads 

 Loading propellants and otherwise preparing for launch 

 Maintaining the vehicle ready for launch 

 Performing any prelaunch checks 

 Safely handling the vehicle exhaust 

The launch site will be similar in scale to existing small-vehicle launch sites such as Pad 0B at 

Wallops Island (which was built to launch the Minotaur I, a solid fuel vehicle roughly the same size as the 

HX vehicle, but much heavier), except that the HX launch site may be inland, and will have at least two, 

preferably three or four launch stands, so that multiple vehicles can be prepped and launched in quick 

succession. 

The launch stands may have small-scale versions of chemical rocket facilities, such as “flame” 

trenches to conduct hot exhaust away from the vehicle and, if hydrogen propellant is used at liftoff, to 

safely burn off the hydrogen exhaust. 

3.1.5.2.5 Location Constraints for Ground Facilities 

3.1.5.2.5.6 Site positions 

As noted previously, the launch system will require at least three sites: the launch site, the boost laser 

site, and the main laser site. Nominally, these would be on a straight line along the vehicle ground track, 

with the two laser sites at 20 and 400 km downrange from the launch site. In practice, the actual locations 

would be fairly flexible. 

The boost laser site will be constrained in distance to the launch site: propagation losses over a 

horizontal path and the need for line-of-sight to the launcher will limit the maximum range, and the angleof-view to the ascending vehicle will limit the minimum range. The boost laser could, however, be far off 

the trajectory ground track, by at least 45, and possibly more; the vehicle could roll to present the HX to 

the boost laser initially and then to the main laser. 

The main laser downrange distance will need to be optimized for a specific set of laser properties and 

vehicle trajectory, but in general should be as far downrange as possible without causing a large loss of 

power early in the trajectory; 400 km is probably shorter than the optimum distance. Changes of 50 or 

even 100 km, however, would have only modest effect on the overall system performance. Similarly, the 

main laser could be offset from the trajectory ground track by 50 km with very small effect. Indeed, as 

discussed later, the trajectory will be shifted to vary the inclination of the final orbit, so the main laser site 

could be anywhere in a zone of at least 100 by 100 km, and probably larger. 

3.1.5.2.5.7 Altitude 

Altitude is not a direct constraint for any of the laser launch facilities, and an all-sea-level launch 

system (including a ship-based launch system) is feasible. Some laser wavelengths in the ranges of 
NASA/TM—2012-217014 253 

interest are significantly absorbed by water vapor, so high-altitude laser sites may have slightly wider 

ranges of usable wavelengths, but windows exist for both direct-diode and fiber laser wavelengths. Figure 

3.98 shows the absorption versus wavelength over a 20-km horizontal path at sea level and at 3 km. 

(millimeter-wave or far-infrared (10.6 m) sources are much more strongly affected by water vapor in the 

lower atmosphere, so those beam sources must be at high altitude.) Note that the propagation model used 

for the COMPASS trajectory modeling was a simple geometric model and did not model the effects of 

water vapor. 

Placing the launch site at high altitude would slightly reduce air drag, gravity loss, and the impact of 

ambient pressure on thrust, but those are minor effects for altitudes below several kilometers. Very high 

elevation sites (both laser and vehicle launch) may have greater availability because they are above lowaltitude cloud layers and precipitation. 

Figure 3.98.—Atmospheric transmission versus wavelength for a 20-km horizontal path at 

midlatitude during winter (molecular absorption/scattering only; no particulates). 

3.1.5.2.5.8 Line of Sight 

The boost laser array requires a line of sight to the launch site, and does not need to track the vehicle 

more than slightly past zenith, which implies that the optimum site for the boost array may be on a slope 

facing the launch site. The main laser requires visibility to 10 to 20 above the geometric horizon along 

the trajectory axis, which may rule out some sites adjacent to high ridges. 

3.1.5.2.5.9 Particulates and Turbulence 

Particulates will affect scattering losses and also will vary with location and season. The largest effect 

will be on the boost-laser to launch-site path, which is roughly parallel to the ground; the boost laser may 

have a slightly shorter wavelength than the main laser (0.8 to 1.0 m versus 1.03 to 1.07 m), which will 

make it more sensitive to scattering because scattering losses generally vary with –4. For example, a 

standard value for particulate scattering loss between an altitude of 0 and 1 km is 0.06 km–1 at 860 nm and 

0.045 km–1 at 1060 nm (Ref. 37). It may be necessary for the boost laser, the launch pad, or both to be 

elevated at least somewhat above the intervening terrain. High scattering losses may rule out some 
NASA/TM—2012-217014 254 

otherwise appealing sites. More generally, scattering should be evaluated at prospective sites and factored 

into the overall system cost. This study assumed scattering losses of 10% (0.9 transmission) at zenith, 

which is good but not exceptional for high-altitude sites. Poor sites may have zenith losses as large as 

50%, which would more than double laser power requirements. 

Adaptive optics requirements for the main laser modules will depend on the levels of atmospheric 

turbulence, known in astronomy as the “seeing,” at and above the laser site. Seeing can vary considerably 

depending on local wind patterns and on season and time of day. Thus, there is some advantage to 

choosing a site with better-than-average seeing. If further analysis of adaptive optics requirements shows 

that the system can operate without Rayleigh beacons, or even without adaptive optics entirely, for some 

specified turbulence parameter r0, then there may be a significant advantage to finding a site that has low 

turbulence (large r0) most or all of the time. 

Near-ground turbulence due to thermal gradients can severely degrade seeing and make horizontal 

beam paths within a few meters of the ground unusable even over distances as short as a few hundred 

meters, so some minimal elevation (on the order of 10 m) of the launch stand and/or the boost laser will 

probably be needed if they are on flat ground. 

3.1.5.2.5.10 Access, Infrastructure, and Resources 

All sites will require road access for construction equipment (graders, concrete mixers, etc.) and 

heavy trucks delivering beam modules and other equipment. Traffic to the laser sites will be minimal 

during normal operation: primarily fuel (if delivered by truck) and beam modules or other equipment 

leaving the site for maintenance or being replaced. 

The launch site will need large-truck access during normal operation for vehicle delivery, unless 

vehicle fabrication is done onsite, simply because each vehicle will fill a standard trailer. The launch site 

should also have, or have access to, a runway and facilities able to handle at least medium cargo aircraft, 

for air delivery of vehicles and payloads. (Air delivery would also accommodate future larger vehicles or 

vehicle components difficult to deliver by truck.) Rail and ship access will not be required. 

Both laser sites will require either low- to medium-voltage transmission lines (e.g. 26 or 138 kV) if 

they do not derive power from onsite generators. (If they do not use energy storage on-site, a highervoltage transmission line would be needed to deliver ~500 MW). Total energy use will be low enough (at 

~1000 launches/year) that truck delivery of generator fuel would be acceptable if generators are used, 

although a pipeline would be preferable. The launch site will not have substantial power or fuel 

requirements unless hydrogen is generated on-site. 

All sites will require a water supply, for cooling at the laser site and for noise and fire suppression at 

the launch site. The quantities needed are small by industrial standards—on the order of 1000 m3

/day—

but water supply may be an issue for desert or mountaintop sites. 

3.1.5.3 CONOPS—Frequency of Launch and Repeat time 

3.1.5.3.1 CONOPS 

3.1.5.3.1.11 Prelaunch 

Vehicles will be delivered from one or more manufacturing facilities to the launch site, either 

complete (preferred) or broken down for ease of transport with final assembly onsite. HXs or complete 

vehicles would typically be given a final integrated test on a static test stand or on a launch stand (with 

hold-downs), using the boost laser array for power. Note that the HX vehicle will able to be fully tested 

on the pad, with the exception of one-shot (pyro or similar) mechanisms for releasing drop tanks or 

shrouds. 

Most launches will be scheduled weeks or months in advance, allowing vehicle deliveries to be 

prescheduled—for example, for a few days before launch (which, at several launches per day, would still 

imply 10 or more vehicles onsite at a time). One or more vehicles (and possibly even some payloads) may 
NASA/TM—2012-217014 255 

be stored onsite, however, to allow for short-notice (possibly <24-hr notice) launches or to replace 

vehicles that fail onsite testing. Launches would be scheduled within a given day based on 

 Orbit requirements—Rendezvous with objects in orbit or injection into specific orbital planes will 

involve launch windows of a few minutes, usually twice per day. 

 Priority/time criticality—Some launches may need to be timed to fixed external events (eclipses) 

or unpredictable short-duration events (solar flares, military missions, etc.). 

 Predicted weather—Non-time-critical launches would presumably always have multiple launch 

dates in case of weather outages. 

 Safety or operational restrictions—Although the launch system will generally not damage 

satellites exposed to the beam, such exposures are obviously undesirable, and coordination with 

other space activities will be needed. 

Many details of the prelaunch operations are still undetermined: for example, whether vehicles are 

moved to the launch stands upright or horizontally, and whether payloads are loaded before or after 

vehicles are moved to a launch stand. Given the expected launch rate (up to several vehicles per day), 

vehicles would typically move to a launch stand less than a day before launch and would be unloaded and 

moved back to a storage/preparation area if a launch was aborted because of weather or technical 

problems. 

Dense propellant could be loaded any time prior to launch. Liquid hydrogen would typically be 

loaded shortly (<1 hr) before launch to minimize boiloff. A swing-away jacket or housing would cover 

most of the vehicle, except the HX itself and the nozzle to provide additional insulation for the hydrogen 

tank. During and after hydrogen loading, the jacket would have dry gas flowed through it to prevent ice 

formation on the tank. (Depending on the tank insulation and coating, this may need to be hydrogen or 

helium to prevent cryopumping nitrogen or oxygen through the tank’s own insulation.) 

Safety and weather monitors would give the final approval for launch. 

3.1.5.3.1.12 Launch 

The general CONOPS for a launch is shown in Figure 3.86. A nominal timeline follows: 

T – 30 s.—The boost laser modules aim at the vehicle on the launch stand and begin tracking. Lowpressure hydrogen gas flows through the pump and HX inlet manifold to cool them to operating 

temperature. 

T – 10 s.—The boost laser array turns on at low power (~1% to 10% power), preheating the HX, and 

a small amount of hydrogen flows to bring the hot propulsion components up to temperature and provide 

an initial hot gas supply for the piston pump or turbopump. This will also provide an opportunity to check 

the pointing of the boost laser array. 

T – 2 s.—The boost laser turns on at full power, and hydrogen is pumped through the HX. 

Liftoff.—Dense propellant flow is turned on, and the vehicle lifts off and accelerates vertically at 

~0.5g acceleration (5 m/s2

) 

T + 40 s.—The vehicle reaches ~Mach 0.9. Dense propellant flow is turned off temporarily to reduce 

thrust and keep the vehicle subsonic. This reduces the peak dynamic pressure and avoids wasting 

propellant on air drag. (Dynamic pressure limits were not included in the COMPASS team’s OTIS 

trajectory calculations, so this throttling was not performed, and the rest of the timeline is somewhat 

compressed). 

T + 80 s.—The vehicle reaches an altitude of ~20 km; dense propellant flow turns back on, and 

vehicle goes supersonic. The vehicle is now high enough for the main laser array to acquire the vehicle 

beacon and begin tracking, although atmospheric absorption and turbulence effects are still too high for 

efficient power transmission. Depending on various trades (power cost, laser life, and atmospheric 

properties) the main laser may switch on now or wait until the vehicle is higher. 

If drop tanks are used, they would be dropped into a disposal zone close to the launch site. The DRM 

vehicle as defined by the COMPASS team will use two drop tanks released at T ~ 75 s, while still 
NASA/TM—2012-217014 256 

accelerating nearly straight up, which would put their nominal impact point very close to the launch site. 

A small sideways or downrange velocity increment would ensure that they fall a safe distance from the 

actual launch facility. It may be better to use the dense propellant somewhat more slowly, both for 

dynamic pressure control and to optimize the trajectory. In this case, the tanks could be released later and 

fall farther from the launch site, although still landing in a small disposal area. The tanks will be small 

and, when empty, lightweight: under 70 kg in the COMPASS design, which for simplicity uses a pressure 

feed system built into the tanks. (Their mass would be even less if they were low-pressure tanks feeding 

propellant to a pump.) The residual propellant will be nontoxic and minimally hazardous. We thus expect 

that the tanks can safely free fall into the disposal area, without needing any recovery system, but this will 

need to be considered in future range safety analyses if drop tanks are used. 

T + 100 to 400 s.—The vehicle flies a conventional trajectory, approximating a gravity turn, although 

the turn starts at somewhat higher altitude than typical for conventional rockets, giving a “lofted” 

trajectory and keeping the vehicle positioned and oriented so that the boost laser has a reasonable 

incidence angle on the HX. Power comes mainly from the boost laser array until roughly T + 200 s and an 

altitude of 100 km. After that, both lasers provide power (with the main laser power adjusted to keep the 

total below the HX’s operating limits) until, somewhere between 300 and 400 s after launch, the boost 

laser beam has diverged far enough that it is not delivering significant power, and the boost array 

shuts off. 

The lofted trajectory will serve two functions: 

(1) It will keep the vehicle well above the horizon, as seen from the main laser site, out to the 

effective range of the laser array. Although some power can be delivered until the vehicle actually goes 

below the laser horizon, scattering losses and beamspread due to turbulence (or due to reaching the 

correction limits of practical adaptive optics) increase rapidly as the vehicle gets close to the horizon. 

(2) It will allow the vehicle to release its payload directly into a stable orbit without a separate 

circularization burn. For the DRM system, the target orbit was a 400-km circular orbit, but payloads can 

reasonably be injected directly into essentially any circular or mildly elliptical orbit that has 

apogee > 300 km or perigee < 500 km and an accessible inclination (see Section 3.3.3.3—Accessible 

Orbits). This will eliminate the cost and complexity of keeping the vehicle stable and powered for a halforbit, or equipping the payload with propulsion and attitude control for the circularization burn. In many 

cases, the circularization burn is the only maneuver required by a satellite, so injecting directly into the 

final orbit eliminates an entire propulsion subsystem. 

Even if the payload cannot be injected directly into its final orbit, a stable initial orbit with a 

reasonable lifetime allows the payload to begin operations—for example, deploy solar panels or initialize 

attitude control systems—before maneuvering, and to use low-thrust propulsion (solar- or beam-powered) 

to reach its final orbit, if desired; this is generally not feasible without a circularization burn if the initial 

perigee is below ~200 km. 

T + 400 to 600 s.—The vehicle accelerates horizontally, powered by the main laser, until it reaches 

nearly its orbital injection altitude. 

T + 600 to 720 s.—The vehicle pitches increasingly nose-down, typically to between 10 and 30

below horizontal at burnout. This again serves two purposes: 

(1) The lofted trajectory tends to give the vehicle some excess upward velocity, on the order of 

100 m/s, which must be cancelled to bring the trajectory tangential to the final orbit. The nose-down 

maneuver provides a small downward component to the final part of the acceleration, with minimal 

impact on the total launch V. 

(2) The downward pitch increases the angle between the vehicle axis and the laser beam, increasing 

the projected area of the HX and thus the collected power. 

T + 720 s.—Orbit insertion and main laser shutdown—The main laser beam modules begin to slew 

telescopes back toward launch site. 

T + 900 s.—The vehicle opens its shroud (if it has not already been jettisoned) and releases the 

payload. The vehicle then aims its nozzle forward and vents residual hydrogen through the propulsion 

system, both to lower its orbit and to ensure that it will not rupture the propellant tank. It may further 
NASA/TM—2012-217014 257 

rotate to a stable high-drag orientation to minimize the time before its orbit decays and it reenters. 

Alternatively, the vehicle can stay in orbit with the payload attached. For example, this might be preferred 

for a short-duration mission that could save cost by using the vehicle’s electronics and attitude control 

system. 

3.1.5.3.1.13 Reentry 

Because the empty vehicle has a very low areal density (low beta), it will not necessarily burn up on 

reentry, and there may be a debris hazard if it reenters over land. If necessary, vehicles can be configured 

with enough residual propellant and power to control the approximate time and general area of reentry. If 

the vehicle is oriented with the HX toward Earth (or even better, toward the Sun), it can be kept at ~300 K 

or higher, so that residual hydrogen gas can deliver impulse at an Isp of ~300 s. Alternatively, a small laser 

array—even a single beam module—can raise the HX temperature for low-thrust maneuvers that are 

regularly done over a fixed location. 

3.1.5.3.2 Launch Rate 

As noted previously, the launch facility would have at least two launch stands, preferably more, so 

that multiple vehicles could be prepared for launch at once (and also so that a problem with one vehicle or 

launch stand would not shut down operations). 

The minimum time between launches would then be limited by the time needed to complete the first 

launch and move the laser array optics back to their starting positions. The boost laser array will typically 

operate for ~400 s, and could certainly reset in <100 s. The main laser array will operate from (nominally) 

T + 80 s for start of tracking through T + 720 s. If it can slew back to the starting position in <80 s, the 

next vehicle could launch at T + 720 s (and be acquired by the main laser array at T + 800 s), 

corresponding to 12 min between launches, sustainable as long as there is sufficient power and cooling 

for the lasers. 

If the launch rate was critical (e.g., for launching multiple vehicles to a single orbital plane), the main 

laser array could begin tracking later and could cut off slightly earlier, since it would deliver relatively 

little power at the extremes of the trajectory. It may also be designed for faster slewing. Time between 

launches might thus be reduced to 10 min or less, with slightly reduced payload. 

If the system is not designed for continuous rapid launch, then once a launch or burst of launches is 

complete, there would need to be a delay to recharge the energy storage system and cool down the 

thermal sink, as well as to move new vehicles onto the launch stands. For example, with energy storage 

and heat sink capacity for two launches, and an 8-hr recharge time, a system could maintain a rate of one 

launch every 4 hr, or a pair of launches 10 min apart once every 8 hr, until either personnel or launchready vehicles and payloads were exhausted. 

3.1.5.3.3 Accessible Orbit 

As shown in Figure 3.99, a single launch site and main laser array define a preferred launch azimuth 

by their relative positions, but small deviations from the preferred azimuth do not have much effect on the 

system performance. A 15 deviation reduces the acceleration path by only 2%. Larger deviations will 

incur progressively larger performance penalties as the effective acceleration path gets shorter, but a 

single continental-U.S. launch site could easily access orbits from ~30 inclined (straight eastward launch 

from 30 latitude) to at least 60 inclined. 

The accessible azimuth range might also be limited by range safety concerns, at least until the 

expected low probability and modest impact of launch failures was demonstrated and accepted. Launch 

trajectories with impact-point paths that cross major population centers might never be permitted. 

Similarly, impact-point paths that cross other countries or their territorial waters will require appropriate 

international agreements. 

Individual cities or smaller hazard zones could be avoided by allowing small doglegs in the trajectory, 

as also shown in Figure 3.101. If done while the vehicle horizontal velocity was low, these doglegs would 
NASA/TM—2012-217014 258 

cost very little V, although a full analysis is needed to see how much deflection is feasible. Doglegs 

would also allow two or more vehicles to launch to the same inclined orbital plane in a single launch 

window, and allow azimuth-offset trajectories to pass closer to the laser array, potentially making up for 

the extra V needed. 

The main laser array will be the most expensive part of the launch system, so if a wider range of 

azimuths is needed, it may be possible to build two or more launch sites and boost laser sites that share a 

single main laser site (e.g., one west of the main array for equatorial launches and one north or south for 

polar launches). This does require the main laser array beam modules to point over a wide range of 

angles, which may raise the cost of the telescope mounts and the land area needed for the array (so that 

lines of sight are not blocked by neighboring beam modules). 

Figure 3.99.—Variable launch azimuth and dogleg trajectory offset capability with a single launch site and laser site. 

The COMPASS DRM analysis considered only eastward launches, and there will be some reduction 

in payload for polar or retrograde launches. However, because of the high Isp of the propulsion system, the 

effect will be small in comparison to the effect on a single- or even two-stage chemical rocket. 

A partly or wholly sea-based launch system would allow launches to (almost) any azimuth, as well as 

having advantages for range safety, minimum orbit inclination, and so on. A partially sea-based system 

would put the vehicle launcher and boost laser array on ships, with the main laser array on land (possibly 

an island). Sea-based laser launch is not ruled out by atmospheric transmission (although some laser 

wavelengths would experience higher losses) or by any other fundamental issues we know of. However, 

the operating and maintenance costs would be substantially higher than for a land-based system. 

Similarly, an air-dropped vehicle using chemical strap-ons to climb into view of the main laser array 

could be launched from any point, and thus to any azimuth allowed by safety limits and array pointing 

capability. 

3.1.5.3.3.14 Geosynchronous Transfer and Upper Stage Capability 

The DRM vehicle does not have sufficient single-stage delta-V for direct injection of payloads into 

geosynchronous transfer orbit (GTO), but adding a solid- or liquid-fuel upper stage (using roughly 50% 

of the normal payload mass) would allow launching small payloads to GTO. Because the laser launch 

mass scales rapidly with final velocity (roughly as vf

-3) using such a small upper stage can also allow 
NASA/TM—2012-217014 259 

launching significantly (~2 times) heavier individual payloads to LEO, although presumably at higher 

marginal launch cost per kilogram. This is similar to chemical single-stage-to-orbit (SSTO) designs, 

which can deliver much larger payloads to LEO if operated as “nearly SSTO” suborbital vehicles carrying 

payloads with attached kick stages. 

3.1.5.3.4 Availability 

Any ground-based laser (GBL) (or millimeter-wave) beam source will be blocked by any visible 

cloud, and may even be unacceptably attenuated by subvisible clouds. Laser sources may also have 

problems with unusually turbulent air or high ground-level winds. Selected sites in the southwest United 

States can have acceptable conditions for laser launch nearly continuously for large parts of the year, and 

on average for 70% to 80% of the year. For applications where operations can be limited during some 

parts of the year, or where individual launches can accept delays of up to several days, a single wellchosen launch site/boost laser array and a single main laser site can probably provide acceptable 

availability. 

A second launch site and boost laser array (sharing a single main laser array) will slightly increase 

availability if the two launch sites are some distance apart, since one may be clear when the other is 

blocked, but unless the two launch sites serve different orbits (e.g., polar vs. low inclination), the weather 

at two sites is likely to be correlated. 

Substantially higher probability of access is possible if two (or more) complete launch systems are 

built, far enough apart to have relatively uncorrelated local weather. In this case, the chances of both 

being unavailable at any given time can be as low as ~10%, with the chance of an extended (e.g., >24 hr) 

outage being even lower. A very-high-availability launch system (e.g., for military use) would thus 

require at least two fully independent launch systems, and possibly three or more. A multisite system 

might have each site optimized for one type of launch (most obviously polar vs. low inclination) with 

“backup” capability for other types of launches: for example, using chemical boosters rather than a boost 

laser array. The backup capability would be used only for time-critical launches when the corresponding 

primary site was down because of weather or other problems. An alternative for high availability would 

be a ship-based launch system, which could move to clear areas (typically equatorial) and move out of the 

paths of storms. 

Finally, it is possible that, at the power levels of interest, a suitable laser array could actually clear at 

least light cloud cover. Unfortunately, the main and boost laser array wavelengths will be chosen to 

minimize water absorption, so they could not do their own cloud clearing. Most cloud-clearing 

experiments have used 10.6-m carbon dioxide (CO2) lasers, but other wavelengths may work. 

The modular laser array will allow very close to 100% availability with respect to hardware failures, 

by providing a modest number of “spare” modules (and associated energy storage and cooling capacity). 

Redundant prime power (e.g., grid tie plus local generators) would eliminate downtime due to power 

failures. With appropriate design (e.g., duplicate control facilities) a laser launch system could be 

essentially immune to single-point failure or physical damage, including (nonnuclear) attack. 

3.1.5.4 Costs 

3.1.5.4.1 Vehicle 

3.1.5.4.1.15 General Comments on the Cost of HX Vehicles 

“If you do the same things you’ve always done, you’ll get the same results you’ve always gotten” 

-- Jeff Greason, CEO, XCOR Aerospace 

In theory, a chemical rocket able to launch payloads to Earth orbit should be simple and cheap to 

build. In practice, many attempts to build “cheap” launch vehicles have either failed (OTRAG, Beal, 
NASA/TM—2012-217014 260 

Rotary Rocket, Kistler, etc.) or have succeeded in building launchers only modestly cheaper than previous 

ones, at a higher than originally projected cost (Orbital Sciences and SpaceX). 

If expendable HX vehicles are built with the same programmatic approaches as past launch vehicles, 

they will be expensive, although marginally less so (by perhaps a factor of 2) than existing small 

expendables. They do have the advantage of being single-stage (or nearly so) versus at-least-two stage 

chemical rockets. However, single-stage expendable chemical rockets are feasible, and many have been 

proposed. HX vehicles are simple, but there have been many designs for equally simple chemical rockets, 

including solid (no valves), hybrid (one valve per stage), and pressure-fed liquid (two valves per stage) 

rockets. HX vehicles have better mass ratios than chemical launch vehicles—higher dry mass/gross liftoff 

weight (GLOW) than single-stage launchers and higher payload mass/dry mass than two-stage 

launchers—but not overwhelmingly so. They will still require advanced materials and lightweight 

construction. 

The true advantage of HX vehicles is that they can be built like other industrial products, not like 

aerospace vehicles. They would be built in sufficient quantities to support assembly-line and massproduction techniques wholly unfeasible for production rates of a few per year, and also in quantities that 

would support development and production by two or more competing suppliers. Production quantities 

are sufficient, and testing costs low enough, to support development and flight testing of new components 

and subsystems rather than depending on heritage hardware wherever possible. 

HX vehicles can be fully integrated and then tested on the ground before flight (unlike multistage 

vehicles and solid rockets). Components and subsystems can be certified and then built to print, as with 

most aircraft parts. They should not require extensive—and expensive—testing of each individual 

component and subsystem for each vehicle. Indeed, since launch failures will pose only modest financial 

risk (given the small vehicle size and ease of relaunching payloads) and no risk of loss of life, the 

certification and component tracking needed should be substantially less than for commercial or military 

aircraft. 

Cars cost ~$20/kg. Exotic, handbuilt cars, built in smaller numbers than we propose for HX launch 

vehicles, cost $200/kg to $500/kg. Light piston aircraft, which do kill people when they fail, cost 

~$500/kg.5

 Tomahawk Block IV cruise missiles have a dry weight (less fuel and explosive) of roughly 

800 kg; the 2004 Block IV production contract was for 2200 missiles at a unit cost of $730,000. Since the 

unit cost presumably includes somewhat more than the actual missile (at a minimum, its solid booster) the 

specific cost is less than $900/kg (Ref. 39). 

The author of this section is not in a position to do a detailed vehicle cost estimate for this report. The 

cost of the HX in particular is very difficult to estimate, since neither the final design nor the fabrication 

approach is known. However, unless the HX vehicle has some specific component (such as the HX) that 

is extremely costly, the production cost at quantities on the order of 1000/yr should fall in this range: 

$1000/kg worst case and $500/kg nominal, with $200/kg a reasonable goal. The corresponding complete 

vehicle cost would be nominally $530 K (dry mass of 1060 kg, including drop tanks, at $500/kg), with a 

worst case of $1060 K and a goal (with no drop tanks and a slightly more optimistic vehicle dry mass of 

800 kg) of $160 K. 

3.1.5.4.1.16 COMPASS Cost Estimate Versus Past Launch Vehicles 

The COMPASS cost estimate of ~$12 million per vehicle for the DRM–1B vehicle is extremely high, 

and inconsistent with all prior launch vehicle experience, both government and commercial. This appears 

to be an artifact of using spacecraft costing rules. 

A very quick analysis of a few existing expendable launch vehicles is given in Table 3.21. (Much 

better and more extensive data are available, but locating, compiling, and analyzing it was outside this 

author’s scope of work.) The general trend is evident even from this limited data set: launch vehicles tend 

 

5

For example, the Cessna 400 (Corvalis TT) composite aircraft, which was introduced in 2004—empty weight 

of 1134 kg, list price of $644,500 (Ref. 38). 
NASA/TM—2012-217014 261 

to cost $3000/kg to $6000/kg dry mass, with vehicles developed primarily for commercial launches on 

the low end of the range. (Ariane 5 is unusually low because ~80% of its dry mass is solid booster 

casings). The Falcon 1E and Delta IV Heavy differ by a factor of almost 30 in dry mass, but they have 

similar cost per kilogram. All of these vehicles were built in quantities of 10s, although generally with 

considerable variation so that any given configuration might launch only a few times. Note that this is the 

launch price, not the actual cost of manufacture, and it includes the vehicle provider’s development cost 

amortization and profit. 

At the Falcon 1E cost of $3250/kg, the DRM–1B vehicle (non-drop-tank configuration, dry mass of 

920 kg) would cost $3.0 million each. If we assume that this is for 20 vehicles, scaling to ~3000 vehicles 

(3 years of 1000/yr) would imply 7 doublings of production, and, with a 0.85 learning curve, a marginal 

vehicle cost of $3.0 million  0.857

 = $0.96 million, roughly $1050/kg. 

We thus estimate that the COMPASS cost estimate is about tenfold too high, even in comparison to 

existing launch vehicle production techniques and procedures. With all respect to the efforts of the 

COMPASS team in developing a cost estimate under difficult time and resource constraints, this author 

must recommend that the COMPASS cost estimates be considered to be invalid. 

(As a check on the preceding cost estimate, we note that the Space-X Falcon 1E is a two-stage launch 

vehicle with each stage comparable to the laser-HX vehicle in complexity. A firm launch price of $9.1M 

was quoted in 2008, so the vehicle itself presumably cost significantly less. It is difficult to understand 

how the HX vehicle could cost more than the Falcon 1E even in small quantities, unless the HX itself is 

exorbitantly expensive.) 

TABLE 3.21.—LAUNCH VEHICLE COST PER KILOGRAM DRY MASS 

Launch vehicle Dry mass Cost,a

$M 

Cost per mass,a

$/kg 

Falcon 1E 3419 11.1 3252 

Pegasus XL 3705 18.1 4891 

Ariane 4L 32015 131 4092 

Ariane 5G 97000 236 2431 

Delta IV heavy 85770 302 3524 

Saturn V 181965 1110 6100 

a

2011 dollars. 

3.1.5.4.2 Ground System Capital Cost 

3.1.5.4.2.17 Laser Costs and Cost Trends 

Diode lasers.—The ground system costs will almost certainly be dominated by the cost of the lasers 

themselves. Projecting laser costs is difficult because there is relatively little published data on highpower laser prices, and the total market for high-power lasers is small (compared to, e.g., computer 

systems) and highly diverse. 

For semiconductor diode lasers, prices for small quantities of packaged high-power multimode diodes 

or diode bars have been roughly stable for the past decade at ~$10/W (then-year dollars, so decreasing 

slowly in constant dollars), but the performance of the devices has improved considerably; in particular, 

the lifetime of diode bar stacks has more than doubled, from <10 000 to >20 000 hr. There does not seem 

to be any reason for this to be a limit, however. Figure 3.100 shows the decline in laser diode bar prices 

with quantity estimated by Lawrence Livermore National Laboratory (LLNL, Ref. 40) in connection with 

inertial confinement fusion laser designs. These would be very similar to the diode lasers needed for 

either direct use (boost laser) or pumping fiber or other solid-state lasers for a launch system, except for 

operating pulsed (which allows somewhat higher peak power per diode). 

A launch system boost laser array would require ~2106

 diode bars at 100 W/bar; this is similar to the 

LLNL Mercury laser project, for which LLNL was quoted $0.35/W several years ago. Even if we assume 
NASA/TM—2012-217014 262 

a factor of 3 for diode power and packaging differences for CW versus pulsed lasers, this would put the 

cost of the boost laser array diodes at ~$1/W, with no further progress in diode laser technology. 

Figure 3.100.—Lawrence Livermore National Laboratory Laser diode cost estimates. 

Although we were unable to find good price-versus-time data for laser diode bars, Figure 3.101 

(Ref. 41) shows data for a reasonably good proxy: red light-emitting diodes (LEDs). This plot illustrates 

Haitz’s Law, analogous to Moore’s Law, which states that LED price per lumen falls 10 times per decade 

(~20% per year), whereas the performance, in lumens per device, increases 20 times per decade. This has 

been driven in part by the adoption of LEDs for mass-market applications requiring high power, such as 

traffic signals and automobile taillights, but also by steady improvements in fabrication and packaging 

technology. If we assume that there will be comparable improvements in high-power diodes (i.e., a 

resumption of the “normal” development curve, possibly stimulated by the prospect of a substantial new 

market) the 2021 price would be ~$1/W. 

As a worst-case estimate on a simple quantity-discount basis, if the “price break” for laser diode bars 

is 20% for each factor of 10 in quantity, the current price for ~2 million bars would be about 40% of the 

small-quantity (100-bar) price, or $4/W. 

An unusual example of price variation with quantity for SOA photonic devices is available online: 

Sensor Electronic Technology publishes a price list for their ultraviolet (240- to 366-nm) LEDs for 

quantities from 1 to 10 through 1 000 000 units (Ref. 42). Figure 3.102 shows prices for two typical 

devices (255 and 345 nm) in two packages (flat window and ball lens). Note that in all cases, there is a 

sharp break in the slope of price versus quantity at a few thousand units. For flat-window devices (where 

the packaging cost is low) the ratio between the 100-unit and million-unit price is roughly 20:1. Although 

we would not expect that large a ratio for commodity laser diodes or diode bars, it would be quite 

appropriate for current low-production devices such as single-mode or frequency-locked high-power 

diodes, or unusual wavelengths. For ball-lens packaging, the cost of the lens and alignment drops much 

more slowly with quantity, from a ~$30/unit premium over a flat window for small quantities to $5/unit to 

$8/unit for a million units. This is consistent with an overall drop in price where packaging is a dominant 

cost of a factor of 4 to 6 over 6 decades, or about 25% to 30% per decade, but with most of the drop at 

higher quantities. 
NASA/TM—2012-217014 263 

We, therefore, have four projections: ~$1/W, ~$1/W, ~ $2/W (if packaging dominates), and a worst 

case of ~$4/W. There would be an additional amount for wavelength- and polarization-combining optics 

(currently ~$1/W for small quantities. This also would presumably drop by ~fourfold for million-unit 

quantities). For system cost estimates, we assumed $2/W for fiber pumps and $1.50/W for boost laser 

diode arrays. 

(With current assembly techniques, fiber-coupled laser diodes and diode arrays are substantially more 

expensive than “free-space” diodes, in the range of $25/W to $50/W retail. This contributes to the cost of 

fiber laser designs using fiber-coupled pump diodes (IPG and others). Fiber-coupled laser diodes also 

have somewhat lower radiance than free-space diodes. The baseline boost-laser beam module concept 

thus does not use fiber coupling but mounts the diodes directly on the telescopes.) 

Figure 3.101.—Haitz’s Law for light-emitting diode (LED) price and performance versus time. 
NASA/TM—2012-217014 264 

Figure 3.102.—Example of scaling ultraviolet light-emitting diode (LED) prices 

over 6 decades of quantity. 

Fiber lasers.—Data on fiber lasers are even more limited than on high-power diodes. Current costs 

for high-power (>1-kW) multimode fiber lasers, which are manufactured in small quantities, are 

estimated at $40/W to $50/W (Ref. 43). Discussions with IPG suggest moderately higher real prices for 

their large systems, ~$75/W, but with projected prices of $35/W for a large order circa 2013. 

Single-mode fiber lasers (and indeed all types of single-mode lasers) are somewhat more expensive 

than comparable multimode lasers. The IPG YLS–10000–SM 10-kW fiber laser, although it is proof of 

technology readiness, is a relatively high-cost design and not a good cost data point. One source estimates 

a cost of $1000/W for the prototype (Ref. 43). We do not have a good reference for kilowatt-class singlemode fiber lasers but have recently been quoted $135/W for a smaller (100-W) fiber laser for 100 units 

(i.e., 10 kW). We take the current cost for single-mode lasers as between 1.5 and 2 times the cost of 

multimode (within the power ranges where both are available). 

Fiber lasers have generally scaled in price with the price of their pump lasers, with multimode fiber 

lasers costing ~3 times the pump laser cost. However, the production volume for high-power fiber lasers 

is very low, ~$500M/year for all fiber lasers (Ref. 44), or ~5 MW of laser power, only a fraction of which 

is for kilowatt-class lasers. Kare and Parkin (Ref. 26) estimated $10.50 for fiber lasers in the 150-MW 

quantity, assuming $100/W for a single 10-kW laser (versus actual 2005 prices of $200/W to 500/W) and 

an 85% learning curve. This is consistent with extrapolated pump diode pricing ($2/W  3 for single 

mode and  5 for multimode). 

Fiber laser prices appear to have dropped at least threefold to fourfold in 6 years (since 2005), which 

is consistent with a tenfold/10 year drop similar to Haitz’s law for LEDs. This would give a price of 

~$10/W to $15/W for single-mode lasers in small quantities by 2021. Alternatively, a ~40-times increase 

in total production, from ~5 to ~200 MW, would give a sevenfold price drop for a 70% learning curve or 

a fifteenfold drop for the 60% learning curve estimated by LLNL for laser diodes. 

We therefore still project a price for single-mode fiber lasers of $10/W for quantities from ~150 to 

200 MW, circa 2021, with the caveat that this could easily be in error by a factor of 2 in either direction. 
NASA/TM—2012-217014 265 

A comment on lasers versus millimeter-wave sources.—The author of this section believes that 

lasers are cost competitive with millimeter-wave gyrotrons on a system basis, even though a naïve look at 

current prices ($1 million buys a 1-MW gyrotron tube, or a 10-kW fiber laser) would suggest that lasers 

have an insurmountable price disadvantage. There are many reasons for this. Partly, this 100:1 ratio is not 

a valid comparison: 

 The preceding comparison is between a bare tube and a complete industrial laser system (the 

correspondingly complete millimeter-wave system would cost between $2 and $4 million for 

1 MW). 

 The microwave system has a much shorter range, and therefore requires several times more 

power per unit mass to orbit. With equivalent assumptions about vehicle design and component 

masses, the higher power/vehicle mass ratio means that the millimeter-wave vehicle will also 

have a smaller payload (or require even higher power). The alternative of scaling the microwave 

system aperture to give laser-system-equivalent range appears to be more expensive and 

technically challenging. 

 However, there are also issues relating to cost and technology projections: 

 Lasers for a launch system would be built in much larger quantities (>10 000 units) than gyrotron 

systems (a few  100 units), so the learning-curve discount will be much greater; 

 High-power fiber lasers are a very new technology (developed circa 2001) and have multiple 

competing technologies (disk lasers, diode pumped alkali vapor lasers, and grating-combined 

wavelength-stabilized diode arrays). The cost/performance ratio is thus likely to improve much 

faster than for gyrotrons. 

 Kilowatt-class lasers have a substantial industrial market and make use of technologies (optical 

fibers and laser diodes) that have very large markets and large R&D communities. Megawattclass gyrotrons currently have a single specialized market (heating fusion plasmas) and use 

technologies from a specialized field (high-power microwave tubes). 

 Laser costs are still dominated by the cost of the laser hardware itself and are far from being 

limited by mature supporting technologies (power supplies, connectors, and structural supports). 

millimeter-wave gyrotrons are already approaching that limit and thus are less likely to see 

substantial system-level cost reductions. 

Some of these aspects have not yet been researched and were not within the scope of this study, but 

they should be addressed (e.g., through analysis of development funding and cost histories for these and 

comparable past technologies) before any decisions on source technology are made. 

3.1.5.4.2.18 Main Laser Array Cost 

A rough capital cost estimate for the main laser site is given in Table 3.22and Table 3.23. Table 3.22 

gives the cost of a single beam module, and Table 3.23 gives the overall site cost, assuming 2500 beam 

modules; 2500 beam modules would give 150 MW of laser output power, allowing for 20% total losses 

between the laser output and the maximum vehicle absorbed power. 
NASA/TM—2012-217014 266 

TABLE 3.22.—CAPITAL COST FOR A NOMINAL 

60-KW MAIN-LASER BEAM MODULE 

Component Cost, 

$K 

Basis 

Single-mode fiber lasers 600 6  10 kW at $10/W 

Telescopes 120 6  40 cm at $20K 

 Telescope mount 60 

 Pointing and tracking 20 

Adaptive optics 90 6  15Ka

 AO beacon 20 1  30K 

Assembly, transport, and installation 80 

TOTAL $990 K a

Thorlabs AO system ($24K)  2 for a high-power mirror 

 0.32 for an 85% learning curve. 

TABLE 3.23.—CAPITAL COST FOR MAIN LASER SITE WITH 2500 BEAM MODULES 

Element Cost, 

$M 

Basis 

Beam modules Number 2475 2500 at $0.99M each 

 Telescopes, etc. 1500 

 Lasers 500 

 AO and beacon 275 

 Assemble, test, and integrate ( A T & I) 200 

Prime power and cooling 20 20 MW at $1/W 

Energy storage 80 Three launches at ~$26Ma

Control and safety systems 40 

Land and infrastructure 80 800 acres at $100K/acre 

Planning, permitting, etc. 100 

TOTAL $2795 M 

a

See Section 3.1.5.2.2—Main Laser Site Design 

Note that this estimate is for the lowest likely beam module power (60 kW). Several cost elements 

(land, site preparation, and optics and tracking) are roughly proportional to the number of beam modules, 

so higher module power would somewhat reduce costs. 

3.1.5.4.2.19 Boost Laser Array Cost 

An approximate capital cost for the boost laser site is given in Table 3.24 (beam modules) and Table 

3.25(overall site), with the assumption of 1200 beam modules at 200 kW each. This would give 240 MW 

of laser output power. 
NASA/TM—2012-217014 267 

 TABLE 3.24.—CAPITAL COST FOR A NOMINAL 200-KW BOOST-LASER BEAM MODULE 

Component Cost, 

$K 

Basis 

Laser diode arrays 300 200 kW at $1.50/W 

 Power supplies 70 350 kW at $0.20/W 

 Beam-shaping micro-optics 30 200 kW at $0.15/W 

Telescopesa 200 10 m2 at $20 K/m2

 Telescope mount 50 

 Pointing and tracking 20 

Adaptive optics, AO beacon N/Ab 

Assembly, transport, and installation 80 

TOTAL $750 K a

Plate-glass honeycomb or similar low-cost optics. b

Not applicable. 

 TABLE 3.25.—CAPITAL COST FOR BOOST LASER SITE WITH 1200 BEAM MODULES 

Element Number Cost, 

$M 

Basis 

Beam modules 900 1200 at $0.75M each 

 Lasers, etc. 480 

 Telescopes, etc. 324 

 A, T, & I 96 

Prime power and cooling 16 20 MW at $0.80/Wa

Energy storage 72 Three launches at $24Ma

Control and safety systems 40 

Land and infrastructure 80 400 acres at $100K/acre 

Planning, permitting, etc. 100 

TOTAL $1208 M a

Smaller cooling load and slightly lower total energy than main laser array because of higher 

efficiency of laser diodes versus fiber lasers. 

The boost laser array cost is somewhat less dominated by the cost of the lasers themselves, and is thus 

more sensitive to other costs, especially optics. 

3.1.5.4.2.20 Launch Site Cost 

A token cost breakdown for the launch site is given in Table 3.26.The launch site can be a 

comparatively modest facility because of the small scale of the vehicles. The vehicles are roughly the 

same size as existing small launch vehicles (e.g., Minotaur or Pegasus) but much lower in both dry and 

wet mass. A new launch pad and gantry for Minotaur I was built at Wallops Island for approximately 

$10 million. We take this as a typical cost for each of four launch stands. However, except for the cooling 

jacket and hydrogen plumbing, the launch stands will be very simple. 
NASA/TM—2012-217014 268 

TABLE 3.26.—CAPITAL COST FOR LAUNCH SITE 

Element Cost, 

$M 

Basis 

Launch stands 40 Four stands at $10 M each 

Launch control and range safety 80 

Static test stand 20 Launch stand and test equipment 

Vehicle preparation facility 30 12 bays at $2.5M eacha

Hydrogen storage and handling 10 

Land and infrastructure 40 

TOTAL $220 M 

a

~500 m2 per bay, $5000/m2

The largest single cost is assumed to be the launch control facility, including actual launch control 

and communications as well as safety systems (perimeter sensors and radar), optical and radar vehicle 

tracking, prelaunch vehicle monitoring, and vehicle test controls. The number given (for the laser sites as 

well as the launch site) is very much an order-of-magnitude estimate, but we note that the intent is to keep 

the total crew involved in a launch to a small number, on the order of 10, with most of the capital cost 

being in sensors and automation. 

Other launch site infrastructure will depend on how much vehicle assembly and launch preparation is 

done onsite. Ideally, vehicles will be delivered from the vehicle plant, by truck or aircraft, essentially 

launch-ready except for installing the payload and closing the shroud. In keeping with the overall theme 

of high-volume “production-line” operation, there will be few or no high-cost facilities such as 

cleanrooms and very limited onsite storage for payloads or vehicles beyond those being actively prepped 

for launch. 

We assume liquid hydrogen is delivered by truck (roughly one truckload per launch) at a cost of 

~$0.30/kg (Ref. 44) so there will be no onsite hydrogen production or bulk storage. If storage is needed, a 

300 000-kg capacity tank (~100 launches) would be ~$10 million. 

The launch facility cost may be higher if road or (especially) runway construction is required. 

3.1.5.4.2.21 Cost of Other Ground Facilities 

We are not including the vehicle manufacturing facilities in the ground facilities; vehicles are 

presumed to be purchased from one or more vendors. 

One possible large cost is for a full-scale test facility for HX propulsion. If a dedicated full-scale test 

facility is deemed to be necessary, however, we assume that an existing engine test facility could be used 

in conjunction with a short-range laser-diode beam source. “Cheap” low-brightness laser diodes and small 

optics would be sufficient for this, and we estimate the full cost at ~$400 million for a 200-MW source. 

However, we would strongly recommend using the boost laser array and launch facility for full-scale 

testing. This would require only building a suitable thrust test stand with the ability to burn off ~10 kg/s 

of hydrogen exhaust safely. We allocate ~$20 million for this within the launch site. 

3.1.5.4.2.22 Total Capital Cost and Cost Scaling 

The estimate in Table 3.22 to Table 3.26 gives a total system capital cost of approximately $4.2 

billion. This is somewhat higher than the prior estimate by Kare and Parkin (Ref. 26) of $2.85 billion. The 

increase is due mainly to three factors: 

(1) Incorporating the boost laser site, which, at $1.2 billion, is nearly 30% of the total cost: This cost 

is particularly high because of the very high power level—220 MW—used for the COMPASS design 

effort. We expect that an 80- to 100-kg payload will in fact require no more than ~150 MW of boost laser 

power, reducing this cost to $800 million or less. Note that $1.2 billion, amortized over 10 000 launches, 

is $120,000 per launch. If the per-launch cost of using, for example, solid booster rockets for vertical 

ascent to 100 km is substantially less than this. The DRM system would be more economical without the 

boost laser. 
NASA/TM—2012-217014 269 

(2) Including many system-level costs not covered in the earlier estimate: These include land and 

infrastructure, launch site, control and safety systems, plus allowance for assembly, transport, and 

installation of beam modules. (Kare and Parkin included only the actual lasers and optics, plus power and 

cooling.) These costs total over $500 million (excluding the costs relating to the boost laser). They are 

extremely rough estimates and should be considered largely as placeholders for more detailed studies. 

However, they serve to give an idea of the potential costs. 

(3) Including adaptive optics and Rayleigh beacons in the beam modules: This adds ~$250 million, 

mostly for the adaptive optics themselves. This would be lower if the power per beam were higher than 

10 kW. 

Offsetting these costs are a slightly lower laser cost estimate ($10 vs. $10.50) and lower power and 

cooling costs ($100 million vs. $240 million) because of the low launch rate assumed for the DRM 

system. 

3.1.5.4.3 Operating Costs and Cost Per Launch 

The operating and maintenance costs for large laser arrays is not well known. We have seen estimates 

ranging from 2% to 20% of capital costs per year (the high end being the entire operating budget for 

research facilities such as national laboratories). We assume an intermediate value of 5% of capital cost 

for operation and maintenance (O&M) for both laser sites and the launch site, or approximately $200 

million/year. (Note that this does not include depreciation or cost of capital.) Much of this would be for 

routine maintenance of mechanical and electronic hardware, and periodic cleaning of exposed optics. On 

the basis of industrial and telecommunications experience, fiber lasers themselves are extremely reliable, 

with essentially no maintenance needed over the launch system lifetime. 

Marginal launch costs are dominated by the vehicle cost, which we assume (per the discussion in 

Section 3.1.5.4.1—Vehicle) will be on the order of $1 million or less, possibly as low as $200,000. If we 

assume a lifetime of >10 000 hr for both diode and fiber lasers, a launch rate of 1000/yr will use only 

~1% of the laser life per year, so the lasers will not need to be replaced over the life of the launch facility. 

At much higher launch rates, laser replacement costs could become significant. 

Power costs will be essentially negligible. A launch will require roughly 400 MW  0.2 hr = 

80 MW-hr. Even at $0.10/kWh, this would be only $8000. 

Propellant costs will be similarly small. Even at a conservative cost for liquid hydrogen of $5/kg, the 

cost per launch would be $15,830 for the COMPASS design, plus a negligible cost for water/hydrogen 

peroxide or liquid nitrogen. 

Actual vehicle preparation and launch operations are expected to take on the order of 1 to 2 personmonths (e.g., a four-person team per vehicle, working for 1 to 2 weeks) for checkout and payload 

attachment, plus <1 day of a ten-person launch operations team. At $10K/person-month, this would be 

$10,000 to $20,000. 

Thus, if the system capital cost is treated as a sunk cost, but including O&M costs and marginal 

launch costs, the nominal cost per launch is given in Table 3.27. 

TABLE 3.27.—COST PER LAUNCH FOR 1000 LAUNCHES/YEAR, 80-KG PAYLOAD 

Type of cost Cost per launch, 

$K 

Cost per kg, 

$ 

Vehicle 200 to 1000 2500 to 12,500 

O&Ma at 5% of capital cost 200 2500 

Marginal operations cost 20 250 

Propellant 20 250 

Power 8 100 

TOTAL 448 to 1248 5600 to 15,600 a

Operations and maintenance. 
NASA/TM—2012-217014 270 

The costs are driven by vehicle cost and O&M cost. Both can be lowered directly, by lowering the 

vehicle manufacturing cost and the launch system O&M cost. Both can also be lowered substantially by 

assuming a higher launch rate. For example, at 10 000 launches per year, the vehicle cost would be about 

60% as much (for an 85% learning curve), and although the total O&M cost might be higher, the cost per 

launch would be much less. A cost per launch of $88,000, plus a vehicle cost between $120,000 and 

600,000, would be a reasonable goal, making the total cost per launch between $208,000 and $688,000 

($1540/kg to $8600/kg). Even lower launch costs would be attainable with true mass production of 

vehicles at significantly lower unit cost or with the development of reusable vehicles. 

3.1.5.5 Feasibility Assessment 

3.1.5.5.1 Roadmaps and Technology Readiness Levels 

3.1.5.5.1.23 Technology Readiness Levels (TRLs) 

TRLs are, by their nature, somewhat subjective, particularly for intermediate levels. We caution 

against directly comparing TRL levels as estimated by different authors, even within this report. Table 

3.28 gives a breakdown of the nominal TRLs for various aspects of the launch system. 

TABLE 3.28.—TRL LEVELS AND BASIS FOR HX LAUNCH SYSTEMS 

Key technology Current 

TRL 

Rationale and proof of estimation Time to reach TRL 6 

HX 2 to 3 Theory understood, but new operating 

regime requires testing and engineering for 

flight 

18 months for megawatt class

3 to 4 years for full scale

Piston pump 4 XCOR: Bench test with LH2, flight 

operations with LOX/RPa

24 months 

Expansiondeflection nozzle 5 E-D nozzles flown with other propellants; 

new design for LH2 

18 months for megawatt class

3 to 4 years for full scale 

Other vehicle 

components 

4 to 8 Per COMPASS study; lowest TRLs are 

refractory materials and “advanced” 

batteries (LiPO) 

N/A for suborbital tests

2 to 4 years for full scale 

hardware 

Fiber laser 8 / 6 60-kW module: TRL 8—10-kW IPG laser

200- to 500-kW module: TRL 6—IPG laser 

+ Bragg grating combining or JHPSSL

N/A 

Diode array 

(boost laser) 

6 Commercial multi-kilowatt, multiwavelength diode sources (e.g., TRUMPF 

Tru-Diode) 

N/A 

Optics/tracking 6 COTSb telescopes/pointing systems (1-m 

aperture)

ABL exceeds all requirements (at high cost) 

N/A for technology, but 18 to 

24 months for “low-cost” 

hardware demonstration 

Adaptive 

optics/guidestar 

5 ABL and Starfire Optical Range (SOR) 

exceed all requirements for single aperture; 

need to validate multi-aperture system and 

COTS hardware 

18 months for low-power tests

3 to 4 years for high-power 

multi-aperture 

a

Liquid oxygen/rocket propellant. 

b

Commercial off the shelf. 
NASA/TM—2012-217014 271 

Vehicle.—The laminar-flow HX itself is at TRL 2 to 3. To our knowledge, no other device has been 

demonstrated with a similar combination of thermal, gas flow, and optical properties, which would imply 

TRL 2. However, somewhat similar structures have been made in a variety of materials, including test 

laminar-flow HXs in nickel (Ref. 7) and various porous or small-channel ceramic structures for 

biochemical applications. The basic laminar-flow heat transfer technique is well established for liquid, 

gas, and two-phase flow (see, e.g., Ref. 45), and the other physical and engineering phenomena involved 

(optical absorption, reradiation, thermal conduction in solid SiC, and thermal failure modes of SiC) are 

generally well understood. The major technological issues are discussed in Section 3.3.1.2.4 (HX issues). 

 Fortunately, as discussed in the following section, it is straightforward to test the HX design and 

(assuming the tests are successful) reach at least TRL 4. 

Turbulent-flow HXs are more common and have been built in a wide range of high-temperature 

materials and tested with high thermal fluxes applied to one surface (albeit usually via hot gas, rather than 

laser flux), so a turbulent-flow version of the laser HX can be classed as at least TRL 3 and possibly 

TRL 4 or higher. (See Ref. 46 for an example.) 

Other than the HX, the vehicle can be built using standard aerospace materials and technologies, with 

varying amounts of development of actual components and subsystems. Reuse of existing flight-proven 

hardware (TRL 8 to 9) may be possible for some subsystems, but it would not in general be appropriate, 

given the differences between the laser launch system and past launch systems and spacecraft, and given 

that the development and testing of optimized hardware can be amortized over, by launch vehicle 

standards, a very large number of units. For example, existing space-qualified electronics could be used 

for control and data handling, attitude determination, and so forth., but with a penalty of as much as 30 kg 

in mass and $1M or more in cost per launch in comparison to a new design. Similarly, even though an 

existing TRL 9 small turbopump might be adapted to the HX vehicle, a piston pump, currently at TRL 4, 

would almost certainly be much less expensive over hundreds to thousands of launches and would offer 

performance characteristics better matched to the HX propulsion system. 

Laser arrays.—The reference example for the main laser is the IPG YLS-10000-SM, a 10-kW, 

single mode (M2 < 1.2) fiber laser operating at 1070 nm. This is a commercial product and would be 

usable as is in a launch system. Provided the beam modules have environmental control (heating/air 

conditioning) comparable to that in heavy manufacturing plants (e.g., steel fabricators and automobile 

assembly plants), the lasers would be operational in the expected environment6

 and would thus be at 

TRL 8. However, the current IPG laser is almost certainly not the best choice for a launch system. Even if 

we assume the use of only current (2011) laser technology, wavelength-combining of several 1- to 3-kW 

lasers would probably give lower cost and higher efficiency. However, because such wavelengthcombined systems are not in production or regular use, they are only at TRL 6. 

In reality, a fundamental advantage of the modular laser array is that the actual lasers are quite small 

and have an almost perfectly “clean” interface to the rest of the system: alternating-current (AC) or DC 

power and cooling water as input, and photons within a specified wavelength range as output. The 

technology freeze date for the lasers could, thus, be quite late, essentially as of the start of volume 

production of lasers. Indeed, new laser technology could be introduced during production. Given the rate 

of advance of laser technology (and the fact that it is being driven by industrial and military applications), 

it is very likely that the actual laser systems that would be used in a launcher to be deployed in, say, 2022, 

are currently at TRL 4 or lower. 

This is even more true for direct-diode arrays as used in the DRM system for the boost laser. Highpower arrays and wavelength-combining and beam-shaping optics which could be used for a launch 

system are in production and commercial use (TRL 8), but better arrays and optics are in development 

(one example is given in Ref. 47) and will almost certainly be at TRL 8 well before a launch system 

would be built. 

 

6

 IPG announced in 2009 that it has delivered 5-kW lasers using the same design to an industrial customer. They 

have not identified the customer or announced actual delivery of any10-kW lasers. 
NASA/TM—2012-217014 272 

3.1.5.5.1.24 Development Roadmap 

Figure 3.103 shows a roadmap for the development and construction of a DRM-capable launch 

facility. 

Figure 3.103.—Roadmap for launch system development and deployment. 

A basic advantage of the modular laser/HX launch system is that nearly all aspects of the system can 

be developed and demonstrated at much less than the final vehicle scale and with less-expensive 

substitutes for the missing system elements. 

3.1.5.5.2 Ground and Flight Demonstrations 

There are several possible demonstration paths and options for the laser/HX system, depending on 

which aspects are regarded as most critical or highest risk. Figure 3.104 shows one possible set of ground 

and flight demonstrations for the HX vehicle. 

We have not given costs and schedules because the costs will depend heavily on how the tests and 

demonstrations are defined, and particularly on what types of beam sources are used, versus what may be 

available, either from unrelated programs or from separate development activities within an overall 

beamed-energy-propulsion (BEP) program. For example, it may be decided to build a multi-megawatt 

diode laser or millimeter-wave source as a user facility, rather than as a single-purpose installation for this 

program. 
NASA/TM—2012-217014 273 

Figure 3.104.—Possible ground (blue) and flight (green) demonstration program for HX vehicle. 

3.1.5.5.2.25 Static Demonstrations of Propulsion System 

To the extent that the highest technical risk is associated with the HX, two possible milestone 

demonstrations are as follows: 

100-kW HX/benchtop thruster 

Objective: Demonstrate basic subscale HX thruster 

Facilities: JHPSSL (100 kW) or ~100-kW diode array facility, wind tunnel, or blowdown air tank 

Test unit: ~200-cm2

 SiC HX (2  100-cm2 -panels + manifolds) with low-expansion nozzle 

Success criterion: Stable operation with nitrogen gas flow at flight conditions (nominally 2000 K gas 

temperature, ~5 g/s gas flow) for >100 s, including ~10-s operation with air flow (simulating 

thermal and surface-erosion conditions during ascent) 

Megawatt HX/LH2 thruster 

Objective: Demonstrate subscale flight-performance HX thruster 

Facilities: 200- to 1000-kW laser diode array, wind tunnel, or blowdown air tank; hydrogen handling 

Test unit: ~200- to 1000-cm2

 SiC HX with low-expansion nozzle 

Success criterion: Stable operation with hydrogen gas flow at flight conditions (nominally 2000 K gas 

temperature, ~5 g/s gas flow) for >500 s, including ~60-s operation with air flow (simulating 

thermal and surface-erosion conditions during ascent) 

This demonstration might alternatively use the proposed Ames 1-MW gyrotron source, which would 

allow demonstration of a larger (>1000-cm2

) HX segment, but this would require demonstration of a 

microwave absorber layer that would adequately reproduce the thermal effects of laser absorption without 

changing the HX thermomechanical properties. This might be a thin layer of resistive material. The 

microwave experiment would not test surface erosion and could be done in air or vacuum. 
NASA/TM—2012-217014 274 

Full-scale propulsion system 

Objective: Demonstrate full sounding-rocket-scale propulsion system (directly scalable to launch 

vehicle) 

Facilities: Multi-megawatt heat source; vacuum chamber (>2 m diameter), engine test facility sized 

for 10,000-lbf thrusters 

Test unit: Integrated propulsion system (LH2 tank, pump, HX, and nozzle) 

Success criterion: Startup, stable operation, and shutdown of HX, nozzle, and propellant flow system 

This demonstration would require a large heat source (at least several megawatts), which could be a 

laser diode array or microwave tube array if a suitable array were available, but would more likely be 

either a high-temperature resistive heater array or tungsten-filament radiator. An alternative would be to 

use a combustion heater (essentially a rocket combustor) with the exhaust flowing either directly over the 

HX face or through a second HX (probably a simpler turbulent-flow design) coupled to the test HX faceto-face. 

3.1.5.5.2.26 Vehicle Flight Demonstrations 

Bottle rocket 

Objective: Demonstrate integrated propulsion system and beam-source tracking of accelerating 

vehicle 

Facilities: JHPSSL or 100-kW class diode array with steering mirror and tracking system; sled track 

or suspended cable at 2- to 5-km range 

Test unit: 100-kW-scale propulsion system using water or LN2 propellant, including tank, nozzle, and 

propulsion controls, and vehicle beacon (typical dry mass, 10 kg) 

Success criterion: Track and accelerate bottle rocket continuously over available track length 

Sounding rocket 

Objective: Demonstrate integrated propulsion system and beam-source tracking of accelerating 

vehicle 

Facilities: ABL or megawatt-class main laser (fiber, etc.) subarray; test range (e.g., WSMR) 

Test unit: Megawatt-scale propulsion system using pressure-fed water or LN2 propellant—including 

tank, nozzle, and propulsion controls—integrated into a partially or fully guided stage, either airdropped or launched to an altitude of several kilometers by a small solid booster 

Success criterion: Track and accelerate HX rocket stage over 10+ km and 30+ s; demonstrate 

acquisition and self-guidance including roll control; may demonstrate supersonic flight 

This would be the near-term option for demonstration using ABL. The solid booster stage would be 

used (in preference to ground launch or air drop) to provide a plume for initial ABL acquisition and 

tracking. Because of the small size (0.1 to 1 m2

) of a megawatt-class HX, this demonstration probably 

would require a high-quality beam source such as ABL or a small number of main-array beam modules 

(e.g., 20  60-kW modules). A boost array module would only be able to focus on the HX up to a range of 

~10 km. (A section of the boost array could, of course, be used for preliminary tests, including ground 

launches of the sounding rocket stage.) 

Suborbital launch 

Objective: Demonstrate near-full-scale vehicle launch, transatmospheric flight, and acceleration 

Facilities: 10- to 100-MW boost laser subarray deployed at either a test range or the eventual launch 

range 

Test unit: Subscale “heavy” vehicle (10% to 50% of DRM vehicle gross liftoff mass, but built using 

off-the-shelf tanks and other hardware, a la DC-X); carries propellant (LH2 and dense propellant) 
NASA/TM—2012-217014 275 

for 100- to 200-s flight time, including at least some high-Isp operation with straight LH2

propellant; vehicle may be recoverable/reusable 

Success criterion: Vehicle flight to TBD altitude and velocity, typically powered flight to 50 km 

altitude/1 km/s. 

This would be a series of tests for validation of the vehicle design before a commitment was made to 

produce full-scale vehicles. It would also provide full validation of the boost array design and could 

validate operation of at least parts of the main array; that is, even if the siting of the tests and the state of 

completion of the main array did not allow using the main array modules for significant propulsion, it 

would at least be possible to test the vehicle acquisition and tracking and the laser pointing and beam 

properties of main-array beam modules. 

To the extent that suborbital launches would approximate the conditions of an orbital launch (except 

for total burn duration), these launches could also flight-test avionics and other hardware for the full-sized 

vehicle. 

On-orbit propulsive test 

Objective: Demonstrate main laser array tracking and beam quality at high power/long range and full 

angular rates; validate vehicle design and operation at burnout and orbit insertion 

Facilities: ~10-MW main laser subarray, deployed at a test range or (preferably) at the main laser site 

Test unit: Subscale “light” vehicle (dry mass 10% to 20% of the DRM vehicle dry mass) configured 

to fit on top of a small conventional launch vehicle; carries a small hydrogen tank (e.g., 20% to 

50% of wet mass is LH2); HX vehicle is launched into LEO by the conventional stage and is 

powered by the laser array as it passes over the array; post-acceeration path may be a stable orbit 

or may reenter 

Success criterion: Achieve planned delta-V 

This demonstration could be done at a lower power level, possibly even using ABL, if a low-cost 

launch opportunity were available. However, to validate the beam module and HX vehicle design and 

operation, the demonstration probably needs to be done at fairly high power, using actual beam module 

prototypes. 

An alternative to maneuvering a spacecraft already in a stable orbit would be to launch a laser-HX 

“upper stage” on a suborbital trajectory and boost it into orbit. However, this would be higher risk (only 

one chance to perform the maneuver) and would require either launching a conventional rocket on a 

trajectory over land or installing a sizeable laser array temporarily on an island or other site where it could 

see an appropriate portion of the trajectory of a suborbital launch from an existing launch range. 

3.1.5.5.2.27 Beam module and Laser Array Demonstrations 

Finally, there are also several possible milestone demonstrations that would involve only the beam 

modules, including tracking test targets (conventional rockets or satellites), delivering power from a 

single module to an instrumented test target, and demonstrating simultaneous operation of multiple 

modules at various total power levels. These are not addressed in any detail here, but should be 

considered in developing a full demonstration roadmap. 

3.1.5.5.3 Issues and Resolutions 

3.1.5.5.3.28 Atmospheric Transmission 

The beam module array has no significant issues with thermal blooming because of the relatively low 

power in a single beam. Thermal blooming close to the vehicle, where individual beams overlap, may be 

a minor issue at launch, when the vehicle velocity is small. This may set a lower limit on the overall area 

of the boost laser array. 
NASA/TM—2012-217014 276 

As noted earlier, at the wavelengths of interest, there is very little absorption for beam paths 

originating at an altitude of 2 km or higher. For lower altitudes, some wavelengths (notably 900 to 980 

nm for the boost laser, and 1.06 to 1.1 m for the main laser) have significant absorption because of water 

vapor, but laser wavelengths can be chosen such that absorption remains small. Scattering will be the 

dominant loss mechanism and will depend on the aerosol properties at the laser sites. 

3.1.5.5.3.29 Environmental Issues 

There are no known environmental issues associated with the vehicle exhaust, which will be mostly 

water or nitrogen (dense propellant) plus a small amount of hydrogen in the atmosphere. The HX surface 

or the burning of the hydrogen in air may be hot enough to generate some oxides of nitrogen. There are 

no chlorides or other long-lived ozone-depleting species in the exhaust. 

To our current knowledge, the vehicle will not use substantial amounts of toxic or hazardous 

materials, and the total mass of structural material in reentering vehicles will be small (<1000 tons/year 

for the reference system). Silicon carbide is a widely used industrial abrasive, and carbon composites are 

also widely used. However, a review of possible hazards from scattered or vaporized vehicle structure 

will be needed when the vehicle and HX compositions are known. 

High launch rates of small vehicles could cause a significant additional orbital debris hazard if 

vehicles do not reenter quickly, but that is primarily an engineering and operations issue. Orbital debris 

concerns may force the inclusion of some type of backup reentry mechanism such as the Terminator 

Tether (Ref. 48). 

High-power laser beams of the type and power used for the HX launch system have no direct effect 

on the atmosphere or environment; they do not cause chemical reactions or heat the atmosphere enough to 

affect weather. The one potential environmental hazard from the laser arrays is injury to birds, which may 

suffer eye damage or minor burns if they fly directly through a beam (the beam intensity is generally too 

low to cause severe burns except close to the vehicle). Individual laser beams can be switched off by 

safety systems if birds enter the beam path, without affecting launch operations, but large flocks flying 

around or above the laser arrays might cause unacceptable power loss. (They would also tend to decrease 

the optical quality of the output apertures). This is unlikely if the arrays are at high altitude, but the impact 

on threatened and endangered birds will need to be analyzed and perhaps tested with prototype arrays. 

Measures similar to those used at airports can be used to keep birds away from the launch site and the 

laser arrays. 

Finally, vehicles lifting off will be point noise sources, with much lower total power than chemical 

launch vehicles but more frequent operation. This may be an issue in the selection of launch sites, or 

impose some limits on, for example, hours of operation, although both people and wildlife around active 

launch facilities appear to become accustomed to the noise; the vertical ascent will limit vehicle noise to 

the immediate vicinity of the launch site. 

3.1.5.5.3.30 Laser Safety 

Laser safety hazards are generally characterized as being due to the direct beam, specular reflections, 

or diffuse reflections. 

In general, the direct and specular-reflection hazards from the modular laser array are much lower 

than the hazards from a single beam of similar total power because, except for a small region close to the 

vehicle, the individual module beams do not overlap. Thus, direct and specular hazard estimates can be 

based on the power of a single module (60 to 200 kW) rather than the full array. 

Direct viewing.—We assume that the space directly between the laser arrays and the vehicle will 

always be controlled, so that there is minimal risk of anyone or anything actually being hit by a direct 

beam. For the boost laser, this implies a controlled ground area directly under the beam wherever the 

beam is within a few meters of the ground, and a controlled airspace covering the area between the boost 

array and the launch site. For the main array, which never points below ~15 from the horizon (and could 

have mechanical hard stops at that angle), the controlled airspace is nominally a cone or inverted pyramid 
NASA/TM—2012-217014 277 

extending upward from the array site. For example, at an 18-km (60,000-ft) altitude, the controlled region 

would have a major axis (along the launch trajectory) of roughly 140 km. The cross-track width of the 

controlled region would depend on the range of launch azimuths used, but it would generally be much 

smaller: a few kilmeters at the “uprange” end and 10 km at the “downrange” end. A radar system would 

detect aircraft within this controlled space and, if necessary, delay or abort launches. 

In addition, individual beam modules or subsections of the laser array would have “boresight” radars 

that track with the laser beam. These could detect aircraft (or birds) flying into the beam and shut down 

the beam. Since only a single module or a small section of the source would be shut down, this would not 

interfere even with a launch in progress. 

The relatively few satellites orbiting below the trajectory altitude would be at risk of intercepting the 

main laser beam, but a combination of schedule coordination and boresight radar detection would make 

the chance of actually hitting a satellite extremely small. In the worst case, a satellite encountering a 

single module beam would be exposed to <10 kW/m2

 (e.g., a 50-kW beam with a Gaussian full width of 

5 m) for on the order of 1 ms, so the energy deposited on the satellite would be small (<10 J/m2

 or 

1 mJ/cm2

) and thus extremely unlikely to cause physical or thermal damage. The only significant hazard 

would be to optical instruments that might focus the laser energy. (Detection of optical reflection of a 

beam from a satellite would be possible, but the speed-of-light delay in reaction would generally be 

longer than the satellite would be in the beam). 

Beyond the vehicle, there would be a direct-beam hazard due to “spillover” from the main beam or to 

individual beam modules that suffer pointing errors. (Loss of contact with the vehicle beacon, due to 

vehicle failure or gross pointing error, would immediately shut down a beam module in ~1 ms, so 

pointing a beam more than a few meters away from a vehicle would require multiple failures). This would 

be the primary concern for ground and aircraft hazards for the boost laser array. 

We assume that the boost array and launch site would be located such that spillover energy at and 

immediately after liftoff would either not hit the ground at all (launch site altitude > laser altitude), would 

be stopped by nearby terrain (such as a mountain “behind” the vehicle launch site), or would be directed 

at an unpopulated and reasonably controllable area, such as the ocean. The maximum range of concern 

would be ~200 km, since the distance to the horizon for a boost laser at 2 km above sea level is at most 

~225 km, but in practice a typical boost-laser beam (10 kW, 100 rad divergence) would be eye-safe for a 

0.1-s exposure more than ~100 km from the laser, although not necessarily safe for observers with 

binoculars or telescopes. (One additional option would be to choose a wavelength range for the boost 

laser that has significant water-vapor absorption, so that a 20-km path at an altitude of 1 to 2 km would 

have acceptable attenuation (10% to 30%) but that, at lower altitudes and longer ranges, the beam would 

rapidly attenuate in most atmospheric conditions.) 

For aircraft, there is a severe hazard for the first ~50 km beyond the launch site, and this would need 

to be restricted airspace during launches, although the restricted zone could be very narrow in both time 

(5 min per launch) and space (<1 km wide). Beyond 100 km, the combination of beam flux and exposure 

time would be eye-safe, and airplane pilots and passengers are generally not using (or able to use) large 

magnifying optics. At intermediate ranges, there is a marginal hazard (a worst-case exposure <5 times the 

ANSI/IEC maximum permitted exposure) and restricting air space may or may not be necessary. Note 

that in no case is there a risk of physical damage to aircraft that are more than a few kilometers beyond 

the launch vehicle. 

(The ground level and aircraft hazards could be further reduced by using longer-wavelength lasers for 

the boost laser, for example, 1.55 m, but currently the cost and performance of both diode arrays and 

fiber lasers at this wavelength are several times worse than near 1 m. Alternatively, if laser diode prices 

drop sufficiently, a “preboost” laser array located at the vehicle launch site (within 1 to 2 km) could 

supply part or all of the power for the first few kilometers of the trajectory. The preboost lasers would 

need minimal optics and would have highly divergent beams, so they would be eye-safe a few kilmeters 

beyond the vehicle, and the boost laser array would then never operate at high power below ~15 above 
NASA/TM—2012-217014 278 

the horizon, so their beams would go above commercial flight altitudes only 40 to 60 km beyond the 

launch site.) 

Spillover beams from the boost laser array would present minimal hazard to satellites, since the 

beams would be below 10 mW/cm2

 at the lowest satellite altitudes (~20-m beam diameter at 400 km, 

assuming 50-rad divergence, with individual beam powers of ~20 kW). However, coordination of 

satellite passes very close to the boost laser site would be desirable. 

The largest hazard to satellites would be spillover or stray beams from the main laser site. These may 

have fluxes over 1000 W/m2 (100 mW/cm2

) out to an altitude >1000 km. It is also possible (although 

unlikely) that the beams could accidentally match the velocity of a satellite in orbit closely enough to 

stretch the exposure time considerably, possibly to as much as 1 s, or several seconds of intermittent 

exposure as the satellite passed through the array of beams. This would still be unlikely to cause thermal 

damage, but it could conceivably damage solar panels or related electronics, or sensors not aimed directly 

at the beam source. The cross section of the array of beams would also be fairly large, comparable to the 

dimensions of the array on the ground. 

The hazard to LEO satellites will thus probably require coordination with satellite tracking agencies 

(currently done via the Laser ClearingHouse) and some constraints on launch windows. In addition, it 

may be desirable, and eventually necessary, to use a boresight radar system sufficiently powerful to detect 

satellites out to a range of >1000 km and to interrupt the laser beam if a satellite is detected approaching 

the beam. (For most satellites, the interruption would be a fraction of a second and would not cause a 

launch failure.) 

Specular reflections.—Specular reflections from the vehicle are generally a small hazard, because 

the vehicle will be designed to have no flat, polished surfaces. Small reflective surfaces, such as might be 

caused by minor damage in vehicle processing, would reflect power from different beam modules in 

slightly different directions, so the flux reaching any one observer more than a few 10s of meters from the 

vehicle7

 would be from one beam module. 

The specular beam intensity refl from an ideal reflective spot of area A at range R is 

2

in

refl 2 2

A

R

  



where the incident flux in is 

beam in 2

beam

4 P

d

 



that is, the beam module power divided by the beam area at the vehicle. 

For the main array, the worst case corresponds to the vehicle being directly over the laser: roughly 

Pbeam = 60 kW and dbeam ~ 1 m (at a range of ~200 km), so in ~76 kW/m2

. R = 200 km and  = 1.06 m, 

so refl ~1.7106

 A2 W/m2. For A ~ 1 cm2, refl ~ 17 mW/m2 or 1.7 W/cm2, which is several orders of 

magnitude below the maximum permissible continuous exposure of 5 mW/cm2

. Even an observer using a 

sizeable telescope would be unable to exceed the allowed eye exposure. 

For the boost laser array, the potential hazards are somewhat greater. Depending on how the boost 

array is aimed and focused, in will typically be between 10 and 100 kW/m2

, but R could be much shorter, 

 

7

Or a few kilometers from the vehicle for the main laser array. However, since the vehicle only sees the main 

laser array when it is well above 50-km altitude, main-laser reflections are not a significant hazard. 
NASA/TM—2012-217014 279 

and  might be as short as 0.85 m. If we assume that in = 30 kW/m2 and a 1-cm2 reflector at 1.0 m, 

refl ~ 3108 R–2 W/m2. 

This falls below the naked-eye maximum permissible exposure (MPE) of ~4 mW/cm2

 (40 W/m2) at a 

range of approximately 2.7 km. This is a reasonable distance over which to exclude unprotected 

observers. However, observers using binoculars or telescopes may exceed the MPE at distances roughly 

5 to 20 times larger (assuming 50-mm binoculars or a 200-mm telescope with ~3 milliradians of apparent 

blurring due to chromatic aberration between the visible and near-infrared). Excluding all binocular users 

from a 13-km radius would be difficult. Excluding all telescope users from a 50+ km radius may be 

unfeasible. 

Fortunately, the actual hazard is almost certainly much lower. Real surfaces generally do not reflect 

beams even as narrow as a diffraction-limited 1-cm2

 mirror (Ref. 49), nor are they 100% efficient 

reflectors. Conversely, however, the extended array source means that the reflected beam flux is nearly 

constant for a wide range of reflected beam widths once the reflected beams from different sources begin 

to overlap. This area-average intensity is lower than the single-beam reflected intensity by roughly the 

fill-factor of the array source, that is, a factor on the order of several hundred for the nominal main laser 

array (~1-m apertures on ~40-m centers) but possibly a factor of 10 or less for the boost laser array (2- by 

5-m aperture for a single module, with modules spaced, for example, 2.5 by 40 m, as seen from the launch 

site.) If we assume that this reduces the peak beam flux by a factor of 10, the safe radii for 

binoculars/telescopes become 4 and ~16 km, respectively. 

If a given observer sees reflections for less than 10 s, the allowable exposure is higher, increasing as 

(10 s/t)

0.25. Once the vehicle is off the launch stand, typical exposure times are likely to be 0.1 s or less as 

the vehicle attitude changes. This would increase the MPE by ~threefold, and reduce the minimum safe 

range by a factor of ~1.7. For the vehicle on the stand, reflections might be stable for several seconds 

before the vehicle actually begins moving, but the hazard zone could be limited to a narrow “pie slice” 

toward the laser array by putting suitable barriers around the launch stand or launch site. 

Finally, the MPE levels are conservative, and exposure levels up to 5 times higher generally require 

only a suitable warning notice. So whereas a telescope-safe distance may be 10 to 16 km from the launch 

site, an occasional observer who gets within as little as 4 to 5 km (or 7 to 8 km in certain specific 

directions) would still be both actually safe and within reasonable regulatory limits. 

We reiterate, however, that the above safe distances are based on 1-cm2

 specular reflective areas, and 

a safety inspection able to detect reflective spots larger than this will be needed. Ideally, this would be an 

automated scan performed on the launch stand shortly before launch. 

Diffuse reflections.—Diffuse reflections obviously present a less concentrated hazard than specular 

reflections, but because the total laser power is so high, the hazard can be significant. Again, there is no 

hazard from the main laser array because all observers are distant (unless a vehicle passes very close to a 

watching astronaut). 

For the boost laser array, an analysis similar to the above analysis for specular reflections was done. 

We assumed approximately Lambertian scattering (even distribution of light in all directions) and a 

scattered flux of 1 MW/m2

, corresponding to 10% scattering from the HX itself or complete scattering of 

the fringes of the beam (e.g., from the fuel tank). In this case, the exposure dropped below the MPE for 

naked-eye viewing at ~100 m from the vehicle. The scaling for aided viewing was similar, except that the 

probable chromatic error of binoculars or telescope eyepieces outside of the visible range did not reduce 

the hazard. The safe range became ~700 m for 50-mm binoculars and ~3 km for a 20-cm telescope. 

However, it is difficult to ensure that scattering will be uniform, especially from the HX. There may 

also be “hot spots” in illumination or scattering. A conservative assumption, pending more detailed 

analysis, is that the keep-out zone should be at least 1 km for all observers, >3.5 km for users of 

binoculars or cameras with through-the-lens viewing, and >15 km for telescopes in directions where there 

might be peaks in the scattered intensity, primarily toward the laser site. This is consistent with the 

specular-reflection safety ranges. 
NASA/TM—2012-217014 280 

3.1.5.5.3.31 Range Safety 

The present study only minimally addressed range safety other than laser safety, and the differences 

between range safety requirements for HX laser launch in comparison to chemical launchers. We note 

briefly the following points, and recommend that range safety be considered in more detail in a future 

study: 

 The combination of simple single-stage vehicles, extensive “type certification” testing, the ability 

to ground-fire complete vehicles, and a distributed laser array with reserve capacity will provide 

extremely high overall reliability. 

 The laser launch vehicle will have minimal stored energy onboard. It, therefore, will not be able 

to generate significant explosive force on the pad or in flight provided the oxidizer component of any 

dense propellant is kept low. (The hydrogen propellant could readily burn in air, but it would be very 

difficult to mix with air and ignite in such a way as to produce a large detonation.) 

 The vehicle will not generate significant thrust in the absence of the laser beam, so uncontrolled 

flight is essentially impossible. Turning off the laser arrays will provide reliable thrust termination. 

 The total vehicle mass will be small in comparison to any prior launch vehicle and most 

suborbital rockets, including the V-2. The worst-case ground impact or explosion will be comparable to 

the crash of a small airplane. 

 The combination of high launch rates and low unit launch cost will enable extensive ground and 

suborbital testing, plus orbital-launch testing with various deliberate failures, allowing much more 

complete demonstration of reliability and of possible failure modes and their effects than is possible with 

any other launch system. 

3.1.5.5.3.32 Launch System Availability 

See Section 3.1.5.3.4—Availability. 

3.1.5.5.3.33 Power Source 

As noted in Section 3.1.5.2.2—Main Laser Site Design, there are several options for powering the 

launch system using off-the-shelf technology, without imposing unprecedented loads on the national 

electrical grid or requiring special approvals and exceptional lead times (as would be needed, for 

example, for a new nuclear power plant or a major transmission line.) The worst-case option is a pair of 

dedicated gas-turbine power plants (at main and boost array sites) able to supply the entire launch system 

load. If grid power were used, it would be with some type of energy storage system able to spread the 

load in time to match available off-peak generating capacity, combined with a medium-voltage (<100-

kV) transmission line from the nearest generating plant or grid tie point. 

3.1.5.5.3.34 Co-use of Laser/HX Thruster Concept and Launch System 

Co-use of Airborne Laser and future laser weapon systems.—The ABL is certainly compatible 

with an HX propulsion system and could potentially be used for development and demonstration as 

described above. The ABL is probably not capable of powering a full launch to orbit, even of a token 

payload, because of power, range, and run-time limitations. 

Since HX propulsion will work with essentially any laser, it is possible to consider it as a way to 

extend the capabilities of the ABL or any future high-average-power laser with laser-powered projectiles, 

either for R&D purposes (discussed later) or as an additional weapon capability. For velocities up to 

roughly 5 km/s, a laser-powered missile could use water or, for example, LN2 propellant, yielding a 

completely “wooden round” hypersonic missile with no fire or explosion hazard and (especially if 

“fueled” in the field) very low weight. 

Conversely, any laser weapon system with megawatt-class or higher average power is a potential 

beam source for testing HX laser propulsion systems and vehicles. 
NASA/TM—2012-217014 281 

The largest commonality between laser weapon systems and HX launch, however, is that individual 

technologies, including, for example, optics manufacturing or pointing and tracking, developed for either 

one may be applicable to the other. This commonality may be enhanced if development programs have at 

least some connection. For example, a laser weapon program may have comparatively short run-time 

requirements but may select a laser technology that can operate for longer times if the program 

management is aware of the potential application to laser launch. 

Co-use of horizontal takeoff.—Horizontal takeoff appears to be a poor choice for purely laserpowered vehicles using GBL, because both line-of-sight geometry and atmospheric absorption and 

scattering limit the useful laser range to a few 10s of kilometers for vehicles in the sensible atmosphere 

(below 30 km). An HX-heated turbojet aircraft engine could provide fuelless (and pollutionless) 

propulsion up to perhaps Mach 2, but conventional hydrocarbon engines are already very good in this 

regime. At higher velocities, a subsonic-flow HX is a poor choice for heating air,8 and preheating fuel, 

while desirable, can generally be done using engine waste heat or energy from aerodynamic heating. 

However, HX propulsion could potentially be used in connection with horizontal takeoff vehicles in 

at least two configurations: 

 A two-stage launcher, using a horizontal-takeoff suborbital vehicle with conventional 

chemical engines (jet or combined cycle) as a first stage, and a laser HX vehicle—expendable 

or reusable—as an upper stage 

 A single-stage reusable vehicle using chemical propulsion in the atmosphere but switching to 

a laser HX rocket mode for out-of-atmosphere acceleration and orbit insertion 

The two-stage configuration is distinguished from an air-launched single-stage vehicle by the 

assumption of a first-stage vehicle with significantly higher altitude and velocity capability than existing 

aircraft. The two-stage configuration would make sense if the carrier aircraft could be built and operated 

at lower cost than a boost laser array (or possibly several boost arrays, if the first stage vehicle enabled 

launching to different azimuths, or if multiple launch systems were built). This is certainly possible; there 

are several active efforts to build reusable suborbital horizontal takeoff and landing (HTHL) vehicles that 

can carry several people or ton-scale payloads to 100+ km altitude at a price on the order of $1K/kg 

(Ref. 50). 

3.1.5.5.3.35 Other Applications of an HX Launch System or Technology 

Suborbital launch and space tourism.—Laser launch could deliver much larger payloads on 

suborbital trajectories than to orbit. For example, the boost laser array proposed for the DRM launch 

system could lift ~2.5 tons (vehicle plus payload) to 150 km (thrust cutoff at 100 km and 1 km/s) with 

~4 min of zero-gravity trajectory before reentry. The combination of very high reliability, demonstrated 

safety, low environmental impact, and low acceleration, plus extremely high launch rate and low perlaunch cost with a reusable vehicle, would make this mode of operation ideal for suborbital space tourism 

as well as scientific and industrial activities. 

A unique capability would be to launch large numbers of “reentry sondes” that could be targeted to 

reenter at specific locations over a wide area, sampling the atmosphere in all three spatial dimensions and 

in time over an extended period. Sondes could be launched in groups on a relatively large reusable 

suborbital vehicle or, with a subscale version of the main laser array (so as to allow focusing on submeter-sized HXs at 50 to 100 km), as expendable vehicles with ~ 100-kg GLOM. 

Intercontinental parcel delivery.—Fast parcel delivery by an unmanned rocket has been proposed 

many times, and although problematic for various practical reasons (how do you clear customs?), it 

remains a possible application of low-cost, high-volume launches. An HX laser launch system would 

potentially be able to deliver (nonmilitary) payloads anywhere in the world within <1 hr. Laser launch 

 

8

More exotic approaches such as plasma-coupled laser scramjets might be of interest, but these are well beyond 

the scope of this study. 
NASA/TM—2012-217014 282 

would be handicapped, however, by weather availability, which is not readily solved in this case by 

having multiple sites because the extra time to carry packages between launch sites would defeat much of 

the benefit. 

Military applications.—Orbital launch systems would also be able to launch payloads onto 

suborbital trajectories or to launch orbital payloads with reentry capability. The payloads could include 

explosive warheads or kinetic energy impactors. 

A laser launch system is very well suited to such a nonnuclear strike: It has a low cost per launch, has 

a high sustained launch rate (although not as high as some cannon-type launchers), and has sufficient 

flexibility in trajectories to reach most areas of the planet. A laser launch is very easily distinguished from 

an intercontinental ballistic missile (ICBM) launch or any other military activity. A laser array would also 

be robust against many kinds of attack. The main limitation of laser launch is that it could be disabled by 

weather, so having strike capability at all times would require multiple launch sites. 

A launch system could be demilitarized for intercontinental strike either by limiting its range of 

trajectories (e.g., by building the beam director mounts with limited range) or by providing some type of 

international control, as well as by not developing or testing weaponlike payloads. 

A laser launch system would be an efficient launcher of small orbit-resident antisatellite or antimissile 

weapons or decoys. This was the original basis for SDIO support of laser propulsion research in the late 

1980s: as a launcher for Brilliant Pebbles interceptors and decoys (Ref. 51). A launch system could also 

launch direct-ascent antisatellite weapons to LEO or to geosynchronous transfer orbit. 

A laser array with even a fraction of the power of a launch array would be able to promptly destroy 

LEO satellites passing within its field of view, provided that the array was designed and equipped to track 

noncooperative targets. As a large fixed system, however, a launch laser array would be both inefficient 

and conspicuous as an antisatellite laser. 

A full launch-scale laser or laser array capable of delivering >10 MW/m2 at a range of >400 km 

could also deliver >1 kW/m2

 at 40,000 km: that is, GEO. This could potentially allow an array to 

temporarily blind or permanently disable (by overheating) satellites in GEO. Again, a launch system 

would not be particularly efficient for this purpose in comparison to a purpose-built antisatellite laser, but 

it would provide some military capability unless it was explicitly designed to be incapable of striking 

GEO (e.g., by limiting pointing southward) 

Further discussion of military applications is beyond the scope of this report. 

Beam module/laser array technology.—Finally, beam modules as such, and more generally beam 

module technologies for low-cost laser power, pointing and tracking, and adaptive optics, may be relevant 

to nonpropulsive power beaming. Some possible applications include delivering power to satellites in 

eclipse or for peak power needs, annealing satellite photovoltaic panels, and powering high-altitude 

airships or “eternal” UAVs. 

3.1.5.6 References for Laser Thermal Consultant Report—DRM 1–B 

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(2003). 

3.1.6 Laser Thermal Cost 

National Aeronautics and Space Administration

Beamed Energy Propulsion Facility 

ROM Cost Estimate 

March 28, 2011
NASA/TM—2012-217014 286 

National Aeronautics and Space Administration

General BEPS Launch Facility Assumptions

• All cost are rough order of magnitude (ROM) estimates in 

FY11M$.

• Estimates assume all facilities are built new on government 

land, no existing infrastructure is used. Available infrastructure 

will reduce costs.

• Costs do not include remediation of potential significant 

environmental impact issues that may be identified.

• Range estimates are presented to highlight the uncertainty 

associated with the lack of definition and supporting cost 

information available at this time. 

• Range estimates for the launch site also includes the 

uncertainty associated with the selected location and the cost 

impact of materials that may be required to protect against 

environmental conditions.

National Aeronautics and Space Administration

General BEPS Launch Facility Assumptions

• Launch site ROM estimates are based on a NASA KSC launch 

site estimating approach. Microwave and laser sites use the 

same approach for ‘like function’ facilities which are scaled for 

expected differences in size or function.

• All sites include site design and testing and activation costs.

• Launch facility design is estimated at 12% of estimated total site 

costs less testing and activation.

• Launch facility testing and activation is estimated at 30% of the 

capital facilities cost.

• Microwave and laser site design is estimated at 8% of estimated 

total site costs less testing and activation.

• Microwave and laser site testing and activation is estimated at 

15% of the capital facilities cost.
NASA/TM—2012-217014 287 

National Aeronautics and Space Administration

Operations Cost Assumptions

• All costs are estimated for one year of operation in FY11$. 

• Yearly operations cost assume 360 launches per year 

(one/day). Additional study is required to determine the feasible 

number of launches per year from a single launch pad. The 

number of launches per pad is affected by the CONOPS, pad 

turnaround time, launch facility maintenance, weather, and 

launch window requirements.

• Total estimated facility operations costs are assumed to be 

allocated to the beamed launch operations. Facility sharing can 

reduce costs. 

• Facility equipment and maintenance is estimated at 5% of the 

total value of capital assets per year resulting in an estimated 

70-80% of total operations costs.

National Aeronautics and Space Administration

Laser Thermal Rocket Launch Site Assumptions

• Launch Site facility requirements

– General – site work, roads

– Support Facilities/Launch Pad – see details below

– Site Infrastructure – water, fire, sewer, power, communications, HVAC, environmental 

control system

– Ground support equipment

• Key cost estimate assumptions

– New facility on government land in secluded area due to danger associated with 

beamed energy. Launch rate 1-4 vehicles a day, if CONOPS allows.

– Launch pad includes all related hardware/services (scaled from microwave estimate). 

Rocket will be approximately 29 meters in length and less than 5,000 kg total wet mass 

with payload.

– Command center includes all related services.

– Payload integration facility (small payloads <320 kg) can process/integrate up to four 

payloads per day (scaled from microwave estimate).

– Hydrogen storage facility sized for ~30,000 kg with no on-site production.

– Rocket storage facility size for 8 rockets (scaled from microwave estimate).

– Admin facility sized for launch site staffing needs (estimated at 100 people).

– Security facilities/equipment includes fencing, cameras, etc.
NASA/TM—2012-217014 288 

National Aeronautics and Space Administration

Laser Thermal Rocket Boost Laser Site 

Assumptions

• Boost Laser Site facility requirements

– General – site work, roads

– Support Facilities – admin facility (50 people), command center (75% of 

launch site center), guard gate facility

– Site Infrastructure – water, fire, sewer, power, communications, HVAC, 

environmental control system

– Ground support equipment

– Laser/Optics– diode lasers(240 MW), optics (Qty.1200, 200 kWe modules)

– Power Storage/Distribution – energy storage, power supply, power 

distribution

• Key cost estimate assumptions

– Diode Laser – Projected multi-MW price $2/W (ref: J. Kare) used as low 

(ML +50%, high +100%)

– Optics – Projected $200K/module (ref: J. Kare) use as low (ML +25%, high 

+50%)

– Energy Storage - $50K/kWh estimated as most likely based on GRC 

experience with wind tunnels (low -40%, high +20%)

National Aeronautics and Space Administration

Laser Thermal Rocket Main Laser Site 

Assumptions

• Main Laser Site facility requirements

– General – site work, roads

– Support Facilities – admin facility (50 people), command center (75% of 

launch site center), guard gate facility

– Site Infrastructure – water, fire, sewer, power, communications, HVAC, 

environmental control system

– Ground support equipment

– Laser/Optics– fiber lasers(140 MW), optics (Qty.2300, 60 kWe modules)

– Power Storage/Distribution – energy storage, power supply, power 

distribution

• Key cost estimate assumptions

– Fiber Laser – Projected multi-MW price $10/W (ref: J. Kare) used as low 

(ML +50%, high +100%)

– Optics – Projected $200K/module (ref: J. Kare) use as low (ML +25%, high 

+50%)

– Energy Storage - $50K/kWh estimated as most likely based on GRC 

experience with wind tunnels (low -40%, high +20%)
NASA/TM—2012-217014 289 

National Aeronautics and Space Administration

Laser Thermal Rocket Launch Facility 

Development ROM

Low ML High Mean Est Low ML High Mean Est Low ML High Mean Est Low ML High Mean Est

Project Total ROM ($M) $379.0 $640.6 $1,395.6 $805.1 $1,280.2 $1,952.4 $2,887.1 $2,039.9 $2,613.6 $3,851.4 $5,283.6 $3,916.2 $4,272.8 $6,444.3 $9,566.3 $6,761.1

Design $32.8 $55.2 $119.6 $69.2 $83.5 $127.4 $188.4 $133.1 $170.3 $250.9 $344.2 $255.1 $286.6 $433.5 $652.2 $457.4

Testing & Activation $73.1 $125.1 $279.5 $159.2 $152.6 $232.9 $344.1 $243.2 $315.2 $464.5 $636.4 $472.0 $540.8 $822.5 $1,260.0 $874.4

Laser Thermal Rocket Facilities ROM Subtotal $273.2 $460.3 $996.5 $576.7 $1,044.1 $1,592.1 $2,354.6 $1,663.6 $2,128.2 $3,136.0 $4,302.9 $3,189.0 $3,445.4 $5,188.4 $7,654.1 $5,429.3

WBS Item 000 0

Site Work $11.2 $15.4 $18.8 $15.1 $8.4 $11.6 $14.1 $11.3 $8.4 $11.6 $14.1 $11.3 $27.9 $38.6 $47.0 $37.8

Roads $18.4 $27.9 $46.2 $30.8 $18.4 $27.9 $46.2 $30.8 $18.4 $27.9 $46.2 $30.8 $55.3 $83.7 $138.6 $92.5

Support Facilities $37.6 $79.9 $154.2 $90.6 $7.1 $14.9 $36.1 $19.4 $7.1 $14.9 $36.1 $19.4 $51.9 $109.8 $226.3 $129.3

Concrete Administration Facilty $7.4 $10.8 $28.5 $3.7 $5.4 $14.2 $3.7 $5.4 $14.2 $14.7 $21.5 $57.0

8 Rocket Storage Facility $7.0 $9.4 $15.3 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $7.0 $9.4 $15.3

Concrete Command Center $3.3 $10.4 $25.3 $2.5 $7.8 $19.0 $2.5 $7.8 $19.0 $8.3 $26.0 $63.4

Payload Integration Facilty $19.0 $47.6 $82.3 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $19.0 $47.6 $82.3

Guard Gate Facility $1.0 $1.8 $2.8 $1.0 $1.8 $2.8 $1.0 $1.8 $2.8 $2.9 $5.3 $8.4

Pad Structure $16.4 $24.4 $30.4 $23.7 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $16.4 $24.4 $30.4 $23.7

Pad Structure $5.4 $7.8 $10.5 $7.9 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $5.4 $7.8 $10.5 $7.9

Emergency Egress System $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0

Flame Trench $8.4 $10.0 $11.1 $9.8 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $8.4 $10.0 $11.1 $9.8

Ligtning Protection Towers $2.6 $6.6 $8.8 $6.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $2.6 $6.6 $8.8 $6.0

Infrastructure $155.9 $254.1 $613.9 $341.3 $91.3 $142.7 $260.5 $164.8 $91.3 $142.7 $260.5 $164.8 $338.6 $539.5 $1,134.9 $671.0

GSE $33.6 $58.5 $133.1 $75.1 $8.4 $14.6 $33.3 $18.8 $8.4 $14.6 $33.3 $18.8 $50.4 $87.8 $199.7 $112.6

Laser/Optics Equipment $0.0 $0.0 $0.0 $0.0 $720.0 $1,020.0 $1,320.0 $1,020.0 $1,860.0 $2,675.0 $3,490.0 $2,675.0 $2,580.0 $3,695.0 $4,810.0 $3,695.0

Power Storage/Processing Facilities $0.0 $0.0 $0.0 $0.0 $190.4 $360.4 $644.5 $398.4 $134.5 $249.3 $422.8 $268.9 $324.8 $609.7 $1,067.3 $667.3

Launch Site Complex ‐ ROM Subtotal ($M) Boost Laser Site ‐  ROM Subtotal ($M) Laser Main Laser Site ‐  ROM Subtotal ($M) Thermal Rocket Facilities ‐ Total ROM ($M)

Laser Thermal Rocket Launch Facilities ROM (FY11$M)

(Excludes FTE/WYE personnel and O&M costs)

National Aeronautics and Space Administration

Laser Thermal Rocket Launch Facility 

Operations ROM

Low ML High Mean Est

Laser Operations Total ROM ($M/Year) $208.9 $314.9 $455.8 $326.5

Facility/Equipment Maintenance $168.1 $253.3 $373.4 $264.9

Operations Personnel $30.0 $40.0 $50.0 $40.0

Propellant/Energy/Consumables $10.8 $21.6 $32.4 $21.6

Other $0.0 $0.0 $0.0 $0.0

Launch Site Complex ‐ ROM Subtotal ($M/Year) $30.8 $48.0 $82.4 $53.7

Facility/Equipment Maintenance $12.2 $20.8 $46.6 $26.5

Operations Personnel $15.0 $20.0 $25.0 $20.0

Propellant/Energy/Consumables $3.6 $7.2 $10.8 $7.2

Other $0.0 $0.0 $0.0 $0.0

Boost Laser Site ‐  ROM Subtotal ($M/Year) $62.0 $94.8 $138.0 $98.3

Facility/Equipment Maintenance $50.9 $77.6 $114.7 $81.1

Operations Personnel $7.5 $10.0 $12.5 $10.0

Propellant/Energy/Consumables $3.6 $7.2 $10.8 $7.2

Other $0.0 $0.0 $0.0 $0.0

Main Laser Site ‐  ROM Subtotal ($M/Year) $116.2 $172.0 $235.4 $174.5

Facility/Equipment Maintenance $105.1 $154.8 $212.1 $157.3

Operations Personnel $7.5 $10.0 $12.5 $10.0

Propellant/Energy/Consumables $3.6 $7.2 $10.8 $7.2

Other $0.0 $0.0 $0.0 $0.0

Laser Thermal Rocket Operations ‐ Total ROM ($M/Yr)

Laser Thermal Rocket Facilities Operations ROM (FY11$M/Year)

(Does not include launch vehicle)
NASA/TM—2012-217014 290 

National Aeronautics and Space Administration

Beamed Energy Propulsion Study

Laser Thermal Vehicle Cost Estimate

as determined by the COMPASS Team

March 2011

National Aeronautics and Space Administration

Cost Assumptions

• DRAFT Cost Estimate based on COMPASS design

• All costs are in FY11$M

• Estimates represent prime contractor cost plus fee (10%)

• This estimate assumes the following:

– Proto-flight development approach

– Full set of ground spares included

– Model assumes TRL Level 6

• This estimate does not include any cost for technology development

– Represents the mean estimate based on cost-risk simulation results

– Parametric estimate based on mostly mass based CERs from historical cost data

– Planetary systems integration wraps

– Software not included

• Does not include:

– Any insight/oversight costs (by a NASA lead center)

– Reserves (can be as high as 40-50%)

– Ground System Cost (ie. Laser)

– Launch Services Costs (ie. Special launch approval process)

– Technology costs for components lower than TRL-6 
NASA/TM—2012-217014 291 

3.1.7 Millimeter-Wave Thermal Collaborative Modeling for Parametric Assessment of Space 

Systems (COMPASS) Report 

National Aeronautics and Space Administration

BEPS Preliminary Cost ROM 

(Represents estimated Prime Contractor cost) 

All Costs in FY11$M

WBS Description

DDT&E Total 

BY $M

Flight 

HW 

BY $M

DD&FH 

Total 

BY $M

06.1.1 Payload 0.0 0.0 0.0

06.2.1 Attitude Determination and Control 2.6 2.4 4.9

06.2.2 Command and Data Handling 4.3 2.6 6.9

06.2.3 Communications and Tracking 1.1 0.8 1.8

06.2.4 Electrical Power Subsystem 1.0 1.1 2.1

06.2.5 Thermal Control 22.7 4.2 26.9

06.2.6 Propulsion 14.0 6.2 20.2

06.2.7 Propellant 0.0 0.0 0.0

06.2.8 Structures and Mechanisms 15.2 7.6 22.8

Subtotal 60.9 24.8 85.7

Systems Integration 46.5 8.0 54.5

Spacecraft Total 107.4 32.8 140.2
NASA/TM—2012-217014 292 
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4 
NASA/TM—2012-217014 324 

3.1.8 Millimeter-Wave Thermal Consultant Report—Design Reference Mission 1–C 

Over the past 60 years, the time-average power output of high-power microwave sources in the key 

millimeter-wavelength range has increased by over 6 orders of magnitude, for the first time putting Earthto-orbit launch within economic reach using gyrotron technology. Today, a 1-MW gyrotron oscillator can 

be purchased for about $2M. By combining the output of many such oscillators, the power and spot size 

needed to propel a vehicle to orbit are possible. 

 

Figure 3.105.—Left: The advent of gyrotron technology has raised the power output of sources in the key millimeterwavelength range by 3 orders of magnitude since 1970. Right: VGB–8125 gyrotron manufactured in Palo Alto by 

CPI Inc. This tube produces a 2.5-MW beam at 95 GHz. (Image courtesy of Communications & Power Industries 

(CPI); used with permission). 

Central to the microwave thermal rocket is a microwave-absorbent heat exchanger (HX) covering the 

underside of the vehicle. The refractory HX bypasses the specific enthalpy limits of conventional 

combustion, imparting 30- to 40-MJ per kilogram of H2 propellant as opposed to 16 MJ/kg for 

conventional hydrogen peroxide (H2/O2) combustion. This higher specific energy will halve the propellant 

flow rate needed to produce a given thrust and reduce the total propellant needed to put a given dry mass 

into orbit by a factor of 3 (because there is a compound effect as described by the rocket equation). 

Separate liquid hydrogen (LH2) and liquid oxygen (LOX) tanks will be replaced by a single LH2 tank—

likewise for the turbopumps and adjoining pipes. There will be no intertank structure, injector, or 

combustion chamber. 

The downside to an LH2 thermal rocket is the relatively low density of hydrogen, which will make the 

tank and turbopump larger and heavier than for dense propellants. The larger tank will add to the drag 

losses of the vehicle, requiring extra propellant to compensate. From a systems point of view, the key 

question is whether or not the benefits of higher specific energy and fewer components will be negated by 

the extra weight indirectly incurred by the low density of LH2. 

1.E‐05

1.E‐04

1.E‐03

1.E‐02

1.E‐01

1.E+00

1.E+01

1.E+02

1.E+03

1.E+04

1.E+05

1.E+06

1930 1940 1950 1960 1970 1980 1990 2000 2010 2020 2030

Pf2 [MW.GHz2]

Year

Gyrotrons under

development

Gyrotrons

Klystrons

Magnetrons KIT, 170 GHz, 2.2 MW KIT, 170 GHz, 4 MW FZK‐CPI, 140 GHz, 0.9 MW

USSR, 326 GHz, 1.5 kW

USSR, 150 GHz, 22 kW
NASA/TM—2012-217014 325 

Figure 3.106—Microwave thermal launch system and ascent trajectory arising from DRM–1C. 

3.1.8.1 Description of Vehicle 

For the 2-week Design Reference Mission 1–C (DRM–1C) study, a single-stage-to-orbit (SSTO) 

baseline design point was provided by Dr. Parkin, along with detail on the technologies by which various 

aspects of the design could be achieved. Independently, the Collaborative Modeling for Parametric 

Assessment of Space Systems (COMPASS) team chose how to vary that design and made alternative 

technical choices to close the mass budget. In particular, COMPASS required that the mass budget close 

with at least a mass growth allowance (MGA) and dry mass margin (DMM) totaling about 30% of the 

rocket dry mass. As a result, they were inevitably driven to a larger scale SSTO than the baseline and also 

tried a two-stage-to-orbit (TSTO) configuration. Mass budgets converged in both of those cases. In the 

2 weeks available, they did not try a solid rocket first-stage TSTO or a LOX/LH2 SSTO comparison 

design computed under the same methodology. 

 

100 km

0 100 200 km

SUSTAIN 

FACILITY

BOOST 

FACILITY

15

Pitchover at 

40 km

Cutoff

(nosecone 

separation later 

at > 150 km 

altitude)

Launch, 40% 

throttle

Beam handover

Horizontal 

acceleration, 

100% throttle

Max. slew rate170 GHz Gaussian beam
NASA/TM—2012-217014 326 

TABLE 3.29.—THE THREE DESIGN POINTS ASSOCIATED WITH DRM–1C 

 Parkin baseline 

(expendable SSTO) 

COMPASS SSTO 

design point 

COMPASS TSTO 

design point 

Stage 1 

Type 

Initial mass (kg) 

LH2 thermal LH2 thermal LH2 thermal 

2,500 4,734 2,156 

Final mass (kg) 580 1,084 954 

Payload (kg) 45 45 240 

V (m/s) 10,378 10,471 6,050 

Final orbit (km) (optimal perigee) by 400 100 by 400 100 by (optimal apogee) 

HX Power (MW) 450 850 300 

Stage 2 

Type --------- --------- Star 27 

Initial mass (kg) --------- --------- 240 

Final mass (kg) --------- --------- 60 

Payload (kg) --------- ---------- 40 

V (m/s) --------- --------- 3,986 

Final orbit (km) --------- --------- 100 by 400 km 

 

Payload

C&DH

C\SiC Heat 

Exchanger

ACS

Turbopump

H2 Outlet (2200 K)

H2 Inlet (40 K)

7.65 m

3.00 m

0.96 m

10.17 cu-m Heat 

Exchanger Area
NASA/TM—2012-217014 327 

3.1.8.2 Baseline Components and Technologies 

In flowpath order, 

Tank and structures.—The tank will be a 3-m-diameter titanium tank of traditional construction. 

The length will be 6 m for the baseline and will be varied to accommodate changes in wet mass. The 

structural factor of safety is 1.25, and the baseline tank pressure is 3 bars. The structural approach is to 

transmit loads into this pressure-stabilized tank with the minimum of intervening structure, analogous to 

attaching weights to a balloon. 

Turbopump.—A single-stage LH2 turbopump with 117-bars discharge pressure will be used. 

Low-flux HX.—This will be a low-temperature low-flux backing layer that “mops up” stray heat that 

cannot be reflected. In essence, it will shield the LH2 tank from heat loads emanating from the primary 

HX and stray beam flux. The backing layer will be maintained at an average temperature of 300 K via 

cryogenic hydrogen (H2) microtubes running in parallel about a centimeter apart. The distance between 

the microtubes will be maximized by maximizing the reflectance and thermal conductivity of the backing 

layer. Hence, it was originally envisaged to be aluminum foil. The effective thermal conductivity could be 

increased by sticking a layer of commercially available pyrolytic graphite sheeting to the underside. 

Silver foil could be substituted for aluminum to increase the reflectance in areas of particularly high 

flux—for example, under the hot end of the HX. 

 

Figure 3.107.—Pyrolytic graphite was first produced in the late 1800s for lamp filaments. Left: Pyrolytic graphite 

sheeting. Right: Thermal conductivity in watts per meters kelvin of pyrolytic graphite sheet versus other metals 

(Panasonic PGS data sheet). Copyright ©2011 Panasonic Corporation of North America; used with permission. 

 
NASA/TM—2012-217014 328 

High-flux HX.—The high-temperature high-flux HX will sit at the focus of the incoming millimeterwave beam and will be conformal with the tank. For the baseline, it was based on a prototype woven 

composite developed by Teledyne and tested at the NASA Glenn Research Center (Ref. 1). As shown in 

the Figure 3.108, the prototype was woven from carbon fiber and infiltrated with silicon carbide. The wall 

thickness is 0.7 mm and channel diameter is about 1.2 cm. The inside channel roughness is about 40 m. 

Figure 3.108.—(a) Plan and section views of an angle-interlock architecture. In the loom, the warp yarns are moved 

up and down selectively by harnesses, whereas the weft yarns are inserted between them by a flying shuttle or 

rapier. The section views show how the loom motions cause warp yarns to pass right through the thickness of the 

resulting fabric, tying together layers of weft yarns. Thus, the angle-interlock architecture suppresses delamination. 

(b) Schematic of an adaptation of an angle-interlock architecture to form an integral set of rocket nozzle tubes. The 

tube walls have an internal architecture similar to that shown in panel (a), with the tops and bottoms of the tubes 

joined at nexus regions formed in the loom. (c) Formed C-SiC composite structure, showing integrally woven struts 

for connecting to the rocket body and details of the high-density matrix formed within the tube walls (Ref. 1). 
NASA/TM—2012-217014 329 

Figure 3.109.—Left: Conformal sections of woven H2 HX for rocket nozzle cooling. Right: Linear 

section of woven HX undergoing testing at Glenn. The tests entailed connecting the composite 

tubes to metal manifolds for the supply and return of 300-to 400-bar H2 coolant. One side of the 

panel was exposed to a hydrogen/oxygen combustion plume from a small rocket combustion 

chamber delivering a free-stream temperature approaching 3,300 K. Panels survived multiple 

cycles simulating operating conditions of a boost rocket engine (cold-wall heat flux of 

20 MW m−2), with individual cycles as long as 8 min and cumulative times of 40 min. 

Attitude Control System (ACS).—Attitude 

control will be achieved in a similar way to 

conventional “cold” gas attitude control system, 

except that the propellant gas will be either 

warm H2 discharge from the turbopump or hot 

H2 bled from the HX outlet. The ACS subsystem 

controls will roll such that the HX will not rotate 

out of view of the beam, and it will provide a 

pitch-over maneuver that will transition the 

rocket from vertical atmospheric ascent to 

horizontal acceleration once atmospheric drag 

becomes low enough to do so. 

Figure 3.110.—Reaction control system 

(RCS) pods. 

Nozzle.—The baseline nozzle was a conventional bell nozzle. Owing to the low specific impulse 

during atmospheric operation, it became clear that some form of altitude-compensating nozzle was 

needed for the SSTO concept. Options were dual-bell, expansion-deflection, and plug nozzle. Based on 

experimental data suggesting superior performance (Ref. 2) and the recent testing of graphite plug nozzles 

by Garvey (Figure 3.111), it was decided to use a carbon-composite plug nozzle. It should be noted that 

the stagnation temperature of the propellant as it travels through the nozzle is less than or equal to the 

peak achieved at the exit of the HX. Ergo, any uncooled materials scheme used for the HX should also be 

usable for the nozzle; no cooling will be necessary. 

RCS Pods
NASA/TM—2012-217014 330 

Figure 3.111.—Graphite annular aerospike/plug nozzle developed and tested by Calstate Long Beach/Garvey 

Spacecraft Corporation in 2003. Copyright Garvey Spacecraft Corporation; used with permission. 

Command and data handling (C&DH).—For the baseline design, the command and data handling 

subsystem is envisaged to be located in the nosecone under the payload. Recent demonstrations of cell 

phone technology in the context of satellite and rocket applications have highlighted that the C&DH 

subsystem could potentially be built for very low mass, power, and cost. This is not to say that a cell 

phone would actually be used, more that the aerospace industry standard hardware is costly and 

cumbersome compared to what it could be. To facilitate fair comparisons, the baseline C&DH subsystem 

was sized according to COMPASS usual practices. Following techniques used for electromagnetic 

isolation in the Department of Defense (DOD) Transient Electromagnetic Pulse Emanations Standard 

(TEMPEST) shielded-enclosure handbook, the team used filters to remove any 170-GHz beam energy 

that might couple into antenna wires prior to entering the C&DH enclosure. 

Nosecone and payload.—A tangent Ogive nosecone was baselined for two reasons. First, it has 

lower drag and better represents the aerodynamics of this design than the X-33 drag model originally used 

by Parkin. Second, the drag model was represented by simple closed-form equations (Ref. 3). 

The payload will be enclosed by the nosecone, which will also act as a Faraday cage, protecting 

electronics from stray beam energy. Millimeter-wave radiation at 170 GHz is qualitatively similar to far 

infrared in that it is line-of-sight and does not wrap around obstacles in the way that radio waves do. The 

skin depth in aluminum is 0.2 m, meaning that the millimeter waves are completely stopped by even 

ordinary kitchen foil. The only way for 170-GHz radiation to leak into the payload enclosure is via 

unshielded wires (as already discussed) and via edges/seams that do not make proper electrical contact 

with their opposing surface. If the amount of leakage via seams proved to be a problem, the seams could 

be electrically sealed using conductive gaskets or paste. 
NASA/TM—2012-217014 331 

3.1.8.3 Figures of Merit and Margins 

The figures of merit for DRM–1C are closely related to the choice of design margins. Keeping design 

margins builds underperformance into the rocket design in order to guarantee a given payload with a 

certain statistical confidence. This confidence can be traded with performance and ultimately reflects the 

risk tolerance of the customer. The COMPASS team’s ~30% margin on dry mass was based on the AIAA 

S–120–2006 industry standard for an authorization to proceed (ATP) situation and was supported by a 

subsequent study examining historical mass growth (Ref. 4). Under these guidelines, 30% and above 

represents a green light, 19% to 30% a yellow light, and 0% to 19% a red light. 

The decision to proceed on a conventional multistage rocket after a 2- to 5-month conceptual design 

study is not the same as the decision to proceed on an SSTO thermal rocket after a 2-week feasibility 

assessment. Despite this, the microwave thermal rocket will perform well when held to a conventional 

standard. 

TABLE 3.30.—FIGURES OF MERIT FOR THE MICROWAVE THERMAL 

ROCKET VERSUS OTHER CONCEPTS FOR COMPARISON 

[RASCAL, RP1, rocket propellant 1.] Parkin baseline COMPASS SSTO COMPASS TSTO RASCAL 

baseline LOX/LH2SSTO LOX/RP1 SSTO Pegasus Scout G 

Figures of merit 

Stages 1 1 2 3 1 1 4 4 

Type (E = expendable, R = reusable) E E E R 1st R R E E 

Growth allowance on dry mass 

excluding payload (%) -- 

31.5 

15 

0

33.4 

15 15 

30 

15

0

30 

15

0

-- -- 

Wet mass (tons) 2.5 4.7 2.2 7.2a 1,065 1,693 19 a 21 

LEO payload (kg) 40 

40 

189 

338

40 

63 68 

–18,658b

680 b 

21,825 b

-635 b

9000 b 

18,644 b

375 210 

Payload fraction (%) 1.8 

1.0 

4.0 

7.1

2.0 

2.9 0.9 a 0.6 b 

2.3 b

0.5 b 

1.1 b

2 a 1 

Structural mass/payload 13 

23 

4.7 

2.2

18.4 

8.79 15.1 a 15.1 b 

4.6 b

7.1 b 

3.5 b

6.4 a 12.7

 a

Not counting airbreathing first stage. 

b

Payload mass adjusted from original to provide given margin. 

Numbers generated by the COMPASS team are shown in green. The COMPASS SSTO point design 

is similar to the RASCAL point design (not counting the airbreathing first stage) in wet mass (4.7 vs. 7.2 

tons), payload (40 vs. 68 kg), payload fraction (1.0% vs. 0.9%), and the cost-related metric of structure to 

payload mass ratio (23 vs. 15.1). It has the clear advantage of one stage versus three for RASCAL. A 

single stage simplifies range constraints and eliminates the cost, complexity and risk associated with 

staging. 

However, that comparison is not an equal one. Table 3.31 shows the margins typically used for 

concepts that the microwave thermal rocket would be compared to. For example, the RASCAL baseline 

point design carries a 15% dry mass growth allowance, whereas the COMPASS point designs are 

encumbered by more than a 30% dry mass growth allowance. For an “apples-to-apples” comparison, the 

values shown in blue were adjusted such that additional margin beyond 15% was counted as payload. 

With this adjustment, the COMPASS SSTO design clearly outperforms RASCAL, having quadruple the 

payload fraction (4.0% vs. 0.9%) and one-third the ratio of the structural mass to the payload (4.7 vs. 

15.1). 

 
NASA/TM—2012-217014 332 

TABLE 3.31.—MARGINS OF OTHER CONCEPT STUDIES 

 Total margin used 

on dry mass 

excluding payload 

(%) 

Study type Reference 

RASCAL bBaseline 15 Reusable 1st

stage 

5 

RASCAL GT 10 

NASA Access to Space Study (includes SSTO designs) 10 SSTO 6 

Because the growth allowance methodology masks the underlying performance of the microwave 

thermal approach, the values shown in red are adjusted such that all margin is payload. This corresponds 

to a totally risk averse “build it and see if we get a positive payload mass” approach. Under this 

methodology, the payload fraction is predicted to be 7.1%. This is greater than the Saturn V U.S. record 

payload fraction of 4% to low Earth orbit (LEO), lighter than the Saturn V liftoff mass by 640 times, and 

does so using only a single stage instead of three. 

3.1.8.4 Technology Margins 

Even greater performance will be possible if the impact of various technological options not 

considered in the feasibility studies are taken into account. Technology margins are margins that would be 

realized as increases in mass margin if changes in the technologies used were made. The technology 

margins for the microwave thermal rocket are currently greater than for a conventional rocket at this stage 

of a decision to proceed for three reasons: First, microwave thermal rockets are an almost conceptually 

unexplored approach, and the technology options and performance limits are not well understood or 

quantified; second, the COMPASS conceptual design activity was a 2-week activity, as opposed to the 

2- to 5-month activity typical of missile conceptual designs (Ref. 3); third, some higher performance 

approaches were not taken in order to keep the design closer to a documented “real-world” component 

with demonstrated performance in the desired range. This was the case for the choice of tank material, 

HX fabrication methodology, and non-subcooled propellant. 

For the compass SSTO point design in Table 3.30, each 10% gain in dry mass margin in Table 3.32 

can be used in any of the following ways: 

 Kept as increased margin (increased confidence in final payload mass) 

 Increased structural mass for reduced fabrication costs 

 Increased payload by about 100 kg (2% increased payload fraction) 

 Reduced wet mass by roughly 10% (375 kg) and corresponding decrease in beam power and 

scale 

 A combination of the above 

 
NASA/TM—2012-217014 333 

TABLE 3.32.—TECHNOLOGY MARGINS PRIORITIZED BY GREATEST POTENTIAL IMPACT FIRST 

Element 

DMMa gain 

(%) 

From 

To 

Description of potential gain 

Propellant 

<30 

LH2

NH3, N

H2 4 or CH4

A rocket-equation-based scaling analysis conducted by Parkin suggests that some dense 

propellants could provide up to 30% greater dry mass margin than LH2 for missions at 

~10.5-km/s V. The full 30% is for propellants that are heated to 3000 K (i.e., the effect is 

larger as the HXs are able to operate hotter). In addition, hydrazine (N2H4)adds mission 

flexibility as a monopropellant in its own right and methane (CH4) and other nontoxic 

noncryogenic hydrocarbons probably can be heated by a molecular mechanism in the bulk 

propellant itself to achieve temperatures exceeding 3000 K at the expense of fouling concerns 

and wall materials to investigate. Ammonia (NH3) will provide the greatest performance of the 

three for a given temperature. Lithium hydride (LiH) and ammonium fluoborate (NH4BH4) 

have the best performance of the “exotic” propellant possibilities but this class of propellants is 

unlisted because they are assumed to be too toxic or uneconomical. 

Airbreathing 

first stage 

20 

Preliminary investigation by Parkin finds that changing the initial altitude of the rocket from 

2 to 20 km increases the rocket mass at cutoff by about 20%. Other advantages to an 

airbreathing first stage are (1) no boost beam facility is needed (only sustain), (2) it is easier to 

achieve a range of orbital inclinations, (3) oxidation of the HX coating is reduced, (4) a 

conventional nozzle may suffice instead of a plug nozzle, and (5) it should be possible to run 

the 3 Meter Target Microwave Thermal Rocket (MTR) at 100% throttle for its full trajectory, 

so no throttle-down maneuver is needed. 

Propellant 

15 

Not 

subcooled

Subcooled 

For LH2, subcooling from 20 to 14 K densifies the tank mixture into slush with about 20% 

higher density than conventional LH2. Paradoxically, it may reduce the insulation needed on 

the tank because during ascent, more heat leakage is needed into the tank to melt the solid 

phase and maintain tank pressurization. This effective 20% margin on propellant translates to 

about a 15% additional margin on dry mass. 
NASA/TM—2012-217014 334 HX coating ~15 SiC ZrO2

Zirconia (ZrO2) is already used as an oxidation resistant coating in rocket nozzles. For 

example, the Scout missile uses a flame-sprayed zirconia coating to protect its ZTA graphite 

nozzle (Ref. 7). At refractory temperatures, zirconia has similar millimeter-wave absorption 

properties (resistivity) to silicon carbide, but it potentially increases the HX peak temperature 

from >2250 to >2700 K. Continuous-use temperature for fused stabilized ZrO2 coating is 

quoted as 2670 K with good resistance to thermal cycling. We plan to investigate zirconia 

coatings in the current 10-kW and future 1-MW static test series. If borne out by experiment, it 

would raise the Isp from 800 to 900 s, in turn increasing the dry mass margin by about 15%. 

Figure 3.112.—Left: Scanning electron microscope (SEM) image (cross-section) 

of ZrO2 interface coating on carbon fibers (white rings around each fiber) showing 

uniformity of coverage achieved by the ultraviolet chemical vapor deposition 

(UVCVD) technique (Ultramet.com). Right: SEM image showing fiber pullout in 

ceramic matrix composite with ZrO2 interface coating (Ultramet.com). 

Structures 

<15 

Unoptimize

d

Optimized 

Excluding the tank, the next major structural masses for the COMPASS design were incurred 

by the thrust structure and pipes at the aft end of the vehicle. HX supports and nosecone masses 

were secondary to this. In particular the aft structure was heavier than budgeted for in the 

baseline and there was not the time or documentation available to determine why. A thoughtful 

and detailed design could result in up to a 15% saving in overall rocket dry mass. 

Ascent trajectory 

<10 

Nonoptimal 

Optimized 

The ascent trajectories obtained by COMPASS were optimized using the OTIS software, 

saving >400 m/s relative to the quasi-optimized trajectories computed by Parkin. However, the 

trajectories obtained display behavior suggesting that constraints given to the optimizer caused 

it to erroneously use a lower fraction of available throttle, resulting in an oversized propulsion 

system (or equivalently a longer downrange distance than needed). There was not time to 

reconcile the trajectories obtained from OTIS with physical reasoning about what the optimizer 

should be doing at any given stage of ascent. Consequently, it is possible that the accelerations 

required of the propulsion system could be reduced by a few g’s and/or more saved on the Vs, 

with the Parkin baseline being the worst case in this regard. Despite this, trajectory 

improvements are unlikely to contribute more than 10% to the dry mass margin. But they will 

make a large difference to turbomachinery throttling requirements and beam facility cost. 

Tank 

material 

9 

Ti 

Graphiteepoxy (Gr-Titanium (Ti) tank was chosen because Gr-Ep was at less than minimum gauge for a 700-kg 

wet-mass rocket. Depending on assumptions including what the minimum gauge actually is 

(0.76 mm is indicated in some literature), Gr-Ep could be used for designs of 10 tons wet mass 

and greater. In this case and excluding insulation, it would be about 5 times lighter than the 

equivalent Ti tank, adding about 9% to the dry mass margin. 
NASA/TM—2012-217014 335 HX fabrication method <8 Woven textile CVD / other 

The 0.7-mm channel thickness is an order of magnitude greater than needed to contain the 

channel pressure of ~100 bars (i.e., it could be more than 10 times thinner if fabricated by other 

means, such a chemical vapor deposition (CVD) or perhaps lithography or micromachining). 

The design space here is large. In the theoretical limit that HX mass were reduced to zero, then 

at most the dry mass margin would be increased by 8%. 

Enhanced heat transfer 

surfaces 

~7 

Smooth/rough circular 

channels 

Heat-transfer-enhancing 

“tricks” 

Enhanced Heat Transfer is an entire field of techniques used to reduce the temperature 

difference between the bulk flow and channel wall, usually at the expense of greater pressure 

drop across the HX. At the hot end of the HX, enhanced heat transfer techniques have not been 

factored into the conceptual design but will certainly be combined with thinned tube walls and 

“high k” fibers to reduce the temperature differential to < 200 K and at best within 50 K of the 

absorbing material thermal limit. A resulting 200-K increase in propellant temperature 

increases Isp by about 5% and the dry mass margin by about 7%. In addition, the flow regime 

can adjusted to optimize the tradeoff between heat transfer performance and pressure drop. 

This has unquantified performance advantages felt via the increased nozzle chamber pressure 

and/or decreased turbopump pressure requirements. 

HX temperature 

throttling 

<6 

Constant peak 

T

Variable peak 

T

The HX would be operated at reduced temperatures (T; e.g., 2250 K) in low atmospheric 

conditions to minimize oxidation erosion of the absorbing layer, increasing toward the 

melting/decomposition temperature of the layer orbital altitudes. For example with zirconia, the 

melting point is 3000 K, corresponding to a vacuum Isp of at most 950 s. The extra 50 s of Isp

averaged over the ~8 km/s of acceleration occurring in near vacuum corresponds to at most a 

6% increase in dry mass margin above that gained by using zirconia in the first place. 

Nevertheless, it is a gain that will be realized in the natural course of thruster evolution. 

Tank insulation 

<3 

SOFI 

State of the art (SOA) or no 

insulation 

Spray-on foam insulation (SOFI) may be easy to apply, but in principle, it would be 

outperformed by aerogel or fiberglass insulation. Preliminary thermal balances suggest that no 

insulation at all is required during ascent, with insulation only needed on the pad to prevent ice 

buildup. There are schemes by which ice buildup may be prevented on a bare or lightly 

insulated tank: for example, by external heating. Also, recent progress in superhydrophobic 

coatings (Ref. 8) could significantly reduce ice buildup and increase the amount that falls off at 

launch. Finally, a helium balloon/tent could surround the launch vehicle on the pad and could 

be rapidly retracted just prior to launch. There is a liftoff acceleration of >3g, as opposed to a 

more conventional 1.2g, so the initial ice weight might have a comparatively small effect on 

overall performance. Since a lightweight externally heated SOFI scheme was accepted by 

COMPASS for use in the study, the most that remains to be saved is about 3% of dry mass. 

HX wall 

thickness 

2 

0.7 mm 0.5 mm 

This decreases HX mass by 25% (rocket dry mass by about 2%) and temperature drop across 

channel wall by 30% at the expense of slightly higher porosity. The 0.5-mm figure is a lower 

limit given in Jan. 2011 by Teledyne, makers of the 0.7-mm-thickness channel. 

Tank structural safety 

factor

1.5 

1.25 

1.1 

This is a traditional fallback, at the expense of trickier tank fabrication. This would decrease 

tank weight by about 10%, but because the tank itself is very light, it would only increase 

overall rocket dry mass margin by about 1.5%. 

Pyrometry 

<1.5 

Convention

al 

Woven 

thermocoupCOMPASS budget mass for >100 thermocouples and associated wire, primarily for the HX. 

Teledyne advises that tiny thermocouple strands can be woven into the HX. 

Pyrometry 

<1.5 

Thermocou

ples 

Remote 

optical 

If the HX was tracked optically using a telescope that tracks with the beam facility, then 

imaging optical pyrometry could be used to follow the thermal performance of the HX. Such a 

system will be used in static testing and it is envisaged that this will be adapted to short-range, 

and then long-range, flight. This will save most mass and data associated with onboard thermal 

diagnostics. 
NASA/TM—2012-217014 336 Turbopump <1.5 100 bars 160 bars 

Previous experience with microwave thermal rocket system analyses indicates that 

performance is maximized when the maximum pressure rise for a single turbopump stage is 

used. Humble indicates that this limit is 160 bars, athough this figure may be out of date. A 

lower pressure was used for the study baseline because it corresponded to an existing point 

design published for a nuclear thermal rocket LH2 turbopump. As it turns out, the COMPASS 

team member who sized the turbomachinery was a coauthor of that paper. Because full 

flowpath iterations were not really possible in the time allowed, the original pressure 

assumption was not varied. 

Aerodynami

cs 

<1.5 

Worst-case boattail drag was assumed. The aerodynamics has not been optimized. For 

example, neither the impact of the nose fineness ratio nor of the fins (to increase HX area) has 

been investigated. Also, the addition of a nose aerospike similar to that used on Trident 

missiles will reduce supersonic drag. 

Nozzle 

<1.5 

The optimal truncation length for the plug nozzle is not currently known. This also relates to 

boattail drag. 

a

Differential Mass Margin. 

3.1.8.5 Payload

The payload for all cases was defined to be a 40 kg ORBCOMM-like satellite delivered to a 400- by 

400-km orbit. 

Because the microwave thermal ascent trajectory does not lend itself to achieving fully circular 

trajectories, the payload itself provides a small delta V to circularize the orbit. If, like ORBCOMM, the 

payload has no primary propulsion system, a Star 5A solid rocket weighing 4.6 kg could be added to 

provide the necessary impulse over a 30-s burn. 

TABLE 3.33.—PAYLOAD DESCRIPTION FOR EACH OF 

THE DESIGN POINTS ASSOCIATED WITH DRM–1C 

 Parkin baseline 

(expendable SSTO) 

COMPASS SSTO 

design point 

COMPASS TSTO 

design point 

 Type (optional) Star 5A*1 or 2 (optional) Star 5A Star 5A Payload

Initial mass (kg) 45 or 50 45 45 

Final mass (kg) 43 or 46 43 43 

Payload (kg) 40 40 40 

V (m/s) 100-300 87 87 

Init. orbit (km) (optimal apogee) by 400 

km 

100 by 400 km 100 by 400 km 

Final orbit (km) 400 by 400 km 400 by 400 km 400 by 400 km 
NASA/TM—2012-217014 337 

In the baseline design, the circularization V is between 100 and 300 m/s, as opposed to 87 m/s 

because the perigee of the initial orbit is left unconstrained. Though lower payload V appears to be more 

desirable, subsequent analysis by Parkin suggests that it is less optimal for the rocket because an apogee 

of ~ –200 km naturally arises by optimizing a gravity turn followed by horizontal acceleration at full 

throttle. Consequently, the difference in optimizer constraints caused COMPASS’s ascent trajectory 

solutions deviate far from the desired behavior of Parkin’s solution, and they are not optimal with respect 

to maximizing payload mass. 

3.1.8.6 Description of Ground Facility 

The advent of submillimeter-wavelength astronomy has highlighted the existence of locations with 

particularly low atmospheric water content, opening up new microwave transmission windows from 35 to 

300 GHz and sometimes beyond. The atmospheric scale height of water vapor is only 1 to 2 km, allowing 

mountain ranges to create “rain-shadow” regions with comparatively good millimeter-wave propagation. 

As seen in Figure 3.113, the Himalayas hold back the water vapor, making Tibet an ideal location for 

China. The Alps are suitable for Europe, and the Andes suitable for South America, with the Chilean 

Atacama desert particularly prolific for submillimeter-wave astronomy. This resolution is insufficient to 

see particular mountains in Ecuador and Kenya that have both good propagation conditions and are close 

to the equator; these mountains are visible in the larger dataset. Also not easily visible are the islands of 

Hawaii, which are known to have excellent propagation characteristics. 

Figure 3.113.—Satellite data indicating the international locations where atmospheric water vapor content is lowest, 

and therefore millimeter-wave propagation is most efficient. Copyright European Space Agency (ESA); used with 

permission. 
NASA/TM—2012-217014 338 

For the continental United States (CONUS), Figure 3.113 clearly shows that the mountainous 

southwest of the country is better suited than the southeast, with Florida being essentially underwater as 

far as millimeter-wave propagation is concerned. Zooming in to the southwest, Figure 3.114 shows that 

the Sierra mountains east of California’s Central Valley provide some of the best propagation 

opportunities. Ongoing site surveys and millimeter-wavelength projects, such as the Combined Array for 

Research in Millimeter-wave Astronomy (CARMA) in eastern California, are revealing many suitable 

locations for a beam facility on the U.S. mainland. Because the southwest is relatively sparsely populated 

and contains large regions of government land, there are several ways in which a 200+ km-long ascent 

trajectory might be constructed. For polar orbits, Alaska appears to provide many of the best sites. 

Figure 3.114.—Examples of CONUS sites with low losses due to water vapor absorption. Such sites are 

predominantly in the southwest. For Mt. Evans and Pikes Peak, a millimeter-wave beam vertically transmitted to 

space would lose 5% to 10% of its energy to atmospheric absorption at 170 GHz.

The ascent trajectory in this DRM consists of two segments: In the first, at the bottom left of Figure 

3.115, the vehicle is powered by a short-range “boost” beam facility, maintaining vertical ascent of the 

rocket at 40% throttle to minimize drag losses. The first segment transitions to the second at an altitude of 

30 to 60 km. In this time, the boost beam facility hands the rocket over to the “sustain” beam facility as 

rocket levels off and begins to accelerate horizontally at 100% throttle. The acceleration of 9g to 19g 

during this segment of the ascent raises the velocity from 1.5 to 8 km/s in only 60 s and enables the rocket 

to achieve orbital velocity within 125 km range of the beam source. This was to minimize the initial cost 

of the beam facility and could eventually be expanded to allow a gentler ascent trajectory suitable for 

human launch. 
NASA/TM—2012-217014 339 

Figure 3.115.—Beam facility arrangement along an ascent trajectory. The red line is the ascent trajectory. During 

initial ascent, a boost beam facility with a maximum range of 60 km propels the vehicle, which is then handed over 

to a larger sustain facility with a 125-km range and higher power. Programmatically, a subscale demonstration 

beam facility would probably become the boost facility, and the decision to proceed on the sustain facility would be 

based on its successful operation. The blue line is an alternate ascent trajectory considered by the COMPASS 

team. Because twice the range would require twice the source or target diameter, it was not further considered for 

these point designs. 

The size and power of the boost and sustain beam facilities for each of the design points is given in 

Table 3.34. Tsiolkovsky originally predicted that a microwave beam facility would need to be 12 km 

across, but owing to the shorter wavelength and high acceleration ascent trajectory, the beam facilities of 

DRM 1–C will be closer to 100 m in diameter, comparable in size to existing millimeter-wave telescopes. 

 

Boost

Sustain
NASA/TM—2012-217014 340 

TABLE 3.34.—BEAM FACILITY POINT DESIGNS 

 Parkin baseline 

(expendable SSTO) 

COMPASS 

SSTO 

COMPASS 

TSTO 

Boost 

Primary diameter (m) 58 58 58 

Maximum range (km) 60 60 60 

Output power (MW) 350 665 300 

Downrange distance (km) 15 15 40 

Beam spillage (%) 20 20 20 

Sustain 

Primary diameter (m) 122 122 122 

Maximum range (km) 125 125 125 

Output power (MW) 870 1,650 800 

Downrange distance (km) 140 140 40 

Beam spillage (%) 20 20 20 

Although the 1990s-era Green Bank Telescope (GBT) aperture is about the right size for application 

to a microwave thermal rocket, it is the current era of closed-loop active surfaces that provide the 

necessary surface accuracy. In fact, surface accuracies now exceed those needed for the microwave 

thermal rocket by nearly an order of magnitude, albeit for smaller apertures. The key question to answer 

in a beam facility conceptual design will be what combination of structural rigidity and active surfaces 

can correct for changing gravitational distortions and mechanical bending modes as the structure slews; 

sudden thermal deformations could be negated using heaters. If greater mechanical stiffness is required, 

there are Fresnel reflector and even holographic techniques that would be unsuitable for astronomical 

work but may be advantageous for beam direction, provided they turn out to be compatible with the 

spectral-beam-combining approach used in the optical back end. 

Active surfaces aside, the beam facility could provide pointing on three different temporal and 

angular scales to track a target HX through its ascent trajectory. The first scale is provided by the primary 

mirror, with low angular acceleration and slew rate but a range of motion from –40 to +40. The second 

scale of pointing would be provided by moving the secondary mirror relative to the primary, and the third 

and finest scale of pointing could be provided in the feed system for each gyrotron beam by means of a 

fast-steering mirror, such as that pictured on the left of Figure 3.117. Mirrors of this kind are used to 

stabilize sections of the optical train in the ABL shown on the right of the figure and, in principle, could 

be adapted to perform fine pointing of quasi-optical gyrotron beams prior to their transformation through 

the secondary and primary mirrors. 
NASA/TM—2012-217014 341 

Figure 3.116.—Three generations of millimeter-wave telescopes. Top: 1990s-era GBT (left) and secondary (right). 

Middle: 2000s-era Atacama Pathfinder Experiment (APEX) at Llano Chajnantor (courtesy of APEX; used with 

permission). Viewed here at dusk, the panels and retroreflectors of its active surface are clearly visible. The 

telescope was manufactured by VERTEX Antennentechnik in Duisburg, Germany. Bottom: 2010s-era CornellCaltech Atacama Telescope (CCAT) (Courtesy TMT Observatory Corporation; used with permission). 
NASA/TM—2012-217014 342 

TABLE 3.35.—COMPARISON OF APERTURE PARAMETERS 

GBT Atacama Large 

Millimeter Array (ALMA) 

CCAT MTR at 120 km 

Optical 

Arrangement 

Parabolic off-axis 

reflector 

Cassegrain quasi-random 

array Cassegrain Cassegrain 

Diameter of 

primary (m) 100 by 110 12 by 12 and 7 by 7 25 by 25 120 by 120 

FWHM beamwidth 

(arcsec) 5.3 at140 GHz 5.2 at 170 GHz 

Pointing accuracy 

(arcsec) 2.7 (open loop) 2 2-0.5 (blind) 1 (closed loop) 

RMSa surface 

accuracy (μm) 

450 (passive) 

390 (open loop) 

68 (intra-panel) 

17.2 measured by nearfield radiometric 

holography (Ref. 9) 

<12.5 

(closed loop) 

88 

(closed loop) 

Minimum elevation 

angle (deg) 5 5 40 

Maximum slew rate 

(deg/s) 0.3 (elevation) 0.2 to 1 4 

Cost FY1990 $97M 

(~$12 K/m2

) 

FY2010 $1.3Bn 

(~$7.5 K/m2) 

FY2020 

$110M FY2010 <$1Bn 

Construction time 11 years (1990–2001) 2003–2011 6 year < 10 year 

a

Root mean square. 

 

Figure 3.117.—Left: 12-cm-diameter fast-steering mirror for high-power laser applications (image courtesy of ATA; 

used with permission). A 30-cm-diameter model features >1-kHz open-loop bandwidth, >300 rad/s2

 angular 

acceleration, <0.2-μm surface accuracy, and 8-mrad mechanical travel with <10-μrad (2-arcsec) pointing 

accuracy. Right: YAL–1A ABL prototype in flight, 1.5-m telescope, megawatt-class 1.3-m laser with undisclosed 

range . 
NASA/TM—2012-217014 343 

Figure 3.118.—Millimeter-wave beam facility optical arrangement.

For the most part, the frequencies needed for microwave thermal launch are readily generated using 

gyrotron or gyroklystron microwave sources. Gyrotron development has been spurred by its application to 

electron cyclotron heating systems in fusion reactors, and in particular the goal of a 4-MW continuouswave (CW) gyrotron oscillator at 170 GHz for the ITER fusion reactor. The advent of edge-cooled 

diamond windows in the late 1990s marked a significant advance in these high-power CW systems, and 

in March 2005 the commercially available 140-GHz CPI VGT–8141 gyrotron demonstrated 30-min 

pulses of nearly 0.9-MW output power. This continues a trend that has seen time-average power output of 

millimeter-wave sources increase by 6 orders of magnitude over the past 60 years. 

Prospects for the long term are focused on increasing the CW power output of coaxial gyrotrons together 

with greater reliability. Present design efforts for “super-power” tubes aim to achieve 4- to 5-MW CW 

output power at frequencies of 95, 130, 140, and 170 GHz. Diamond windows become a limiting factor at 

such high power levels, and future generations of devices are expected to use multiple diamond output 

windows. 

From a practical perspective, the present commercially available gyrotrons cost $2M/MW to 

$3M/MW and have a lead time of 2 years, depending on how much of the supporting equipment (e.g., 

power supply, optics, and cooling) is included. In low-production quantities (1 to 10 units), the costs 

mentioned earlier in this section include the gyrotron oscillator itself, a diamond output window 

(~$100K), a superconducting magnet system (~$500K), and a power supply (~$300K). 

A key parameter affecting the ultimate payload price of microwave thermal launch is gyrotron service 

life and maintenance overheads. Although the present generation of gyrotrons is commercially available, 

they are still undergoing refinement for high reliability. In this regard, the experiences of General 

Atomics with three different gyrotrons for fusion work give some indication of the present state of affairs: 

SIDE

TOP

PIVOT POINT (ABOUT WHICH DISH + ‘PIVOT REFLECTOR’ ROTATE AS 

SOLID BODY OUT OF PLANE OF DIAGRAM)

SECONDARY

GYROTRON HALL

GRATING

WAVEGUIDES

Sequence:

1. Each beam begins in a gyrotron

2. Emitted into evacuated corrugated waveguide 

(65W/m measured loss at 1 MW) and propagates to 

vicinity of grating

3. Each beam ‘launched’ into free-space Gaussian 

beam and exits waveguide via window 

1. (option of adding fast steering mirror at this 

stage)

4. Each beam of slightly different frequency hits 

grating at slightly different angle

5. Grating combines ‘rainbow’ into single Gaussian 

beam (grating can theoretically combine up to 2500 

sub-beams; will actually need O(100))

6. Beam hits pivot reflector (pivots as primary slews 

so gyrotrons don’t have to rotate with the dish)

1. Option of adding adaptive optics here

7. Beam hits secondary

1. Option of adding adaptive optics here

8. Beam hits primary

9. Beam hits target heat exchanger (not shown)

In general:

• Ohmic loss from each reflection is 0.1% for Al.

• Grating loss expected to be 1-5% (2-beam loss of a 

simple millimeter wave grating has been measured 

at <0.5%, n-beam loss experiments and analysis 

planned)

• To provide for various orbital inclinations, the entire 

arrangement including dish gyrotron hall will be 

built on a platform that can be rotated. This 

platform will be mounted on circular rails to hold 

and distribute the weight and provide power

• Depending on detailed analysis, the primary and/or 

secondary may be collapsed down to flat fresnel

reflectors
NASA/TM—2012-217014 344 

Of the three development gyrotrons, two repairable failures are described by Lohr et al. (Ref. 10). The 

first was a filament short that occurred after 5000 hr of operation, and the second was a braze failure 

causing a loss of vacuum. In both cases, the gyrotrons were repaired for about 10% of their original cost. 

Aside from these early problems, there is little fundamental difference between gyrotrons and other 

high-power vacuum devices such as klystrons as far as reliability and lifetime are concerned, and a mean 

time between failures of 30,000 hr is a reasonable expectation as the products mature. Lifetimes of only a 

few thousand hours were anticipated during the proposal to construct the Stanford Linear Accelerator; 

however, actual experience led to a mean time between failures now exceeding 50,000 hr. 

3.1.8.7 Concept of Operations (CONOPS)—Frequency of Launch and Repeat Time

There was insufficient time to investigate the operational and economic questions surrounding a 

microwave thermal launch system, in particular the flight rates needed. From an economic point of view, 

a CONOPS resembling airport operations with frequent routine operations is clearly better than a 

traditional spaceport with infrequent, one-off “spectacular” operations. ORBCOMM was chosen as the 

type of payload for which a large number of routine launches may be needed. The launch sequence for an 

expendable microwave thermal launch of an ORBCOMM is given in Table 3.36. 

TABLE 3.36.—CONOPS FOR ORBCOMM LAUNCH 

Time Event

T – 1 day 50-kg payload is integrated into rocket on pad

T – 30 min Rocket is enclosed in helium-filled tent and LH2 tank is filled

T – 5 s Tent and umbilicals are retracted

T + 100 s Rocket reaches 40-km altitude, transitions from boost to sustain beam facility, 

vertical ascent to horizontal acceleration, and throttles up from 40% to 100% 

T + 200 s Rocket cuts off into –360- by +400-km transit orbit

T + 10 min Nosecone separated and payload ejected

T + 45 min Payload circularization burn to 400 by 400 km (230 m/s, < 10 kg solid rocket)

3.1.8.8 Costs—Comments 

It is elasticity of demand that provides the incentive for launch prices to go down. If reducing the 

price of launch from $10,000/kg to $5,000/kg does not at least double demand, then there is no incentive 

to do it, and indeed this has not happened in 40 years. 

$10,000/kg is a metastable price level, with 1990s-era economic analyses predicting $600/kg to be the 

next stable price level because it corresponds to a predicted spike in demand and flight rate. Of course, 

$600/kg could only be attained if the cost price was less. This is not the case for conventional rockets, the 

cheapest foreign examples of which probably cost in the region of $2000/kg to make. Not even the 

economies of scale achieved in Minuteman production achieved $600/kg, and these quantities of 

production are unlikely to be reached again. 

If the economic barrier to space access is ever to be broken, it will most likely be by the first concept 

to achieve reusability, enabled by the first technology that achieves the highest margins at the smallest 

scale. According to the results of DRM 1–C and specifically Table 3.30, microwave thermal rockets are a 

candidate. 

The logical first step is an expendable small satellite launcher, which corresponds to a scale and 

power level that minimizes initial beam facility cost. There is yet some uncertainty about where this 

minimum lies and how low it goes, but it is certainly less than the cost of developing a heavy launcher. 
NASA/TM—2012-217014 345 

Currently, small launchers cost substantially more per unit mass than their larger counterparts. If 

improvements in vehicle technical performance can be translated into improvements in cost performance, 

this should have the greatest impact the soonest for small satellites. 

In reading the cost estimates associated with the DRM 1–C vehicle and beam facility, it is important 

to consider what the same estimation methodology applied to a conventional alternative might be, and 

whether the cost estimation methodology is likely to capture the efficiencies of a microwave thermal 

launch approach versus a conventional alternative. 

3.1.8.8.1 Beam Facility Costs 

If a 1-MW gyrotron and its power supply costs $1M plus an equal amount in lifetime maintenance 

and lasts for 20,000 hr of operation, then its average cost is 10 cents/kWh. In October 2010, year-to-date 

average U.S. retail prices of electricity were 9.91 cents/kWh. 

In other words, the cost of converting electricity to a millimeter-wave beam is comparable with the 

cost of the electricity itself. Rolled into the electricity price is the cost of the generator, step-up 

transformer, transmission line, step-down transformer and the personnel to maintain it. The gyrotron and 

its power supply are of similar complexity. 

The GBT is about the size needed for the microwave thermal beam facility. Its fiscal year 2007 

(FY2007) operating budget was $10.4M, which included 109 full-time equivalents (FTEs). 

3.1.8.8.2 Minimum Beam Facility Cost 

It is possible to estimate the size of a rocket that minimizes beam facility cost. This minimum cost 

occurs at a balance between two opposing expenses: one that gets smaller as the rocket gets larger, and 

one that gets larger as the rocket gets larger. 

The first expense is optics. The smaller the rocket to aim at, the larger and more expensive the dish 

needed on the ground. In fact, the relationship is inversely proportional: If the HX diameter is doubled 

(e.g., from 3 to 6 m), then the ground dish diameter is halved (e.g., from 120 to 60 m). The second 

expense is millimeter-wave power. The larger the rocket, though easier to aim at, the more power needed 

to propel it. 

When the cost of the aperture area equals the cost of the millimeter-wave power, the beam facility 

cost is minimized. On the basis of the assumptions in Table 3.37and if it is assumed that the HX size is 

equal to the tank size, the beam facility cost for DRM 1–C can be estimated versus vehicle wet mass 

(Figure 3.119, top). 

TABLE 3.37.—ASSUMPTIONS USED IN THE BEAM FACILITY COST CALCULATION 

Energy loss assumptions 

Gaussian beam spillage 20% 

Atmospheric absorption 10% 

Dish scattering/absorption 10% 

HX reflection/reradiation 20% 

Geometric assumptions 

Beam wavelength 1.76 mm (170 GHz) (λ/20 = 88 m) 

(HX length)/(tank length) 1 

(HX width)/(tank width) 1 

Tank length to diameter (L/D) ratio 1 (elliptical end caps) 

Tank ullage 10% 

Tank propellant density 68.9 kg/m3 (LH2, 22 K) 

Maximum beam tilt angle 60 along-track, 10 cross-track 

Maximum target range 125 km (at cutoff) 

Other assumptions 

Trajectory average specific impulse Isp 724 s (LH2) 

Vacuum specific impulse 800 s (LH2) 

V 10,378 m/s 

Peak acceleration 19g (just before cutoff) 
NASA/TM—2012-217014 346 

 

Figure 3.119.—Top: Beam facility cost based on assumptions in Table 3.37. Bottom: Effect of chines and greater 

spillage. 

10

1

10

2

10

3

10

4

10

5 10

8

10

9

10

10

Vehicle wet mass (kg)

10

1

10

2

10

3

10

4

10

5 10

8

10

9

10

10
NASA/TM—2012-217014 347 

The minimum beam facility cost is found by choosing an aperture cost ($k/m2) on the top and a 

gyrotron cost ($M/MW) on the bottom Figure 3.119, and following the line that connects them until a 

minimum is found. This minimum corresponds to a given wet mass and beam facility cost. The choices of 

gyrotron and aperture cost are taken from Table 3.38. 

The worst-case gyrotron cost is taken to be the current cost, and the best is super-power tubes 

combined with the effect of increased production. The lower bound on aperture cost is the materials cost, 

and GBT is the closest analogy for a 100 meter-class dish. The relation between technology factors and 

overall dish cost is not yet well understood. If a duplicate of the GBT were built today, it might be 

significantly cheaper because it is no longer cutting edge and could make use of carbon composites, but 

water cooling of the panels would increase cost as might upgrading the active surface to 80-m accuracy. 

It is not yet known whether the higher CCAT areal cost is due to its increased surface accuracy relative to 

GBT or because it is a single small dish and there is fixed overhead independent of size. The ALMA areal 

cost, which averages over fifty-four 12-m dishes and twelve 7-m dishes, suggests that economy of scale 

more than surface accuracy that drives these areal costs. 

TABLE 3.38.—ASSUMPTIONS USED IN THE BEAM FACILITY COST CALCULATION 

Gyrotron cost FY2010 $M/MW 

Current commercial off-the-shelf cost for ITER and others (Ref. 11) 3 

Assumed cost of 2- to 4-MW super-power tubes circa 2014 to 2020 1 

Cost of 100th gyrotron; initial 1M$/MW with 85% learning curve 0.25 

Assumed lower limit to vacuum tube approach: Retail magnetron cost 0.04 

Aperture cost FY2010 $k/m2

CCAT (2012, 25-m, 8-m closed-loop active surface) 225 

Millimeter-wave telescope rule of thumb 40 

GBT (1990, 100- by 110-m, 390-m open-loop active surface) 20 

ALMA (2010, many 12- by 12-m and 7- by 7-m, 17.2-m active surface) 7.5 

Fresnel reflector flat primary (feasibility analysis needed) 5 

25-m surface accuracy (ground) Al 6060 T6 panels 0.86 

Unground Al 6060 T6 panels 0.28 

As an example that corresponds to the 2500-kg baseline design, choosing the blue GBT aperture line 

at the top of Figure 3.119 combined with a $1M/MW estimate for gyrotron bulk cost gives a beam facility 

cost of $1.2Bn. This corresponds to 880 MW of millimeter-waves transmitted from a 128-m aperture to a 

2.8-m target. 

Pushing the envelope further, the bottom of Figure 3.119 shows how a more exotic vehicle might 

substantially decrease the minimum beam facility cost and the wet mass at which that occurs. The curves 

are generated using the assumptions of Table 3.38 except for a tank L/D of 6, (HX width)/(tank width) of 

3, and beam spillage of 40%. This corresponds to a longer, thinner tank attached to a high-pressure rigid 

HX that is triple the width of the tank. If we assume a GBT-class aperture cost and a decade of gyrotron 

power increases and cost reductions, the beam facility minimum cost drops to $150M for a vehicle of 

only 600-kg wet mass, corresponding to 280 MW of millimeter-waves transmitted from a 64-m dish onto 

a 4.4-m target. A rocket at this scale would probably require an HX areal density better than that of the 

woven composite example baselined for DRM 1–C. This example suggests that over time, better 

exploration of the design tradespace combined with evolutionary technology improvements in gyrotrons, 

millimeter-wave telescopes and HX will correspondingly reduce the initial cost and scale at which a 

microwave thermal launch system can be reduced to practice. 
NASA/TM—2012-217014 348 

3.1.8.9 Feasibility Assessment 

The following is a list of issues which need to be addressed 

1. What are the attenuation effects of high flux levels (MW/m2) at the frequencies of interest on the 

atmosphere? 

2. What would be the expected availability and transmittance of the BEP ground terminal at a given 

frequency, given aerosol and precipitation conditions at a proposed ground site? 

3. How large of a keep-away zone must be maintained around and beyond an ascending craft to 

accommodate reflection, refraction and scattering components of the beam exceeding substantial 

irradiances, and what level is to be considered substantial? Considerations should be given to 

maximum permissible exposure for humans, and damage thresholds for ground, airborne and 

space-based assets within proximity of the BEP activity. 

4. For a ground based terminal operations, how much grid power over what duration is required? 

To meet these requirements, what would the generation, storage and conversion system look like? 

5. How would a micro/mm-wave BEP system work within the spectrum constructs of the 

NTIA/FCC? 

6. Does it make sense to collocate beamed energy propulsion and power technologies on the craft, 

for each phase of a proposed mission? 

7. Would there be a benefit to developing a hybrid aperture to handle multiple BEP frequencies 

depending upon range to target, atmospheric conditions, etc.? 

8. Can retro-directive control be utilized to aid in the beam pointing, acquisition and tracking of the 

craft? 

9. What is the most practical way of performing high power beam combination using COTS 

technology for the frequencies of interest? What are the potential near term accomplishments that 

could influence this capability? 

10. What are the cost/benefit tradeoffs in considering multiple BEP transmission locations, both at 

the ground range and at an intermediate relay point in space? 

11. Could deployable array technologies by implemented to decrease the weight of a BEP aperture, 

and increase the allowable size that can be launched? 

3.1.8.10 Roadmaps and Technology Readiness Levels 

Part of the task in developing a technology roadmap has been to produce a way to best represent the 

technology readiness for DRM 1–C. It is particularly important to identify feasibility issues and to 

communicate the level of technical risk in each choice, so comprehensive technology profiles are 

important. 

In the TRL summary that follows Figure 3.120), the system elements of DRM 1–C are laid out by 

work breakdown structure (WBS). Alternate possibilities are denoted using A, B, C notations, where A is 

baseline and B, C, and so forth are fallback technologies ordered by preference, usually with lower 

performance than the baseline by some metric. To make the baseline plan clearer, the fallback options are 

grayed out. 

For many of the WBS elements, it was necessary to revert their TRLs back to TRL 3 even on highly 

developed technologies like mirrors simply because they will be operating in a high-power microwave 

environment. Since the TRL alone would incorrectly give the impression nobody has built a millimeterwave telescope before (for example), the format of the sheet has been modified to represent the degree to 

which analogy plays a role. 
NASA/TM—2012-217014 349 

Analogy is important because the process of choosing a technological path usually involves finding 

the closest analogy to the unit wanted and assessing what has to be changed in order to fit the new 

purpose. That change can be quantitative in nature, such as scaling-up a power supply to higher power. Or 

it can be qualitative in nature, such as changing the surface accuracy of a large dish from 5000 to 5 m, 

requiring a qualitatively different approach. It matters if a close analogy to the propellant tank has flown 

to LEO. So, an extra dimension has been added to the TRL summary by introducing a three-symbol 

system: 

x A unit fulfilling the application criteria exactly has passed this milestone. 

o A unit substantially analogous to the exact unit has passed this milestone. 

? Unclear at this time 

Also, there are “Quantitative” and “Qualitative” analogy fields to indicate what kind of analogy is 

being used. In general, the sheet is laid out so that gaps in key milestones are obvious and suggest a 

definite and traditional course of action, such as a detailed conceptual design, lab prototype effort, subscale demonstration, further survey of the literature and so forth. As technologies become ready for 

application to the microwave thermal rocket, the “bleeding edge” red blocks should disappear roughly 

from left to right. 
NASA/TM—2012-217014 350 

Figure 3.120.—TRL, risk, and uncertainty summary for the microwave thermal rocket. FSM, fast steering mirror; 

FADIS, Fast Directional Switch; KKV, kinetic kill vehicle; CVD, chemical vapor deposition; PSI, Pressure Systems, 

Inc.; and TPS, thermal protection system.

Notes

Concept formulatedProof-of-principlePerformance est'd.Comprehensive modelsProof-of-principleBreadboardSubsystem proto.System proto.Ground ops<1 km1 - 60 km> 60 km (suborbital)> 100 km (LEO)QualitativeQuantitativeQual'd for exact purposeFormal TRL

1 Beam Facility

1.1A Power supply: Pulse via flywheels x o o o o o o o ? - - - - ? x 3 60 kV to be demonstrated

1.1B Power supply: Pulse via NiCd cell array x x x o o o o o o - - - - ? x 2 Active denial uses more expensive Li-ion cells

1.1C Power supply: Grid via solid state regulator x x x x x x x x x - - - - - - x 9 In use by CPI and others

1.2A mm-wave sources: Gyrotron oscillator xxxxxxxx - - - - - - 6

1.2B mm-wave sources: Gyroklystron amplifier xxxxxxxx? - - - - - -? 6

1.3A Power combining: Spectral via diffractive element(s) x x x o o - - - - ? x 2 Analogy insufficient for 100 beam combiner

1.3B Power combining: Phase-locked oscillators xoo o - - - - xx 2

1.3C Power combining: Master oscillator power amplifier (MOPA) x o o o o o o - - - - x 6 Haystack RADAR

1.4A Optical front end: Single large cassegrain telescope x x x o o o o o o x x 3 GBT/ALMA/CCAT

1.4A.1 Secondary optic: Al hyperbolic reflector x o o o o o o o o x 3 GBT/ALMA/CCAT

1.4A.2A Primary optic: Al Fresnel reflector xoooooo ? x 3

1.4A.2B Primary optic: Al hyperbolic reflector x o o o o o o o o x 3 GBT/ALMA/CCAT

1.4A.3 Wavefront control: Segmented active surface (closed-loop) x o o o ? ? ? ? x 3 CCAT

1.4A.4 Coarse beam steering: Mechanical x o o o o o o o o ? x 3 GBT/ALMA/CCAT

1.4A.5A Fine beam steering: Mechanical, fast steering mirror(s) x o o o o o o o o o o ? ? 3 FSMs used on ABL at 1.3 microns

1.4A.5B Fine beam steering: Spectral / diffractive x o o ? ? ? ? ? ? x 3 Analogy is FADIS

1.4B Optical front end: Multiple small cassegrain telescope x x x o o o o o o x 3 Analogy is ALMA. Fed by phase-locked or split s

1.4B.1 Secondary optic: Al hyperbolic reflector x o o o o o o o o ? 3 GBT/ALMA/CCAT

1.4B.2 Primary optic: Al hyperbolic reflector x o o o o o o o o ? 3 GBT/ALMA/CCAT

1.4B.3 Wavefront control: Segmented active surface (open-loop) x o o o o o o o o ? x 3 Analogy is GBT/ALMA

1.4B.4 Coarse beam steering: Mechanical x o o o o o o o o ? x 3 GBT/ALMA/CCAT

1.4B.5A Fine beam steering: Mechanical phase shifters x o o ? o ? ? ? ? x 3 Was used in the past according to B.Bridges

1.4B.5B Fine beam steering: Mechanical fast steering mirrors x o o o o o o o o o o ? ? 3 FSMs used on ABL at 1.3 microns

1.5A Pointing knowledge: Coherent doppler transponding xoooooooooooo x 3

1.5B Pointing knowledge: Closed-loop IR optical tracking x o o o o o o o o o o o o ? x 3 KKV

1.5C Pointing knowledge: Retrodirective beam control x o o o o o o o o o o o o ? x 3 1990's JPL experiment at Goldstone

2 Vehicle

2.1 Tank: Ti x x x ? ? ? ? ? ? ? ? ? ? ? x 3 PSI Ti Tanks

2.2A Heat exchanger: Textile C/SiC x o o o o o ? ? ? ? ? ? ? x x 3 Teledyne prototypes

2.2B Heat exchanger: Extruded/Sintered/CVD SiC segments x x x o ? ? ? x 3 Reaction Engines SiC HX

2.3A Nozzle: Uncooled C/SiC xooooooo ? ? ? ? ? x 3

2.3B Nozzle: Regeneratively-cooled C/SiC xo ? ?o ? ? ? ? ? ? ? ? x 3

2.4A Pressurization: Turbopump, autogeneous xxxoooooooooo x 3

2.4B Pressurization: Blowdown, autogeneous xxooooo ?? ??? ?? x 3

2.5A Avionics: Cell phone-derived x o o o x x 2 PhoneSAT

2.5B Avionics: Conventional xoooooooooooo x 3

2.6A Attitude control: LH2 warm/cold gas xooooooo ? ? ? ? ? ? x 2

2.6B Attitude control: Gas generator xoooooooooooo x 3

2.7A Aeroshell: Expendable x ? ? ? ? ? ? ? ? ? ? ? ? x 3 Anything else that needs to be added

2.7B Aeroshell: Reusable x o o o o o o o o o o o o x x 3 Analogies are shuttle, X33, X34, X37

2.7B.1A TPS: Woven load-bearing C/SiC shell xo ? ? ? ? ? ? ? ? ? ? ? xx 1

2.7B.1B TPS: Carbon foam panels x o o o o o ? ? ? ? ? ? ? ? 3 Ultramet C foam panel tested at GRC

2.7B.1C TPS: Conventional best-practise xoooooooooooo ? x 3

2.7B.2 TPS: Integrated cryo-insulation x o o ? o o ? ? ? ? ? ? ? ? x 2 Listed as TRL 2 in SOA review paper 2001. See als

2.7B.3A Vehicle recovery: Landing gear xoooooooooooo x 3

2.7B.3B Vehicle recovery: Parachute xoooooooooooo x 3

2.7B.3C Vehicle recovery: Precision thruster docking above ground x o o o o o o o o o 1 Analogy is Pixel

2.8 Payload storage and release mechanism: TBD ???????????????

Relevant 

environment

MICROWAVE THERMAL ROCKET 

TECHNOLOGY RISK AND 

UNCERTAINTY SUMMARY

Theory Practise

Diff. 

from 

analogy

1) Alternate options are denoted by A, B, 

C etc. and presented in order of baseline, 

fallback plan B, plan C etc.

x = exact unit capable of passing qual

o = analog of unit not capable of pass qual

- = not applicable
NASA/TM—2012-217014 351 

Figure 3.121.—Microwave thermal launch roadmap.

3.1.8.11 Ground and Flight Demonstrations 

To date, the development approach has been a series of tests progressing to higher power levels and 

greater realism. The primary objective of these tests is to identify and retire risks as early as possible in 

the program, and the corresponding milestones are given in the sections that follow. The secondary 

objective is use high-profile milestones amenable to media coverage to attract, involve, train and motivate 

a cadre of talented engineers and institutions needed to implement the full-scale microwave thermal 

launch system. Such milestones are listed in Table 3.39. Prestige and momentum are the keys. 

TABLE 3.39.—KEY FACILITIES AND MILESTONES 

Years Power 

level 

Type Goals and milestones 

2005 to 

2007 

200 W Static  Microwave thermal principle demonstrated for first time. 

 Nuclear-thermal class wall temperature of ~2000 K demonstrated with H2 in 

single channel. 

2010 to 

2012 

20 kW Static  First operation of credit-card-size multichannel microwave thermal thruster 

 Beat game-changing peak wall temperature of 2250 K, held for at least 3 min 

2012+ 1.2 MW Static/free flight  Year 1: Beat Myrabo’s beamed-energy altitude record of 71 meters (set in 

year 2000) 

 Year 2+: Grand Challenge to beat Goddard’s curve (and each other) using 

beamed-energy rockets 

2014+ 2 MW Static/free flight  Year 1: “Breakthrough” in combining power of gyrotrons into single beam 

 Year 2: Green light given for 20-MW facility. 

2015+ Megawatt 

class 

Exoatmopheric 

free flight using 

ABL 

 Year 1: ABL locks onto target; target generates thrust. 

 Year 2+: ABL-propelled beamed-energy rocket reaches edge of space (e.g., 

30-, 50-, or 100-km apogees). Picosatellite placed in orbit? 

2016+ 20 MW Static/free flight  Year 1: 20-MW beam power achieved; exceeds most powerful laser ever 

constructed. 

 Year 2: Beam propels largest rocket yet from California to Nevada; delivers 

bowl of petunias and airmail. 

2019+ 200 MW Static/free flight  Year 1: 200-MW milestone achieved. 

 Year 2+: Air-launched prototype achieves orbit for first time. 
NASA/TM—2012-217014 352 

It is quite likely that the game-changing peak wall temperature of 2250 K will be achieved by 2012, 

so the rough-order-of-magnitude (ROM) cost for this first important milestone is at most $250K. Beating 

Myrabo’s altitude record entails constructing the beam facility, developing the steerable optics, and 

developing a small-scale pressure-fed test vehicle. Only the gyrotron portion of the 1-MW beam facility 

has been funded, and it lacks a power supply, container integration, cooling, steering optics and most 

important, all engineering personnel with expertise in these areas and a project manager. Therefore, the 

ROM cost for this milestone is estimated to be more than $1M and less than $10M. The necessary 

elements are already fairly well-defined and uncoupled, so they could be shared into parallel tasks starting 

in summer 2011 and integrated back into a final record attempt by early 2013 if no major obstacles are 

encountered.

3.1.8.11.1 200-W Static Test Stand 

An initial test stand was built at Caltech by Parkin in 2005 and used to demonstrate the basic 

operating principle of microwave thermal propulsion for the first time. Numerical methods were 

developed to simulate the arrangement and predictions were shown to be consistent with measurements. 

If the hottest part of the tube seen in Figure 3.122 were attached to a nozzle, the vacuum specific impulse 

Isp would be about 700 s. 

Figure 3.122.—Laboratory demonstration of microwave thermal heat-exchange principle in a 2.45-GHz resonant 

cavity (Ref. 12). Single mullite thermocouple tube with H2 propellant. Dh, 1.19 mm; length of tube, L, 36.6 cm; inlet 

power, Pinlet, 1.1 bars; inlet Reynolds number, Reinlet, 280; and peak wall temperature, Twall, (1800 ± 100) K. The 

end of the tube was dipped in brine to make the hydrogen plume visible.

3.1.8.11.2 10-kW Static Test Stand 

The next evolutionary step has been to increase the power level from 200 W to 10 kW, move from a 

resonant cavity to a free-space power-beaming geometry, move from a single channel to multichannel 

HX, and add a simple nozzle. Owing to various administrational constraints at Carnegie Mellon and the 

NASA Ames Research Center (ARC), it was necessary to substitute a floodlight for a microwave source 

and helium for the hydrogen propellant. As part of his graduate research at Stanford, David Murakami has 

used the test stand to bring the microchannel aluminum HX shown on the right of Figure 3.123 up to its 

melting point. In the next few months, he will remake the HX from graphite and run to refractory wall 

temperatures. In particular, this arrangement will be used to explore the temperature limits and oxidation 

rates of SiC and ZrO2 coatings. A similar specification of facility is under construction at Caltech by 

Escape Dynamics Inc (http://www.escapedynamics.com). It is expected to feature magnetrons instead of a 

floodlight. 

The most successful HX approaches from these test stands will be carried forward and tested using 

the next generation of test stand. 
NASA/TM—2012-217014 353 

Figure 3.123.—Left: The 10-kW test stand uses a large parabolic mirror to concentrate the energy from a 20-kW 

tungsten light bulb onto a credit-card-sized HX (yellow box). Right: Aluminum proof-of-concept HX, soon to be 

attempted in graphite. Such small channel sizes and a credit-card size are necessary at this power level. 

3.1.8.11.3 1.2-MW Portable Static/Free-Flight Test Facility 

Many of the most important technical uncertainties surrounding a microwave thermal launch system 

can be addressed using a gyrotron beam facility. Most questions that have been raised in the course of 

DRM 1–C and others that were raised beforehand have been translated into a corresponding experiment 

in Table 3.40. , which is the preliminary experiment roster for the gyrotron beam facility. The outcome of 

most of these experiments is believed to be known, nevertheless they affirm important assumptions and 

need to be demonstrated. 

ARC has procured a $2M gyrotron for this facility, and it is due to arrive in April 2012. The gyrotron 

facility CONOPS is based on Figure 3.124, a scene from the 2009 Space Elevator Games in which Jordin 

Kare and Lasermotive won $900K for propelling a laser-powered vehicle up a helicopter-suspended 

ribbon. In observing the proceedings, it was most striking that every subsystem involved was 

containerized and mobile, from the test control van to the beam source and power supply. 

The “organizational template” for this beam facility is that of a wind tunnel: a facility provided by the 

government for academic and industrial users to visit. It has been too expensive for any one team of 

developers to buy and maintain, yet it is needed by all for progress to be made. This is a type of publicprivate partnership with a clear interface and a proven track record. 

 
NASA/TM—2012-217014 354 

TABLE 3.40.—PRELIMINARY EXPERIMENT ROSTER FOR THE GYROTRON BEAM FACILITY 

Experiment name Test type Description 

Dirt Static Effect of dirt on optics. Measure absorption coefficient as function of time. Determines 

cleaning interval to prevent unacceptable absorption loss/temperature buildup. 

Coherent doppler Static Investigate operation of coherent Doppler transponder in a high-particle medium 

(HPM) environment. Measure downrange position and determine methods to mitigate 

interference, should it arise. 

Bandwidth Static Measure actual bandwidth of gyrotron and compare with theory to determine whether 

theory is adequate to predict bandwidth of 170-GHz models. Will be needed for 

spectral beam combining design. 

Propagate Dynamic Measure atmospheric scintillation and absorption properties if relevant studies similar 

to Etcheverry do not exist 

Striate Dynamic Investigate ionospheric effects such as striation on beam using reflection for orbital test 

spheres. Only do this if studies in relevant regime do not exist. 

Fresnel strip Static Investigate design of Fresnel reflector primary and/or secondary by assembling a 0.5-

m-wide cross section on the ground and measuring optical characteristics in one 

dimension. 

Active panel Static Investigate various active panel concepts at full-scale and power loading for final 

aperture. 

HXa Static 1-MW HX tests—measure peak wall temperature with imaging pyrometer, mass flow 

rate, thrust, and material reflectance versus temperature versus angle using ellipsometer 

HX vehiclea Dynamic Dynamic tests of 1-MW vehicle, beginning with one-dimensional ribbon-guided 

leading up to unconstrained ascent to < 150 m. Longer distances may be attempted with 

larger and faster optics 

Surface breakdown Static/dynamic Define stability boundary of surface breakdown 

Thermal blooming Static/dynamic Analog of intensity 

Tracking Dynamic Verify the tracking accuracy of a large aperture slewed to a fast-moving target. 

HX Static Measure the temperature of the HX with multiple low-power beams incident on its 

surface. Also measure any intermodulation products produced at the transmitter or the 

HX. 

Given that ARC is located in the heart of Silicon Valley and can only support static tests, the 

containerized CONOPS enables the whole facility to be periodically relocated to support free-space tests 

at China Lake Naval Air Weapons Station (NAWS) and other locations as needed. Preliminary 

discussions with China Lake have affirmed that short-range testing is relatively straightforward and can 

be accomplished with the gyrotron(s) operated under delegated National Telecommunications and 

Information Administration (NTIA) authority. Sky-facing beams would be scheduled via the laser 

clearing house at Vandenberg, as has been the procedure for ABL, the Space Elevator Games, and so 

forth. A preliminary visit to the Bircham Mesa site formerly considered for the 1995 SELENE power 

beaming project revealed the presence of Indian petroglyphs in that area. For this reason and others, 

Junction Ranch is a more likely initial test location within the China Lake NAWS base. 
NASA/TM—2012-217014 355 

Figure 3.124.—Scene from the 2009 Space Elevator Games, with prize money provided by the NASA Centennial 

Challenges program. The containerized, mobile, and flexible CONOPS was a great success.

A grand-challenge competition is possible using the 1-MW beam facility. At this power level, the 

rocket sizes and attainable altitude are comparable with those of Goddard’s early flights. A competition 

could leverage the prestige associated with being a rocket pioneer, this time for beamed-energy rockets. 

Figure 3.125 shows the altitude attained by Goddard’s test launches from 1925 to 1936, with two recordbreaking launches in particular defining a rate of progress to beat. The timer might be started by the first 

team to beat Myrabo’s record of 71 m, set over a decade ago. Thereafter, teams would compete against 

Goddard’s rate of progress and each other. 
NASA/TM—2012-217014 356 

Figure 3.125.—Goddard’s progress in rockets and altitude from 1925 to 1936 suggests the scale, timeframe, number 

of trials and success rate at which a microwave thermal rocket might be expected to develop. Each data point is a 

launch attempt; some more successful than others. 

From the perspective of achieving an operational microwave thermal launch system, there are several 

benefits to a prize-based approach at this stage of development. First, it attracts and trains the workforce 

subsequently needed to implement a full-scale microwave thermal launch system. Second, it enables a 

more thorough exploration of the heat-exchanger/vehicle tradespace prior to vehicle scale-up. Third, it 

allows all the concepts to be measured to the same standard at roughly the same time. Fourth, if it is 

judged by outsiders to be a definitive stepping-stone, it is likely to bring in outside investment equivalent 

to about 10 times the prize purse. And fifth, the spirit of competition could be a powerful accelerator of 

the development roadmap presented here. 

3.1.8.11.4 2-MW Portable Static/Free-Flight Test Facility 

This facility is an evolution of the 1.2-MW facility to include a second gyrotron of nearly the same 

frequency combined with the first into a single beam via a diffraction grating. If increased power for 

short-range testing is of limited interest at this stage, it might be possible to accomplish the objectives of 

this stage by combining with one or more gyrotrons from an external facility such as at General Atomics 

in San Diego. 

3.1.8.11.5 MW-class Exoatmospheric Free Flight Test Capability (ABL) 

After the 2-MW beam facility demonstrates beam combining, but before the decision to proceed with 

a 20-MW facility, it will be advantageous to use ABL to retire various risks associated with 

transatmospheric operation. 

0

500

1000

1500

2000

2500

3000

Dec‐25

Apr‐27

Sep‐28

Jan‐30

May‐31

Oct‐32

Feb‐34

Jul‐35

Nov‐36

Apogee (meters)

Date
NASA/TM—2012-217014 357 

Figure 3.126.—Thermal image of the ABL (right) engaging a dummy target (left) graphically demonstrates that a 

shaky, slewing platform can lock onto and track a distant target ascending to orbit. What if the uncooperative 

dummy were replaced by a cooperative thermal rocket? HEL, High Energy Laser; .

In particular, ABL could be used in combination with a small-scale demonstrator vehicle to validate 

HX performance at altitudes of about 10 to 100 km. Depending on the particular power, spot size, and 

tracking accuracy of ABL, it may be possible to reach “space” (100 km) or even to place a picosatellite in 

orbit at this relatively early stage of the program. 

3.1.8.11.6 20-MW Static/Free-Flight Test Facility 

The purpose of the 20-MW facility will be to provide a subscale facility that is qualitatively similar to 

the large-scale facility in every way, such that the only problems anticipated in scale-up are those of 

resizing components according to established design rules. In addition to enabling sub-orbital hops, 

completion and successful operation of a beam facility at this scale will signal readiness for scale-up to a 

full-scale prototype. 

3.1.8.11.7 200-MW Static/Free-Flight Test Facility 

Construction of the full-scale prototype facility should be executed by a larger and more 

“operational” style team according to design rules established by earlier research-style teams. This facility 

will be at a dish diameter and power level sufficient to acquire an air-launched microwave thermal rocket 

and propel it from an altitude of about 20 km through to LEO. 

Because the dish size will be slightly larger than needed for a higher power vehicle, the remaining 

step to transition from a prototype vehicle with small payload to a larger vehicle should mainly be to add 

more gyrotrons to the optical back end. It should not entail a new dish. 

3.1.8.12 Issues and Resolutions 

Minimum gauges.—Scaling a rocket down to these small sizes correspondingly decreases the 

thickness of the HX walls and Gr-Ep LH2 tank walls to the point where the minimum gauges (minimum 

fabrication thicknesses) are reached. For the HX, the extra HX mass due to thicker-than-necessary walls 

was accepted. For the tank, the baseline material was changed from Gr-Ep to Ti, making the tank 
NASA/TM—2012-217014 358 

relatively heavy but cheaper to make, and giving it considerable room to become smaller without reaching 

minimum gauge.

Thermal blooming.—In verifying propagation characteristics of the millimeter-wave beams, 

Manning found that thermal-blooming effects were not negligible as previously assumed for the size and 

power level of beams considered in this study. The calculations show that corrections will probably need 

to be made for the beam bending into the wind (which can be corrected by a tracker) and, possibly a few 

orders of aberration. This will add some extra complexity to one of the mirrors in the optical back-end. 

Nozzle.—For the HX pressures under consideration, the conventional nozzle was found to have 

unacceptably low Isp at low altitudes. This was resolved by replacing it with a plug nozzle. 

Heat transfer.—The COMPASS study predicted temperature differences between the HX channel 

wall and bulk flow of up to 800 K based on the relatively poor heat transfer of smooth-walled tubes. 

Though only briefly mentioned in the COMPASS report, this number may generate concern, so it is 

addressed here: The thermal engineer did not have time to take into account surface roughness, which 

considerably improves heat transfer by disrupting the laminar sublayer surrounding the turbulent core 

flow, essentially destroying a layer of insulation. For example, Parkin predicts that the 40-m surface 

roughness of the Teledyne woven channels reduces the temperature difference between inner channel 

wall and flow to less than 100 K in the regime of interest. Various additional mitigations could be used 

including narrowing the channels, patterning the channel surface to rotate the flow, periodically weaving 

“high-k” carbon fibers across the flow channel to conduct heat from the surface directly into the core 

flow, refractory foam / porous flow heat exchange for the final 200 K of heating, and various other 

mitigations. Data from the heat exchange experiments at 200 W was very carefully compared to theory. A 

similar process will soon be underway at 10 kW, and static tests at 1, 20, and 200 MW will afford many 

opportunities to make sure that a minimal temperature drop between the channel wall and bulk flow is 

realized. 

Birds and lost aircraft.—The vicinity of the beam will be monitored by radar, if something was 

about to enter the beam, the beam could momentarily dodge the obstacle and power could be reduced or 

be turned off if needed. If a bird flew toward the center of the beam near the ground and did not turn 

around as the intensity increased, it would eventually experience about 20 suns equivalent of heating and 

presumably die quickly of effects related to that. It is something that we would want to avoid as much as 

birds hitting cars, turbines, windmills, and so forth. Thankfully, the beam facilities will be located at dry, 

high-altitude sites where birds tend not to live, and data quantifying the probability of birds interrupting a 

launch is relatively easy to gather. Migrating birds can achieve remarkable altitudes, so for planning 

purposes, we would consult maps of North American flyways—the seasonal migration routes taken by 

particular species. 

Horizontal and low-angle propagation.—The atmospheric scale height of atmospheric water vapor 

(distance over which the concentration of water vapor decreases by a factor of e) is about 1.5 km because 

water is a relatively heavy molecule. This means that most of the beam absorption that occurs does so 

close to the ground. For this reason, horizontal propagation over more than a few kilometers from a boost 

beam facility only looks attractive for high optical quality sites such as Atacama, and not from China 

Lake, California. One potential mitigation is to reduce the frequency of the boost facility, reducing 

horizontal attenuation but increasing the boost dish diameter. Another option is to use just a single sustain 

beam facility (Figure 3.127) in combination with an airbreathing first stage (or solid rocket, or N2H4

propellant dual mode approach). The key question in sizing the first stage cutoff altitude is the angle at 

which propagation of the beam becomes erratic (adversely steered by water vapor gradients) or 

attenuation reduces the beam energy to less than 30%. This issue needs further analysis prior to an 

airbreathing first-stage study. 
NASA/TM—2012-217014 359 

Figure 3.127.—A single beam facility sized for a 130-km range could, in principle, illuminate a rocket throughout the 

entire ascent trajectory. The areas within the beam facility range are shaded purple, and the baseline ascent 

trajectory is the red (lower) line. A key unknown is the minimum practical elevation angle at which the beam could 

acquire the rocket on liftoff (ALT, altitude). 

Primary aperture size, slew rate, and surface accuracy. In short, the beam facility primary 

aperture needs to be about the size of the GBT, at a surface accuracy better than GBT but worse than 

ALMA, and at a slew rate about 4 times the maximum of CCAT. High slew rates are not needed for 

astronomical dishes. Because the GBT active surface already corrects for gravitational sag as the dish 

slews, it is reasonable to expect that at low slew rates the more modern active surface control approach of 

ALMA, if applied to a larger dish, could still maintain the same surface accuracy across it. If any factor 

prevents the realization of high surface accuracy across a large dish, it will most likely to be the transient 

response of the dish structure to the high slew rate. If the transient response cannot be fully compensated 

using the existing active surface, additional mitigations could be passive or active. Passive mitigations 

include a more rigid structure, more extensive used of carbon fiber composites, and collapsing the dish 

down into a flat Fresnel reflector, which would probably reduce cost also. Active mitigations include a 

second level of larger displacement actuators behind the dish surface to correct for bulk motion. There 

may also be lessons to be learned from the ABL optical stabilization system. Ideally, a beam facility 

conceptual design will be needed that involves the engineers who designed the GBT, ALMA, and/or 

CCAT telescope structure and active surfaces. 

Millimeter-wave electrical breakdown. Electrical breakdown due to high power radiofrequency 

(RF) fields (i.e., a gigawatt millimeter wave) is possible if the electric field strength exceeds the 

breakdown strength of the surrounding media. Air breakdown with pressures in range of 76010–6 torr 

have been studied and are fairly understood. The most widely accepted formula that combines the effects 

of pressure and frequency is the W.C. Taylor work (Ref. 12). This was experimentally confirmed by 

several researchers and was found to be quite accurate over a large parameter space. Figure 3.128 shows 

experimental data from Massachusets Institute of Technology’s 110-GHz gyrotron oscillator experimental 

facility. 
NASA/TM—2012-217014 360 

Figure 3.128.—Comparison of Taylor theoretical equation with experimental data (Ref. 13). 

Extrapolating both the theoretical and experimental data, results in a minimum RF breakdown 

strength for bulk air of 2.5 to 3 MW/m2

 at 70 torr and 170 GHz This extrapolation is for bulk air only and 

does not account for surface flashover effects, field enhancements (surfaces and structures), hightemperature surfaces (>2500 C), or aerosols (dust, etc). Each of these adds additional complexity to the 

breakdown problem. Although the effect of each of these additional parameters has been studied, 

extrapolating the data to frequency and power levels required for this program is difficult. 

The secondary parameter most studied for its effect on RF breakdown is surface flashover. Many 

studies have been performed on surface flashover in the 2- to 140-GHz range, but few, if any, have been 

published in the 170-GHz range. The main effect seen in the lower frequency range has been a flattening 

of the breakdown curve in the low-pressure regime (i.e., breakdown remains at the Paschen minimum 

even as the pressure is reduced from 10s of torrs to 10–6 torr. Some studies also suggest a possible 

reduction of the overall breakdown strength at a surface by a factor of 2, but it is not clear from this 

research whether the reduction is due to unique surface physics or simply to field enhancement effects. 

Figure 3.129 shows experimental data illustrating this phenomenon. In any event, the dominant role of 

surface flashover is in the low-pressure regime, on the left-hand side of the curve. In the high pressure 

regime (right-hand side of the curve), the role of surface flashover seems to be at the minimum and is 

most likely due to surface defects and roughness causing localized field enhancements. 

Figure 3.129.—Pressure versus RF breakdown field strength 

at 2.4 GHz (Ref. 14). Solid symbols denote bulk air, and 

open symbols denote with surface. 

4

8

12

16

Electric Field (kV/cm)

10 -4 10 -3 10 -2 10 -1 10 0 10 1 10 2 10 3

Pressure (torr)
NASA/TM—2012-217014 361 

Much less data are available in the 170-GHz range of the effects of aerosols and/or dust on electrical 

breakdown strength. This topic has been well studied and documented in the optical range, but no data 

has been published in the millimeter-wave regime. From a theoretical perspective and extrapolating 

optical data and trends, the effect of dust on electrical breakdown was predicted to be of minimal impact 

until the particle size approaches a significant fraction of a wavelength. At smaller sizes, the localized 

electric field (on the surface of the individual particle) remains unperturbed and relatively small. 

Extrapolation and physical arguments are less certain and more complex for aerosols, where possible 

chemical reactions and polar molecules are likely to be found. In this environment, wavelength-dependent 

effects can be seen in single molecules. Although studies at optical wavelengths have shown the effects of 

aerosols to be of secondary concern, some testing laboratory will be required at 170 GHz to determine 

safe operating conditions with a high confidence level. 

The effect of high-temperature surfaces on millimeter-wave breakdown is the least studied and least 

understood. Only a single reference (Ref. 15, M. Gilden et al.) has been identified that contains some 

experimental data at microwave frequencies. The experimental data in this paper was limited to the 

effects on breakdown of a hot wire at 1200 C in a 5-GHz cavity (Figure 3.130). Combining the observed 

decrease in breakdown strength with other papers and later theory indicates that the primary effect was 

due to variations in gas density near the hot surface. Electron emission is also thought to play a role, but 

additional data will be required to quantify what role this emission will play and in what regime it will be 

the most significant. 

Figure 3.130.—Results of breakdown measurements with a hot wire in the 

presence of gas flow (Ref. 15). 

3.1.8.13 Launch Window Analysis 

Launch opportunities are constrained by the weather and the potential for beam conjunction with a 

LEO satellite. The weather at potential beam facility sites has been relatively well studied by astronomical 

site surveys. Table 3.41 shows that weather at the CONUS site examples given in Figure 3.131 is 

sufficiently good, with more than 50% probability of suitable weather at any given time. 

 
NASA/TM—2012-217014 362 

TABLE 3.41.—WEATHER AT THE CONUS SITE EXAMPLES GIVEN IN FIGURE 3.131) 

[Only nighttime data were published because the data were taken for an astronomical site survey.] 

 % of night with clear sky % of night with transitional sky 

Pikes Peak, CO 64 11 

Mt. Evans, CO 54 15 

Jelm Mountain, WY 32 18 

A beam-satellite conjunction analysis was very kindly carried out by Dr. James Mason at ARC. His 

results, listed shortly, show that space is big—you could launch continuously and only once or twice a 

day would you hit a satellite with more than 0.1 suns of illumination intensity. Even then, it would last for 

only about 0.2 s because of the large relative velocities between the beam and satellite. In reality, those 

conjunctions would be predicted and avoided. 

3.1.8.14 Conjunction Analysis Results for Equatorial Launch 

Launch into 36° inclination from China Lake. Launches were simulated every 4 min for 8 days. Each 

intersection between the microwave beam and background satellites was recorded. On average, there 

were 5.25 intersections per day, each for 0.2 s. 

Figure 3.131.—Launch time versus range of intersecting satellite. Each color represents a single day of launches 

(i.e., the 8 days are overlaid). The North American Aerospace Defense Command (NORAD) satellite ID is listed 

along with the intersection duration and range. 
NASA/TM—2012-217014 363 

Figure 3.132.—Launch time versus inverse range squared (1/R2) of intersecting satellite. Each color represents a 

single day of launches (i.e., the 8 days are overlaid). 

Figure 3.133.—Probability of an intersection happening in each half hour period, over the 8 days. 
NASA/TM—2012-217014 364 

Figure 3.134.—Histogram of clear launch windows—periods that have no intersections. Over the 8 days there were 

29 periods with more than 1 hr of clearance. 

Figure 3.135.—Histogram showing the distribution of intersection maximum intensity, Imax, during the 8 days. Note 

that most days have Imax less than 100 W/m2

 (i.e., 8% of solar constant). 
NASA/TM—2012-217014 365 

Figure 3.136.—Histogram of the maximum energy incident on a satellite per square meter during an intersection. This 

was calculated using Imax  intersection duration. 

3.1.8.15 Conjunction Analysis Results for Sun Synchronous Orbit 

Direct launch was into a 98° Sun-synchronous inclination from China Lake. Launches were simulated 

every 4 min for 8 days. Each intersection between the microwave beam and background satellites was 

recorded. On average, there were 12.75 intersections per day, each for 0.03 s. 

Figure 3.137.—Launch time versus range of intersecting satellite. Each color represents a single day of 

launches (i.e., the 8 days are overlaid). The NORAD satellite ID is listed along with the intersection 

duration and range. 
NASA/TM—2012-217014 366 

Figure 3.138.—Launch time versus range of intersecting satellite. Each color represents a single day of launches 

(i.e., the 8 days are overlaid). The NORAD satellite ID is listed along with intersection duration and range. 

Figure 3.139.—Probability of an intersection happening in each half-hour period, over the 8 days. 
NASA/TM—2012-217014 367 

Figure 3.140.—Histogram of clear launch windows—periods that have no intersections. Over the 8 days, there were 

41 periods with more than 1 hr of clearance. 

Figure 3.141.—Histogram showing the distribution of intersection maximum intensity during the 8 days. Note that 

most have Imax less than 200 W/m2

 (i.e., 15% of solar constant). 
NASA/TM—2012-217014 368 

Figure 3.142.—Histogram of the maximum energy incident on a satellite per square meter during an intersection. This 

was calculated using Imax  t, where t is the intersection duration.

3.1.8.16 Beam-Combining Analysis 

There are several ways of combining the power output of millimeter-wave sources, including 

 Master Oscillator Power Amplifier (MOPA) beam combining 

 Phase-locked oscillators 

 Temporal beam combining 

 Spectral beam combining 

The first and second approaches require the millimeter-wave sources to have a systematic phase 

relationship between them (mutual coherence). For MOPA, this is achieved by a single oscillator 

producing a wave that is amplified by a large number of amplifiers. This approach is not as financially 

attractive as combining the output of a large number of oscillators because amplifiers are more expensive 

per watt of output. 

The high cost of amplifiers would make phase-locking oscillators together look more attractive. The 

phase locking of two oscillators is achieved by leaking some of the signal from oscillator A into 

oscillator B and vice versa, with the exact conditions described by Adler’s relation. Unfortunately, 

gyrotron oscillators are highly overmoded structures of many wavelengths across resonating far from 

their fundamental frequency, so this type of phase locking would require fundamental research to 

accomplish in a reliable way. 

Temporal beam combining is best suited to sources that are optimized for pulsed operation. Since the 

available gyrotrons are CW, it was not considered here or elsewhere. 

Spectral beam combining is the reverse of splitting white light into a rainbow using a prism or 

diffraction grating. This technique is commonly used in the telecommunications industry. A reflective 

metal grating is used to combine the gaussian beams from a large number of gyrotron oscillators of 

slightly different frequencies and incident angles into a single gaussian output beam (which can 

subsequently be directed to the rocket via a telescope). The physics of diffraction gratings has been well 

understood for over 200 hundred years. 
NASA/TM—2012-217014 369 

The maximum number of gyrotron beams that can be combined this way is equal to the spectrometer 

resolution R, which is a function of their frequency f and bandwidth Δf. For ~170-GHz step-tunable 

gyrotrons with a bandwidth of 70 MHz, this means that 

2428 gyrotrons 170 GHz

70 MHz   





  f

f R

This number would be reduced by practical considerations and unusable bands in the atmospheric 

absorption peaks. Nevertheless, a 20-GHz band from 150 to 170 GHz (Figure 3.114—Figure 3.131) could 

in principle support 285 gyrotrons, each with its own 70 MHz of bandwidth. For 4-MW gyrotron 

oscillators, that corresponds to more than 1 GW of power in total. If more power is needed, lower and 

higher wavelengths and/or additional dishes could be used, depending on which option minimized cost. 

One of the difficulties with diffraction gratings at optical wavelengths has been fabricating the fine 

grooves, which are spaced approximately a wavelength apart, and achieving high efficiency. At 170 GHz, 

the grooves could be made by an end mill and are 1.8 mm apart. For combining ~2500 gyrotron beams, 

3092 grooves total would be needed, making the grating length about 5.7 m. The grating would not need 

to be a single piece and could be milled in sections and then mounted to a water-cooled backing plane. 

Because it is possible to achieve high tolerances and reflectance, the beam-combining losses are expected 

to be only 1 to 2 percent of the incident energy. Belousov et al. (Ref. 16) measure less than 1% energy 

loss in combining a 70- and 79-GHz beam. This kind of performance should be demonstrated with many 

simultaneous frequencies. Earlier it was suggested that this could be demonstrated at low power using 

IMPact ionization Avalanche Transit-Time (IMPATT) diodes. 

3.1.8.17 References for Millimeter-Wave Thermal Consultant Report for DRM 1–C 

1. Marshall; and Cox: Integral Textile Ceramic Structures, 2008. 

2. Wasko, Robert A.: Performance of Annular Plug and Expansion-Deflection Nozzles 

Including External Flow Effects at Transonic Mach Numbers. NASA TN–D–4462, 1968. 

3. Fleeman: Tactical Missile Design, 2001. 

4. Thompson: 2009. 

5. Young; and Olds: Responsive Access Small Cargo Affordable Launch (RASCAL) 

Independent Performance Evaluation, AIAA 2005–3241, 2005. 

6. Access to Space Study. Summary Report, Office of Space Systems Development, NASA 

Headquarters, Washington D.C., 1994. 

7. National Research Council: High-Temperature Oxidation-Resistant Coatings, 1970. 

8. U.S. PTO patent application 2008/0286473, 2008. 

9. Mangum: 2007. 

10. Lohr, et al.: Operating Experience on Six 110 GHz, 1MW Gyrotrons for ECH Applications, 2008. 

11. Kare, J.T.; and Parkin, K.L.G.: Comparison of Laser and Microwave Approaches to CW 

Beamed Energy Launch, Beamed Energy Propulsion. AIP Conference Proceedings 830, 

K. Komurasaki, ed., Melville, NY, 2006, p. 388. 

12. Parkin, K.L.G.: The Microwave Thermal Thruster and Its Application to the Launch 

Problem. Ph.D. thesis, June, 2006. 

13. Temkin, Richard: W-Band Breakdown Research. Kick-Off Meeting Counter High Power 

Microwaves, Albuquerque, NM, 2010. 

14. Hemmert, David, et al.: High-Power Microwave Window Breakdown Under Vacuum and 

Atmospheric Conditions. Proc. SPIE 4031, vol. 90, 2000. 

15. Gilden, M.; and Pergola, J.: Microwave Breakdown Near a Hot Surface. PTGMTT National 

Symposium Digest, 1963, pp. 39–44. Belousov, V.I.; Denisov, G.G.; and Peskov, N. N. Yu: 

Quasi-Optical Multiplexer Based on Reflecting Diffraction Grating, Int. J. Infrared 

Millimeter Waves, vol. 12, no. 9, 1991. 
NASA/TM—2012-217014 370 

3.1.9 Millimeter-Wave Thermal Cost Estimates 

National Aeronautics and Space Administration

Beamed Energy Propulsion Facility 

ROM Cost Estimate 

March 28, 2011

National Aeronautics and Space Administration

General BEPS Launch Facility Assumptions

• Launch site ROM estimates are based on a NASA KSC launch 

site estimating approach. Microwave and laser sites use the 

same approach for ‘like function’ facilities which are scaled for 

expected differences in size or function.

• All sites include site design and testing and activation costs.

• Launch facility design is estimated at 12% of estimated total site 

costs less testing and activation.

• Launch facility testing and activation is estimated at 30% of the 

capital facilities cost.

• Microwave and laser site design is estimated at 8% of estimated 

total site costs less testing and activation.

• Microwave and laser site testing and activation is estimated at 

15% of the capital facilities cost.
NASA/TM—2012-217014 371 

National Aeronautics and Space Administration

General BEPS Launch Facility Assumptions

• All cost are rough order of magnitude (ROM) estimates in 

FY11M$.

• Estimates assume all facilities are built new on government 

land, no existing infrastructure is used. Available infrastructure 

will reduce costs.

• Costs do not include remediation of potential significant 

environmental impact issues that may be identified.

• Range estimates are presented to highlight the uncertainty 

associated with the lack of definition and supporting cost 

information available at this time. 

• Range estimates for the launch site also includes the 

uncertainty associated with the selected location and the cost 

impact of materials that may be required to protect against 

environmental conditions.

National Aeronautics and Space Administration

Operations Cost Assumptions

• All costs are estimated for one year of operation in FY11$. 

• Yearly operations cost assume 360 launches per year 

(one/day). Additional study is required to determine the feasible 

number of launches per year from a single launch pad. The 

number of launches per pad is affected by the CONOPS, pad 

turnaround time, launch facility maintenance, weather, and 

launch window requirements.

• Total estimated facility operations costs are assumed to be 

allocated to the beamed launch operations. Facility sharing can 

reduce costs. 

• Facility equipment and maintenance is estimated at 5% of the 

total value of capital assets per year resulting in an estimated 

70-80% of total operations costs.
NASA/TM—2012-217014 372 

National Aeronautics and Space Administration

Microwave Launch Site Assumptions

• Launch Site facility requirements

– General – site work, roads

– Support Facilities/Launch Pad – see details below

– Site Infrastructure – water, fire, sewer, power, communications, HVAC, environmental 

control system

– Ground support equipment

• Key cost estimate assumptions

– New facility on government land in secluded area due to danger associated with 

beamed energy. Launch rate 1-4 vehicles a day, if CONOPS allows.

– Launch pad includes all related hardware/services. Rocket will be small, approximately 

9 meters in length, and less than 3,000 kg total wet mass with payload.

– Command center includes all related services.

– Payload integration facility (small payloads <200 kg) can process/integrate up to four 

payloads per day.

– Hydrogen storage facility sized for ~30,000 kg with no on-site production.

– Rocket storage facility size for 8 rockets.

– Admin facility sized for launch site staffing needs (estimated at 100 people).

– Security facilities/equipment includes fencing, cameras, etc.

National Aeronautics and Space Administration

Microwave Booster Station Assumptions

• Microwave Booster Station facility requirements

– General – site work, roads

– Support Facilities – admin facility (50 people), command center (75% of 

launch site center), guard gate facility

– Site Infrastructure – water, fire, sewer, power, communications, HVAC, 

environmental control system

– Ground support equipment

– Beam Generation/Pointing – gyrotron field (300 MW + 10% reserve), antenna 

system (45m), beam combiner

– Power Storage/Distribution – energy storage, power supply, power distribution

• Key cost estimate assumptions

– Gyrotron/Power Supply cost – Current market $2M per 1 MW based on NASA 

ARC information used as high, projected market information $2M per 4 MW 

used as low, most likely estimated as $2M per 2 MW

– Energy Storage - $50K/kWh estimated as most likely based on GRC 

experience with wind tunnels (low -40%, high +20%)

– Antenna System - $50K/m2 (Ref: J. Kare, $40K/m2 escalated) used as low (ML 

+50%, high +100%)
NASA/TM—2012-217014 373 

National Aeronautics and Space Administration

Microwave Sustaining Station Assumptions

• Microwave Sustaining Station facility requirements

– General – site work, roads

– Support Facilities – admin facility (50 people), command center (75% of launch 

site center), guard gate facility

– Site Infrastructure – water, fire, sewer, power, communications, HVAC, 

environmental control system

– Ground support equipment

– Beam Generation/Pointing – gyrotron field (600 MW + 10% reserve), antenna 

system (120m), beam combiner

– Power Storage/Distribution – energy storage, power supply, power distribution

• Key cost estimate assumptions

– Gyrotron/Power Supply cost – Current market $2M per 1 MW based on NASA 

ARC information used as high, projected market information $2M per 4 MW 

used as low, most likely estimated as $2M per 2 MW

– Energy Storage - $50K/kWh estimated as most likely based on GRC 

experience with wind tunnels (low -40%, high +20%)

– Antenna System - $50K/m2 (Ref: J. Kare, $40K/m2 escalated) used as low (ML 

+50%, high +100%)

National Aeronautics and Space Administration

Microwave Launch Facility Development ROM

Low ML High Mean Est Low ML High Mean Est Low ML High Mean Est Low ML High Mean Est

Project Total ROM ($M) $324.5 $553.9 $1,224.4 $701.0 $851.4 $1,446.7 $2,605.7 $1,634.6 $1,781.0 $3,088.3 $5,316.4 $3,395.2 $2,956.9 $5,088.9 $9,146.5 $5,730.8

Design $28.2 $47.9 $105.1 $60.4 $55.6 $94.5 $170.1 $106.7 $116.1 $201.2 $346.4 $221.2 $199.9 $343.6 $621.6 $388.4

Testing & Activation $61.6 $106.8 $243.3 $137.2 $100.3 $171.2 $309.8 $193.8 $213.7 $371.4 $640.4 $408.5 $375.5 $649.4 $1,193.5 $739.5

Microwave Launch Facilities ROM Subtotal $234.8 $399.2 $876.0 $503.3 $695.5 $1,181.0 $2,125.8 $1,334.1 $1,451.2 $2,515.6 $4,329.6 $2,765.5 $2,381.5 $4,095.8 $7,331.4 $4,602.9

WBS Item 000 0

Site Work $11.2 $15.4 $18.8 $15.1 $8.4 $11.6 $14.1 $11.3 $8.4 $11.6 $14.1 $11.3 $27.9 $38.6 $47.0 $37.8

Roads $18.4 $27.9 $46.2 $30.8 $18.4 $27.9 $46.2 $30.8 $18.4 $27.9 $46.2 $30.8 $55.3 $83.7 $138.6 $92.5

Support Facilities $32.4 $68.5 $134.7 $78.5 $7.1 $14.9 $36.1 $19.4 $7.1 $14.9 $36.1 $19.4 $46.7 $98.4 $206.8 $117.3

Concrete Administration Facilty $7.4 $10.8 $28.5 $3.7 $5.4 $14.2 $3.7 $5.4 $14.2 $14.7 $21.5 $57.0

8 Rocket Storage Facility $5.6 $7.5 $12.2 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $5.6 $7.5 $12.2

Concrete Command Center $3.3 $10.4 $25.3 $2.5 $7.8 $19.0 $2.5 $7.8 $19.0 $8.3 $26.0 $63.4

Payload Integration Facilty $15.2 $38.1 $65.8 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $15.2 $38.1 $65.8

Guard Gate Facility $1.0 $1.8 $2.8 $1.0 $1.8 $2.8 $1.0 $1.8 $2.8 $2.9 $5.3 $8.4

Pad Structure $13.1 $19.5 $24.3 $19.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $13.1 $19.5 $24.3 $19.0

Pad Structure $4.3 $6.2 $8.4 $6.3 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $4.3 $6.2 $8.4 $6.3

Emergency Egress System $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0

Flame Trench $6.7 $8.0 $8.9 $7.9 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $6.7 $8.0 $8.9 $7.9

Ligtning Protection Towers $2.1 $5.3 $7.0 $4.8 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $0.0 $2.1 $5.3 $7.0 $4.8

Infrastructure $132.8 $221.1 $545.5 $299.8 $88.5 $138.3 $254.2 $160.3 $88.5 $138.3 $254.2 $160.3 $309.7 $497.7 $1,053.8 $620.4

GSE $26.9 $46.8 $106.5 $60.1 $6.7 $11.7 $26.6 $15.0 $6.7 $11.7 $26.6 $15.0 $40.4 $70.2 $159.8 $90.1

Beam Generation/Pointing Equipment $0.0 $0.0 $0.0 $0.0 $368.5 $593.1 $1,023.5 $661.7 $926.5 $1,544.2 $2,502.1 $1,657.6 $1,295.0 $2,137.2 $3,525.5 $2,319.3

Power Storage/Processing Facilities $0.0 $0.0 $0.0 $0.0 $197.8 $383.5 $725.2 $435.5 $395.5 $767.0 $1,450.4 $871.0 $593.3 $1,150.6 $2,175.7 $1,306.5

Launch Site Complex ‐ ROM Subtotal ($M) Booster Station Site ‐  ROM Subtotal ($M) Sustaining Station Site ‐  ROM Subtotal ($M) Microwave Launch Facilities ‐ Total ROM ($M)

Microwave Beamed Energy Launch Facilities ROM (FY11$M)

(Excludes FTE/WYE personnel and O&M costs)
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Microwave Launch Facility Operations ROM

Low ML High Mean Est

Microwave Operations Total ROM ($M/Year) $155.7 $260.3 $439.7 $285.2

Facility/Equipment Maintenance $114.9 $198.7 $357.3 $223.6

Operations Personnel $30.0 $40.0 $50.0 $40.0

Propellant/Energy/Consumables $10.8 $21.6 $32.4 $21.6

Other $0.0 $0.0 $0.0 $0.0

Launch Site Complex ‐ ROM Subtotal ($M/Year) $28.9 $45.0 $76.4 $50.1

Facility/Equipment Maintenance $10.3 $17.8 $40.6 $22.9

Operations Personnel $15.0 $20.0 $25.0 $20.0

Propellant/Energy/Consumables $3.6 $7.2 $10.8 $7.2

Other $0.0 $0.0 $0.0 $0.0

Booster Station Site ‐  ROM Subtotal ($M/Year) $44.5 $74.3 $126.6 $81.8

Facility/Equipment Maintenance $33.4 $57.1 $103.3 $64.6

Operations Personnel $7.5 $10.0 $12.5 $10.0

Propellant/Energy/Consumables $3.6 $7.2 $10.8 $7.2

Other $0.0 $0.0 $0.0 $0.0

Sustaining Station Site ‐  ROM Subtotal ($M/Year $82.3 $141.0 $236.8 $153.4

Facility/Equipment Maintenance $71.2 $123.8 $213.5 $136.2

Operations Personnel $7.5 $10.0 $12.5 $10.0

Propellant/Energy/Consumables $3.6 $7.2 $10.8 $7.2

Other $0.0 $0.0 $0.0 $0.0

Microwave Launch Operations ‐ Total ROM ($M/Yr)

Microwave Launch Facilities Operations ROM (FY11$M/Year)

(Does not include launch vehicle)

National Aeronautics and Space Administration

Cost Assumptions

• DRAFT Cost Estimate based on COMPASS design

• All costs are in FY11$M

• Estimates represent prime contractor cost before fee

• This estimate assumes the following:

– Proto-flight development approach

– No ground spares are included

– Model assumes TRL Level 6

• This estimate does not include any cost for technology development

– Represents the mean estimate based on cost-risk simulation results

– Parametric estimate based on mostly mass-based CERs from historical cost data

– Launch vehicle systems integration wraps

– Software not included

• Does not include:

– Any insight/oversight costs (by a NASA lead center)

– Reserves (can be as high as 40-50%)

– Ground System Cost (ie. Microwave Station, Launch Pad, Booster Station, etc...)

– Launch Services Costs (ie. Special launch approval process)

– Technology costs for components lower than TRL-6 
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Microwave Launcher

Preliminary Cost ROM 

(Represents estimated Prime Contractor cost) 

All Costs in FY11$M

WBS Description

DDT&E 

Total 

BY $M

Flight 

HW 

BY $M

DD&FH 

Total 

BY $M

06.1 Payload 0.0 0.0 0.0

06.2.1 Attitude Determination and Control 6.4 4.8 11.2

06.2.2 Command and Data Handling 4.7 2.7 7.4

06.2.3 Communications and Tracking 1.8 1.3 3.0

06.2.4 Electrical Power Subsystem 1.7 1.2 2.9

06.2.5 Thermal Control 12.6 2.3 14.8

06.2.6 Propulsion 14.6 2.5 17.1

06.2.7 Propellant 0.0 0.0 0.0

06.2.8 Structures and Mechanisms 11.0 5.6 16.6

Subtotal 52.7 20.3 73.0

Systems Integration 57.0 8.1 65.2

Spacecraft Total 109.7 28.4 138.1
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3.2 Space Design Reference Missions 

3.2.1 Low-Earth-Orbit to Geosynchronous-Earth-Orbit Collaborative Modeling for Parametric 

Assessment of Space Systems (COMPASS) Report 
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National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov

Baseline for Comparison

• SOA Geostationary Delivery

– Falcon 9 launched from Kwajalein, 9 

° inclination

– Single ~2375 kg (not including Biprop

Apogee system) payload S/C

– Falcon 9 upper stage takes S/C to 

GTO, onboard biprop system performs 

final delivery to GEO

• Laser Propelled Transport

– Falcon 9 launched from Kwajalein to 

400 km circular (for better ground laser 

visibility)

• Two ~1970 kg (less than SOA) GEO 

Payload S/C plus single water tank 

– Separate launch (dry) Transport 

vehicle

• Falcon 9

• Used 10 times

• Transport (dry) should be ~ 600 kg (with 

margin)

• Objective: reduced launch costs 

approaching 50%... 11

1600kg MMH/NTO prop load. Results are 4 

propellant tanks ~0.9m dia, two GN2 pressurant

tanks (the 0.42m by 0.67m tank by ATK is 

basically an exact fit), and a SFT feed 

system. Total Mass: 146kgTotal 

Inert: 133kgTotal Dry: 96kg Item Mass 

(kg)

Q/D valves 2

Filter/Lines 6

Regulators 2

Valves 6

HP Valves 1.5

N2 Mass 37

Ox Mass 1000

Fuel Mass 605

N2 Tank 20

Ox Tank 26

Fuel Tank 26

Engine 6
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National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov

400km: Communication Times to 

Equatorial Ground Stations

• The analysis was conducted over a 1 year period.

• On average, the spacecraft is visible from each ground site ~2.3 

minutes per pass.
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National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov

Propellant Selection

• Water (or other dense 

propellant- Ar?) leave more 

room for payload

– But 800 seconds with H2O is 

TRL 1-2

• LH2 propellants much less 

dense (~68 kg/m^2) but 900 

seconds performance (TRL4?)

– From Impulsive case

• 5000 kg LH2 would take up 

about half of volume

• 5000 kg of Water should fit

27

LH2v

Falcon 9 five m shroud

GEOsat

GEOsat
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National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov

Copernicus: 400km LEO to 3000km 

Circ Mission

ΔVp,1, ΔVp,2, ΔVp,3, 

ΔVp,4, ΔVp,5

ΔVa,1

400km LEO, i=9.0°

3000km Circ., i=9.0°
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National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov

Copernicus: 3000km Circ to GEO 

Mission

3000km Circ., i1=9.0°

GEO, i5=0.0°

ΔVp,1, ΔVp,2 ΔV , ΔVp,3 a,1, ΔVa,2

i2=9.0°

i3=9.0°

i4=5.88°
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National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov

BEP GNC

• Requirements

– Deliver spacecraft to GEO in less than 10 days

– Point the reflector with 0.5 degrees of accuracy 

– Rendezvous and dock with the propellant stage in a LEO 

circular orbit

– Single fault tolerant

• Assumptions

– BEP Lightcraft has zero products of inertia

• Navigation Design

– Eight Sun sensors for course attitude determination

– One Northrop Grumman internally redundant SIRU

• Fault tolerant, redundant HRGs, accelerometers and electronics

• Spaceflight heritage – Deep Impact, MESSENGER, Cassini, 

NEAR and others

– Two Selex-Galileo Avionica Autonomous Star Trackers (A-STR)

• Includes processing electronics and data/command interfaces

• Spaceflight heritage – MESSENGER, STEREO, MRO, SDO

• Up to 10 tracked stars

• Bias error < 10 arcsec, 3 sigma

– Two Surrey Satellite Technology GPS receivers

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Pre-decisional. For NASA use only. www.nasa.gov

BEP GNC

• Control

– Four Teldix RSI 12-75/60 reaction wheels for fine pointing accuracy and 

momentum storage

– Thirteen 5 lbf (22.2 N) RCS thrusters for momentum management, 

rendezvous with propellant stage and proximity operations during docking

• 2 pods of 4

• 1 pod of 5

• Automated Rendezvous and Docking

– 2 Next Generation Advanced Video Guidance Sensors (NGAVGS) for 

Automated Rendezvous and Docking (AR&D)

• Derivative of the AVGS used on Orbital Express

• Two sets of laser diodes which operate at nominal wavelengths of 

806 and 845 nm and a mirror through which the lasers fire

• Camera that images the return from the lasers

• hardware, software, and firmware that process the returned images 

into relative position and attitude data

• Designed to interact with retro-reflectors placed on the target 

spacecraft target retro-reflectors are arranged in a pattern known to 

the AVGS software

• Data is output from the sensor and fed to the spacecraft Guidance 

and Navigation System

• Breadboard version of the NGAVGS was built at MSFC

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National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov

RCS 

• COTS Membrane Tank on Both 

Transport and Tank Stages

• Gaseous Nitrogen Pressurant

• 13 Aerojet MR-50S Thrusters

– 228s ISP

– 5lbf Thrust

• Located in 3 Locations

– 2 Pods of 4

– 1 Pod of 5

• Unique Placement Required to 

Avoid Possible Exposure to Beam

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Pre-decisional. For NASA use only. www.nasa.gov

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Thermal Subsystem - Overview

• Thermal assumptions In orbit

– Heat dissipation of 90 w each for both the engine and payload sections

– Earth view factor of 0.5 

– Solar array view factor of 0.15

– 30° sun angle

• Thermal assumptions Laser Receiver

– A off-set parabolic surface will be used to receive the laser beam and concentrate it into the engine plasma chamber. 

– Mirror reflectivity 0.9999 

– Back side of the mirror can be used to radiate excess heat to space.

– The mirror consists of a single reflective surface with a secondary concentrator, structure, drive motors, shield and launch structure

– Laser mirror shields are placed on the bottom half of the vehicle (both engine and payload sections), in the direction facing the laser 

to reflect any inadvertent laser impact from hitting the vehicle. 

• The thermal subsystem components

– MMOD Shielding

– Cold plates and heat pipes to transfer from the electronics to the radiator section.

– Heaters, sensors, data acquisition and controller

– MLI is used to insulate the outside of the vehicle. 

– Heaters, sensors, data acquisition and controller

– MLI is used to insulate the outside of the vehicle. 

• Environmental models

– LEO environmental conditions were used to determine the heat transfer while in orbit. 

• Thermal System Mass: 182.66 kg

• Trades considered in analysis 

– 2 mirror shapes were considered, a standard parabolic Cassigrainan arrangement and an off-set parabola. 

– Mirror arrangement to shield the vehicle and the use of additional mirrored laser shields. 

• Options to Reduce mass and recommendations:

– The mirror surface (for both the laser receiver and shielding) is vulnerable to micro meteor impacts. These impacts can cause a 

failure of the mirror if its surface reflectivity is reduced. Mitigation 

approaches should be investigated to minimize this impact. 

– Further analysis and experimentation should be performed to determine the amount of mirror degradation that will occur (if any) 

during operation. 
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National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov

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Thermal Analysis: Radiator Sizing

•2 Radiator panels were sized for use on the vehicle. One for 

the engine section and one for the payload section. 

•A surface mount radiator was used for both locations.

•The radiator model was based on a first principles analysis of 

the area needed to reject the identified heat load to space. 

From the area a series of scaling equations were used to 

determine the mass of the radiator. Worst case thermal 

environment of LEO was used to size the radiator. 

•No louvers were utilized

•Heat Pipes were used to conducted heat to the radiator 

surface. 

Assumptions used

Variable Value

Radiator Solar Absorptivity 0.14

Radiator Emissivity 0.84

Radiator Sun Angle 70°

Radiator Operating 

Temperature

320 K

Total Radiator Dissipation 

Power

90 W

View Factor to Earth 0.5

View Factor to the Solar 

Array

0.15
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National Aeronautics and Space Administration

Pre-decisional. For NASA use only. www.nasa.gov

Laser Reflector and Concentrator

•The laser reflector is used to collect the incoming laser radiation and direct and concentrate it 

into the plasma engine. 

•Two types of laser reflectors were considered: 

o A parabolic Cassigranian configuration 

o A off-set parabolic configuration.

•Both utilize a secondary reflector to further concentrate and direct the laser into the plasma 

engine. 

Cassigranian Arrangement

Incoming Laser 

Secondary Mirror

Primary Mirror

Focus to 

Engine

Primary Mirror

Secondary Mirror or 

Direct Focus to 

Engine

Incoming Lase

Off-Set Parabola Arrangement

•The Off-set Parabola was selected due to the entrance angle limitations of the laser into the plasma engine. 

The cassigrainian design required entrance angles of up to 60°, this was beyond the capabilities of the 

refractive optics on the plasma engine. 
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3.2.2 Low-Earth-Orbit to Geosynchronous-Earth-Orbit Consultant Reports—Design Reference 

Mission 2–A 

The Design Reference Mission 2–A (DRM 2–A) low-Earth-orbit to geosynchronous-Earth-orbit 

(LEO–GEO) beamed-energy transport tug will cut GEO lost costs in half. This savings will be enabled 

by the elimination of the Centaur upper stage and circularization engine required in current GEO 

launches. Instead, two 2000-kg GEO satellites will be launched into LEO on a single launch vehicle, 

rendezvous with the tug, and be transported to their circular GEO orbits in 32 hr. The tug will then return 

to LEO for another transport. In this manner, two GEO satellites could be placed into their orbits on a 

single launch. 

3.2.2.1 Description of Vehicle 

The DRM 2–A LEO–GEO tug was designed around a 5000-N thermal plasma engine. At 6000 K and 

a chamber pressure of 50 psia, the engine achieves a specific impulse Isp of 813 s with a steam (H2O) 

monopropellant. Forty-four megawatts of beamed power must reach the thrust chamber to achieve the 

required 6000 K temperature. To realize this energy level, power will be beamed from the ground and 

focused into the thrust chamber through primary and secondary reflectors and a large sapphire window. 

The primary reflector will actively track the laser emitters. Figure 3.143 shows a concept drawing of the 

tug and propellant stage. 

The beamed-energy tug propulsion stage shown in Figure 3.144 is 5.78 m long and cylindrical, 

consisting of a 4.0- by 5.6-m elliptical primary reflector cantilevered off the side of the vehicle on a ±60

gimbal. Two degrees of freedom will be required to track the ground-based emitter, and either the entire 

vehicle will be rolled during a burn or another degree of freedom will be added to the gimbal (although a 

two-degree-of-freedom gimbal creates complications because the rotation does not occur about the focal 

point). The secondary reflector will be fixed. Both reflectors and shielding to protect the entire vehicle 

from the laser flux will be 97% efficient, 0.9999 reflective carbon fiber composite. A 1.5- by 3.3-m solar 

array will supply 560 W of nominal operating power, and a 2000-W-hr lithium-ion battery pack will 

supply an additional 1625 W during 5- to 15-min burns to run pumps, tracking, and so forth. The main 

engine will provide 5000 N of thrust with 50-psia pressure in a 30-cm spherical combustion chamber. The 

engine will use water heated to 6000 K to provide 813-s Isp. The plasma ball will be electromagnetically 

suspended in the chamber, and the nozzle will be regeneratively cooled or film cooled. The beamed 

energy will pass through a 10-cm-diameter sapphire refractory window into the combustion chamber. The 

nozzle will have an area ratio of 50, will be regeneratively cooled, and will be gimbaled which will allow 

for orientation for communication pointing on a screw or ball mechanical motor. 

The thermal plasma engine was selected because of the potential to provide sufficient thrust to 

perform the LEO–GEO transfer in a short time at a high enough Isp to reduce fuel mass to enable the 

launch of two GEO satellites per launch. 
NASA/TM—2012-217014 434 

Figure 3.143.—Conceptual tug and propellant stage. RCS, Reaction Control System. 

Figure 3.144.—Conceptual tug, propellant stage, and GEO satellites. 

3.2.2.2 Description of Ground Facility 

The DRM 2 mission as defined requires two types of laser ground facilities, a 2- to 3-m-diameter 

aperture Perigee Facility for burns for altitudes between 400 and 3000 km and a much larger diameter 

Apogee Facility for burns at GEO. Both types of laser facilities will be required to deliver about 50 MW 

of laser energy to the spacecraft. On the basis of atmospheric losses, each of the ground facilities will 

need an 80MW laser to deliver the required power to the tug. The apertures for these ground facilities will 

resemble large astronomical telescopes such as the Starfire Optical Range (SOR) and the proposed 

30-Meter Telescope (TMT). 
NASA/TM—2012-217014 435 

All of the facilities will have two main components, a Beam Generation system and a Beam Control 

System. In addition, each facility will require sufficient power and heat dissipation capacity for the 

operation of the laser. 

3.2.2.2.1 Beam Generation 

The Beam Generation system will create the photons at the power levels needed for beamed-energy 

propulsion (BEP). The laser could be solid state, chemical, or electrical, but the wavelength must be 

optimized for transmission through the atmosphere. The 80-MW laser required by this DRM is a 

technological challenge in that currently the chemical oxygen iodine laser (COIL) on the Airborne Laser 

Test Bed (ALTB) is the only megawatt-class laser currently in use. Solid state and electrically driven 

lasers are rated at about 100 kW. 

The Solid State Laser’s (SSL’s) current efficiency of power output to all power consumed (wall-plug 

efficiency) is about 8% to 10%. There are significant variations to the above numbers (IPG claims ~25% 

wall-plug efficiency), but none of those devices offer a path to the power goal as coherent light. 

Chemical lasers such as COIL offer a less risky path for the power goal, but these devices require 

pressure recovery systems that would eliminate any possibility of transportability at that power level. 

Also, the environmental opposition to the chemicals involved would pose a high political price for any 

program. COIL has shown that its power is directly scalable by adding additional laser modules in series. 

The research has also shown that COIL’s performance could be dramatically improved with modified 

laser fuels and an improved singlet oxygen generator. The COIL runs at a very low internal pressure and 

requires the use of large vacuum reservoirs. It is feasible that enough laser modules could be linked 

together to generate the needed 80 MW. At this power (flux) level, the critical technologies would be the 

vacuum management, the optics, and the ability to maintain the integrity of the coatings. 

An electrically driven 80-MW laser will require about 200 MW of electrical wall-plug power to 

operate. Laser run times are projected to be 5 to 15 min, and this power could be provided from batteries 

or flywheel energy storage systems instead of being drawn directly from the grid. Less developed than 

SSL, Diode-Pumped Alkali Laser (DPAL) gain generators use rubidium or cesium gas as the lasing 

species with narrow-band diode lasers for the pump. Their efficiency will be similar to SSL since both 

suffer from the poor efficiency of the diode laser pumps estimated today at about 50% (reported higher 

efficiencies usually ignore the losses due to poor spectral matching with the stimulated species). To date, 

no electric laser has been able to show direct scalability to 80 MW. At this power level, the laser would be 

2 orders of magnitude larger than any current state-of-the-art electrically driven laser. Current research on 

DPALs shows promise and scalability at a basic physics level, but current efforts have not been able to 

demonstrate the potential in the lab. If the efficiency of SSLs and DPALs were improved in terms of total 

efficiency and wall-plug efficiency, the DPALs would be the better choice because of the thermal effects 

of the SSL medium. 

A high risk for DPALs is the design of a resonator that can produce 80 MW of DPAL power. The 

highest power generated to date is in the hundreds of watts. The high gain per centimeter of the DPAL 

gain medium requires innovative designs similar to the efforts proposed for chemical devices in the 

1980s. A possible strawman device would use a short-length annular gain cell with half-symmetric axicon 

mirrors to compact the annular resonator, such as the one shown in Figure 3.145. 

Figure 3.145.—Half-symmetric (HS) axicon mirror annular resonator for possible DPAL emitter. 
NASA/TM—2012-217014 436 

DPAL resonators have been predicted with power densities of 40 kW/cm2. Therefore, achieving the 

required 80 MW leads to an annular ring on the order of 1.5 m in diameter with a length (longitudinal 

axis) of 5 to 8 cm and a radial cross section of 5 cm. Packaging of the diode pumps and gas flow for 

cooling would clearly be an engineering challenge. 

Obvious challenges for the laser source and beam path include improvement in wall-plug efficiency 

10% to 50% (dominated by pump diodes), improvement in the reliability of very low absorption (VLA) 

coatings to handle power densities without thermal distortion, and improvements in beam path 

conditioning to avoid thermal blooming in the code path. 

3.2.2.2.2 Beam Control 

The Beam Control system will resemble the optics of a large telescope and will provide the 

acquisition, tracking, and pointing required to beam the energy to the orbiting spacecraft. It also will have 

the safety systems required to shut down the laser in the event of an optical coating failure, beam 

misalignment, or a predictive avoidance issue. In addition, the Beam Control system will use adaptive 

optics such as those at the Starfire Optical Range (Figure 3.146) for atmospheric compensation and fast 

steering mirrors for fine beam alignment and jitter control. It may also be possible to use radiant intensity 

at the target, known as gradient descent optimization (GDO) wave-front control for the atmospheric 

compensation measurement. 

Figure 3.146.—Starfire optical range (SOR). 

3.2.2.2.3 Perigee Facilities 

Three Perigee Facilities spaced evenly around the globe on or near the equator should provide the 

best viewing opportunities to enable perigee burns on nearly every orbit. DRM 2–A uses the Kwajalein 

Atoll as the launch site, which is located at a latitude of 9o. The Perigee Facilities will require only 2-m 

apertures to compensate for beam diffraction, but the ability of the optical coatings to handle the 80-MW 

power levels may drive larger optics to lower the flux. The Perigee Facilities will most likely resemble the 

3-m telescope at SOR at Kirtland Air Force Base (see Figure 3.143) or the Advanced Electro-Optical 

System (AEOS) 3.67-m telescope on Maui. The near equator locations of the Perigee Facilities implies 

they could be ship based. Lasers based on ships similar to the now-decommissioned Virginia-class 

cruisers with two 150-MW reactors or a nuclear-powered aircraft carrier would provide sufficient power 

and mobility to maximize the burn duration for each orbit of the tug. If some or all of the Perigee 

Facilities are ship based, the lasers may require additional power to compensate for the additional water 

content in the air due to the marine environment. 
NASA/TM—2012-217014 437 

3.2.2.2.4 Apogee Facility 

There will be more flexibility on the location, but the Apogee Facility will need to be at least 30 m in 

diameter to focus the spot on the spacecraft’s reflector at GEO. The facility could be located below a 

latitude of 45o and still provide good lasing opportunities at both the 3000-km altitude (and so could be 

used for perigee burns) and GEO. The TMT (Figure 3.147) being constructed at Mauna Kea, Hawaii, is a 

good representation of a baseline Apogee Facility. The laser requires the same 80-MW power as the 

Perigee Facility and for practical reasons could be identical. 

 

Figure 3.147.—30-Meter-Telescope (TMT). Courtesy TMT Observatory Corporation; used with permission. 

3.2.2.3 CONOPS—Frequency of Launch and Repeat Time 

The LEO–GEO tug described here will halve the number of launches required to place satellites in 

GEO (Figure 3.148). A standard CONOPS will involve the launch of two 1970-kg GEO satellites and 

4895 kg of water propellant on a Falcon 9 Block 2 from Kwajalein into a 400-km circular orbit with an 

inclination of 9. The tug will rendezvous with the stack of two GEO satellites and propellant and begin 

orbit-raising burns within 3 days of launch. Six 5- to 8-min burns will be required to raise the initial orbit 

to a 3000-km, 9o-inclined circular orbit. Three 15-min burns at 3000 km will be required to raise the 

apogee to GEO, one 15-min apogee burn will be required to change planes to a 0o

 inclination, and two 

more 15-min apogee burns will be required to circularize. The total transit time to GEO, if we assume that 

there are burns every orbit, will be 32 hr, requiring 4.2 km/s v. 

Figure 3.148.—Tug rendezvous with propellant and GEO satellites. 

GEO Satellite LEO-GEO Tug Propellant 

Stage GEO Satellite 
NASA/TM—2012-217014 438 

Following undocking from the GEO spacecraft, the tug will return to a 400-km, 9o inclined orbit, 

using residual water propellant that was launched with the two GEO spacecraft, and will be ready for the 

next transit in 24 hr. We assume that the tug itself will have been launched separately. The overall 

concept is summarized in Table 3.29, and orbital diagrams are shown in Figure 3.149. Figure 3.150 shows 

the tug stowed for launch. 

TABLE 3.42.—CONOPS FOR LEO–GEO LASER TRANSPORT 

Test step Beaming Receiving Control and operation 

0 Launch of tug on Falcon 9 from Kwajalein Atoll 

1 Launch of two 2000-kg GEO satellites and 

5000-kg propellant on Falcon 9 from Kwajalein Atoll 

1a Rendezvous and docking with tug 

2 to 7 Perigee Tug + two GEO satellites 400-km circular orbit with 9o inclination to 

3000-km circular orbit with 9o

 inclination 

8 to 10 Perigee Tug + two GEO satellites 3000-km circular orbit with 9o inclination to 

3000- by 36,000-km orbit with 9o

 inclination 

11 Apogee Tug + two GEO satellites 3000- by 36,000-km orbit with 9o inclination to 

0o

 inclination 

12 to 13 Apogee Tug + two GEO satellites 3000- by 36000-km orbit with 0o inclination to 

36,000-km circular orbit with 0o

 inclination 

14 Apogee and perigee Tug Return to LEO 

Other operational scenarios include the tug rendezvousing with existing GEO spacecraft for either 

servicing or return to LEO. 

Figure 3.149.—LEO–GEO Laser Transport Orbital Transfer. 
NASA/TM—2012-217014 439 

Figure 3.150.—LEO–GEO laser transport stowed for launch. 

3.2.2.4 Costs—Comments 

3.2.2.4.1 Recurring Costs 

The DRM 2–A LEO–GEO beamed-energy transport tug will cut recurring GEO lost costs in half. 

This savings will be enabled by the elimination of the Centaur upper stage and circularization engine on 

the satellite required in current GEO launches. Instead, two 2000-kg GEO satellites could be launched 

into LEO on a single Falcon 9, rendezvous with the tug, and be transported to their circular GEO orbits in 

32 hr. The tug would then return to LEO for another transport. In this manner, two GEO satellites could 

be placed into their orbits on a single Falcon 9 launch. 

Additional recurring costs would include the negligible liquid water fuel cost and the cost to operate 

the laser. For the purposes of this project, a usage cost of 0.12 $/kW-hr can be assumed. Although this 

cost is not negligible, it is less than 10% of the launch cost. 

Therefore, with the combination of (1) the reduction in cost versus current launch vehicles of the 

Falcon 9 and (2) the ability to stack two satellites on a single launch, recurring costs of GEO launch will 

be cut in half. 

3.2.2.4.2 Nonrecurring Costs 

The Mauna Kea TMT provides a good parametric comparison to the laser ground facilities required 

for the DRM 2 mission. The cost estimate for building a Giant Segmented Mirror Telescope such as the 

TMT is $1.1 to $1.4B. The TMT is scheduled for completion in 2018. The ABL, which is currently the 

only megawatt-class laser in use, cost about $5B to develop and build. The ABL laser and beam control 

systems have been estimated to cost about $1B each. The High Energy Laser Joint Technology Office 

estimates that it currently costs about $1000/W for a solid-state laser. If we assume that there will be 

significant reductions in cost in the coming decades, it will likely cost about $10B to build the required 

80-MW laser. Because of the limited usage time of the lasers (maximum 30-min burn), substantial costsharing could be realized with other applications. 

Moore’s law of doubling results in about 10 steps to get to 80 MW from the current 100-kW electric 

lasers. If laser power doubles occurs every 3 years, ground lasers should be available around 2038. 

3.2.2.5 Feasibility Assessment 

The following is a list of issues which need to be addressed 

1. What are the attenuation effects of high flux levels (MW/m2

) at the frequencies of interest on the 

atmosphere? 

2. What would be the expected availability and transmittance of the BEP ground terminal at a given 

frequency, given aerosol and precipitation conditions at a proposed ground site? 

3. How large of a keep-away zone must be maintained around and beyond an ascending craft to 

accommodate reflection, refraction and scattering components of the beam exceeding substantial 

irradiances, and what level is to be considered substantial? Considerations should be given to 
NASA/TM—2012-217014 440 

maximum permissible exposure for humans, and damage thresholds for ground, airborne and 

space-based assets within proximity of the BEP activity. 

4. For a ground based terminal operations, how much grid power over what duration is required? 

To meet these requirements, what would the generation, storage and conversion system look like? 

5. How would a micro/mm-wave BEP system work within the spectrum constructs of the 

NTIA/FCC? 

6. Does it make sense to collocate beamed energy propulsion and power technologies on the craft, 

for each phase of a proposed mission? 

7. Would there be a benefit to developing a hybrid aperture to handle multiple BEP frequencies 

depending upon range to target, atmospheric conditions, etc.? 

8. Can retro-directive control be utilized to aid in the beam pointing, acquisition and tracking of the 

craft? 

9. What is the most practical way of performing high power beam combination using COTS 

technology for the frequencies of interest? What are the potential near term accomplishments that 

could influence this capability? 

10. What are the cost/benefit tradeoffs in considering multiple BEP transmission locations, both at 

the ground range and at an intermediate relay point in space? 

11. Could deployable array technologies by implemented to decrease the weight of a BEP aperture, 

and increase the allowable size that can be launched?

12. What are some near term demonstrations that may be accomplished to establish a 

maturity level for critical technologies, and support the overall feasibility of the BEP 

concept?

3.2.2.5.1 Roadmaps and Technology Readiness Levels 

The five technology drivers for DRM 2 are the 80-MW laser, high-flux optical coatings, the beam 

director, the 5000-N laser plasma engine, and the reflective surfaces for the laser mirrors. Table 3.43 

summarizes the current Technology Readiness Levels (TRLs) and time required to reach TRL 6. 

TABLE 3.43.—TECHNOLOGY READINESS 

3.2.2.5.2 Ground and Flight Demonstrations 

One possible path to developing the 5000-N, 50-MW laser plasma engine is the series of ground 

demonstrations and tests summarized in Figure 3.151. The first step will be to use the solar thermal 

plasma chamber to measure temperatures and thermal loads in the plasma chamber. This small-scale 

thermal plasma engine testing will test the ability to create and maintain a plasma ball and could take 

place at the Satellite Propulsion Complex at Edwards Air Force Base. These tests will require an engine 

test section and a high-energy laser. The primary objectives will be to create and control the plasma ball 

and to measure engine thrust and Isp. This testing should be able to be completed for approximately $5M. 

Key technology Current 

TRL 

Rational and proof of estimation Time to reach 

TRL 6 

Laser 80 MW 2 Approximately 100 kW is the largest electric laser available. 2038 

High-flux optical coatings 2 Current optics accommodate current laser power. 2038 

Beam director (30-m mirror) 3 The TMT is scheduled to be complete in 2018. 2018 

Thermal laser engine chamber 2 Some solar plasma work has been accomplished but not at this 

energy level or thrust. 

2025 

Laser engine reflectors 3 The secondary reflector is very-high-energy-loaded optic, beam 

control, and thermal control critical. 

2025 
NASA/TM—2012-217014 441 

Figure 3.151.—Proposed technology roadmap. 

The second step will be to build and test the performance of a full-scale engine at the 100-kW laser 

level. This engine would be a small-scale version the final 50-MW engine, and it would use current 

100-kW lasers as the power source. Whereas the first chamber only created the plasma and measured its 

temperature, the second one will be a complete engine. It will need a propellant feed system and an 

expansion nozzle to measure the thrust. This engine should be used to fully understand the issues of 

plasma containment, propellant feed, and thrust control. It should also be space qualified. These tests will 

again use the Satellite Propulsion Complex at Edwards Air Force Base, and should be completed for 

approximately $30M. 

The first flight test should demonstrate the thermal plasma engine in space. It will require a highenergy laser-beaming facility capable of beaming to LEO, such as SOR. The space-qualified engine will 

be integrated onto an Expendable Evolved Launch Vehicle (EELV) Secondary Payload Adapter (ESPA)-

class satellite to demonstrate engine performance in space and perform impulsive orbit-raising 

maneuvers. Because the performance of the engine is scalable with the input power, a full-power 80-MW 

facility will not be required. A power-beaming facility on the order of 0.5 to 1.0 MW will prove the 

technology. This flight test could be completed for $50M. 

When the full-power laser does become available, another ground demonstration should be run with a 

fully functioning version of the required flight engine. In the 2035 time frame, lasers may able to provide 

50 MW of power, and this engine should be ready take advantage of those new lasers. It will take a few 

more years before the lasers are at the 80 MW of power needed to actually demonstrate the engine 

in space. 

The development of an 80-MW laser and its optical systems will also be a challenge. There is a fair 

amount of high-power laser work going on within the Department of Defense (DOD), and hopefully, this 

work will progress to the power levels required for this mission. 

3.2.2.5.3 Issues and Resolutions 

The development of the laser plasma engine and the 80-MW laser are the two key technology issues 

that must be addressed. The technology roadmap described earlier provides a possible resolution. There 

will also be geopolitical issues with the construction of multiple 80-MW laser ground stations because 

other nations will consider the ground facilities to counter-space weapons. One path to resolve this 

political issue would be to build the ground facilities as a multiuse international effort for spacecraft 

Development Year

0

20

40

60

80

100

120

140

2010 2013 2016 2019 2022 2025 2028 2031 2034 2037 2040

Laser Power (MW)

Required 80MW

In ~2038

Current Electric 

Lasers 100kw 

1 MW Electric Laser

Solar Plasma

Chamber 100 kw Laser

Plasma Engine

50MW Laser

Plasma Engine

Electric Laser Power

3yr Doubling Cycle

Development Year

0

20

40

60

80

100

120

140

2010 2013 2016 2019 2022 2025 2028 2031 2034 2037 2040

Laser Power (MW)

Required 80MW

In ~2038

Current Electric 

Lasers 100kw 

1 MW Electric Laser

Solar Plasma

Chamber 100 kw Laser

Plasma Engine

50MW Laser

Plasma Engine

Electric Laser Power

3yr Doubling Cycle
NASA/TM—2012-217014 442 

propulsion, debris removal, and possibly planetary defense. The other path will be to openly state that the 

ground stations are part of our national defense infrastructure but will be used primarily for on-orbit 

operations. 

The U.S. Strategic Command Satellite Clearing House is responsible for determining when U.S. 

lasers can be fired into space so as not to cause damage to satellites. There may be a scheduling issue 

dictating when the laser can be fired, especially for the GEO burns. This issue could be resolved for the 

most part during mission planning, with contingencies for missed burns. 

3.2.2.6 Bimodal Operation 

If the ground facilities are designed for dual use, there will be the advantage of cost sharing between 

the DOD, NASA, and possibly even international partners. There will be very little difference between a 

ground-based laser used for propulsion and one used for national defense. Both the national defense and 

debris removal modes will require the ability to find and track uncooperative targets. If the DOD pursues 

the dual-use model, they may also invest in a space-based relay system. A space-based relay system 

would allow other missions, such as missile defense, in addition to debris removal and space propulsion. 

3.2.3 Low-Earth-Orbit to Geosynchronous-Earth-Orbit Consultant Reports—Design Reference 

Mission 2–B 

Design Reference Mission 2–B (DRM 2–B) is a mid-term capability based on DRM 2–A that 

replaces the multiple 80-MW ground-based beam emitters with a single 1-MW emitter if ground-based 

(or a 0.5-MW emitter if space-based). The design of the tug is identical to DRM 2–A. The modifications 

to the DRM 2–A mission are the concept of operations (CONOPS) of the power transfer, the orbital 

transfer burns, and the amount of water propellant required for a single transfer. In this manner, two GEO 

satellites could be placed into their orbits on a single launch in approximately 20 days. 

3.2.3.1 Description of Vehicle 

The low-Earth-orbit to geosynchronous-Earth-orbit (LEO–GEO) tug would be identical to that 

discussed in Section 3.2.2.1. 

3.2.3.2 Description of Ground Facility 

Two options are available for providing an in-space emitter capability: a 0.5-MW in-space GEO laser 

emitter and a 1.0-MW ground-based emitter and a GEO relay mirror. In the first option, 0.5 MW would 

be received at the tug (versus 50 MW in DRM 2–A) to power the 50-N laser plasma engine (versus 

5000 N in DRM 2–A). For the laser to emit 0.5 MW, if we assume a 30% laser efficiency, 1.667 MW of 

input power be will required; 4065 m2

 of solar array area will be required to generate that much power if 

we assume that the triple-junction cells used will be 30% efficient. In addition, 4600 m2

 of radiator area 

will be required to dissipate the excess heat from the laser. For full power to be realized at a range of 

81,000 km, a 57-m-diameter emitting aperture will be required. These requirements to realize a 0.5-MW 

GEO-based emitter do not represent midterm capability, and so a ground-based emitter with a GEO relay 

mirror also was considered. 

The most significant technology challenge of the DRM 2–A mission, which would likely require 

more than a decade of technology development to achieve, is the 80-MW ground facility. A more nearterm 1-MW emitter would have approximately the same configuration as the 80-MW apogee facility 

described in Section 3.2.2.2.4, including a 10-m-diameter aperture to maintain beam quality at GEO. A 

single-relay 7-m-diameter James Webb Space Telescope-type mirror at GEO would then steer the beam 

from the ground to the tug, providing continuous power. 

The tug would be in view of the relay mirror for approximately 85% of its orbit; however, because of 

spot size growth, it would receive the full 0.5 MW of power for less than half of its orbit. This reduction 

in power will correspond to a reduction in thrust from 50 N, resulting in a longer transfer time—on the 

order of 20 days. 
NASA/TM—2012-217014 443 

3.2.3.3 CONOPS—Frequency of Launch and Repeat Time 

The change from 50-MW of power provided to the tug for multiple impulsive burns in DRM 2–A 

enabling a 32-hr transit time from LEO to GEO to 0.5 MW of continuous power for DRM 2–B will 

necessitate a longer transfer time. If we assume that the engine is identical engine to that of DRM 2–A, 

the 2 orders of magnitude reduction in received power will correspond to a 2 order-of-magnitude 

reduction in thrust (from 5000 to 50 N). This linear scaling would be achieved by maintaining the 6000-K 

plasma ball temperature by decreasing the propellant flow rate and thereby maintaining the 810-s specific 

impulse Isp. A 50-N engine does not provide enough thrust for a Hohmann transfer on this size of a 

vehicle; therefore, a continuous-thrust spiral transfer will be used. Changing from the Hohmann to the 

spiral transfer will require an additional 1.8 km/s v (approx 6.0 vs. 4.2) and, therefore, will require 

approximately 6400 kg of water propellant instead of 4895 kg to transfer two 1970-kg GEO satellites and 

return to LEO. The total transfer time will also be greater—approximately 20 days instead of 32 hr. This 

CONOPS is summarized in Table 3.44. 

TABLE 3.44.—CONOPS FOR LEO–GEO LASER TRANSPORT 2–B 

Test 

step 

Beaming Receiving Control and operation 

0 Launch of tug and 8000-kg propellant on Falcon 9 from Kwajalein Atoll 

1 Launch of two 2000-kg GEO satellites and 

5000-kg propellant on Falcon 9 from Kwajalein Atoll 

1a Rendezvous and docking with tug 

2 0.5-MW laser 

continuous 

Tug + two 

GEO satellites 

Spiral transfer from LEO to GEO in 20 days 

3 0.5-MW laser 

continuous 

Tug Return to LEO 

Note that the spiral transfer will require 6400 kg of propellant, exceeding the capacity of the Falcon 9 

launch vehicle by 1400 kg. Additional propellant will be launched with the tug, enabling five transfers 

(10 satellites delivered to GEO). For additional transfers, it is assumed that the required propellant will be 

supplied by other launches or a depot. 

3.2.3.4 Costs—Comments 

3.2.3.4.1 Recurring Costs 

The DRM 2–B LEO–GEO beamed-energy transport tug will cut recurring GEO lost costs in half. 

This savings will be enabled, as in DRM 2–A, by the elimination of the Centaur upper stage and the 

circularization engine on the satellite required in current GEO launches. Instead, two 2000-kg GEO 

satellites could be launched into LEO on a single Falcon 9, rendezvous with the tug, and be transported to 

their circular GEO orbits in 20 days. The tug would then return to LEO for another transport. In this 

manner, two GEO satellites could be placed into their orbits on a single Falcon 9 launch. 

Additional recurring costs would include the negligible liquid water fuel cost and the cost to operate 

the laser. For the purposes of this project, a usage cost of 0.12 $/kW-hr can be assumed. Although this 

cost is not negligible, it is less than 10% of the launch cost. 

Therefore, with the combination of the reduction in cost versus current launch vehicles of the 

Falcon 9, and the ability to stack two satellites on a single launch, recurring costs of GEO launch will be 

cut in half. 

3.2.3.4.2 Nonrecurring Costs 

In the first power transfer option, a 0.5-MW emitter will be placed in GEO. On the basis of the 

DRM 2–A sizing of solar arrays and radiators, 8300 kg in solar arrays would be required if we assume an 

efficiency of 200 W/kg (ISS arrays are 32 W/kg efficient) and 20,000 kg in radiators would be required if 

we assume that they would have twice the mass of the solar arrays. This option will also require the 
NASA/TM—2012-217014 444 

in-space construction of a 57-m-diameter aperture. The largest GEO transfer capacity currently available 

is on a Delta IV heavy launch vehicle at 5000 kg. If we assume a bulk purchase of 10 launches costing 

$200M each (Delta IVs are currently estimated at $1B each), $2B would be incurred. The launches break 

down as follows: two launches for the solar arrays, four launches for the radiators, and four launches for 

the assembled or inflated aperture. The cost for the hardware itself is estimated at $10B. 

The second power transfer option involved a 1-MW ground-based emitter and a GEO relay mirror. 

The 10-m-diameter ground-based aperture and 1-MW emitter are estimated to cost $500M. If we assume 

that the GEO relay mirror is a copy of the James Webb Space Telescope, it would have an estimated cost 

of $500M. Therefore $1B in nonrecurring total infrastructure is estimated. The nonrecurring tug launch 

on a Falcon 9 is assumed to cost $30M. 

3.2.3.5 Feasibility Assessment 

The GEO relay mirror option is feasible in the mid-term, requiring approximately 10 years for initial 

operational capability. 

3.2.3.5.1 Roadmaps and Technology Readiness Levels 

The five technology drivers for DRM 2–B are the 1-MW laser, the GEO relay mirror, the beam 

director, the 50-N laser plasma engine, and reflective surfaces for the laser mirrors. The current 

Technology Readiness Levels (TRLs) and time required to reach TRL 6 are summarized in Table 3.45. 

TABLE 3.45.—TECHNOLOGY READINESS FOR DRM 2–B

Key technology Current 

TRL 

Rational or proof of estimation Time to reach 

TRL 6 

1-MW laser 2 100-kW lasers are approximately the largest electric lasers available. 2015 

GEO relay mirror 4 The James Webb Space Telescope 6.5-m-diameter mirror is scheduled for 

launch in 2014. 

2018 

Beam director 

(10-m mirror) 

6 30-m Telescope (TMT) is scheduled to be complete in 2018. 2018 

Thermal laser engine 

chamber 

2 Some solar plasma work has been accomplished, but not at this energy level or 

thrust. 

2018 

Laser engine 

reflectors 

3 The secondary reflector is very-high-energy-loaded optic, beam control, and 

thermal control critical. 

2018 

3.2.3.5.2 Ground and Flight Demonstrations 

Ground and flight demonstrations will be substantially similar to those discussed in Section 3.2.2.5.2 

scaled to the 50-N laser plasma engine and 1.0-MW ground-based emitter. 
NASA/TM—2012-217014 445 

3.2.4 Low-Earth-Orbit to Geosynchronous-Earth-Orbit Cost 
NASA/TM—2012-217014 446 
NASA/TM—2012-217014 447 

3.2.5 Deep Space Mission Design Lab (MDL) Report 
NASA/TM—2012-217014 448 
NASA/TM—2012-217014 449 
NASA/TM—2012-217014 450 
NASA/TM—2012-217014 451 
NASA/TM—2012-217014 452 
NASA/TM—2012-217014 453 
NASA/TM—2012-217014 454 

Attitude Control, p10

Presentation Version

M i s s i o n D e s i g n L a b o r a t o r y

Do not distribute this information without permission

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Date 11/29 to 12/3

BEPS

Handling Time of Flight Lag

• At start of burn:

– Range = 11900 km

– Round-trip light time of flight = 0.08 sec

– Tracking rate = 0.025 deg/sec

– Pointing direction must “lead” observed position by 7 arcsec

• At end of burn:

– Range = 44200 km

– Round-trip light time of flight = 0.3 sec

– Tracking rate = 0.005 deg/sec

– Pointing direction must “lead” observed position by 5 arcsec

If you see it here...

...aim here
NASA/TM—2012-217014 455 
NASA/TM—2012-217014 456 
NASA/TM—2012-217014 457 
NASA/TM—2012-217014 458 
NASA/TM—2012-217014 459 

Communications p8

Final Version

M i s s i o n D e s i g n L a b o r a t o r y

Do not distribute this information without permission

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Nov 29-Dec 3, 2010

BEPS-DMR 3

Signal Margins

BEP Module Link Margin (dB) Comments 

S-Band telemetry 

to ground 

30 kbps 3.2 Omni to 11 meter 

100,000 km 

S-band telemetry to 

ground 

4 Mbps 22.1 Omni to 11 meter 

1000 km 

S-band to Galileo 50 kbps 0.8 Omni to Omni 

50 Km range 

S-band uplink 

 from ground 

20 kbps 12.9 11 meter to omni 

100,000 km 

TDRSS Return 

Link 

10 kbps 0.1 Omni to TDRSS 

TDRSS Forward 

Link 

2 kbps 2.0 TDRSS to Omni 

Note: Applicable Near Earth Network (NEN) station include Hawaii 13-m USN, 

Wallops 11-m, Santiago `12/7-m or 9-m, Dongara, Aus.13-m USN. 
NASA/TM—2012-217014 460 
NASA/TM—2012-217014 461 
NASA/TM—2012-217014 462 
NASA/TM—2012-217014 463 

Electrical Power, p7

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M i s s i o n D e s i g n L a b o r a t o r y

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Date 11/29 to 12/3

BEPS

EPS Block Diagram

Battery 

Module

28vDC Payload Bus Module

1553

28vDC Battery Bus Module

I Battery

T

V

ASRG (Advance 

Stirling Radioactive 

Generator)

Payload Module 

(Science Instruments) 

+ ASRG Ring = 

Galileo II Spacecraft 

Bus

Propulsion Module 

Loads

Control Module (CM), FPGA & 

PSE Specific software, PSE H/W & 

Spacecraft I/F.

Control Module (CM), FPGA & 

PSE Specific software, PSE H/W & 

Spacecraft I/F.

1553

To CM

To C&DH

To CM

1553 To C&DH

ASRG (Advance 

Stirling Radioactive 

Generator)

ASRG (Advance 

Stirling Radioactive 

Generator)

ASRG (Advance 

Stirling Radioactive 

Generator)

1553 To CM

1553 To CM

1553 To CM

1553 To CM

1553 To CM

Remainder of the 

Galileo II Spacecraft 

EPS not specified.

1553 To CM

Propulsion Module Launch Interface 

Payload Module Separation Interface, 

Remaining configuration becomes the 

Galileo II Spacecraft. 

ASRG Ring

Payload Module Launch Interface 
NASA/TM—2012-217014 464 

Electrical Power, p9

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M i s s i o n D e s i g n L a b o r a t o r y

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Date 11/29 to 12/3

BEPS

Electrical Power Curve

400.00

450.00

500.00

550.00

600.00

0 365 730 1095 1460 1825 2190 2555 2920 3285 3650

P

o

w

e

r

i

n

W

a

t

t

s

Mission Time (Days)

GRC-BEPS Mission Over 10 Yr Life With ASRG'S; Average Load During 

Day=376.714W; Average Load During Night=360.0909W

Daylight Average ASRG Power Provided

Daylight Average ASRG Power Required

400.00

450.00

500.00

550.00

600.00

0 5 10 15 20 25 30 35 40 45 50 55 60 65 70

P

o

w

e

r

i

n

W

a

t

t

s

Mission Time (Days)

GRC-BEPS Mission Over 10 Yr Life With ASRG'S; Average Load During 

Day=376.714W; Average Load During Night=360.0909W

Daylight Average ASRG Power Provided

Daylight Average ASRG Power Required

3 ASRG’s

4 ASRG’s

Prop 

Module 

Load

Galileo II 

Spacecraft 

Load

27.6 

watts 

Margin 

after 

10 

Years
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NASA/TM—2012-217014 467 

Flight Dynamics, p4

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BEPS

Jupiter Transfer Trajectory Design

• Jupiter Transfer Trajectory Insertion (TTI) Burn

– Epoch: 19 Dec 2016 12:19:49.716 UTCG

– Duration: 100.4 minutes

– Departure C3: 80 km2/s2, V: 8.9 km/s

– V: 6.7 km/s

– Propellant Used: 3716 kg

– Altitude range over burn: 11327 km – 43919 km

– Flight Path Angle range over burn: 35.64 – 66.43 deg

– Post-Burn Mass: 3183 kg

• BEPS Dry Mass Plus BEPS Payload: 3100 kg

• Residual BEPS Propellant: 83 kg • “

• BEPS Propulsion Stage Dropped

– Spacecraft mass on Jupiter transfer trajectory : 2223 kg

• Dry Mass: 1300 kg

• Propellant Mass: 923 kg

• Jupiter Transfer Orbit

– Transfer time is approximately 2.5 years

– No accounting for trim maneuvers

• Jupiter Orbit Insertion (JOI) Burn

– 600 kg propellant allocated (per Galileo)

– Epoch: 31 May 2019 06:59:56.224 UTCG

– Duration: 70.2 minutes

– Arrival C3: 33.8 km2/s2, V: 5.8 km/s

– V: 971.8 m/s

– Capture Orbit: 200,000 km x 12,000,000 km -- Period: 98 days

Flight Dynamics, p5

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Transfer Trajectory Insertion (TTI) 

Ignition

Parking Orbit

TTI Maneuver

Coast

Geosynchronous 

Altitude
NASA/TM—2012-217014 468 

Flight Dynamics, p6

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BEPS

Transfer Trajectory Insertion (TTI) 

Burnout

Parking Orbit

TTI Maneuver

Coast

Geosynchronous 

Altitude

Flight Dynamics, p7

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BEPS

Laser Site Angles/Rates During TTI Burn
NASA/TM—2012-217014 469 

Flight Dynamics, p8

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BEPS

Laser Site Range/Range Rate During TTI 

Burn
NASA/TM—2012-217014 470 
NASA/TM—2012-217014 471 
NASA/TM—2012-217014 472 

I n t e g r a t e d D e s i g n C e n t e r / M I s s I o n D e s I g n L a b o r a t o r y

N A S A G O D D A R D S P A C E F L I G H T C E N T E R

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Mechanical

Sara Riall David Peters

November 29-December 03, 2010

“I believe there is no philosophical high-road in science, with 

epistemological signposts. No, we are in a jungle and find our way 

by trial and error, building our road behind us as we proceed. “

Max Born (1882-1970) German Physicist. Nobel Prize, 1954.

Discipline, p2

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BEPS

Galileo & BEPS
NASA/TM—2012-217014 473 

Discipline, p3

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BEPS

Launch Configuration of the 

Propulsion and Propellant Stage

Discipline, p4

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BEPS

Launch Configuration of Galileo 

Payload

Note: Dimensions correspond closely 

with Galileo 
NASA/TM—2012-217014 474 

Discipline, p5

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BEPS

Deployed Configuration 

Note: Dimensions correspond closely 

with Galileo 

Discipline, p6

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Propulsion Stage Deltas From DRM-2

Original Layout

1) Added Propellant Stage in Launch Configuration

2) Removed Solar Array

3) Moved docking to Propellant stage

Current Layout

Note Main Reflector in DRM-2 version removed for simplicity

Secondary reflector
NASA/TM—2012-217014 475 

Discipline, p7

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BEPS

Propellant stage deltas

Original Layout

1) Moved from Payload to Propulsion for Launch

2) Added 3 ASRGs

3) Added Separation ring to Tank/Structure

4) Expanded Tank Section to 3.3 m

Current Layout

M i s s i o n D e s i g n L a b o r a t o r y

Mass Rack-up

1. Structures and mechanisms Propellant stage (COMPASS) 81.44

2. Structures and mechanisms Propulsion Stage (COMPASS) 62.97

3. Structure mass increase to support 3 ASRG’s (MDL) 13.45

4. Addition of separation system (MDL) 22.32

180.18
NASA/TM—2012-217014 476 

Discipline, p9

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BEPS

Docking and Separation Ring

Note: 

1) The Docking Adapter was moved from the Propulsion 

Stage/Propellant Stage interface to the Propellant/Payload 

interface.

2) A separation system was added to the propellant stage to 

separate the assembly after the trajectory burn. This interface 

allows the fuel tank to be jettisoned with the propulsion stage 

while permitting the ASRG’s to stay with the payload.

Docking Adapter

2 1

Discipline, p10

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Mission Timeline
NASA/TM—2012-217014 477 

Discipline, p11

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OVERVIEW

Mission Operations, p1

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I n t e g r a t e d D e s i g n C e n t e r / M i s s i o n D e s i g n L a b o r a t o r y 

N A S A G O D D A R D S P A C E F L I G H T C E N T E R

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Mission Operations

Deborah Knapp

November 29 – December 3, 2010

Beamed Energy Propulsion 

System (BEPS)
NASA/TM—2012-217014 478 

Mission Operations, p2

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Topics

• Study Activities

• Customer Requirements

• Concept of Operations

• Overall Lifecycle

• Operations Timeline

• Design Requirements & Assumptions

• Ground System Functional Architecture

• Technology Required

• Risk/Issues/Concerns

• Additional Trades

Mission Operations, p3

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Study Activities

• Provide the BEPS customer with an Operations 

Concept

• Provide the BEPS customer with basic Ground 

System Architecture diagram

• Mission Operations costs are not required
NASA/TM—2012-217014 479 

Mission Operations, p4

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Customer Requirements

• Spacecraft data rates:

– NENS

• 20 kbps command

• 30 kbps, 4 Mbps telemetry downlink

– TDRSS

• 2 kbps command

• 10 kbps telemetry downlink

Mission Operations, p5

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Concept of Operations

• Launch BEPS & Galileo on separate launch 

vehicles into LEO

• BEPS & Galileo rendezvous in LEO (800 km circ)

• Transfer from LEO to 3000 km circular orbit

• Transfer from 3000 km circular orbit to 3000 x 

35000 km orbit (Jupiter Departure Orbit (JDO))

• ~100 min burn to transfer from JDO to Jupiter 

Intercept Trajectory

• BEPS & Galileo separate

– BEPS performs collision avoidance maneuver

• BEPS Mission Ends and BEPS flies away

• Galileo mission continues to Jupiter
NASA/TM—2012-217014 480 

Mission Operations, p6

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Overall Lifecycle

Launch

Launch BEPS 

& Galileo on 

separate 

launch 

vehicles to 

LEO (800 km)

~100 min 

Burn

From JDO to 

Jupiter Insertion 

Trajectory

BEPS & Galileo 

Rendezvous & 

Dock

BEPS & 

Galileo 

Separate

BEPS performs 

collision 

avoidance 

maneuver

Raise orbit 

LEO to 3000 km circ

3000 km circ to JDO

Start of 

Galileo 

Mission

Jupiter or Bust

End of 

BEPS 

Mission

Mission Operations, p7

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BEPS

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Operations Timeline

Launch

Rendezvous 

in LEO

~100 min 

Burn

End of 

Mission

Nominally 46 days

LEO to 

3000km circ to 

3000km x GEO

Separation

Collision avoidance 

maneuver

30 days 15 days <1 day
NASA/TM—2012-217014 481 

Mission Operations, p8

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Design Requirements & Assumptions

• Mission Operation Control (MOC):

– Provides “standard” set of functional components to support Mission Operations (e.g., S/C 

commanding, mission planning/scheduling, and RT TLM monitoring, HK trending and 

analysis).

– Provides level zero processing (LZP)

• Level 0 Products: Time-ordered, quality annotated data sets produced for each 

contact

• Space-Ground contacts

– Near Earth Network (NEN) stations for telemetry & command

• 10 meters or greater

– Tracking and Data Relay Satellite System (TDRSS) for telemetry & command during 

rendezvous and docking

– Laser at Mauna Kea, HI for laser burns

• Space-Space contacts

– BEPS will communicate with the Galileo during rendezvous and docking with an omni antenna at Galileo data rates/frequencies

• Spacecraft data rates:

– NENS

• 20 kbps command

• 30 kbps, 4 Mbps telemetry downlink

– TDRSS

• 2 kbps command

• 10 kbps telemetry downlink

• Data Latency Requirements:

– Real time for laser burn

• Data Recovery:

– Assume 98% recovery requirement (end-to-end)

• Orbit Determination via FDF

Mission Operations, p9

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Ground System Functional 

Architecture

NEN

10+ meter

White Sands 

Complex

TLM

CMD

Mission Ops Center

Mission planning & scheduling

Orbit determination/control

Network & contact scheduling

Commanding

S/C monitor/control

RT health/safety processing

Trending/Analysis

Instrument data handling

Level 0 product processing

Level 0 Data Archive Mauna Kea, HI

Laser

Needed for 

Rendezvous 

& Dock

SSA 

TLM: 10kbps

CMD: 2 kbps

S-band 

TLM: 30kbps, 

4 Mbps

CMD: 20 kbps

Legend:

CMD = Commanding

TLM = Telemetry data
NASA/TM—2012-217014 482 

Mission Operations, p10

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Technology Required

• Use COTS/GOTS-based MOC as basis for MOC implementation

– Software packages are available to satisfy MOC required functionality:

• ITOS, ASIST, EPOCH 2000, ALTAIR are commercially available today and provide required 

functionality for Spacecraft Command/Control and Level Zero (LZ) Processing.

– Low data volume so no special/new technology required

– Most required technologies have been at least demonstrated; many in currently 

operational systems.

• Technology Complexity: Low

• Technology Readiness Level: 9

Mission Operations, p11

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Risk/Issues/Concerns

• Risks

– Burn shutdown early

• Issues

– none

• Concerns

– none
NASA/TM—2012-217014 483 

Mission Operations, p12

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Additional Trades

• Co-locate BEPS MOC with either Galileo MOC or at Laser site
NASA/TM—2012-217014 484 
NASA/TM—2012-217014 485 

Orbital Debris, p4

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BEPS

Number of Impacts to the PM Optical 

Surface 

• The number of impacts to the PM optical surface is obtained from:

h = F x A x T

where: h = number of impacts

F = particle flux (meteoroid or orbital debris), in impacts\m2\year. 

A = cross-sectional area of the surface of interest, in m2

T = time, in years 

• F is obtained from environment models.

– Orbital debris models:

• ORDEM 2000 (Orbital Debris Environment Model)

– Still the most widely used OD model, although it predates the FY-1C ASAT and Iridium-Cosmos 

collisions; tends to underestimate the current OD population. 

– Major limitation: Model is limited to the LEO region (<= 2000 km)

• ORDEM 2008 Beta

– Not released for official use yet. Expected to be released as ORDEM 2010.

– Covers up to GEO (35,786 km).

– Micrometeoroid models:

• Grün model

– Still used for near-Earth assessment of the micrometeoroid population. Simple and fast for quick 

computations.

• MEM 1c (Meteoroid Engineering Model, version 1c)

– Latest NASA micrometeoroid environment model.

– Requires detailed ephemeris inputs.

– Longer running time.
NASA/TM—2012-217014 486 

Orbital Debris, p7

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Micrometeoroid Environment

• For this initial assessment, the Grün model of the Micrometeoroid 

Environment at 1 AU will be used. 

– The model assumes an omnidirectional particle flux in Earth’s vicinity; the flux is a 

function of the micrometeoroid mass.

– Shielding and focusing effects due to Earth are taken into consideration.

– Limited to particles < 10 g.

– The model might underestimate the number of impacts, but even in that case the 

effect of micrometeoroids is expected to be low compared to that of orbital debris.

Reference: Natural Orbital Environment Guidelines for Use in Aerospace Vehicle Development, NASA TM-4527, pages 7-1 to 7-4
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Thermal, p5

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Component Layout
NASA/TM—2012-217014 500 

Thermal, p6

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BEPS

Waste Heat of BEPS 

Laser power distribution

Laser power nominally arriving at craft 50

Primary mirror 97% efficient, .9999 reflective 

Scattered radiation MW 1.5

Absorbed radiation kW 5

Laser energy to secondary mirror 48.5

Secondary mirror 98% efficient, .9999 reflective

Scattered radiation MW 0.97

Absorbed radiation kW 4.85

Laser energy to window 47.53

Saphire window 99% efficient .9999 transmissive

scattered or reflected, MW 0.4753

Absorbed radiation kW 4.753

Laser energy into chamber 47.0547

Chamber walls 97% efficient, .9999 reflective

scattered or reflected, MW 1.411641

absorbed radiation, kW 4.70547

Chamber efficiency 80% (20% radiation losses out window and nozzle)

Laser radiation losses 9.12767071

Power available to form plasma, MW 36.5106828

Plasma absorbtion efficiency 75% (25% losses into chemical reaction, ionization, etc

plasma inefficiency losses 9.12767071

Power available for propulsive thrust, M 27.3830121

Overall laser propulsion efficiency 54.7660242

Total 20 kWatts

MW
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Thermal, p13

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BEPS

Chemical Propulsion 

• heater control system

• MLI fuel tanks internal 

• MLI module (external) 

• MLI thrusters 

• lines controlled and MLI

• Estimated heater power

– 20 watts per tank (2)

– 5 Watts per thrusters (6)

– 20 Watts lines

Total 90 Watts
NASA/TM—2012-217014 504 

Thermal, p14

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BEPS

Mass Budget for Thermal

Payload/Propellant Stage

Active Thermal Control 3.34

Passive Thermal Control 36.76

Payload Laser Shield 26.64

66.74

Thermal Control (Non-Propellent)

Active thermal Control 2.65

Passive thermal control 19.61

Mirror System 93.66

Additional pocketed back side of 

Primary reflector

50.00*

Total 232.66

* Modification to DRM-2 baseline design
NASA/TM—2012-217014 505 

3.2.6 Deep Space Consultant Report—Design Reference Mission 3 

Several possible applications of beamed-energy propulsion (BEP) were explored for deep space 

missions. The primary goal of this portion of the study was to find an application of BEP that would 

either enable a mission not otherwise achievable or measurably enhance myriad possible missions. 

Early work on the direct propulsion of gossamer bodies such as the Forward-Landis Star Wisp 

microwave-propelled interstellar craft and the success of the Japanese Interplanetary Kite-craft 

Accelerated by Radiation Of the Sun (IKAROS) solar sail mission to Venus led to the consideration of 

laser-powered sails to propel payloads to the outer planets. 

The concept finally adopted for Design Reference Mission 3 (DRM 3) was to show how BEP as used 

in DRM 2 for an orbit-raising tug could improve the performance of a planetary or deep space mission. 

The study was performed in collaboration with the Mission Design Laboratory at the NASA Goddard 

Space Flight Center. The approach was to use the results of NASA Glenn Research Center’s Compass 

team design for the orbit-raising tug with the minimal changes necessary to adapt the design to provide 

acceleration for the one-way trip to Jupiter; the object being to investigate performance enhancements 

possible with BEP. The Galileo Mission to Jupiter (1989 to 2003) was chosen for the comparison. 

Galileo was launched from STS–34 Atlantis in October 1989 on a trajectory to Jupiter that included 

three planetary gravity well boosts (Venus-Earth-Earth) and took just over 6 years to reach its target. It 

was thought that BEP might provide for a direct Earth-to-Jupiter transit instead of the time-consuming 

triple-gravity-well boost trajectory necessitated by the use of a traditional chemical rocket upper stage as 

employed by the Galileo Mission. 

The Galileo payload requirements (size, mass, electric power, and thermal management) were taken 

to be those of the payload for the comparison mission. The mission vehicle will be a compound of two 

modules, a modified Galileo payload and the propulsion module derived from the orbit-raising tug of 

DRM 2. The simulated Galileo payload and the propulsion module will be launched on separate Falcon 9 

class launch vehicles and joined in an on-orbit docking maneuver as in DRM 2. The combined 

payload/booster will be put into a 3000-km by geosynchronous Earth orbit (GEO) and finally boosted into 

a direct Jupiter Transfer Orbit with an approximately 100-min BEP burn as shown in Figure 3.152. 

Figure 3.152.—Graphic representation of BEPS DRM 3 transfer trajectory from Earth 

orbit to Jupiter Direct Transfer Orbit showing path of BEP acceleration. 

With BEP, it will be possible to gain enough velocity to arrive at Jupiter via a direct transfer orbit in 

about 2.5 years instead of the more than 6 years required for the original Galileo mission. An additional 
NASA/TM—2012-217014 506 

benefit of the BEP-enhanced mission is that the launch opportunities for flights using similar direct 

transfer orbits will occur approximately every 400 days. 

3.2.6.1 Description of Vehicle 

The vehicle design for DRM 3 was adapted directly from the orbit-raising tug designed by Glenn’s 

COMPASS design team for DRM 2. Like DRM 2, the payload and propulsion vehicles are launched 

separately and joined on orbit. However, whereas the orbit-raising tug was designed to be reusable, to 

deliver self-contained payloads to geostationary orbit, the propulsion vehicle in DRM 3 is designed to be 

an expendable booster. For the orbit-raising tug to be reusable under the DRM 2 concept of operations 

(CONOPS), the propellant would have to be replaced for each mission and be manifested as part of the 

satellite payload launch. In the DRM–3 configuration, the propulsion module includes the water 

propellant for the laser plasma engine for near-Earth orbit-raising and Jupiter Transfer Trajectory 

Insertion burns. The propulsion module is by far the heavier payload for the two Falcon-9-class launches 

used in DRM 3. 

The Galileo payload analog for the study was updated from the original Galileo mission by replacing 

the original Radioisotope Thermoelectric Generators with Advanced Stirling Radioisotope Generators 

(ASRGs). Otherwise, all power, mass, and size specifications from the original Galileo vehicle were used 

to define the payload. Three of the four ASRGs needed to power the Galileo-like science payload will be 

launched with the propulsion module. Including these ASRGs with the propulsion module will eliminate 

the need for the solar panel used for power on the DRM 2 orbit-raising tug and will have the advantage of 

additional mass fraction being available on the payload launch vehicle, should it be needed. 

Figure 3.153.—Drawing of stowed DRM 3 

propulsion module in Falcon 9 fairing. 
NASA/TM—2012-217014 507 

The propulsion module will easily fit within the 5.2-m Falcon 9 fairing (See Figure 3.153). The 

stowed module will have an overall length of about 8 m. 

The major components of the propulsion module folow: (1) the BEP system, consisting of the beam 

collecting and focusing optics, the thrust chamber and nozzle, and propellant storage and delivery; (2) the 

electrical energy system consisting of the ASRGs, a battery, and power management and distribution; (3) 

the attitude and control system consisting of reaction wheels for fine pointing, an autonomous docking 

system, and monopropellant thrusters with fuel for maneuvering; (4) the thermal management system; and 

(5) the communications system. 

The optical components for the BEP system are essentially the same as for the orbit-raising tug for 

DRM 2 and will consist of a highly reflective, 97%-efficient 4- by 6-m elliptical primary collecting mirror 

that is mounted on an arm and gimbal that will enable it to track the laser beam and focus it onto a 98%-

efficient, 0.5-m-diameter secondary turning mirror that will direct the beam through a 99%-efficient, 10-

cm sapphire window into the thrust chamber. The reflectivity of the mirrors and the transmissivity of the 

sapphire window to the micron laser beam were taken to be 0.9999. 

The laser engine will be a 30-cm spherical chamber operating at 50 psi with regenerative cooling of 

the chamber wall by the propellant. The chamber wall will be 97% efficient with 0.9999 reflectivity. The 

design efficiency of the chamber will be 80%, with 20% radiation losses out of the window and nozzle. 

The plasma absorbtion efficiency will be 75% (with losses in chemical reactions, ionization, etc.), 

resulting in as overall laser propulsion efficiency of greater than 50%. The nozzle will be regeneratively 

cooled with an area ratio of 50. 

Propellant tanks will be provided on the propulsion module for the laser engine water propellant, the 

hydrazine monopropellant for the maneuvering thrusters and for pressurization gas. The payload module 

will carry its own hydrazine and pressurization tanks for maneuvering and for the Jupiter orbit insertion 

engine. 

Electric power for the propulsion module will be provided by three ASRG units and a rechargeable 

lithium-ion (Li-Ion) battery that will be launched with the propulsion module. A fourth ASRG will be 

launched with the Galileo-like payload and added when the payload and propulsion modules are docked. 

The ASRGs (with beginning-of-life power of 143 W each) will be used for primary power for all 

functions except propulsion, when the 100-Ah Li-Ion battery will provide supplemental power for the 

propellant pumps during laser engine operation. For a typical 10-min orbit-raising propulsion firing, the 

battery depth of discharge will be 8.3%; and for the Jupiter Transfer Trajectory Insertion firing of 100 

min, the depth of discharge will be 83%. The power system will be a 28-V-battery-dominated bus, with 

separate 28-V buses on the payload and propulsion modules. The ASRGs will be retained by the payload 

during separation of the payload from the expended booster and be used for power during the remainder 

of the mission. 

The attitude control system will include autonomous rendezvous and docking capability. Reaction 

wheels will be used for attitude control and fine pointing of both the propulsion and payload modules. 

Hydrazine monopropellant systems will be provided on the propulsion and payload modules to provide 

maneuvering thrust and thrust to unload the reaction wheels. The hydrazine system on the payload also 

provides the propellant for decelerating the payload for Jupiter orbit insertion. 

The propulsion module thermal management system has both passive and active elements. Cooling of 

the water propellant pump and the walls and window of the laser engine chamber and the nozzle will be 

accomplished by pumping the water propellant through tubes attached to the outer wall of the chamber 

and nozzle prior to injecting the water into the laser engine chamber. Thermal energy from the cool side 

of the ASRG will be used to maintain fluid temperatures in the propellant tanks. Electronic and battery 

temperature will be maintained through passive cooling. The approximately 5 kW of energy absorbed by 

each of the laser mirrors will be radiated from the backside of the mirror. 

The communications system will be S-Band microwave, with direct communication to Earth through 

the Near Earth Network, with the Tracking and Data Relay Satellite System (TDRSS) available for backup. 

The major physical changes between the DRM 2 orbit-raising tug and the DRM 3 booster are shown 

in Figure 3.154 (the orbit-raising tug) and Figure 3.155 (the booster). The most obvious difference is that 
NASA/TM—2012-217014 508 

the solar panel has been removed and replaced with ASRGs to provide power, initially for the propulsion 

module, then for the combined payload and propulsion module, and finally, after separation, for the 

science payload on its path to Jupiter and its mission in the plane’s environs. In addition, the 100-min 

Jupiter Transfer Trajectory Insertion burn increases the stored-energy requirement for the propulsion 

module and a larger capacity (100-Ah) battery is substituted for the DRM 2 design. A second difference is 

that the propellant for orbit raising and the Jupiter transit acceleration will be launched as part of the 

propulsion module rather than with the payload, as is the case with the reusable orbit-raising tug. The 

third major difference is the change of the separation point when the propulsion module and payload are 

separated at the conclusion of propulsion activity (achieving geostationary orbit for DRM 2 and Jupiter 

Transfer Orbit for DRM 3). For the DRM 3 Galileo analog mission, the separation ring will be added 

such that the propellant tanks are jettisoned with the propulsion unit and the ASRGs continue on to 

Jupiter as part of the Galileo spacecraft bus for the science mission. 

Figure 3.154.—Drawing of the DRM 2 orbit raising-tug showing the major 

vehicle systems. The primary reflector is not shown for clarity. 

Figure 3.155.—Drawing of the DRM 3 expendable propulsion booster showing the 

major vehicle systems. 
NASA/TM—2012-217014 509 

A minor difference between the DRM 2 orbit-raising tug and the DRM 3 expendable propulsion module 

is that the chemical monopropellant thruster pods will be replaced with six single-nozzle, articulated 

thrusters mounted on Canfield joints. 

One feature that was added during the DRM 3 study and that should be applied to the DRM 2 orbitraising tug is an enhanced method for the guidance and control of the ground-to-spacecraft laser beam 

link. Retroreflectors will be strategically positioned on the face of the primary collecting mirror, and 

photovoltaic sensors will be placed around the sapphire window into the laser engine chamber (see Figure 

3.156). More complete descriptions of the laser system design and operation are given in the next two 

sections. 

Because the payload and propulsion module will be launched separately and joined on orbit, there are 

necessary system redundancies, among which are attitude control, communications, and power 

management for the electric systems. A topic for future investigation would be whether or not installing a 

water propellant tank on the mission payload vehicle and moving a portion of the water propellant to the 

mission payload launch would result in a sufficient mass fraction margin gain for the propulsion module 

launch or enough increased mission capability to offset any added spacecraft design complexity. In this 

scheme, the propellant on the propulsion module would be used for orbit raising from LEO to the Jupiter 

Departure Orbit, and the propellant on the Galileo mission spacecraft used for the 100-min Jupiter 

Transfer Trajectory Insertion burn. 

3.2.6.2 Description of Beaming Station 

The laser-beaming stations are the same as defined and described for the orbit-raising tug in DRM 2. 

An 80-MW laser will be required to provide approximately 50 MW at the spacecraft. Two laser director 

systems will be used for power transmission to the spacecraft, depending on the spacecraft altitude. For 

orbit raising from the initial 800-km rendezvous circular orbit achieved through the Falcon 9 launch to the 

3000-km circular intermediate orbit and the 3000-km by GEO Jupiter Departure Orbit, perigee burns (and 

3000-km circularization apogee burns) will be accomplished with a small-aperture laser beam director 

that is 0.75 m in diameter. This laser aperture located on the Earth will provide a laser spot size of slightly 

less than 3 m in diameter at 800 km and about 3.2 m in diameter at 3000 km. 

The 100-min Jupiter Transfer Trajectory Insertion burn from the Jupiter Departure Orbit to the Jupiter 

Direct Transfer Orbit (see Figure 3.156) will require the 30-m-diameter laser beam director to keep the 

beam spot within the area of the 4- by 6-m elliptical spacecraft primary collecting mirror. 

Operation of the laser propulsion system will require a sophisticated interactive laser guidance and 

control system to maintain the ground-to-spacecraft link. The system will have to be able to acquire and 

maintain the high-power laser beam for spacecraft propulsion over a large flight path angle and reflector 

range of motion, with widely varying tracking rate requirements and varying laser time of flight lag, with 

a high degree of pointing accuracy. In addition it will have to account for anomaly detection and 

response. 
NASA/TM—2012-217014 510 

Figure 3.156.—Schematic of laser tracking and guidance system. 

The components of the system as shown in Figure 3.156 include a laser tracker on the ground station, 

retroreflectors located on the spacecraft primary collecting mirror, and photovoltaic sensors around the 

sapphire window of the laser plasma engine. The ground-based portion of the system includes the 30-maperture transmitter (which will be able to be steered in azimuth and elevation and will be located on 

Mauna Kea in Hawaii to be above as much atmosphere and cloud cover as possible), a low-power laser 

guidance beam, the main propulsion beam, and a laser-sensitive beam-tracking system. Components on 

the satellite, which will collect and focus the beam into the laser plasma engine, include a primary 

collector and secondary reflector to direct the laser beam into the laser plasma chamber. The primary 

collecting reflector will have a 4-m aperture with a retroreflector pattern on the surface to provide a return 

signal to the laser-tracking sensor mounted on the ground transmitter. This return signal will affirm 

acquisition of the laser guide beam by the spacecraft. A secondary sensor set will consist of germanium 

photovoltaic cells arrayed around the rim of the plasma chamber window to provide fine control for 

reflector steering. The cells will be sensitive enough to detect the low-power guidance beam and will be 

sufficiently robust to withstand the full-power propulsion beam. 

Laser link acquisition will be accomplished with the power beam off. The Global Positioning System 

will be used to provide position knowledge within 100 m. The guidance beam will be defocused to cover 

a large region. When the retroreflectors are detected, the guidance beam will be progressively focused 

until the beam is captured on the primary collecting reflector. Spacecraft roll attitude and reflector 

steering will be adjusted until the guidance beam is centered on the plasma engine chamber window, as 

determined by the photovoltaic sensor ring surrounding the window. When the guidance beam is centered 

on the engine window, the power beam may be switched on. The ground transmitter tracking and steering 

system will use the retroreflector return signal to keep the power beam centered on the primary collecting 

reflector and the spacecraft will use the photovoltaic sensor ring information for control of the spacecraft 

attitude and for reflector steering. 

Any anomaly in the operation of the high-power laser will have to be detected and corrected 

instantaneously to prevent potential thermal damage to the spacecraft and sensitive instruments or optical 

damage to especially sensitive components, such as star trackers. Because laser steering and especially 

laser cutoff could occur more rapidly than reflector steering and spacecraft attitude control, fail-safe will 

be with the ground laser. If the power beam loses lock on the primary reflector as detected by loss of lock 

by the ground laser tracker, the ground station will command shutdown of the power beam and then 
NASA/TM—2012-217014 511 

initiate reacquisition of the satellite with the defocused guidance beam. If the power beam loses lock on 

the engine chamber window as detected by the photovoltaic cells around the window, an alert will be sent 

by microwave communication to the ground, after which the ground station will command shutdown of 

the power beam and then initiate reacquisition of the satellite with the defocused guidance beam. 

Figure 3.157.—Drawing of critical laser elevation, flight path 

angle and reflector angle at beginning and end of Jupiter 

Transfer Trajectory Insertion burn. 

The elevation of the laser beam above the horizon is shown in Figure 3.157 along with the flight 

angle path and relative reflector angle for the start and finish of the 100-min Jupiter Transfer Trajectory 

Insertion burn, showing the reflector mirror angle travel to be well within the design swing of 60 from 

normal. 

For the Jupiter Transfer Trajectory Insertion burn, the laser transmitter must be capable of smoothly 

tracking the spacecraft at a rate that varies from essentially 0/min to 1.7/min. In addition, at the end of 

the Jupiter Transition Trajectory Insertion burn, the range to the spacecraft will be 44,200 km, the laser 

tracker will have to resolve an angle of 4.6 milliarcsec (the angle subtended by 1 m) to see the 

retroreflector pattern, and the transmitter will have to steer to an equal accuracy to keep the beam on the 

reflector. To compensate for the round-trip time of flight of light, the beam will have to “lead” the 

spacecraft by varying amounts along the acceleration trajectory. At the start of the Jupiter Transfer 

Trajectory Insertion burn, with a range of 11,900 km, the round-trip time of flight for light will be 0.08 s. 

The laser tracking rate will be 0.025/s, and the beam will have to “lead” the spacecraft’s observed 

position by 7 arcsec. At the end of the Jupiter Transfer Trajectory Insertion burn, the range will be 

44,200 km and the round-trip time of flight for light will be 0.3 s. The laser tracking rate will be 0.005/s, 

and the beam will have to “lead” the spacecraft’s observed position by 5 arcsec. 

3.2.6.3 CONOPS—Frequency of Launch and Repeat Time 

Differences between the two missions that lead to design changes in the propulsion module design 

and spacecraft systems allocations (e.g., photovoltaic panel replaced by ASRGs, greater battery capacity, 
NASA/TM—2012-217014 512 

and water propellant launched as part of the propulsion module) have been dealt with in a previous 

section. The notional operational timeline (from launch) is shown as Figure 3.158. 

Figure 3.158.—Sequence of operations for DRM 3: BEP-enhanced Galileo mission. 

The initial launch and orbit-raising sequence closely mirror that of DRM 2 and will employ two 

Falcon 9 launch vehicles (one for the propulsion module and one for the Galileo payload). The propulsion 

module and Galileo mission spacecraft will be launched separately from Kwajelein Atoll into an 800-km, 

9-inclined circular low Earth orbit (LEO). 

The propulsion module and Galileo mission spacecraft will be joined on orbit using Next Generation 

Advanced Video Guidance Sensors and automated rendezvous and docking hardware, firmware, and 

software. The time allocated for the launch, docking, and spacecraft diagnostics is 1 month. 

Orbit-raising protocols from DRM 2 will be followed to first transfer the combined spacecraft from 

an 800-km, 9-inclined circular orbit to a 3000-km, 9-inclined circular orbit and then to a 3000- by 

35,000-km, 9-inclined Jupiter Departure Orbit. The time allocated for the orbit change maneuvers is 15 

days. 

The final phase of near-Earth operation will be the 100-min laser-beam-powered Jupiter Transfer 

Trajectory Insertion acceleration from Jupiter Departure Orbit to Jupiter Transfer Orbit, followed by 

separation of the propulsion module from the Galileo mission craft with a collision avoidance maneuver. 

Here, the BEP phase of the mission will end with the propulsion module flying away and the Galileo 

craft continuing to Jupiter, where it will use onboard hydrazine monopropellant to perform a Jupiter Orbit 

Insertion burn of 70 min and enter the highly elliptical Galileo Jupiter Orbit of 200,000- by 

12,000,000-km with a 98-day period. 

Because BEP provides the possibility of a direct Jupiter transfer trajectory, flight opportunities will 

not be limited by a requirement that inner solar system planets be positioned for gravity-well-flyby 

acceleration assists. Thus, flight opportunities will recur almost yearly with the 399-day Earth-Jupiter 

synodic period; however, flight energy requirements will vary with each opportunity. 
NASA/TM—2012-217014 513 

3.2.6.4 Costs—Comments 

The design of the propulsion module and the CONOPS for DRM 3 were adapted with minimal 

changes from DRM 2, and cost estimates for the BEP-enhanced Galileo mission can be derived from 

DRM 2 estimates. 

BEP can open new opportunities for deep space exploration and science missions. The cost and time 

to develop the 80-MW ground station will be the most significant obstacles to the realization of the 

potential of beamed-power propulsion to this deep space application. 

3.2.6.5 Feasibility Assessment 

The following is a list of issues which need to be addressed 

1. What are the attenuation effects of high flux levels (MW/m2

) at the frequencies of interest on the 

atmosphere? 

2. What would be the expected availability and transmittance of the BEP ground terminal at a given 

frequency, given aerosol and precipitation conditions at a proposed ground site? 

3. How large of a keep-away zone must be maintained around and beyond an ascending craft to 

accommodate reflection, refraction and scattering components of the beam exceeding substantial 

irradiances, and what level is to be considered substantial? Considerations should be given to 

maximum permissible exposure for humans, and damage thresholds for ground, airborne and 

space-based assets within proximity of the BEP activity. 

4. For a ground based terminal operations, how much grid power over what duration is required? 

To meet these requirements, what would the generation, storage and conversion system look like? 

5. How would a micro/mm-wave BEP system work within the spectrum constructs of the 

NTIA/FCC? 

6. Does it make sense to collocate beamed energy propulsion and power technologies on the craft, 

for each phase of a proposed mission? 

7. Would there be a benefit to developing a hybrid aperture to handle multiple BEP frequencies 

depending upon range to target, atmospheric conditions, etc.? 

8. Can retro-directive control be utilized to aid in the beam pointing, acquisition and tracking of the 

craft? 

9. What is the most practical way of performing high power beam combination using COTS 

technology for the frequencies of interest? What are the potential near term accomplishments that 

could influence this capability? 

10. What are the cost/benefit tradeoffs in considering multiple BEP transmission locations, both at 

the ground range and at an intermediate relay point in space? 

11. Could deployable array technologies by implemented to decrease the weight of a BEP aperture, 

and increase the allowable size that can be launched? 

12. What are some near term demonstrations that may be accomplished to establish a maturity level 

for critical technologies, and support the overall feasibility of the BEP concept? 

3.2.6.6 Roadmaps and Technology Readiness Levels 

See DRM 2 LEO–GEO, Section 3.2.3.5.1. 

3.2.6.7 Ground and Flight Demonstrations 

See DRM 2 LEO–GEO, Section 3.2.3.5.2. 

3.2.6.8 Issues and Resolutions 

The baseline design for thermal management for the high-power laser system mirrors and sapphire 

lens is passive, with the back surface of each mirror providing its own radiative cooling and the sapphire 

lens being cooled by contact with the plasma chamber. For the short-duration beamed-energy thrusts 
NASA/TM—2012-217014 514 

during orbit raising, the design should be adequate. However, for the 100-min Transfer Transition 

Insertion burn, the high operating temperature (918 °C) calculated for the secondary (0.5-m diameter) 

mirror could present a material problem. Pocketing the radiator surface to increase area would reduce the 

operating temperature. Alternatively, the water loop could be extended to the mirror or a heat pipe could 

be used as a bridge from the secondary mirror to the water loop. If additional cooling was required for the 

center of the sapphire lens, the lens could be segmented with thin cooling channels. 

The development of high-power, efficient, continuous-wave (CW) lasers capable of satisfying the 

requirements for DRM 2 or DRM 3 is beyond the scope of NASA funding and will undoubtedly require 

many years. A nearer term alternative to the 80-MW ground-based laser-power-beaming station for 

DRM 2 and DRM 3 also was examined. The two potential solutions are an approximately 1.5-MW laserpower-beaming station on the ground capable of delivering 1 MW to a redirecting mirror in geostationary 

orbit or a 1-MW laser-power-beaming station in space in geostationary orbit. 

The laser plasma thruster was assumed to be capable of linear operation over a range from 500 kW to 

greater than 50 MW. Thus, with a 1-MW beam originating at geostationary orbit (either from an orbiting 

power-beaming station or from a reflector in geostationary orbit fed from the ground), approximately 

100 N of thrust could be generated from the engine. This would be adequate for a spiral orbit raising from 

LEO to geostationary orbit for DRM 2, although it will require more propellant than would be needed 

with the higher power 80-MW system. 

For the critical Jupiter Transition Trajectory Insertion burn, calculations show that it would require 

more than 1.5 million km to achieve the velocity change of 6.7 km/s required for a direct Jupiter Transfer 

Orbit. At this range, if we assume a 30-m beaming aperture or reflector and a coherent beam, the laser 

beam spot will be greater than 100 m. 
NASA/TM—2012-217014 515 

4.0 Concluding Remarks 

This study took an in-depth look at the feasibility of beamed-energy propulsion (BEP) and concluded 

that although the technology is definitively game-changing and technically achievable given the appropriate development resources, the infrastructure costs would be extremely high, making it comparably 

unattractive at this time (Figure 4.1). 

Figure 4.1.—BEP evaluation summary. LEO to GEO, low Earth orbit to geosynchronous 

Earth orbit; TRL, Technology Readiness Level; OCT, Office of the Chief Technologist; 

BAA, Broad Agency Announcement.

Of course, it was hoped that beamed-energy technology would provide the “silver bullet” solution to 

low-cost access to space and enable less-expensive deep space missions, or that it would at least find a 

niche where the benefits would be great. The authors of this report would like to make it clear that, as 

with all studies, if the ground rules change, it follows that the conclusions will change. The approach used 

to determine costs was the method used by the NASA Glenn Research Center. This report recognized that 

vehicle costs for a commercial venture would be lower than for a government endeavor. A bottoms-up 

commercial cost approach could deliver a much lower cost not only for the facilities but also for the 

vehicles. A similar process to what is currently happening to the International Space Station (ISS) 

cargo/tourism launch vehicles could also be applied to the small payload area. 

At the end of the study, it became apparent that the costs to create any new launch or space propulsion facility would range in the billions of dollars. Likewise, the costs to replace the NASA Kennedy 

Space Center would surely be in the same range. So it should not have come as a surprise to find high 

infrastructure costs for such a capability. Historically, any change to the current way business is done has 

involved a high initial investment. The enticement for the investment has been the potential of a valuable 

market. The definition of this possible market was outside of the scope of this study. But the findings 

indicate that the technology could be effectively used for the launch of payloads in the range of 1 to 

10 kg, or “CubeSats” or “nanosats” as they are known. Investigations should be conducted on the market 

for small payloads. 

During this study, two applications of BEP technology were revealed. They should be pursued since 

they could prove to be very beneficial to both DARPA and NASA. The path forward for BEP could start 

with a small, low-cost ground demonstration that would show the key components of the technology, 

achieving the next readiness level. However, it is the conclusion of the authors that sufficient equipment 

is currently available to enable an in-space experiment. The jump to this level would have a major impact 

on the spacecraft design community and would open up other possibilities. 

NASA/TM—2012-217014 517 

Appendix A.—Acronyms and Definitions 

ABL Airborne Laser 

ACS attitude control system 

AEOS Advanced Electro-Optical System 

AFB Kirtland Air Force Base 

AFRL Air Force Research Laboratory 

AIMD additive increase/multiplicative decrease 

ALRU Amplifier Line Replaceable Unit 

AO adaptive optic 

ASRG Advanced Stirling Radioisotope Generators 

BEPS Beamed-Energy Propulsion Study 

C carbon 

C&DH command and data handling 

CCD charge-coupled device 

CMOS complementary metal-oxide-semiconductor 

CO2 carbon dioxide 

COMPASS COllaborative Modeling for Parametric Assessment of Space Systems 

CONOPS concepts of operation 

COTS commercial off the shelf 

CubeSat miniature satellite that is usually a 10-cm cube 

CVD chemical vapor deposition 

CW continuous wave 

DARPA Defense Advanced Research Projects Agency 

DC direct-current 

DDT&E design, development, testing and engineering 

DOD Department of Defense 

DPSSL Diode-Pumped Solid-State Laser 

DRM Design Reference Mission 

E–D expansion-deflection 

EEE electrical, electronic, and electromechanical 

EMI electromagnetic interference 

EPS electrical power system 

ESA European Space Agency 

ETO Earth-to-orbit 

FADIS Fast Directional Switch 

FOM figures of merit 

FPA flight path angle 

FPGA field-programmable gate array 

FSM fast steering mirror 

GBL ground-based laser 

GEO geosynchronous orbit 

GLOM gross liftoff mass 

GLOW gross liftoff weight 
NASA/TM—2012-217014 518 

GNC guidance, navigation, and control 

GPS Global Positioning System 

H&S health and safety? DRM-1A 

HiDVE High Delta-V 

H2O2 hydrogen peroxide 

He helium 

HX heat exchanger 

IFR ideal final result 

I/O input/output 

IMU inertial measurement unit 

IP Internet Protocol 

ISUS Integrated Solar Upper Stage 

JHPSSL Joint High Power Solid-State Laser 

KKV Kinetic Kill Vehicle 

LLNL Lawrence Livermore National Laboratory 

LED light-emitting diodes 

LEO low Earth orbit 

JPL Jet Propulsion Laboratory 

LLC Laser lightcraft 

LTD Lightcraft Technology Demonstrator 

LH2 liquid hydrogen 

LiFePo lithium-iron-polonium 

LLNL Lawrence Livermore National Laboratory 

LN2 liquid nitrogen 

LSD laser-supported detonation 

MDL Mission Design Lab 

MEL Master Equipment List 

MEMS microelectromechanical systems 

MLI multilayer insulation 

MOPA Master Oscillator Power Amplifier 

N2 nitrogen 

NaS sodium sulfide 

NiCd nickel cadmium 

NiMH nickel metal hydride 

O&M operation and maintenance 

OCT Office of the Chief Technologist 

O/S operating system 

ORU orbital replacement unit 

OTIS Optimal Trajectories by Impact Simulation 

OTIS Optimal Trajectories by Implicit Simulation 

P&S protocols and standards 

PD pulse detonation 

PDE pulsed detonation engine 

PEL power equipment list 
NASA/TM—2012-217014 519 

PI&D propellant injection and distribution 

PLVTS Pulsed Laser Vulnerability Test System 

PMAD power management and distribution 

PMD propellant management device 

P-POD Poly-Pico Satellite Orbital Deployer 

PRF pulse repetition frequency 

PSI Pressure Systems, Inc. 

PV photovoltaics 

R&D research and development 

RCS reaction control system 

ROM rough order of magnitude 

ROSA Roll-Out Solar Array 

S/C spacecraft 

SABLE Scaled Atmospheric Blooming Experiments 

SA solar array 

SOR Starfire Optical Range 

SDIO Strategic Defense Initiative Organization 

SHARP Stationary High Altitude Relay Platform 

SiC silicon carbide 

SLOC source lines of code 

SOA state of the art 

SSDP solid-state, diode-pumped 

SSME Space Shuttle Main Engine 

SSTO single stage to orbit 

SPO systems program office 

TBD to be determined 

TBR to be resolved 

TDRSS Tracking and Data Relay Satellite System 

TEA laser transversely excited atmospheric-pressure laser 

TMT 30-Meter Telescope 

TPS thermal protection system 

TRL Technology Readiness Level 

TVC thrust vector control 

UAV unmanned aerial vehicles 

UV ultraviolet 

VAB Vehicle Assembly-processing Building 

Δv velocity 

VHDL very-high-speed integrated circuits (VHSIC) hardware description language 

VHF very high frequency 

VMJ vertical multi-junction 

WSMR White Sands Missile Range 

NASA/TM—2012-217014 521 

Appendix B.—Consultant Resumes 

Jonathan T. Black 

Assistant Professor of Aeronautics and Astronautics 

Air Force Institute of Technology, AFIT/ENY 

2950 Hobson Way, Bldg. 640 

Wright-Patterson AFB, Ohio 45433 

jonathan.black@afit.edu 937–255–3636, Ext. 4578 Fax: 937–656–7053 

Education 

 Ph.D. in Mechanical Engineering, University of Kentucky, December 2006 

 Support: University of Kentucky Dissertation Year Fellowship, NASA Graduate Student Researchers 

Program, Department of Mechanical Engineering 

 M.S. in Mechanical and Aerospace Engineering, George Washington University, August 2003 

 Support: Joint Institute for Advancement of Flight Sciences Graduate Research Scholar Assistantship 

 B.S. in Industrial Engineering, with Honors, International Minor in French Studies, University of 

Illinois at Urbana-Champaign, May 2001 

Professional History 

Air Force Institute of Technology, Dayton, Ohio 

Jan. 2007 – present, Assistant Professor, Department of Aeronautics and Astronautics 

University of Kentucky, Lexington, Kentucky 

Jul. – Dec. 2006, University of Kentucky Dissertation Year Fellow, Dynamic Structures and Controls 

Laboratory, Department of Mechanical Engineering 

Jul. 2004 – Jun. 2006, NASA Graduate Student Researchers Program Fellow, Dynamic Structures 

and Controls Laboratory, Department of Mechanical Engineering 

Aug. 2003 – Jun. 2004, Graduate Research Assistant, Dynamic Structures and Controls Laboratory, 

Department of Mechanical Engineering 

Air Force Research Laboratory, Albuquerque, New Mexico 

Summer 2004 and 2005, Space Scholars Program, Space Vehicles Directorate, Kirtland Air Force 

Base 

NASA Langley Research Center, Hampton, Virginia 

Aug. 2001 – Jul. 2003, Graduate Research Scholar Assistant, Joint Institute for Advancement of 

Flight Sciences, Structural Dynamics Branch 

Licenses, Registrations, and/or Certifications 

 Member of the American Institute of Aeronautics and Astronautics (AIAA) 

 Member of the Society of Experimental Mechanics (SEM) 

 Member of the American Society of Mechanical Engineers (ASME) 

 Engineer in training, State of Kentucky, no. 12668 

Awards and Honors 

 AFOSR Young Investigator Research Program Grant, 2007, 1 of 29 awardees nationally of 215 

applicants, first ever AFIT recipient 

 AIAA Associate Fellow, 2010 

 Southwestern Ohio Council for Higher Education, Faculty Excellence in Teaching Award, 2010 

 National Society of Black Engineers Aerospace Systems Conference Best Paper Award 

(co-author), 2010 
NASA/TM—2012-217014 522 

 AIAA National Career Enhancement Award, First Place, Very Large Category, 2008 

 University of Kentucky Dissertation Year Fellowship, 2006/2007, 1 of 12 awardees universitywide 

 AIAA Foundation Willy Z. Sadeh Award in Space Sciences and Space Engineering, 2004/2005 

 Department of the Air Force, Special Act or Service Citation, 08/2004 and 08/2005 

 NASA Graduate Student Researchers Program Fellowship, 2004/2005, 2005/2006 

 Kentucky Graduate Scholarship, 2003-2006 

 AIAA Region I Student Paper Competition, Graduate Division, 2nd Place, 2003 

(1) Swenson, E.D., Wiesel, W., Black, J.T., Cobb, R.G., “Space Vehicle Research,” Air Force 

Research Laboratory Space Vehicles Directorate, AFRL/RV – AFIT/ENY Memorandum of 

Agreement, FY12-FY17 $1,250,000 (25%), awarded but not yet received. 

(2) Black, J.T., “Beamed Energy Propulsion Study: LEO – GEO Tug,” DARPA/NASA Joint Phase I 

Study ($2,000,000 total study), FY11, $70,000 (100%) 

(3) Black, J.T., “Characterizing MAV Wings In Flight,” AFOSR AFRL/AFIT MOA Small Grant 

Program, FY11 – FY13 $135,000 (100%), awarded and first year funding received. 

(4) Black, J.T., “Risk and Reward Metric for Space Systems,” Air Force Research Laboratory Space 

Vehicles Directorate, FY11-FY12 $130,000 (100%), awarded and first year funding received. 

(5) Black, J.T., “OhioSAT CubeSat Senior Design Project Workshop,” Ohio Space Grant Consortium, 

NASA Education Program, FY11 $55,000 (100%). 

(6) Black, J.T., Cobb, R.G., Swenson, E.D., “Novel Multifunctional Imaging Chromotomographic 

Spectrometer Flight Experiment (CTEx),” AS&T Outreach Program, FY10 – FY12 $200,000 

(34%). 

(7) Cobb, R.G., Black, J.T., Raquet, J., “Dynamic Two-Way Time Transfer Flight Experiment,” AS&T 

Outreach Program, FY11 – FY12 $110,000 (45%). 

(8) Black, J.T., Swenson, E.D., “Space Hardware Qualification Course,” AFIT Distance Learning 

Tuition Re-imbursement, FY11 $45,000 (50%) 

(9) Swenson, E.D., Lawrence, T., Black, J.T., “Dynamic Responsive Orbital Networking Experimental 

Satellite (DRONES) Engineering Model Research,” Air Force Space Command Space and Missiles 

Center, FY11 $15,000 (10%) 

(10) Swenson, E.D., Lawrence, T., Black, J.T., “FalconSat-7 Engineering Model Research,” US Air 

Force Academy, FY11 $10,000 (10%) 

(11) Cobb, R.G., Black, J.T., and Swenson, E.D., “Space Telescope Control System,” Air University 

Base Procured Investment Equipment Grant, FY11 $295,000 (33%) 

(12) Black, J.T., “3D Video Camera Measurement System,” AFIT End-of-Year Supplemental Funding, 

FY10 $177,000 (100%) 

(13) Black, J.T., “Hybrid Laser/Video 3D Non-Contact Motion Capture and Analysis,” AFOSR Young 

Investigator Research Program, FY09 – FY11 $360,000 (100%) 

(14) Cobb, R.G., Black, J.T., Collins, P., “Colony I CubeSat,” AS&T University Outreach, FY10 – 

FY11 $250,000 (50%) 

(15) Cobb, R.G., Black, J.T., Raquet, J., “Dynamic Two-Way Time Transfer Flight Experiment,” 

SAF/FMB, FY10 – FY11 $110,000 (45%) 

(16) Reeder, M.F., Cobb, R.G., Black, J., “Design and Testing of Flapping-Wing Micro Air Vehicles,” 

Air Force Research Laboratory Air Vehicles Directorate (AFRL/RB), FY09 – FY11 $155,000 

(33%) 

(17) Black, J.T., “Experimental Fin Tips for Reusable Launch Vehicles (ExFiT),” Air Force Research 

Laboratory Air Vehicles Directorate (AFRL/RB), FY09 – FY10 $31,600 (100%) 

(18) Black, J.T., “Experimental Fin Tips for Reusable Launch Vehicles (ExFiT),” Department of 

Defense Space Test Program (STP), Space Experiments Review Board (SERB), FY10 $10,000 

(100%) 
NASA/TM—2012-217014 523 

(19) Simmons, R., Black, J.T., “Experimental Fin Tips for Reusable Launch Vehicles (ExFiT),” AFIT 

Faculty Research Council, FY10 $25,000 (50%) 

(20) Hartsfield, C., Black, J.T., “CubeSat Education and Flight Program,” AFIT Faculty Research 

Council, FY10 $30,000 (50%) 

(21) Black, J.T., “Dynamic Surface Mapping of Lightweight Aerospace Structures in Motion,” Ohio 

Space Grant Consortium, FY09 $30,000 (100%) 

(22) Petersen, G., Black, J.T., “Semi-Autonomous Lead Vehicle for Convoy Operations,” Air Force 

Research Laboratory Center for Rapid Product Development, FY09 $100,000 (15%) 

(23) Cobb, R.G., Swenson, E.D., Black, J.T., “Attitude Path Planning for Responsive Spacecraft,” Air 

Force Research Laboratory Space Vehicles Directorate (AFRL/RV), FY09 $35,000 (25%) 

(24) Black, J.T., Cobb, R.G., “Space Imaging Power Laboratory,” Air University Base Procured 

Investment Equipment Grant, FY09 $256,000 (100%) 

(25) Black, J.T., “Multifunctional Chromotomography Imager,” AFIT Faculty Research Council, FY09 

$30,000 (100%) 

(26) Swenson, E.D., Black, J.T., Cobb, R.G., “Experimental Validation of Spacecraft Integration and 

Test Requirements for ORS Satellites,” Operationally Responsive Space Office, FY09 $18,000 

(25%) 

(27) Black, J.T., “Dynamic In-Flight Test and Measurement System,” AFIT End-of-Year Supplemental 

Funding, FY08 $189,000 (100%) 

(28) Black, J.T., Swenson, E.D., “Satellite Model Refinement Using High-Density Experimental Data,” 

AFIT Faculty Research Council, FY08 $25,000 (100%) 

(29) Black, J.T., “3D Non-Contact Surface Motion Capture and Analysis Video System,” AFIT New 

Faculty Seed Funding Award, FY07 $30,000 (100%) 

Student Grants 

(1) Simmons, J., Dayton Area Graduate Studies Institute (DAGSI) Fellowship, FY08 – FY11 $105,000 

(2) Black, J.T., Ohio Space Grant Consortium (OSGC) CRADA for Summer Intern, FY08 $6,000, 

FY09 $6,500 and FY10 $7,000 

Space Experiments 

(1) Black, J.T., Cobb, R.,G., Swenson, E.D., Hartsfield, C., and M. Hawks, “Space-Based 

Chromotomography Experiment (CTEx),” DoD Space Experiments Review Board (SERB), 

collaborative flight experiment with AFIT Engineering Physics department, 2013 scheduled launch. 

(2) Black, J.T. and R.G. Cobb, “AFITSAT I – Carbon Nanotube Field Emission Array Experiment,” 

Colony I Cubesat, collaborative flight experiment with AFIT Electrical Engineering department, 

2012 scheduled launch. 

(3) Cobb, R.G. and J.T. Black, “AFITSAT II – Precision Timing and Ranging Experiment,” Colonly II 

Cubesat, collaborative flight experiment with AFIT Electrical Engineering department, 2013 

scheduled launch. 

(4) Black, J.T., “Experimental Fin Tips for Reusable Launch Vehicles (ExFiT),” Department of 

Defense Space Test Program (STP), Space Experiments Review Board (SERB), May 2010 flight. 

(5) Cobb, R.G., Black, J.T., and E.D. Swenson, “Rigidizable Inflatable Get-Away-Special Experiment 

(RIGEX),” Department of Defense Space Test Program (STP), Space Experiments Review Board 

(SERB), STS-123, Space Shuttle Endeavour, Mar. 2008 flight. 

Personnel Supported on Grants or Contracts 

(1) Stoval, A., ENV PhD Student, FY11 – FY12 

(2) Blandino, J., ENY Visiting Faculty, FY10 

(3) Allen, C., ENY Research Associate, AD-21, FY08 – FY11 

(4) Jennings, A., ENY Research Associate, AD-21, FY11 

(5) Kaczmerek, J., ENY Research Associate, GS-5, FY10 – FY11 

(6) Simmons, J., DAGSI PhD Student, FY08 – FY11 
NASA/TM—2012-217014 524 

(7) Magree, D., DAGSI / SOCHE MS Student, FY09 – FY10 

(8) Jennings, A., SOCHE PhD Student, FY08 – FY10 

(9) Stein, E., SOCHE Research Associate, FY11 – FY12 

(10) Lomonno, C., SOCHE Research Associate, FY11 

(11) Briggs, G., SOCHE Research Associate, FY11 

(12) Alerding, J., SOCHE Summer Intern, FY11 

(13) Benson, M., SOCHE Summer Intern, FY11 

(14) Buhrman, N., SOCHE Summer Intern, FY11 

(15) Pace, T., ENY Summer Intern, FY10 

(16) Trinh, A., ENY Summer Intern, FY10 

(17) Rippl, M., ENY Summer Intern, FY09 – FY11 

(18) Jameson, M., ENY Summer Intern, FY09 

(19) Rodgers, D., ENY Summer Intern, FY09 

(20) Pollock, S., ENY Summer Intern, FY09 

(21) Simpkins, J., ENY Summer Intern, FY08 

(22) Sollars, R., ENY Summer Intern, FY08 

Publications 

Key: 

* Student advised by J. Black 

# Students advised by others 

A AFIT non-faculty researchers 

F Other AFIT faculty 

E External collaborators 

Archival Journal Articles 

(1) Magree, D.*, Black, J.T., Jennings, A.A, Briggs, G.*, and C. AllenA, “Pan-Tilt-Zoom Hybrid 

Camera System for Dynamic Tracking and Measurement,” AIAA Journal, accepted for publication, 

Dec. 2010. 

(2) Swenson, E.D.F, and J.T. Black, “Measuring and Modeling 3D Mode Shapes of FalconSAT-5 

Structural Engineering Model,” Journal of Experimental Mechanics, Vol. 51, No. 6, Jul. 2011, 

pp.933-945, DOI: 10.1007/s11340-010-9421-8. 

(3) Black, J.T., Cobb, R.C. F, Swenson, E.D. F, and B.J. Cooper*, “Rigidizable Inflatable Get-AwaySpecial Experiment Space Flight Data Analysis, Journal of Spacecraft and Rockets, Vol. 48, No. 3, 

May – June 2011, pp. 447-487, DOI: 10.2514/1.50939. 

(4) Jennings, A.A, Black, J., Allen, C.A, Simpkins, J. A, Sollars, R. A, “Vibrometer Steering System 

for Dynamic In-flight Tracking and Measurement,” Journal of Experimental Mechanics, Vol. 51, 

No. 1, Jan. 2011, pp. 71-84, DOI: 10.1007/s11340-010-9337-3. (for P&T consideration) 

(5) Cobb, R.G. F, Black, J.T., and E.D. SwensonF, “Design and Flight Qualification of the Rigidizable 

Inflatable Get-Away-Special Experiment,” Journal of Spacecraft and Rockets, Vol. 47, No. 4, Jul. – 

Aug. 2010, pp. 659-669, DOI: 10.2514/1.48636. 

(6) Black, J.T., Pitcher, N.A. #, Reeder, M.F. F, and R.C. MapleF, “Videogrammetry Dynamics 

Measurements of a Lightweight Flexible Wing in a Wind Tunnel,” Journal of Aircraft, Vol. 47, No. 

1, Jan. – Feb. 2010, pp. 172-180, DOI: 10.2514/1.44545. (for P&T consideration) 

(7) Black, J.T., Leifer, J.E, and S.W. SmithE, “Global Static Testing and Model Validation of Stiffened 

Thin-Film Polyimide Panels,” Journal of Spacecraft and Rockets, Vol. 45, No. 6, Nov. – Dec. 2008, 

pp. 1319-1324, DOI: 10.2514/1.37131. 

(8) Black, J.T., Smith, S.W. E, Leifer, J. E, and L.J. BradfordE, “Local Testing and Reduced Model 

Validation of Thermal-Formed Thin-Film Polyimide Panels,” Mechanical Systems and Signal 

Processing, Vol. 22, Iss. 6, Aug. 2008, pp. 1412–1426, DOI:10.1016/j.ymssp.2007.11.010. 
NASA/TM—2012-217014 525 

(9) Black, J.T., Smith, S.W. E, Leifer, J. E, and L.J. BradfordE, “Measuring and Modeling the 

Dynamics of Stiffened Thin-Film Polyimide Panels,” Journal of Guidance, Control, and Dynamics, 

Vol. 31, No. 3, May – Jun. 2008, pp. 490-500, DOI: 10.2514/1.32236. 

(10) Leifer, J. E, Black, J.T., Smith, S.W. E, Ma, N. E, and J.K. LumppE, “Measurement of In-Plane 

Motion of Thin-Film Structures Using Videogrammetry,” Journal of Spacecraft and Rockets, Vol. 

44, No. 6, Nov. – Dec. 2007, pp. 1317-1325, DOI: 10.2514/1.25566. 

(11) Black, J.T., Leifer, J. E, DeMoss, J.A. E, and E.N. WalkerE, “Experimental and Numerical 

Correlation of Gravity Sag in Solar-Sail-Quality Membranes,” Journal of Spacecraft and Rockets, 

Vol. 44, No. 3, May – Jun. 2007, pp. 522-527, DOI: 10.2514/1.20958. 

(12) Pappa, R.S. E, Black, J.T., Blandino, J.R. E, Jones, T.W. E, Danehy, P.M. E, and A.A. 

DorringtonE, “Dot-Projection Photogrammetry and Videogrammetry of Gossamer Space 

Structures,” Journal of Spacecraft and Rockets, Vol. 40, No. 6, Nov. – Dec. 2003, pp. 858-867. 

Archival Journal Articles Submitted 

(1) Jennings, A.A, Black, J.T., Magree, D.*, Briggs, G.*, and C. AllenA, “Accuracy of 

Photogrammetry Texture-Based Surface Reconstruction on Curved Surfaces,” AIAA Journal, 

submitted Oct. 2010. 

Refereed Conference Proceedings on Abstract Review 

(1) Niederhauser, J.D., and J.T. Black, “Characterization and Analysis for Flying COTS Electronics 

On-Orbit,” 25th AIAA Utah State University Conference on Small Satellites, Logan, UT, Aug. 

2011, Paper SSC11-XII-3. 

(2) Debes, J., Howard, N., Harrington, R., Cobb, R., and J. Black, “Rapid Build and Space 

Qualification of CubeSats,” 25th AIAA Utah State University Conference on Small Satellites, 

Logan, UT, Aug. 2011, Paper SSC11-VII-7. 

(3) Ross, J.T., Risbeck, M.R., Simmons, R.J., Lofthouse, A.J., and J.T. Black, “Experimental Fin Tips 

For Reusable Launch Vehicles (ExFiT) Flight Data Validation,” 52nd 

AIAA/ASME/ASCE/AHS/ASC Conference on Structures, Structural Dynamics and Materials, 

Denver, CO, Apr. 2011, AIAA Paper 2011-2162 

(4) Grigsby, D.A. and J.T. Black, “Satellite Capabilities Mapping – Utilizing Small Satellites,” 24th 

AIAA Utah State University Conference on Small Satellites, Logan, UT, Aug. 2010, Paper SSC10-

VII-1. 

(5) Book, T.A., Starr, W.J., Morse, A.J., Miller, S.D., Black, J.T., Swenson, E.D., Cobb, R.G., and C.R. 

Hartsfield, “A Design Overview of a Space-Based Chromotomographic Hyperspectral Imaging 

Experiment,” 24th AIAA Utah State University Conference on Small Satellites, Logan, UT, Aug. 

2010, Paper SSC10-I-1. 

(6) Starr, W.J., Book, T.A., Morse, A.J., Miller, S.D., Swenson, E.D., Cobb, R.G., and J.T. Black, 

“Target Acquisition/Tracking for the Space-Based Chromotomographic Hyperspectral Imaging 

Experiment,” AIAA/AAS Astrodynamics Specialist Conference, Toronto, Canada, Aug. 2010, 

AIAA Paper 2010-7655. 

(7) Yates, J.M., Spanbauer, B.W., and J.T. Black, “Geostationary Orbit Development and Evaluation 

for Space Situational Awareness (GODESSA),” AIAA/AAS Astrodynamics Specialist Conference, 

Toronto, Canada, Aug. 2010, AIAA Paper 2010-7528. 

(8) Bellows, C.T., Keller, N.M., and J.T. Black, “Mission Feasibility Study for Space Based Wireless 

Power Transfer,” AIAA/AAS Astrodynamics Specialist Conference, Toronto, Canada, Aug. 2010, 

AIAA Paper 2010-7522. 

(9) O'Dell, D.C., Bostick, R., Hawks, M.R., Swenson, E.D., Black, J.T., Cobb, R.G., and G.P. Perram, 

“Chromotomographic Imager Field Demonstration Results,” 7th SPIE Airborne Intelligence, 

Surveillance, Reconnaissance (ISR) Systems and Applications Conference, Apr. 2010, Proc. SPIE, 

Vol. 7668, 766804 (2010), doi:10.1117/12.849702. 
NASA/TM—2012-217014 526 

(10) Swenson, E., Black, J., and R. Cobb, “Correcting the Effects of Orthogonalized Measured Modes 

When Tuning Finite Element Models,” 51st AIAA/ASME/ASCE/AHS/ASC Conference on 

Structures, Structural Dynamics and Materials, Orlando, FL, Apr. 2010, AIAA Paper 2010-2546. 

(11) Jennings, A., Black, J., Magree, D., Briggs, G., Allen, C., and M. Jameson, “Effect of Camera 

Setup on Photogrammetry Texture-Based Surface Reconstruction,” 51st 

AIAA/ASME/ASCE/AHS/ASC Conference on Structures, Structural Dynamics and Materials, 

Orlando, FL, Apr. 2010, AIAA Paper 2010-2749. 

(12) Magree, D., Briggs, G., Allen, C., Jennings, A., Pollock A., and J. Black, “Pan Tilt Zoom Camera 

System for Dynamic In-Flight Tracking and Measurement,” 51st AIAA/ASME/ASCE/AHS/ASC 

Conference on Structures, Structural Dynamics and Materials, Orlando, FL, Apr. 2010, AIAA 

Paper 2010-2804. 

(13) Swenson, P., Thomas, G., Cobb, R., and J. Black, “Experiment Deployment Testing of a One Meter 

Reflector From a CubeSat,” 51st AIAA/ASME/ASCE/AHS/ASC Conference on Structures, 

Structural Dynamics and Materials, Orlando, FL, Apr. 2010, AIAA Paper 2010-2906. 

(14) Thomas, G., Swenson, P., Cobb, R., Swenson, E., and J. Black, “Prototype Development and 

Dynamic Characterization of Deployable CubeSat Booms,” 51st AIAA/ASME/ASCE/AHS/ASC 

Conference on Structures, Structural Dynamics and Materials, Orlando, FL, Apr. 2010, AIAA 

Paper 2010-2907. 

(15) Miller, S., Book, T., Morse, A., Swenson, E., Cobb, R., and J. Black, “Structural Design and 

Analysis of a Novel Space-Based Chromotomographic Spectrometer,” 51st 

AIAA/ASME/ASCE/AHS/ASC Conference on Structures, Structural Dynamics and Materials, 

Orlando, FL, Apr. 2010, AIAA Paper 2010-2935. 

(16) Trottier, M., Baghal, L., Swenson, E., Black, J., and C. Finley, “Accurate Dynamic Response 

Predictions of Various Plug-and-Play SAT I Configurations,” 51st AIAA/ASME/ASCE/AHS/ASC 

Conference on Structures, Structural Dynamics and Materials, Orlando, FL, Apr. 2010, AIAA 

Paper 2010-2958. 

(17) Jameson, M., Rogers, D., Allen, C., Blandino, J., Jennings, A., Magree, D., Pollock, S., and Black, 

J., "Motion Capture and Photogrammetry System Hybridization for Dynamic In-Flight Tracking 

and Measurement", 1st National Society of Black Engineers Aerospace Systems Conference, Feb. 

2010, Conference Best Paper. 

(18) McFarland, D., Swenson, E., Black, J., Cobb, R., and Fosbury, A., “Near Real-Time Closed-Loop 

Optimal Control Feedback for Spacecraft Attitude Maneuvers,” AIAA Modeling and Simulation 

Technologies Conference, Chicago, IL, Aug. 2009, AIAA Paper 2009-5814. 

(19) Doupe, C.C., Swenson, E.D., George, L.E., and Black, J.T., “Finite Element Model Tuning with 

Varying Experimental Data Density,” AIAA Modeling and Simulation Technologies Conference, 

Chicago, IL, Aug. 2009, AIAA Paper 2009-6038. 

(20) Kahraman, M.O., Swenson, E.D., and J.T. Black, “A Constraint Based Approach for Building 

Operationally Responsive Space Satellites,” 4th International Conference on Recent Advances in 

Space Technologies, June 2009. 

(21) Cooper, B., Cobb, R., and J. Black, “Rigidizable Inflatable Get-Away-Special Experiment Post 

Flight Analysis,” 50th AIAA/ASME/ASCE/AHS/ASC Conference on Structures, Structural 

Dynamics and Materials, Palm Springs, CA, May 2009, AIAA Paper 2009-2157. 

(22) Doupe, C., Swenson, E., George, L., and J. Black, “Finite Element Model Tuning with 3D Mode 

Shapes from FalconSAT- 5,” 50th AIAA/ASME/ASCE/AHS/ASC Conference on Structures, 

Structural Dynamics and Materials, Palm Springs, CA, May 2009, AIAA Paper 2009-2636. 

(23) Owens, J., Cobb, R., and Black, J., “Design and Flight Qualification of the Rigidizable Inflatable 

Get Away Special Experiment” 50th AIAA/ASME/ASCE/AHS/ASC Conference on Structures, 

Structural Dynamics and Materials, Palm Springs, CA, May 2009, AIAA Paper 2009-2155. 
NASA/TM—2012-217014 527 

(24) Pitcher, N.A., Black, J.T., Reeder, M.F., and R.C. Maple, “Videogrammetry Measurements of a 

Mini-UAV Wing in a Wind Tunnel,” 50th AIAA/ASME/ASCE/AHS/ASC Conference on 

Structures, Structural Dynamics and Materials, Palm Springs, CA, May 2009, AIAA Paper 2009-

2416. 

(25) Simpkins, J., Sollars, R., Allen, C., Jennings, A., and J.T. Black, “Dynamic Calibration and 

Performance of a Laser Steering System for Dynamic In-Flight Tracking and Measurement,” 50th 

AIAA/ASME/ASCE/AHS/ASC Conference on Structures, Structural Dynamics and Materials, 

Palm Springs, CA, May 2009, AIAA Paper 2009-2159. 

(26) Quarles, W., Blandino, J., and J. Black, “Development of a Laser Positioning System for On-Orbit 

Characterization of Deployable Booms,” 50th AIAA/ASME/ASCE/AHS/ASC Conference on 

Structures, Structural Dynamics and Materials, Palm Springs, CA, May 2009, AIAA Paper 2009-

2176. 

(27) Simpkins, J., Sollars, R., Jennings, A., Allen, C., and J. Black, “Calibration and Performance of 

Laser Steering System for Dynamic In-Flight Tracking and Measurement,” SPIE Smart Structures, 

Materials, and Non-Destructive Evaluation Conference, San Diego, CA, Mar 2009, SPIE Paper 

7292-131. 

(28) Simmons, J., Deleon, A., Black, J., and E. Swenson, “Aeroelastic Analysis and Optimization of 

FalconLaunch Sounding Rocket Fin,” 47th AIAA Aerospace Sciences Meeting, Orlando, FL, Jan. 

2009, AIAA Paper 2009-515. 

(29) Black, J.T., Swenson, E.D., and L. George, “Extracting 3D Mode Shapes of FalconSAT-5 

Structural Engineering Model,” 49th AIAA/ASME/ASCE/AHS/ASC Conference on Structures, 

Structural Dynamics and Materials, Schaumburg, IL, Apr. 2008, AIAA Paper 2008-1851. 

(30) Black, J., Leifer, J., and S.W. Smith, “Global Static Testing and Model Validation of ThermalFormed Thin Film Polyimide Panels,” 49th AIAA/ASME/ASCE/AHS/ASC Conference on 

Structures, Structural Dynamics and Materials, 8th Gossamer Spacecraft Forum, Schaumburg, IL, 

Apr. 2008, AIAA Paper 2008-2137. 

(31) Black, J., Leifer, J., and S.W. Smith, “Global Static Testing and Model Validation of ThermalFormed Thin Film Polyimide Panel-Arrays,” 49th AIAA/ASME/ASCE/AHS/ASC Conference on 

Structures, Structural Dynamics and Materials, 8th Gossamer Spacecraft Forum, Schaumburg, IL, 

Apr. 2008, AIAA Paper 2008-2215. 

(32) Black, J., Smith, S.W., Leifer, J., and L. Bradford, “Experimental Characterization and Modeling of 

Global Behavior of Semi-Rigid Thin Film Polyimide Panels,” 48th AIAA/ASME/ASCE/AHS/ASC 

Conference on Structures, Structural Dynamics and Materials, 8th Gossamer Spacecraft Forum, 

Waikiki, HI, Apr. 2007, AIAA Paper 2007-1833. 

(33) Black, J.T., Smith, S.W., and J. Leifer, “Reduced Model Validation of Thermal-Formed Polyimide 

Panels,” 25th International Modal Analysis Conference, Model Validation Methods Special 

Session, Orlando, FL, Feb. 2007. 

(34) Black, J.T., Whetzal, J.A., deBlonk, B.J., and J.J. Massarello, “Deployment Repeatability of 

Testing of Composite Tape Springs for Space Optics Applications,” 47th 

AIAA/ASME/ASCE/AHS/ASC Conference on Structures, Structural Dynamics and Materials, 7th 

Gossamer Spacecraft Forum, Newport, RI, May 2006, AIAA Paper 2006-1905. 

(35) Leifer, J., Smith, S.W., Black, J.T., Ma, N., and J.K. Lumpp, “Measurement of In-Plane Motion of 

Thin-Film Structures Using Videogrammetry,” 47th AIAA/ASME/ASCE/AHS/ASC Conference 

on Structures, Structural Dynamics and Materials, 7th Gossamer Spacecraft Forum, Newport, RI, 

May 2006, AIAA Paper 2006-1805. 

(36) Black, J.T., Smith, S.W., and J. Leifer, “Investigating Semi-Rigid Kapton Panels for use in Space 

Telescopes,” 56th International Astronautical Congress, Fukuoka, Japan, Oct. 2005, IAC-05-

C2.1.B.05. 

(37) Campbell, J.E., Black, J.T., and S.W. Smith, “Toward Field Videogrammetry of Rivulets on Bridge 

Stay Cables,” EURODYNE, 6th European Conference on Structural Dynamics, Paris, France, Sep. 

2005. 
NASA/TM—2012-217014 528 

(38) Black, J.T., deBlonk, B.J., Patrick, B., and S. Chodimella, “A Composite Support Structure for a 

Membrane Mirror,” 46th AIAA/ASME/ASCE/AHS/ASC Conference on Structures, Structural 

Dynamics and Materials, 6th Gossamer Spacecraft Forum, Austin, TX, Apr. 2005, AIAA Paper 

2005-2195. 

(39) Smith, S.W., Song, H., Baker, J., Black, J.T., and D. Muheim, “Flexible Models for Solar Sail 

Control,” 46th AIAA/ASME/ASCE/AHS/ASC Conference on Structures, Structural Dynamics and 

Materials, 6th Gossamer Spacecraft Forum, Austin, TX, Apr. 2005, AIAA Paper 2005-1801. 

(40) Campbell, J.E., Reichenbach, P.E., Black, J.T., and S.W. Smith, “Photogrammetric 

Characterization of Laboratory Rivulets for Cable Stayed Bridges,” 23rd International Modal 

Analysis Conference, Orlando, FL, Feb. 2005. 

(41) Black, J.T. and R.S. Pappa, “Photogrammetry and Videogrammetry Methods for Solar Sails and 

Other Gossamer Space Structures,” 45th AIAA/ASME/ASCE/AHS/ASC Conference on Structures, 

Structural Dynamics and Materials, 5th Gossamer Spacecraft Forum, Palm Springs, CA, Apr. 2004, 

AIAA Paper 2004-1662. 

(42) Black, J.T., Leifer, J., DeMoss, J.A., Walker, E.N., and W.K. Belvin, “Experimental and Numerical 

Correlation of Gravity Sag in Solar Sail Quality Membranes,” 45th AIAA/ASME/ASCE/AHS/ASC 

Conference on Structures, Structural Dynamics and Materials, 5th Gossamer Spacecraft Forum, 

Palm Springs, CA, Apr. 2004, AIAA Paper 2004-1579. 

(43) Blandino, J.R., Sterling, J., Baginski, F., Steadman, E., Black, J.T., and R.S. Pappa, “Optical Strain 

Measurement of an Inflated Cylinder with Application to Scientific Balloons,” 45th 

AIAA/ASME/ASCE/AHS/ASC Conference on Structures, Structural Dynamics and Materials, 5th 

Gossamer Spacecraft Forum, Palm Springs, CA, Apr. 2004, AIAA Paper 2004-1500. 

(44) Pappa, R.S., Black, J.T., and J.R. Blandino, “Photogrammetric Measurement of Gossamer 

Spacecraft Membrane Wrinkling,” Society for Experimental Mechanics Annual Conference, 

Charlotte, NC, Jun. 2003. 

(45) Leifer, J., Black, J.T., Belvin, K.W., and V. Behun, “Evaluation of Shear Compliant Boarders for 

Wrinkle Reduction in Thin Film Membrane Structures,” 44th AIAA/ASME/ASCE/AHS/ASC 

Structures, Structural Dynamics and Materials Conference, 4th Gossamer Spacecraft Forum, 

Norfolk, VA, Apr. 2003, AIAA Paper 2003-1984. 

(46) Blandino, J.R., Pappa, R.S., and J.T. Black, “Modal Identification of Membrane Structures with 

Videogrammetry and Laser Vibrometry,” 44th AIAA/ASME/ASCE/AHS/ASC Structures, 

Structural Dynamics and Materials Conference, 4th Gossamer Spacecraft Forum, Norfolk, VA, 

Apr. 2003, AIAA Paper 2003-1745. 

(47) Johnston, J.D., Blandino, J.R., Black, J.T., and R.S. Pappa, “Structural Analysis and Testing of a 

1/20th Scale NGST Sunshield Membrane Layer,” 44th AIAA/ASME/ASCE/AHS/ASC Structures, 

Structural Dynamics and Materials Conference, 4th Gossamer Spacecraft Forum, Norfolk, VA, 

Apr. 2003, AIAA Paper 2003-1742. 

(48) Black, J.T., “Photogrammetry and Videogrammetry Methods for Solar Sails and Other Gossamer 

Structures,” AIAA Region I Student Conference, College Park, MD, Apr. 2003. 

(49) Black, J.T. and R.S. Pappa, “Videogrammetry Using Projected Circular Targets: Proof-of-Concept 

Test,” 21st International Modal Analysis Conference, Orlando, FL, Feb. 2003. 

(50) Pappa, R.S., Black, J.T., Blandino, J.R., Jones, T.W., Danehy, P.M., and A.A. Dorrington, “Dot 

Projection Photogrammetry and Videogrammetry of Gossamer Space Structures,” 21st 

International Modal Analysis Conference, Orlando, FL, Feb. 2003. 

(51) Pappa, R.S., Jones, T.W., Black, J.T., Walford, A., Robson, S. and M. Shortis, “Photogrammetry 

Methodology Development for Gossamer Spacecraft Structures,” 43rd 

AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference, 3rd 

Gossamer Spacecraft Forum, Denver, CO, Apr. 2002, AIAA Paper 2002-1375. 
NASA/TM—2012-217014 529 

NASA Technical Memoranda

(1) Black, J.T., “Photogrammetry and Videogrammetry Methods Development for Solar Sail 

Structures,” Masters Thesis, Joint Institute for Advancement of Flight Sciences, NASA Langley 

Research Center, Sep. 2003, NASA/CR-2003-212678. 

(2) Black, J.T. and R.S. Pappa, “Videogrammetry Using Projected Circular Targets: Proof-of-Concept 

Test,” Feb. 2003, NASA/TM-2003-212148. 

(3) Pappa, R.S., Black, J.T., Blandino, J.R., Jones, T.W., Danehy, P.M., and A.A. Dorrington, “Dot 

Projection Photogrammetry and Videogrammetry of Gossamer Space Structures,” Feb. 2003, 

NASA/TM-2003-212146. 

(4) Pappa, R.S., Jones, T.W., Black, J.T., Walford, A., Robson, S., and M.R. Shortis, “Photogrammetry 

Methodology Development for Gossamer Spacecraft Structures,” Jun. 2003, NASA/TM-2002-

211739. 

Other 

(1) Anderson, J., Douple, C., Black, J., and E. Swenson, “Reach for the Sky: Small Satellite Finite 

Element Model Optimization Using Laser Vibrometry,” Polytec InFocus, Iss. 01, 2011, pp. 11-13. 

(2) PhotoModeler Inc., PhotoModeler Video - Industrial Example, online at 

http://www.photomodeler.com/applications/scientific_engineering/pmv_muav.htm 

(3) Black, J.T., “New Ultra-lightweight Stiff Panels for Space Apertures,” PhD Dissertation, 

Department of Mechanical Engineering, University of Kentucky, Dec. 2006, online at 

http://lib.uky.edu/ETD/ukymeen2006d00532/BlackDis.pdf 

(4) Black, J.T., “Photogrammetry and Videogrammetry Methods Development for Solar Sail 

Structures,” Masters Thesis, Joint Institute for Advancement of Flight Sciences, NASA Langley 

Research Center, George Washington University, Gelman Library Special Collections, Stacks, Call 

Number: AS36.G3 2003, Sep. 2003. 

(5) Pappa, R.S., Jones, T.W., Black, J.T., Walford, A., Robson, S., and M.R. Shortis, “Photogrammetry 

Methodology Development for Gossamer Spacecraft Structures,” Sound and Vibration, Aug. 2002, 

pp. 12-21. 

Presentations 

(1) Black, J.T., “(YIP08) Hybrid Laser/Video 3D Non‐contact Motion Capture and Analysis,” Air 

Force Office of Scientific Research Structural Mechanics and Non-Destructive Evaluation Program 

Review, Destin, FL, Aug. 2010. 

(2) Black, J.T., “Freely Deploying Structures Analysis,” AFOSR Workshop on Improved Precision for 

Space Structures, May 2010. 

(3) Black, J.T., and R.F. Beach, “Space Research and Education in Ohio,” Ohio State University 

Department of Aeronautics Graduate Symposium, Nov. 2009, Invited Presentation. 

(4) Black, J.T., “Dynamic In-Flight Test and Measurement,” US-Korean Conference on Science, 

Technology, and Entrepreneurship, July 2009, Invited Presentation 

(5) Jennings, A., Allen, C., and J.T. Black, “Calibration and Development of Laser Steering for InFlight Noncontact Measurements,” 34th AIAA Dayton-Cincinnati Aerospace Science Symposium, 

Dayton, OH, Mar. 2009. 

(6) Allen, C., Jennings, A., and J.T. Black, “Dynamic In-Flight Tracking and Measurement Labs: An 

Overview,” 34th AIAA Dayton-Cincinnati Aerospace Science Symposium, Dayton, OH, Mar. 

2009. 

(7) Cooper, B., Black, J., and R. Cobb, “Post Space Flight Analysis of the Rigidizable Inflatable GetAway-Special Experiment,” 34th AIAA Dayton-Cincinnati Aerospace Science Symposium, 

Dayton, OH, Mar. 2009. 

(8) Sheirich, P. and J. Black, “An Engineering Trade Space Analysis for a Space Based Hyperspectral 

Chromotomographic Scanner,” 34th AIAA Dayton-Cincinnati Aerospace Science Symposium, 

Dayton, OH, Mar. 2009. 
NASA/TM—2012-217014 530 

(9) Doupe, C., Swenson, E., and J. Black, “FalconSAT 5 Finite Element Model Tuning Based on Laser 

Vibrometry,” 34th AIAA Dayton-Cincinnati Aerospace Science Symposium, Dayton, OH, Mar. 

2009. 

(10) Black, J.T. and R.F. Beach, “In-Space Laser Power Beaming Demonstration,” invited presentation 

at the US Air Force Academy’s FalconSAT-6 Symposium, Feb. 2009. 

(11) Cooper, B., Black, J., and R. Cobb, “Inflatable/Rigidizable Space Structures,” 4th ASME Dayton 

Engineering Sciences Symposium, Dayton, OH, Oct. 2008. 

(12) Black, J.T., “Space Research and Education at AFIT,” invited presentation at Sandia National 

Laboratories, Apr. 2008 

(13) Welty, N.F., Rohe, W.C., Roach, N.R., Black, J.T., and J. Walter, “Systems Engineering Approach 

to Development of a Research Testbed for Spacecraft Dynamics,” 33rd AIAA Dayton-Cincinnati 

Aerospace Sciences Symposium, Dayton, OH, Mar. 2008. 

(14) Black, J.T., Smith, S.W., and J. Leifer, “New Ultra-lightweight Stiff Panels for Space Apertures,” 

32nd AIAA Dayton-Cincinnati Aerospace Science Symposium, Dayton, OH, Mar. 2007. 

(15) Black, J.T., “New Ultra-lightweight Stiff Panels for Space Apertures,” Presentation to NASA 

Sponsors, NASA Langley Research Center, Hampton, VA, Nov. 2006. 

(16) Black, J.T., Smith, S.W., and J. Leifer, “Static Compression Testing and Modeling of Semi-Rigid 

Panels for Orbital Apertures,” 31st Dayton-Cincinnati Aerospace Science Symposium, Dayton, 

OH, Mar. 2006. 

(17) Black, J.T., “New Gossamer Structures for use in Future Space Telescopes,” University of North 

Dakota, Department of Space Studies, Space Studies Colloquium, 28 Nov. 2005. 

(18) Black, J.T., “56th International Astronautical Congress,” University of North Dakota, Department 

of Space Studies, Brown Bag Lunch Seminar, 29 Nov. 2005. 

(19) Black, J.T., “New Gossamer Structures for use in Future Space Telescopes,” Air Force Research 

Laboratory, Space Vehicles Directorate, Summer Seminar Series, 08 Jul. 2005. 

(20) Black, J.T., and W.K. Belvin, “Ultra-lightweight and Inflatable Space Structures,” Air Force 

Research Laboratory, Space Vehicles Directorate, Summer Seminar Series, 15 Jul. 2004. 

(21) Black, J.T., and W.K. Belvin, “Ultra-lightweight and Inflatable Space Structures,” University of 

Kentucky, Department of Mechanical Engineering, AIAA Seminar, 16 Feb. 2004. 

(22) Black, J.T., “Photogrammetry and Videogrammetry: The Second Generation,” NASA Langley 

Research Center, Structural Dynamics Branch, Technical Interchange Meeting, 15 Oct. 2002. 

(23) Black, J.T. and R.S. Pappa, “Photogrammetry and Solar Sails,” NASA Langley Research Center, 

Structural Dynamics Branch, Technical Interchange Meeting, 19 Mar. 2002. 

Service 

Institute 

 Space Systems Certificate Program, Chair, Department of Aeronautics and Astronautics, Air 

Force Institute of Technology, 2008–present 

 Executive Committee Member of the Ohio Space Grant Consortium representing the Air Force 

Institute of Technology, 2007–present 

 AFIT Faculty Council Secretary, 2009–2010 

 AIAA Faculty Advisor, Air Force Institute of Technology, 2007–present 

 DoD Joint Space Academic Working Group (JSAG) member, 2008–present 

 AFIT Inter-departmental Space Working Group (ISWG) member, 2008–present 

 Faculty Search Committee, Chair, Department of Aeronautics and Astronautics, Air Force 

Institute of Technology, 2009, 2010 

 Dean’s Reader, Callaway, D., ENY Dissertation, 2011 

 Dean’s Reader, Hyde, M., ENG Dissertation, 2010 
NASA/TM—2012-217014 531 

 Dayton Area Graduate Studies Institute (DAGSI) Application Evaluator, AFIT/ENY, Jan. 2008, 

Jan. 2009 

 Best Thesis Award Committee, Department of Aeronautics and Astronautics, Air Force Institute 

of Technology, Mar. 2007, Mar. 2008 

Professional 

 36th AIAA Dayton-Cincinnati Aerospace Sciences Symposium (DCASS) General Chair, Mar. 

2012 

 53rd AIAA/ASME/ASCE/AHS/ASC Conference on Structures, Structural Dynamics and 

Materials, Structural Dynamics Technical Chair, Apr. 2012 

 35th AIAA Dayton-Cincinnati Aerospace Sciences Symposium (DCASS) Technical Chair, Mar. 

2011 

 52nd AIAA/ASME/ASCE/AHS/ASC Conference on Structures, Structural Dynamics and 

Materials, Structural Dynamics co-Technical Chair, Apr. 2011 

 AIAA Structural Dynamics Technical Committee, 2008–present 

 AIAA Gossamer Spacecraft Program Committee, 2008–present 

 AIAA Faculty Advisor, Air Force Institute of Technology, 2007–present 

 AIAA Dayton-Cincinnati Section Treasurer, 2008–2010 

 AIAA Journal Reviewer 

 AIAA Journal of Aircraft Reviewer 

 AIAA Journal of Spacecraft and Rockets Reviewer 

 Shock and Vibration Reviewer 

 International Micro Air Vehicle Journal Reviewer 

 Acta Astronautica Reviewer 

 Special Session Organizer, 10th Gossamer Spacecraft Forum, 50th 

AIAA/ASME/ASCE/AHS/ASC Conference on Structures, Structural Dynamics and Materials, 

Palm Springs, CA, May 2009 

 Session co-chair at AIAA/ASME/ASCE/AHS/ASC Conference on Structures, Structural 

Dynamics and Materials, 2009, 2010 

 Session co-chair at Gossamer Spacecraft Forum, AIAA/ASME/ASCE/AHS/ASC Conference on 

Structures, Structural Dynamics and Materials, 2005, 2007–2010 

 33rd AIAA Dayton-Cincinnati Aerospace Sciences Symposium (DCASS) Keynote and Invited 

Sessions Chair, Mar. 2008 

 AIAA Gossamer Spacecraft Program Committee, special project on the history of gossamer 

programs, 2006 

 Dayton Area Graduate Studies Institute (DAGSI) Graduate Fellowship Application Evaluator, 

2008, 2009, 2010 

 Abstract Reviewer, AIAA/ASME/ASCE/AHS/ASC Conference on Structures, Structural 

Dynamics and Materials, 2003 – 2011 

 AIAA Langley Student Branch Executive Board, Hampton, VA, 2002–2003 

 
NASA/TM—2012-217014 532 

James Claude Dickens, Ph.D., P.E. 

Vice Director of the Center for Pulsed Power and Power Electronics 

Professor of Electrical and Computer Engineering 

Licensed Professional Engineer State of Texas, no. 88713 

Texas Tech University, Lubbock, Texas 79409 

james.dickens@ttu.edu 806–742–3533 Fax: 806–742–1245 

8401 County Road 6940 Cell: 806–790–9876 

Lubbock, Texas 79407 TTU: 806–742–1254 

Education 

 Ph.D. in Electrical Engineering, Texas Tech University, 1995 

 M.S. in Electrical Engineering, Texas Tech University, 1993 

 B.S. in Electrical Engineering, Texas Tech University, 1991 

Teaching areas: Electronics, Circuits, Pulsed Power, Electrical Power Distribution Systems, Power 

Electronics, Optics and Optical Systems, Failure Analysis, Fields, Microwave and RF Circuits, 

Communications and Communication Circuits, and Electric Space Propulsion. 

Research areas: Pulsed Power, Pulsed RF Sources and Applications, Pulsed Lasers, Fiber Laser and 

applications, Electrical Failure Forensics, Grounding and Shielding, Power Electronics, High 

Efficiency DC/DC Converters, Explosive Pulsed Power, High Power Microwaves, Arc and Plasma 

Diagnostics, and Electric Space Propulsion. 

Short courses taught: Grounding and Shielding, Power Electronics, Personnel Safety, Power 

Distribution Systems, Pulse Generators, High Voltage Generators, Compact Pulsed Power, Compact 

DC/DC Technology, High Power Microwaves, and Electrical and Optical Diagnostics. 

Professional History 

Texas Tech University 

Charles Bates Thornton Professor, Electrical and Computer Engineering, Sept. 2009 

Professor, Electrical & Computer Engineering, Sept. 2008 

Associate Professor, Electrical & Computer Engineering, Sept. 2001–Aug. 2008 

Assistant Professor, Electrical & Computer Engineering, Jan. 1999–Aug. 2001 

Research Assistant Professor, Electrical & Computer Engineering, May 1996–Dec. 1998 

Consulting 

 Multiple Expert Witness Cases (discovery, depositions, reports and testimony) 

 Primex Aerospace 

 FMV Swedish Defense Department 

 Federal Aviation Administration (FAA) 

 High Tech Consultants 

 BAE Systems 

 L3 Communications 

 Lockheed Martin 

 Boeing Corporation 

 White Sands Missile Range 

 NASA 

Awards, Professional Societies, and Services 

 Technical Program Chair for the 2008 IEEE Power Modulator Conference in Las Vegas, NV 

 Guest editor of IEEE Transactions on Plasma Science Special Issue on Pulsed Power 
NASA/TM—2012-217014 533 

 Lecturer at High Power Microwave Workshop at FOI in Grindsjön, Sweden, Aug. 2005. 

 Presented an invited talk at Kumamoto University’s International Forum on Pulsed Power 

Science. All expenses were paid by the host institution. 

 2005 Chancellor’s Excellence in Research Award from Texas Tech University. 

 Vice-Chairman of 2003 IEEE International Pulsed Power Conference. 

 Civilian Research Defense Foundation (CRDF) Chairman of Physics panel 2003 

 Civilian Research Defense Foundation (CRDF) International onsite reviewer 2000–present. 

 CRDF Physics Review Panel 1999–present. 

 AIAA Electric Propulsion Technical Committee Secretary 2000–2002. 

 Session Organizer at 2002 International Conference of Plasma Science. 

 Session Organizer at the 2002 Power Modulator Symposium. 

 AIAA Electric Propulsion Technical Committee 1999–present. 

 Organizer and lecturer at workshop on Pulsed Power Technology, Bechtel, 2001. 

 Organizer at workshop on Pulsed Power Technology, China Lake. 

 Workshop on Pulsed Power Technology workshop at Texas Tech University. 

 Pulsed Power Science and Technology Committee 1999–present. 

 Session Organizer at 2000 International Conference of Plasma Science. 

 Japanese Society for the Promotion of Science (JSPS) Fellow 1996 

 American Institute of Aeronautics and Astronautics (AIAA), 1993–present. 

 Institute of Electrical and Electronics Engineers (IEEE), 1990–present. 

Grants and Contracts 

TTU ORS Report $ 13,169,836 total award credited to individual PI. (1/20/11) 

 Dr. Dickens has the highest funded research in the COE over the past decade. 

 Dr. Dickens has been P.I / Co-P.I. on three highly competitive MURI grants. 

 Dr. Dickens is P.I. on a highly competitive Navy IED basic research grant. (>400 original 

submissions 8 awarded) 

 Dr. Dickens has been P.I. on five highly competitive DURIP equipment grants. 

Direction of Graduate Students (all students wrote Thesis or Dissertation) 

Student Degree Function Grad date 

Braxton Bragg Ph.D. Chair Active 

Billy Sullivan Ph.D. Chair Active 

Cameron Hettler Ph.D. Chair Active 

Jonathan Parson Ph.D. Chair Active 

John Walter Ph.D. Chair Active 

Shad Holt Ph.D. Chair Active 

Braxton Bragg M.S.E.E. Chair 2010 

Colt James Ph.D. Chair 2010 

Cameron Hettler M.S.E.E.. Chair 2009 

William Sullivan M.S.E.E. Chair 2009 

Jonathan Parson M.S.E.E. Chair 2008 

Yeong-Jer Chen Ph.D. Chair 2008 

Thomas Holt Ph.D. Committee member 2008 

David Belt Ph.D. Co-Chair 2008 

Ryan Karhi M.S.E.E. Committee member 2007 

Colt James M.S.E.E. Chair 2007 

Brent McHale Ph.D. Committee member 2006 
NASA/TM—2012-217014 534 

Student Degree Function Grad date 

Mic Cevallos Ph.D. Chair 2006 

Shad Holt M.S.E.E. Chair 2006 

Juan Carlos Ph.D. Committee member 2004 

Michael Hoffman M.S.E.E Chair 2003 

Heath Keene M.S.E.E Chair 2003 

David J. Hemmert Ph.D Committee member 2003 

Johnathan Blackwell M.S.E.E Chair 2002 

Michael Cevallos M.S.E.E. Chair 2002 

Roberto Izquierdo M.S.E.E Chair 2001 

Juan Carlos M.S.E.E Chair 2001 

Zhan Mei M.S.E.E Chair 2001 

Efren Brito M.S.E.E Chair 2001 

John Walter M.S.E.E Chair 2000 

Michael Cevallos M.S.E.E Chair 2000 

Guruprakash 

Radhakrishna 

M.S.E.E Committee member 2000 

Naveen Anumalla M.S.E.E Committee member 1999 

Publications 

Journals 

(1) “Window Breakdown Caused by High Power Microwaves,” IEEE Trans. Plasma Science, HPM 

Issue, 26 296 (1998) (with A. Neuber, M. Kristiansen, D. Hemmert, H. Krompholz, and L.L. 

Hatfield). 

(2) “High Voltage Subnanosecond Breakdown,” IEEE Transactions on Plasma Science, Special Issue 

on High Power Microwave Generation, 26 874 (1998) (with J. Mankowski and M. Kristiansen). 

(3) “Imaging of High-Power Microwave-Induced Surface Flashover,” IEEE Transactions on Plasma 

Science Special Issue “Images in Plasma Science”, 27 138 (1999) (with A. Neuber, D. Hemmert, 

M. Kristiansen, H. Krompholz, and L.L. Hatfield). 

(4) “High Power Microwave Generation by a Coaxial Virtual Cathode Oscillator,” IEEE Transactions 

on Plasma Science, 27 1538 (1999) (with W. Jiang, K. Woolverton, and M. Kristiansen). 

(5) “Efficiency Enhancement of Coaxial Virtual Cathode Oscillator,” 27 1543 (1999) IEEE 

Transactions on Plasma Science, (with W. Jiang and M. Kristiansen). 

(6) “Pulsed Power Generation Using Ferromagnetic Circuits,” IEEE Transactions on Plasma Science, 

Vol. 28, 2000. (with S.I. Shkuratov, M. Kristiansen, L.L. Hatfield, and R. Martin). 

(7) “High-Current and High-Voltage Pulsed Testing of Resistors,” IEEE Transactions on Plasma 

Science, Volume: 28 , Issue: 5 , Oct. 2000 Pages:1607 - 1614 , (with S.I. Shkuratov, M. 

Kristiansen, L.L. Hatfield, and E. Horrocks). 

(8) “Optical Diagnostics on Helical Flux Compression Generators,” IEEE Transactions on Plasma 

Science, Special Issue on Pulsed Power Science and Technology, Vol. 28, 1445-1450, 2000. (with 

A. Neuber, M. Kristiansen, and H. Krompholz). 

(9) “High Voltage Subnanosecond Corona Inception,” IEEE Transactions on Plasma Science, Special 

Issue on Pulsed Power Science and Technology, Vol. 28, 2000. (with J. Mankowski and M. 

Kristiansen). 

(10) “Microwave Magnetic Field Effects on High Power Microwave Window Breakdown,” IEEE 

Transactions on Plasma Science, Special Issue on High Power Microwaves, Vol. 28, 472-477, 

2000. (with D. Hemmert, A. Neuber, M. Kristiansen, H. Krompholz, and L.L. Hatfield). 

 
NASA/TM—2012-217014 535 

(11) “Electrical Behavior of a Simple Helical Flux Compression Generator for Code Benchmarking,” 

IEEE Transactions on Plasma Science,Volume: 29 , Issue: 4 , Aug. 2001, Pages:573 - 581 , (with 

A. Neuber, M. Kristiansen, J.B. Cornette, K. Jamison, R. Parinson, M. Giesselmann, P. Worsey, J. 

Baird, and M. Schmidt). 

(12) “Subnanosecond Corona Inception in an Ultra Wideband Environment,” IEEE Transactions on 

Plasma Science,Volume: 30 , Issue: 3 , June 2002 Pages:1211 - 1214, (with John Mankowski M. 

Kristiansen, J. Lehr, W. Prather, and J. Gaudet). 

(13) “Compact Explosive-Driven Generator of Primary Power Based on a Longitudinal Shock Wave 

Demagnetization of Hard Ferri- and Ferromagnetics”, IEEE Transactions on Plasma 

Science,Volume: 30, Issue: 5, Oct. 2002 Pages:1681 - 1691, (with S.I. Shkuratov, M. Kristiansen, 

and J.C. Hernendez). 

(14) “Current Mode of Pulsed Power Generation in Moving Magnet Systems”, IEEE Transactions on 

Plasma Science,Volume: 30 , Issue: 5 , Oct. 2002 Pages:1674 - 1680, (with S.I. Shkuratov, M. 

Kristiansen, and J.C. Hernandez). 

(15) “Rapid decontamination of large surface areas,” IEEE Transactions on Plasma Science,Volume: 28 

, Issue: 1 , Feb. 2000, Pages:173 - 179. (with Farrar, L.C.; Haack, D.P.; McGrath, S.E.; O'Hair, 

E.A.; Fralick, J.A). 

(16) “Shock Wave Demagnetization of BaFe12O19 Hard Ferrimagnetics”, Journal of Applied Physics 

91 (2002) 3007-3009 (with S.I. Shkuratov, E.F. Talantsev and M. Kristiansen. 

(17) “Transverse Shock Wave Demagnetization of Nd2Fe14B High-Energy Hard Ferromagnetics, 

Journal of Applied Physics 92 (2002) 159-162 (with S.I. Shkuratov, E.F. Talantsev, and M. 

Kristiansen). 

(18) “Ultracompact Explosive-Driven High-Current Source of Primary Power Badsed on Shock Wave 

Demagnetization of Nd2Fe14B Hard Ferromagnetics”, Review of Scientific Instruments, 73 (2002) 

2738-2742 (with S.I. Shkuratov, E.F. Talantsev, and M. Kristiansen). 

(19) “The Conductivity of a Longitudinal-Shock-Wave-Compressed Nd2Fe14B Hard Ferromagnetics, 

Modern Physics Letters B, 16, No. 12 (2002) 1-11 (with E.F. Talantsev, S.I. Shkuratov, and M. 

Kristiansen). 

(20) “Theoretical Treatment of Explosive Driven Ferroelectric Generators”, IEEE Transactions on 

Plasma Science,Volume: 30 , Issue: 5 , Oct. 2002 Pages:1665 - 1673, (with S.I. Shkuratov, Ya. 

Tkach, E.F. Talantsev, M. Kristiansen, L.L. Altgilbers, and P.T. Tracy). 

(21) “Single Shot, Repetitive and Life-Time High-Voltage Testing of Capacitors”, IEEE Transactions on 

Plasma Science,Volume: 30 , Issue: 5 , Oct. 2002 Pages:1943 - 1949, (with S.I. Shkuratov, E.F. 

Talantsev, L.L. Hatfield, and M. Kristiansen). 

(22) “Studies on a Helical Magnetic Flux Compression Generator,” SAE Transactions, Journal of 

Aerospace, vol. 109, pp. 865-869, 2000 (This Transactions volume was published in 2001 and 

“...contains the best 135 technical papers of all those presented in 2000.”) (with A. Neuber, M. 

Kristiansen, M. Giesselmann, B. Freeman, D. Dorsey, P. Worsey, J. Baird, and M. Schmidt). 

(23) “Longitudinal Shock Wave Demagnetization of High Energy Nd2Fe14B Ferromagnetics,” Appl. 

Phys. Lett. 82, 1248 (2003) (with S.I. Shkuratov, E.F. Talantsev, and M. Kristiansen. 

(24) “Thermodynamic State of The Magnetic Flux Compression Generator Volume,” IEEE Trans. on 

Plasma Science, vol. 30, 1659-1664 (2002). (with A. Neuber, T. Holt, and M. Kristiansen). 

(25) ”Helical MFCG For Driving A High Inductance Load,” accepted for publication in the Special 

Edition of the Journal of EM Phenomenon on FCGs (Oct. 2003) (with Andreas A. Neuber, JuanCarlos Hernández, Magne Kristiansen). 

(26) “Quantification of Ohmic and Intrinsic Flux Losses in Helical Flux Compression Generators,” 

IEEE Transactions on Plasma Science, vol. 32, pp. 1902-1908 (2004) (with Juan-Carlos Hernandez, 

Andreas A. Neuber and Magne Kristiansen). 

(27) “DC Flashover of a Dielectric Surface in Atmospheric Conditions,” IEEE Transactions on Plasma 

Science, vol. 32, pp. 1828-1834 (2004) (with John T. Krile, Andreas A. Neuber, and Hermann G. 

Krompholz). 
NASA/TM—2012-217014 536 

(28) “Magnetic Flux Compression Generators,” Proceeding of the IEEE, vol. 92 number 7, 1205-1215 

July 2004. (with Andreas A. Neuber). 

(29) “Microwave Frequency Determination Mechanisms in A Coaxial Vircator,” IEEE Transactions on 

Plasma Science, vol. 32, pp. 1799-1804 (2004) (with X. Chen, J. Mankowski, L. Hatfield, E. Choi, 

and M. Kristiansen). 

(30) “A Bench Top Railgun With Distributed Energy Sources,” Mankowski, J.; Dickens, J.; 

Giesselmann, M.; McDaniel, B.; McHale, B.; Kristiansen, M., Magnetics, IEEE Transactions on 

Plasma Science, Volume 43, Issue 1, Part 2, Jan. 2007 Page(s):167 – 169. 

(31) “Interface Breakdown During High-Power Microwave Transmission,” Neuber, A. A.; Edmiston, G. 

F.; Krile, J. T.; Krompholz, H.; Dickens, J. C.; Kristiansen, M., Magnetics, IEEE Transactions on 

Plasma Science, Volume 43, Issue 1, Part 2, Jan. 2007 Page(s):496 – 500. 

(32) “High-Power Microwave Surface Flashover of a Gas–Dielectric Interface at 90–760 torr,” 

Edmiston, G.; Krile, J.; Neuber, A.; Dickens, J.; Krompholz, H., IEEE Transactions on Plasma 

Science, Volume 34, Issue 5, Part 1, Oct. 2006 Page(s):1782 – 1788. 

(33) “The Impact of Field Enhancements and Charge Injection on the Pulsed Breakdown Strength of 

Water,” Wetz, D.A.; Mankowski, J.J.; Dickens, J.C.; Kristiansen, M., IEEE Transactions on Plasma 

Science, Volume 34, Issue 5, Part 1, Oct. 2006 Page(s):1670 – 1679. 

(34) “Conduction and breakdown mechanisms in transformer oil,” Butcher, M.; Neuber, A.A.; Cevallos, 

M.D.; Dickens, J.C.; Krompholz, H., IEEE Transactions on Plasma Science, Volume 34, Issue 2, 

Part 3, April 2006 Page(s):467 – 475. 

(35) “A bench top railgun with distributed energy sources” Mankowski, J; Dickens, J; Giesselmann, M, 

et al., IEEE TRANSACTIONS ON MAGNETICS Volume: 43 Issue: 1 Special Issue: SI 

Pages: 167-169 Part: 2 Published: JAN 2007. 

(36) “Interface breakdown during high-power microwave transmission”, Neuber, AA; Edmiston, GF; 

Krile, JT, James Dickens, IEEE TRANSACTIONS ON MAGNETICS Volume: 43 Issue: 1 

Special Issue: SI Pages: 496-500 Part: 2 Published: JAN 2007. 

(37) “Contributing factors to window flashover under pulsed high power microwave excitation at high 

altitude”, Edmiston, G; Neuber, A; McQuage, James Dickens, IEEE TRANSACTIONS ON 

DIELECTRICS AND ELECTRICAL INSULATION Volume: 14 Pages: 783-789 Published: 

2007. 

(38) Low-Jitter Triggered Spark Gap With High-Pressure Gas Mixtures Yeong-Jer Chen; Mankowski, 

J.J.; Dickens, J.C.; Walter, J.; Kristiansen, M.; IEEE Transactions on Plasma Science, Volume 36, 

Issue 5, Part 3, Oct. 2008 Page(s):2546 - 2553 Digital Object Identifier 

10.1109/TPS.2008.2004366. 

(39) Compact Electroexplosive Fuses for Explosively Driven Pulsed Power McCauley, D.R.; Belt, 

D.W.; Mankowski, J.J.; Dickens, J.C.; Neuber, A.A.; Kristiansen, M.; IEEE Transactions on 

Plasma Science Volume 36, Issue 5, Part 3, Oct. 2008 Page(s):2691 - 2699 Digital Object 

Identifier 10.1109/TPS.2008.2004230. 

(40) Utilization of a Nonexplosive Test Bed for Flux-Compression-Generator Electroexplosive Opening 

Switches Belt, D.W.; Mankowski, J.J.; Neuber, A.A.; Dickens, J.C.; Kristiansen, M.; IEEE 

Transactions on Plasma Science, Volume 36, Issue 5, Part 3, Oct. 2008 Page(s):2684 - 2690 

Digital Object Identifier 10.1109/TPS.2008.2001889. 

(41) Secondary Arc Formation Within a Distributed Energy Railgun Karhi, R.W.; Mankowski, J.J.; 

Dickens, J.C.; Kristiansen, M.; Wetz, D.A.; IEEE Transactions on Plasma Science, Volume 36, 

Issue 5, Part 3, Oct. 2008 Page(s):2738 – 2746 Digital Object Identifier 

10.1109/TPS.2008.2004228. 

(42) Imaging of the Explosive Emission Cathode Plasma in a Vircator High-Power Microwave Source 

Walter, J.; Mankowski, J.; Dickens, J.; IEEE Transactions on Plasma Science Volume 36, Issue 4, 

Part 1, Aug. 2008 Page(s):1388 - 1389 Digital Object Identifier 10.1109/TPS.2008.924489. 
NASA/TM—2012-217014 537 

(43) Optimization of a low jitter, 50 kV, 100 Hz triggered spark gap with high pressure gas mixtures, 

Chen, Y.; Dickens, J.; Mankowski, J.; Kristiansen, M.; IEEE Transactions on Dielectrics and 

Electrical Insulation, Volume 16, Issue 4, August 2009 Page(s):971 – 978 Digital Object Identifier 

10.1109/TDEI.2009.5211842. 

(44) Optimization of an FCG-Based High-Power Microwave System Using Nonexplosive Pulsed Power, 

Davis, C. B.; Neuber, A. A.; Young, A.; Walter, J.; Dickens, J. C.; Kristiansen, M.; Plasma Science, 

IEEE transactions on Volume 37, Issue 12, Part 2, Dec. 2009 Page(s):2321 - 2327 Digital Object 

Identifier 10.1109/TPS.2009.2033602. 

(45) Recent Advances in Explosive Pulsed Power, L.L. Algilbers, A.H. Stults, J. Dickens, et. al., Journal 

of Directed Energy, Vol. 3, No. 2, Spring 2009, pg 149. 

(46) High electric field packaging of silicon carbide photoconductive switches Hettler, C.; James, C.; 

Dickens, J.; Pulsed Power Conference, 2009. PPC '09. IEEE June 28 2009-July 2 2009 

Page(s):631-634 Digital Object Identifier 10.1109/PPC.2009.5386380. 

(47) VUV Emission and Streamer Formation in Pulsed Dielectric Surface Flashover at Atmospheric 

Pressure, Rogers, T.G.; Neuber, A.A.; Frank, K.; Laity, G.R.; Dickens, J.C.; IEEE Transactions on 

Plasma Science, Volume: 38 , Issue: 10 , Part: 1 Digital Object Identifier: 

10.1109/TPS.2010.2059714 Publication Year: 2010 , Page(s): 2764 – 2770. 

(48) Design and Evaluation of a Compact Silicon Carbide Photoconductive Semiconductor Switch, 

James, C.; Hettler, C.; Dickens, J.; IEEE Transactions on Electron Devices, Volume: 58 , Issue: 2 

Digital Object Identifier: 10.1109/TED.2010.2089689 Publication Year: 2011 , Page(s): 508 – 511. 

(49) Pulsed magnetic field excitation sensitivity of match-type electric blasting caps, Parson J., Dickens 

J., Walter J., et al., REVIEW OF SCIENTIFIC INSTRUMENTS Volume: 81 Issue: 10 Article 

Number: 105115 OCT 2010. 

Book Chapters 

(1) Explosively Driven Pulsed Power, Helical Magnetic Flux Compression Generators”, A. Neuber 

Edtr., Springer Berlin Heidelberg New York, ISBN 3-540-26051-X, 2005. 

Conference Proceedings and Presentations 

(1) “Insulator and Electrode Mass Erosion and Surface Voltage Holdoff Recovery for a Transient, High 

Current Surface Discharges,” Proceedings of the 6th EML Conference, April 1991 (with T. G. 

Engel and M. Kristiansen). 

(2) “Electrode Performance of a Three Electrode Triggered High Energy Spark Gap Switch,” 

Proceedings of the 9th IEEE International Pulsed Power Conference, Albuquerque, NM, June 21-

23, pp 471-474 (with T. Engel and M. Kristiansen). 

(3) “Characterization of a Novel Single-Gap Transformer Coupled L-C Generator,” Proceedings of the 

20th IEEE Power Modulator Symposium, Myrtle Beach, SC, June 23-25, 1992, p. 293 (with T. 

Engel and M. Kristiansen). 

(4) “Communications Impact of Hall Plasma Thrusters,” AIAA-95-2929, 31st AIAA Joint Propulsion 

Conference. San Diego CA, July 10-12, 1995, 95-2929 (with J. Mankowski, E. O’Hair, and M. 

Kristiansen). 

(5) “Instrumentation and Data Acquisition of the Hall Thruster,” AIAA-95-2930, 31st AIAA Joint 

Propulsion Conference. San Diego CA, July 10-12, 1995, 95-2929 (with J. Mankowski, E. O’Hair, 

and M. Kristiansen). 

(6) “Plume Model of Hall Effect Plasma Thrusters with Temporal Consideration,” XXIV International 

Electric Propulsion Conference, Moscow, Russia, September 19-23, 1995 (with E. O’Hair, and M. 

Kristiansen). 

(7) “Compact Modulator Using Inductive Energy Storage and a Solid State Opening Switch,” 

Proceedings of the 22nd International Power Modulator Symposium, Boca Raton, Florida, June 

1996 (with J. Bridges and M. Kristiansen). 
NASA/TM—2012-217014 538 

(8) “Breakdown at Window Interfaces Caused by High-Power Microwave Fields,” Intense Microwave 

Pulses IV, SPIE Proceedings, Volume 2843, Denver, Colorado, August 1996 (with J. Elliott, L.L. 

Hatfield, M. Kristiansen, and H. Krompholz). 

(9) “Window and Cavity Breakdown Caused by High Power Microwaves,” Presented at the 24th IEEE 

International Conference on Plasma Science, San Diego, CA, May 1997 (with A. Neuber, D. 

Hemmert, H. Krompholz, L.L. Hatfield and M. Kristiansen). 

(10) “High Voltage Subnanosecond Dielectric Breakdown,” Proceedings of the International Workshop 

on High Power Microwave Generation and Pulse Shortening, Edinburgh, Scotland, June 1997 (with 

J. Mankowski and M. Kristiansen). 

(11) “Breakdown at Window Interfaces Caused by High-Power Microwave Fields,” Proceedings of the 

International Workshop on High Power Microwave Generation and Pulse Shortening, Edinburgh, 

Scotland, June 1997 (with A. Neuber, D. Hemmert, L.L. Hatfield, M. Kristiansen, and H. 

Krompholz). 

(12) "Evaluation of a Russian SOS Diode for Use in a Compact Modulator System," Proceedings of the 

11th IEEE International Pulsed Power Conference, Baltimore, Maryland, June 29 - July 02, 1997 

(with M. Kristiansen, M. Giesselmann and J.G. Kim). 

(13) “Window and Cavity Breakdown Caused by High-Power Microwave Fields,” Proceedings of the 

11th IEEE International Pulsed Power Conference, Baltimore, Maryland, June 29 - July 02, 1997 

(with A. Neuber, D. Hemmert, L.L. Hatfield, M. Kristiansen, and H. Krompholz). 

(14) “Streamer Discharges by Pulsed Power on a Spiral Transmission Line,” Proceedings of the 11th 

IEEE International Pulsed Power Conference, Baltimore, Maryland, June 29 – July 02, 1997 (with 

H. Akiyama, T. Sueda, S. Tsukamoto, Y. Nishihashi, S. Katsuki, M. Hagler and N. Inoue). 

(15) “A Subnanosecond High Voltage Pulser for the Investigation of Dielectric Breakdown,” 

Proceedings of the 11th IEEE International Pulsed Power Conference, Baltimore, Maryland, June 

29 - July 02, 1997 (with J. Mankowski and M. Kristiansen). 

(16) “Development of Improved Triggered Vacuum Switches,” Submitted to International Workshop on 

Electromagnetic Launchers, Edinburgh, Scotland, 1998 (with T. Warren, M. Kristiansen, et.al.). 

(17) “High-Power Microwave Generation by Coaxial Vircator,” Proc. 12th IEEE International Pulsed 

Power Conference, Monterey, CA, June 27-30, 1999 (with W. Jiang, K. Woolverton, M. 

Kristiansen). 

(18) “Energy Efficiency Analysis of an Inductive Storage System,” Proc. 12th IEEE International Pulsed 

Power Conference, Monterey, CA, June 27-30, 1999 (with J.G. Kim, J Zhang, M. Giesselmann, M. 

Kristiansen, and J. Mankowski). 

(19) “The Design of a Compact Pulse Transformer,” Proc. 12th IEEE International Pulsed Power 

Conference, Monterey, CA, June 27-30, 1999 (with J. Zhang, M. Kristiansen, M. Giesselmann, J. 

Mankowski, D. Castro, and D. Garcia). 

(20) “Pulsed Power Conditioning with a Transformer for an Inductive Energy Storage System,” Proc. 

12th IEEE International Pulsed Power Conference, Monterey, CA, June 27-30, 1999 (with J. 

Zhang, M. Giesselmann, M. Kristiansen, J. Mankowski, D. Castro, and D. Garcia). 

(21) “Pulsed Power Generation Using Ferromagnetic Circuits,” Proc. 12th IEEE International Pulsed 

Power Conference, Monterey, CA, June 27-30, 1999 (with S.I. Shkuratov, M. Kristiansen, L.L. 

Hatfield, and R. Martin). 

(22) “Pulsed, High Energy Testing of Resistors,” Proc. 12th IEEE International Pulsed Power 

Conference, Monterey, CA, June 27-30, 1999 (with S.I. Shkuratov, M. Kristiansen, L.L. Hat-field, 

and E. Horrocks). 

(23) High Voltage Subnanosecond Corona Inception,” Proc. 12th IEEE International Pulsed Power 

Conference, Monterey, CA, June 27-30, 1999 (with J. Mankowski, M. Kristiansen, J. Lehr, W. 

Prather, and J. Gaudet). 

(24) Optical Diagnostics on Helical Flux Compression Generators,” Proc. 12th IEEE International 

Pulsed Power Conference, Monterey, CA, June 27-30, 1999 (with A. Neuber, M. Kristiansen and 

H. Krompholz). 
NASA/TM—2012-217014 539 

(25) “Simulation, Design and Test of a MOV Pulse Shaping Device for High Power Microwave 

Generators,” Proc. 12th IEEE International Pulsed Power Conference, Monterey, CA, June 27-30, 

1999 (with M. Giesselmann, M. Kristiansen, and E. Kristiansen). 

(26) “High Power Microwave Window Breakdown Under Vacuum and Atmospheric Conditions”, 

SPIE’s International Symposium on AeroSense, (Intense Microwave Pulsed VII”, 24-28, April 

2000, Orlando, FL (with D. Hemmert, A. Neuber, M. Kristiansen, H. Krompholz, L.L. Hatfield). 

(27) “Fundamental Studies of a Helical Magnetic Flux Compression Generator,” 13th International 

Conference on High-Power Particle Beams, Nagaoka, Japan, June 25-30, 2000. (with M. 

Kristiansen and A. Neuber). 

(28) “Energy Balance of Shock Wave Ferromagnetic Generators,” Proceedings of 12th Symposium on 

High Current Electronics, Tomsk, Russia, Sept. 24-29, 2000 (with S.I. Shkuratov and M. 

Kristiansen). 

(29) “Dielectric/Gas Interface Breakdown Caused by High Power Microwaves,” 13th International 

Conference on High-Power Particle Beams, June 2000, Nagaoka, Japan (with D. Hemmert, A. 

Neuber, H. Krompholz, M. Kristiansen and L.L. Hatfield). (Invited). 

(30) “Helical Flux Compression Generator for Basic Research,” Proc. 12th Sympoisum on High Current 

Electronics, Tomsk, Russia, Sept. 24-29, 2000 (with A. Neuber, M. Kristiansen, M. Giesselman, B. 

Freeman, P. Worsey, and H. Krompholz). 

(31) “3-D PIC Simulation of a Coaxial Vircator,” 13th International Conference on High-Power Particle 

Beams, June 2000, Nagaoka, Japan (with W. Jiang (Nagaoka University of Technology), and M. 

Kristiansen. 

(32) “High-Power Microwave Generation by a Coaxial Vircator,” 13th International Conference on 

High-Power Particle Beams, June 2000, Nagaoka, Japan (with W. Jiang (Nagaoka University of 

Technology), and M. Kristiansen). 

(33) “Pulsed Power Systems”, Encyclopedia of Physical Science and Technology, Third Edition edited 

by Robert A. Meyers, Vol. 13 (2001), Academic Press, Inc., San Diego, CA (with M. Kristiansen). 

(34) “A Completely Explosive Pulsed Power Mini-System”, Submitted for publication in the Review of 

Scientific Instruments, 2002, RSI MS # A02449 (with E.F. Talantsev, S.I. Shkura-tov, and M. 

Kristiansen. 

(35) “Autonomous Ultra-Compact Explosive-Driven High-Voltage Generator Based on a Trans-verse 

Shock Wave Demagnetization of Nd2Fe14B Hard Ferromagnetics”, Submitted for publication in 

the Review of Scientific Instruments, 2002, RSI MS # A02200 (with S.I. Shkura-tov, E.F. 

Talantsev, and M. Kristiansen. 

(36) “Pre-Breakdown Current Behavior In Dc Volume Breakdown In Transformer Oil,” Proceedings of 

the 14th IEEE Pulsed Power Conference, Dallas, TX, June 15-18, 2003, p. 289-292. (with M. 

Butcher, A. Neuber, and H. Krompholz). 

(37) “Optical Diagnostics Of Liquid Nitrogen Volume Pre-Breakdown Events,” Proceedings of the 14th 

IEEE Pulsed Power Conference, Dallas, TX, June 15-18, 2003, p. 1029-1032. (with M. Butcher, A. 

Neuber, and H. Krompholz). 

(38) “Pulsed And Self Electrical Breakdown In Biodegradable Oil,” Proceedings of the 14th IEEE 

Pulsed Power Conference, Dallas, TX, June 15-18, 2003, p. 1036-1039. (with M. Cevallos, A. 

Neuber, M. Haustein, and H. Krompholz). 

(39) “Physical Efficiency Limits Of Inch-Sized Helical MFCG's,” Proceedings of the 14th IEEE Pulsed 

Power Conference, Dallas, TX, June 15-18, 2003, p. 413-416. (with A. Neuber, T. Holt, J. 

Hernandez, and M. Kristiansen). 

(40) “Small Sized MFCG For Driving A High Impedance Load,” Proceedings of the 14th IEEE Pulsed 

Power Conference, Dallas, TX, June 15-18, 2003, p. 1065-1068. (J. Hernandez, A. Neuber and M. 

Kristiansen). 

(41) “Ferromagnetic and Ferroelectric Materials as Seed Sources for Magnetic Flux Compressors,” 

Proceedings of the 14th IEEE Pulsed Power Conference, Dallas, TX, June 15-18, 2003, 1069-1072. 

( with N. Schoeneberg, J. Walter, A. Neuber, and M. Kristiansen). 
NASA/TM—2012-217014 540 

(42) “Design Criteria For Prevention Of Armature “Turn-Skipping” In Helical Magnetic Flux 

Compression Generators,” Proceedings of the 14th IEEE Pulsed Power Conference, Dallas, TX, 

June 15-18, 2003, p. 1077-1080. (with J. Rasty, X. Lee, A. Neuber, and M. Kristiansen). 

(43) “Conductivity Measurements Of Explosively Shocked Aluminum And OFHC Copper Used For 

Armature Material In A Magnetic Flux Compression Generator,” Proceedings of the 14th IEEE 

Pulsed Power Conference, Dallas, TX, June 15-18, 2003, p. 1073-1076. (with D. Hemmert, J. J. 

Mankowski, A. Neuber, J. Rasty, and M. Kristiansen). 

(44) “Surface Flashover Across Ceramic Disks In Vacuum At Cryogenic Temperatures,” Proceedings of 

the 14th IEEE Pulsed Power Conference, Dallas, TX, June 15-18, 2003, p. 305-308. (with H. 

Keene, A. Neuber, and H. Krompholz). 

(45) Physics Of Dielectric Surface Flashover At Atmospheric Pressure,” Proceedings of the 14th IEEE 

Pulsed Power Conference, Dallas, TX, June 15-18, 2003, p. 285-283. (with J. Krile, A. Neuber, J. 

Dickens, and H. Krompholz). 

(46) “Nanosecond Laser-Triggered Microwave Switch,” Proceedings of the 14th IEEE Pulsed Power 

Conference, Dallas, TX, June 15-18, 2003, p. 309-312. (with M. Mcquage and A. Neuber). 

(47) “Short Pulse Electric Field Sterilization of Liquid Media,” Proceedings of the 14th IEEE Pulsed 

Power Conference, Dallas, TX, June 15-18, 2003, p. 1124-1127. (with D. Wetz, K. Truman, J.J. 

Mankowski, and A. Neuber). 

(48) “Electrical Breakdown in Transformer Oil,” Joint Fall Meeting of the Texas Sections of the APS 

and AAPT and Zone 13 Society of Physics Students, Lubbock, TX, Oct 2003 (abstract only 

published) (with M. Cevallos, A. Neuber, and H. Krompholz). 

(49) G. Edmiston, A. Neuber, H. Krompholz, J. Dickens, J. Krile, "High Power Microwave Surface 

Flashover of a Gas-Dielectric Interface at 90 to 760 Torr," Proceedings of the 15th Int. IEEE Pulsed 

Power Conference, Monterey, CA, June 13-17, 2005. 

(50) J. Krile, A. Neuber, J. Dickens, H. Krompholz, G. Edmiston, “Similarities of Dielectric Surface 

Flashover at Atmospheric Conditions for Pulsed Unipolar and RF Excitation,” Proceedings of the 

15th Int. IEEE Pulsed Power Conference, Monterey, CA, June 13-17, 2005. 

(51) H. Krompholz, L.L. Hatfield, A. Neuber, D. Hemmert, K. Kohl, J. Chaparro,“ Subnanosecond 

Breakdown in Argon at High Overvoltages,” Proceedings of the 15th Int. IEEE Pulsed Power 

Conference, Monterey, CA, June 13-17, 2005. 

(52) J. Qian, R.P. Joshi, J.F. Kolb, K.H. Schoenbach, J. Dickens, A. Neuber, M. Butcher, M. Cevallos, 

H. Krompholz, E. Schamiloglu, J. Gaudet, “Simulation Studies of Liquid Water Breakdown by a 

Sub-Microsecond Pulse,” Proceedings of the 15th Int. IEEE Pulsed Power Conference, Monterey, 

CA, June 13-17, 2005. 

(53) H. Veselka, A. Neuber, J. Dickens, “Shock Induced Conductivity for High Power Switching,” 

Proceedings of the 15th IEEE Pulsed Power Conference, Monterey, CA, June 13-17, 2005. 

(54) M. Butcher, M. Cevallos, A. Neuber, H. Krompholz, J. Dickens, “Investigation of Charge 

Conduction and Self-Breakdown in Transformer Oil,” Proceedings of the 15th Int. IEEE Pulsed 

Power Conference, Monterey, CA, June 13-17, 2005. 

(55) A.A. Neuber, Y.J. Chen, J.C. Dickens, M. Kristiansen, “A Compact, Repetitive, 500 kV, 500 J, 

Marx Generator,” Proceedings of the 15th Int. IEEE Pulsed Power Conference, Monterey, CA, June 

13-17, 2005. 

(56) Y.J. Chen, J.J. Mankowski, A. Neuber, J.C. Dickens, “A Low-Cost Metallic Cathode for a Vircator 

HPM Source,” to be published in Proceedings of the 15th Int. IEEE Pulsed Power Conference, 

Monterey, CA, June 13-17, 2005. 

(57) D. Belt, J. Dickens, J. Mankowski, A. Neuber, “Multistage Helical Flux Compression Generator 

Non-Explosive Test Bed,” Proceedings of the 15th Int. IEEE Pulsed Power Conference, Monterey, 

CA, June 13-17, 2005. 

(58) M.D. Cevallos, M.D. Butcher, J.C. Dickens, A.A. Neuber, H.G. Krompholz, “Bubble Dynamics 

and Channel Formation for Cathode Initiated Discharges in Transformer Oil,” Proceedings of the 

15th Int. IEEE Pulsed Power Conference, Monterey, CA, June 13-17, 2005. 
NASA/TM—2012-217014 541 

(59) M.D. Cevallos, M.D. Butcher, J.C. Dickens, A.A. Neuber, H.G. Krompholz, “Composite 

Shadowgraphy and Luminosity Images of Self Breakdown Discharge Channels in Transformer 

Oil,” Proceedings of the 15th Int. IEEE Pulsed Power Conference, Monterey, CA, June 13-17, 

2005. 

(60) A. Neuber, J. Krile, G. Edmiston, H. Krompholz, J. Dickens, "Limits of High Power Microwave 

Transmission due to Interface Breakdown," presented (oral) at 2005 Tri-Service VED Workshop, 

12-16 September 2005. 

(61) A. Neuber, G. Edmiston, J. Krile, K. Morales, J. Dickens, H. Krompholz, "High Power Microwave 

Breakdown Limits of Dielectric/Gas Interfaces,” 2005 International COE Forum on Pulsed Power 

Science will be held on November 13 – 14 at Kumamoto, Japan. (invited). 

(62) A. Neuber, J. Krile, G. Edmiston, H. Krompholz, J. Dickens, M. Kristiansen ,“Interface Breakdown 

During High Power Microwave Transmission,” 13th EML Symposium, May 22-25, 2006, Berlin, 

Germany. 

(63) Lynn, C.; Neuber, A.; Dickens, J, “Opening Switch Utilizing Shock Wave Induced Conduction in 

PMMA and PVC,” Pulsed Power Plasma Science Conference, 2007. PPPS 2007. IEEE June 17 

2007-April 22 2007 Page(s):556 - 556 Digital Object Identifier 10.1109/PPPS.2007.4345862. 

(64) Karhi, Ryan W.; Mankowski, John J.; Dickens, James C, “A Synchronous Free-Running Arc 

Distributed Energy Railgun,” Pulsed Power Plasma Science Conference, 2007. PPPS 2007. IEEE 

June 17 2007-April 22 2007 Page(s):1016 - 1016 Digital Object Identifier 

10.1109/PPPS.2007.4346322. 

(65) Chen, Hao; Jiang, Chunqi; Kuthi, Andras; Gundersen, Martin A.; Dickens, James, ”Small BackLighted Thyratrons,” Pulsed Power Plasma Science Conference, 2007. PPPS 2007. IEEE June 17 

2007-April 22 2007 Page(s):430 - 430 Digital Object Identifier 10.1109/PPPS.2007.4345736. 

(66) James, Colt; Dickens, James C.; Holt, Shad, “Evaluation of Switch Jitter on a High Pressure 

Coaxial Spark Gap,” Pulsed Power Plasma Science Conference, 2007. PPPS 2007. IEEE June 17 

2007-April 22 2007 Page(s):429 - 429 Digital Object Identifier 10.1109/PPPS.2007.4345735. 

(67) McCauley, D.; Belt, D.; Mankowski, J.; Dickens, J.; Neuber, A.; Kristiansen, M, “ElectroExplosive Fuse Optimization for Helical Flux Compression Generator using a Non-Explosive Test 

Bed,” Pulsed Power Plasma Science Conference, 2007. PPPS 2007. IEEE June 17 2007-April 22 

2007 Page(s):656 - 656 Digital Object Identifier 10.1109/PPPS.2007.4345962 . 

(68) Belt, D.; Walter, J.; Mankowski, J.; Dickens, J,”Modeling of a Single Element Pulsed Ring-Down 

Antenna for Implementation in a Phased Array System,” Pulsed Power Plasma Science Conference, 

2007. PPPS 2007. IEEE June 17 2007-April 22 2007 Page(s):816 - 816 Digital Object Identifier 

10.1109/PPPS.2007.4346122. 

(69) Walter, J.; Dickens, J.; Mankowski, J.; Kristiansen, M,”Theoretical Pulsed Ring Down Antenna 

Array Performance” Pulsed Power Plasma Science Conference, 2007. PPPS 2007. IEEE June 17 

2007-April 22 2007 Page(s):482 - 482 Digital Object Identifier 10.1109/PPPS.2007.4345788. 

(70) Chen, Yeong-Jer; Mankowski, John J.; Walter, John W.; Dickens, James C “Jitter and Recovery 

Rate of a Triggered Spark Gap with High Pressure Gas Mixtures,” Pulsed Power Plasma Science 

Conference, 2007. PPPS 2007. IEEE June 17 2007-April 22 2007 Page(s):255 - 255 Digital Object 

Identifier 10.1109/PPPS.2007.4345561. 

(71) Walter, J.; Dickens, J.; Kristiansen, M,“Electrical and Optical Measurements of Explosively Driven 

Plasma Jets,” Pulsed Power Plasma Science Conference, 2007. PPPS 2007. IEEE June 17 2007-

April 22 2007 Page(s):723 - 723 Digital Object Identifier 10.1109/PPPS.2007.4346029. 

(72) Holt, Shad L.; Krile, John T.; Hemmert, David J.; Hackenberger, Wesley S.; Alberta, Edward F.; 

Walter, John W.; Dickens, James C.; Altgilbers, Larry L.; Stults, Allen H “Testing of New 

Ferroelectric Elements Custom Engineered for Explosively Driven Ferroelectric Generator 

Applications,” Pulsed Power Plasma Science Conference, 2007. PPPS 2007. IEEE June 17 2007-

April 22 2007 Page(s):722 - 722 Digital Object Identifier 10.1109/PPPS.2007.4346028. 

(73) Krile, John T.; Holt, Shad L.; Hemmert, David J.; Walter, John W.; Dickens, James C.; Altgilbers, 

Larry L.; Stults, Allen H “Development of an Ultra-Compact Explosively Driven Magnetic Flux 
NASA/TM—2012-217014 542 

Compression Generator System,” Pulsed Power Plasma Science Conference, 2007. PPPS 2007. 

IEEE June 17 2007-April 22 2007 Page(s):721 - 721 Digital Object Identifier 

10.1109/PPPS.2007.4346027. 

(74) Veselka, H.; Neuber, A.; Dickens, J “Pressure Induced Conductivity for High Power Switching,” 

Power Modulator Symposium, 2006. Conference Record of the 2006 Twenty-Seventh International 

14-18 May 2006 Page(s):219 - 219 Digital Object Identifier 10.1109/MODSYM.2006.365221. 

(75) Opening Switch Utilizing Stress Induced Conduction in Polymethylmethacrylate Lynn, C.; Krile, 

J.; Neuber, A.; Dickens, J.; IEEE International Power Modulators and High Voltage Conference, 

Proceedings of the 2008 27-31 May 2008 Page(s):483 - 486 Digital Object Identifier 

10.1109/IPMC.2008.4743696. 

(76) A 50 kV, 100 Hz Low Jitter Triggered Spark Gap with High Pressure Gas Mixtures Chen, YeongJer; Mankowski, John J.; Dickens, James C.; Walter, John; Kristiansen, Magne; IEEE International 

Power Modulators and High Voltage Conference, Proceedings of the 2008 27-31 May 2008 

Page(s):197 - 200 Digital Object Identifier 10.1109/IPMC.2008.4743614. 

(77) Energy Deposition and Electromagnetic Compatibility Assessment of Electroexplosive Devices 

Parson, J.; Dickens, J.; Walter, J.; Neuber, A.; IEEE International Power Modulators and High 

Voltage Conference, Proceedings of the 2008 27-31 May 2008 Page(s):439 - 442 Digital Object 

Identifier 10.1109/IPMC.2008.4743684. 

(78) Analysis of Mesoband Single Element Pulsed Ring-Down Antennas for Implementation in 

Phased Array Systems Belt, D.; Mankowski, J.; Walter, J.; Dickens, J.; Kristiansen, M.; IEEE 

International Power Modulators and High Voltage Conference, Proceedings of the 2008 27-31 May 

2008 Page(s):152 - 155 Digital Object Identifier 0.1109/IPMC.2008.4743602. 

(79) A Compact, Self-Contained High Power Microwave Source Based on a Reflex-Triode Vircator and 

Explosively Driven Pulsed Power Young, A.; Holt, T.; Elsayed, M.; Walter, J.; Dickens, J.; Neuber, 

A.; Kristiansen, M.; Altgilbers, L.L; Stults, A.H.; IEEE International Power Modulators and High 

Voltage Conference, Proceedings of the 2008 27-31 May 2008 Page(s):147 - 150 Digital Object 

Identifier 10.1109/IPMC.2008.4743600. 

(80) Window Flashover Initiation under Pulsed Microwave Excitation Krile, J.; Edmiston, G.; Dickens, 

J.; Krompholz, H.; Neuber, A.; IEEE International Power Modulators and High Voltage 

Conference, Proceedings of the 2008 27-31 May 2008 Page(s):560 - 563 Digital Object Identifier 

10.1109/IPMC.2008.4743719. 

(81) Scaling and Improvement of Compact Explosively-Driven Ferroelectric Generators Bolyard, D.; 

Neuber, A.; Krile, J.; Walter, J.; Dickens, J.; Kristiansen, M.; IEEE International Power Modulators 

and High Voltage Conference, Proceedings of the 2008 27-31 May 2008 Page(s):49 - 52 Digital 

Object Identifier 10.1109/IPMC.2008.47. 

(82) Shock Wave Simulation of Ferrite-Filled Coaxial Nonlinear Transmission Lines Sullivan, W.; 

Dickens, J.; Kristiansen, M.; IEEE International Power Modulators and High Voltage Conference, 

Proceedings of the 2008 27-31 May 2008 Page(s):517 - 520 Digital Object Identifier 

10.1109/IPMC.2008.4743706. 

(83) High-Current Compact FCG Seed Source Implementing Solid State Switching Elsayed, M.; Holt, 

T.; Young, A.; Neuber, A.; Dickens, J.; Kristiansen, M.; Altgilbers, L.L.; Stults, A.H.; IEEE 

International Power Modulators and High Voltage Conference, Proceedings of the 2008 27-31 May 

2008 Page(s):25 - 28 Digital Object Identifier 10.1109/IPMC.2008.4743567. 

(84) Compact Silicon Carbide Switch For High Voltage Operation James, C.; Hettler, C.; Dickens, J.; 

Neuber, A.; IEEE International Power Modulators and High Voltage Conference, Proceedings of 

the 2008 27-31 May 2008 Page(s):17 - 20 Digital Object Identifier 10.1109/IPMC.2008.4743565. 

(85) Jitter and recovery rate of a 50 kV, 100 Hz triggered spark gap with high pressure gas mixtures 

Yeong-Jer Chen; Mankowski, J.J.; Walter, J.W.; Dickens, J.C.; Plasma Science, 2008. ICOPS 

2008. IEEE 35th International Conference on 15-19 June 2008 Page(s):1 - 1 Digital Object 

Identifier 10.1109/PLASMA.2008.4590809. 
NASA/TM—2012-217014 543 

(86) Jitter and recovery rate of a triggered spark gap with high pressure gas mixtures Chen, Yeong-Jer; 

Dickens, J.; Walter, J.; Kristiansen, M.; Pulsed Power Conference, 2009. PPC '09. IEEE June 28 

2009-July 2 2009 Page(s):244 - 249 Digital Object Identifier 10.1109/PPC.2009.5386309. 

(87) High electric field packaging of silicon carbide photoconductive switches Hettler, C.; James, C.; 

Dickens, J.; Pulsed Power Conference, 2009. PPC '09. IEEE June 28 2009-July 2 2009 Page(s):631 

- 634 Digital Object Identifier 10.1109/PPC.2009.5386380. 

(88) High voltage photoconductive switches using semi-insulating, vanadium doped 6H-SiC, James, C.; 

Hettler, C.; Dickens, J.; Pulsed Power Conference, 2009. PPC '09. IEEE June 28 2009-July 2 2009 

Page(s):283 - 286 Digital Object Identifier 10.1109/PPC.2009.5386303. 

(89) Stand-alone, FCG-driven High Power Microwave system Young, A.; Neuber, A.; Elsayed, M.; 

Walter, J.; Dickens, J.; Kristiansen, M.; Altgilbers, L.L.; Pulsed Power Conference, 2009. PPC '09. 

IEEE. 

(90) June 28 2009-July 2 2009 Page(s):292 - 296 Digital Object Identifier 10.1109/PPC.2009.5386301. 

(91) VUV emission from dielectric surface flashover at atmospheric pressure Rogers, G.; Neuber, A.; 

Laity, G.; Dickens, J.; Frank, K.; Schramm, T.; Pulsed Power Conference, 2009. PPC '09. IEEE 

June 28 2009-July 2 2009 Page(s):855 - 859 Digital Object Identifier 10.1109/PPC.2009.5386373. 

(92) Electrical conduction in select polymers under shock loading Lynn, C.; Neuber, A.; Krile, J.; 

Dickens, J.; Kristiansen, M.; Pulsed Power Conference, 2009. PPC '09. IEEE June 28 2009-July 2 

2009 Page(s):171 - 174 Digital Object Identifier 10.1109/PPC.2009.5386199. 

(93) Performance of a compact triode vircator and Marx generator system Walter, J.; Dickens, J.; 

Kristiansen, M.; Pulsed Power Conference, 2009. PPC '09. IEEE June 28 2009-July 2 2009 

Page(s):133 - 137 Digital Object Identifier 10.1109/PPC.2009.5386184. 

(94) Energy deposition assessment and electromagnetic evaluation of electroexplosive devices in a 

pulsed power environment Parson, J.; Dickens, J.; Walter, J.; Neuber, A.; Pulsed Power 

Conference, 2009. PPC '09. IEEE June 28 2009-July 2 2009 Page(s):892 - 896 Digital Object 

Identifier 10.1109/PPC.2009.5386262. 

(95) A compact 5kV battery-capacitor seed source with rapid capacitor charger Holt, S. L.; Dickens, J. 

C.; McKinney, J. L.; Kristiansen, M.; Pulsed Power Conference, 2009. PPC '09. IEEE June 28 

2009-July 2 2009 Page(s):897 - 901 Digital Object Identifier 10.1109/PPC.2009.5386260. 

(96) Optimizing power conditioning components for a Flux Compression Generator using a nonexplosive testing system Davis, C.; Neuber, A.; Young, A.; Walter, J.; Dickens, J.; Kristiansen, M.; 

Pulsed Power Conference, 2009. PPC '09. IEEE June 28 2009-July 2 2009 Page(s):951 - 955 

Digital Object Identifier 10.1109/PPC.2009.5386110. 

(97) Prediction of compact explosively-driven ferroelectric generator performance Bolyard, D.; Neuber, 

A.; Krile, J.; Dickens, J.; Kristiansen, M.; Pulsed Power Conference, 2009. PPC '09. IEEE June 28 

2009-July 2 2009 Page(s):167 - 170 Digital Object Identifier 10.1109/PPC.2009.5386196. 

(98) Low jitter triggered spark gap with high pressure gas mixtures and Kr85 Yeong-Jer Chen; Dickens, 

J.C.; Walter, J.W.; Kristiansen, M.; Plasma Science - Abstracts, 2009. ICOPS 2009. IEEE 

International Conference on 1-5 June 2009 Page(s):1 - 1 Digital Object Identifier 

10.1109/PLASMA.2009.5227620. 

(99) Vacuum ultraviolet spectroscopy of dielectric surface flashover at atmospheric pressure Laity, G.; 

Frank, K.; Rogers, G.; Kristiansen, M.; Dickens, J.; Neuber, A.; Schramm, T.; Plasma Science - 

Abstracts, 2009. ICOPS 2009. IEEE International Conference on 1-5 June 2009 Page(s):1 - 1 

Digital Object Identifier 10.1109/PLASMA.2009.5227565. 

(100) Spectral analysis of vacuum ultraviolet emission from pulsed atmospheric discharges, Laity, G.; 

Neuber, A.; Rogers, G.; Frank, K.; Hatfield, L.; Dickens, J.; Plasma Science, 2010 Abstracts IEEE 

International Conference on Digital Object Identifier: 10.1109/PLASMA.2010.5534416 Publication 

Year: 2010 , Page(s): 1 IEEE Conferences. 

 
NASA/TM—2012-217014 544 

Laurence Andrew Dobrot 

Executive Profile 

Over 25 years of successful leadership, program management and systems engineering excellence with a 

reputation for building teams and getting the job done. Experienced Air Force leader with assignments in 

both the aircraft and space fields. 

Highlights of Qualifications 

 Security Clearance: Top Secret/SCI (2007) Systems Engineering 

 Certified Defense Program Manager (Lvl 3) Risk Analysis 

 Team Building and Leadership Budget and Cost Control 

 Personnel Development Resource Management 

 Strategic Planning Public Speaking 

 Systems Planning, Research, Development and Engineering (SPRDE Lvl 3) 

Professional History 

Leadership/Team Work 

 Deputy Program Director for the $5B Airborne Laser weapon system, part of the Missile Defense 

Agency 

 Commanded a 200-person team of military, government civilians, support contractors and the 

1,000-member prime contracting team that delivered the largest and most complex airborne 

laser weapon system in the world. 

 The Airborne Laser demonstrated lethal capability in shooting down threat representative 

ballistic missiles. 

 Managed all elements of cost, schedule and system performance. 

 Briefed members of Congress, Senior Defense Staff and Combatant Commanders on program 

plan and progress. 

 Led a 100-member Military Space Plane Integrated Concept Team for Air Force Space Command. 

 Worked closely with NASA and the Air Force Research Labs to define missions, requirements, 

concepts, funding and technologies for the Air Force’s reusable space architecture. 

 The Air Force recently launched an X-37B Orbital Test Vehicle, which is a direct result of 

the team’s development work. 

 Air Force’s Protection Functional Capabilities Boards working group lead. 

 Managed the Protection portfolio and guided the requirement documents through the Joint 

Capabilities Integration System process on their way to the Joint Requirements Oversight 

Council. 

Program Management 

 Managed the Engineering/Manufacturing Development contract for the Stage I rocket motor and 

flight termination system of the Small Inter-Continental Ballistic Missile. 

 Directed the 40-person team that managed the cost, schedule and technical performance. 

 Developed and administered the Air Force’s critical infrastructure program. 

 Authored the Air Force policy for identification and management of critical infrastructure. 

Represented the Air Force at national level Homeland Defense forums and guided 

Department of Defense senior staff in the development of policy. 

 Oversaw the entire military education program for the Afghan Ministry of Defense and the 

Afghan National Army while stationed in Afghanistan. 

 Created from the ground up, a native language computer training center for the Afghan 

Ministry of Defense, tapped into the local underemployed youth to teach. 
NASA/TM—2012-217014 545 

Systems Engineering 

 Mechanical/structural engineer responsible for all aspects of design, repair, modification and 

support of both mechanical and structural systems for the C-5 fleet. 

 Designed and installed repairs for C-5 main frame forgings cracked and damaged during 

Desert Storm, and instituted them into a standard Engineer's Repair Manual for the C-5 fleet. 

 Chief investigating engineer on a Titan-IVB and three C-5 aircraft Class A mishaps. 

 Determined most the probable cause for each event and engineered solutions to ensure these 

accidents would not reoccur. 

 Designed and coordinated the organic modification of a T-43 navigational trainer into the 

European Team travel aircraft, at less than 10% of contractor's proposed costs. 

 Defined and documented requirements for the Evolved Expendable Launch Vehicle. 

 Clearly communicated and documented the needs of the customer in a language that the 

developers and manufacturers could build and test to. 

 Built and staffed requirement documents that were approved by the Joint Requirements 

Oversight Council. 

Problem Solving 

 Masterminded a strategic plan that transitioned three major programs, 850 personnel and closed 

down a 40 year organization. 

 Designed a set of modifications that eliminated fire propagation between the engines and pylons 

on the C-5 aircraft. 

Education and Training 

 Master of Science in Systems Management, University of Southern California 

 Master of Strategic Studies, Army War College 

 Master of Arts in National Security and Strategic Studies With Distinction, Naval War College 

 Bachelor of Science in Mechanical Engineering, Michigan Technological University 

 
NASA/TM—2012-217014 546 

Richard L. Fork 

Optics Building 456 

University of Alabama in Huntsville 

Huntsville, Alabama 35899 

forkr@uah.edu 256–824–2523 256–824–6618 

Education 

Ph.D. Physics: Mass. Inst. of Tech. (MIT), Cambridge, Massachusetts, 1962 

Thesis title: “Role of Excited State Degeneracy in Dispersion Phenomena” 

B.S. Physics: (summa cum laude), Principia College, Elsah, IL, 1957 

Professional Career 

Professor of Electrical and Computer Engineering, UAH, Huntsville, Alabama, 8/1994 to present 

Professor of Physics, Rensselaer Polytechnic Institute, Troy, New York, 1990–1994 

Member Technical Staff, AT&T Bell Labs, Holmdel, New Jersey, 1970–1990 

Member Technical Staff, AT&T Bell Labs, Murray Hill, New Jersey, 1962–1970 

Research Assistant, Physics Department, MIT, 1957–1962 

Research Assistant, Naval Research Laboratory, Washington, D.C., Summer 1956 

Teaching Experience at Universities 

Have taught physics, electrical engineering, optics, and laser-related courses since 1994. Taught 

courses in undergraduate physics, photonics, and optics at RPI for 4 years. Graduated two Ph.D. 

students and one Masters Degree student at RPI. Graduated three Ph.D. students at UAH including 

the first black Ph.D. student in the optics program, currently supervising four Ph.D. students and one 

master’s student at UAH. Developed new advanced optics laboratory course for undergraduate and 

graduate students. Have teaching evaluations consistently higher than average. 

Professional Activities 

Co-Chaired 7th Annual Ultrafast Conference in Huntsville 16–17 Jan 2004. Chair of sessions at 

professional meetings, such as, Optical Society of America, Physics of Quantum Electronics Give 

papers at many laser oriented conferences, such as, Directed Energy Professional Society, American 

Physical Society, International Astronautical Conference, and many others. Average one or more 

invited talks at domestic and international meetings per year, give 5-10 contributed talks or briefings 

talks at professional meetings and NASA, DOD meetings each year. Gave invited talk at 90th 

birthday celebration for Nobel Laureate Willis Lamb. Review journal articles and proposals for 

publications such as Optics Express. Have served on multiple NSF, NSF-SBIR, DOD and other 

review panels. Recently served with a NASA MSFC led discussion group addressing orbital debris 

mitigation. Served recently on a NASA Glenn Research Center led study of beamed energy for 

propulsion. A current focus area is the distribution of energy in space, such as for beamed optical 

energy propulsion, horizontal lift to orbit and beamed optical energy for change of orbit in space. 

Have consulted for SMDC regarding use of lasers for defense applications including both sensing and 

missile defense. 

Research at University of Alabama in Huntsville 

Manage Laser Science and Engineering Laboratory at University of Alabama in Huntsville. Typical 

funding sources to date have been DIA, NSF and NASA. Theme of current work is use of lasers for 

distribution of energy in space for debris removal and beamed energy for propulsion. A current focus 

area is design of laser based energy infrastructure for near Earth space intended to support orbit 

change, lift to orbit, and general distribution of energy in space by optical means. 
NASA/TM—2012-217014 547 

Research at AT&T Bell Laboratories 

Twenty seven years as a Member of Technical Staff at AT&T Bell Laboratories in the 

Communication Sciences Division. Efforts typically addressed the physics and engineering issues of 

time resolved optical phenomena using lasers with an emphasis on solid state materials and devices. 

Much of the work resulted in key original laser related inventions that played major roles in 

advancing capabilities multiple aspects of ultrashort pulse lasers. Demonstration of the first 

modelocked laser, invention of the colliding pulse modelocked laser that produced the first optical 

pulses well into the femtosecond time domain, the first femtosecond optical amplifier, gigawatt 

amplification of laser pulses, the first use of integrated self phase modulation, group velocity 

dispersion, saturable gain, and saturable absorption to produce (at that time) world record 27 fsec 

duration optical pulses, femtosecond white light continuum generation, ultrafast spectroscopy of the 

dynamics of semiconductor materials including the first study of the dynamics of quantum well 

materials, first measurement of velocity overshoot in semiconductors, first direct time resolved 

measurements of: exciton screening, band edge renormalization, and band filling. A number of world 

records were set including the generation of the first 6 femtosecond pulse. That record stood for ten 

years. Produced first femtosecond movie in 1986. Movie showed the ablation of material by intense 

femtosecond pulses. This has been an extremely active field since then and is now emerging as 

relevant to novel means of propulsion in space. 

Honors and Awards 

Fellow, American Physical Society, 1964 

Fellow, Optical Society of America, 1995 

Laser Focus Invention of the Year 1983 "Colliding Pulse Laser" 

Most cited technical paper of 1964 in any journal (scanning interferometry) 

UAH Outstanding Engineering Professor, 1995 

World record shortest pulse, 6 fs, from 1987 to 1997 

Sample Papers (of more than 100 published papers) 

(1) "Locking of He-Ne Laser Modes Induced by Synchronous Intracavity Modulation", L. E. 

Hargrove, R. L. Fork, and M. A. Pollack, Applied Physics Letters 5, 4-5 (1964). Original 

discovery of modelocked lasers. 

(2) "A Scanning Spherical Mirror Interferometer for Spectral Analysis of Laser Radiation", R. L. 

Fork, D. R. Herriott, and H. Kogelnik, Applied Optics 3, 1471-1484 (1964). Most cited paper in 

any journal in 1964. 

(3) "Real Time Autocorrelation Interferometer", R. L. Fork and F. A. Beisser, Applied Optics 17, 

3534-3535 (1978). First real time ultrashort pulse measurement capability. 

(4) "Generation of Optical Pulses Shorter Than 0. 1 Picoseconds by Colliding Pulse Modelocking", R. 

L. Fork, B. I. Greene, and C. V. Shank, Applied Physics Letters 38, 671-672 (1981). First optical 

pulse generation clearly in the femtosecond time regime. 

(5) "Amplification of 70 Femtosecond Optical Pulses to Gigawatt Powers", R.L. Fork, C. V. Shank, 

and R. T. Yen, Applied Physics Letters 41, 761-763 (1982). 

(6) "Femtosecond White Light Continuum Pulses", R. L. Fork, C. V. Shank, C. Hirlimann, and R. T. 

Yen, Optics Letters 8, 1-3 (1983). 

(7) "Femtosecond Optical Pulses", R. L. Fork, C. V. Shank, R. Yen, and C. Hirlimann, IEEE Journal 

of Quantum Electronics 19, 500-506 (1983). 

(8) "Negative Dispersion Using Pairs of Prisms", R. L. Fork, O. E. Martinez, and J. P. Gordon, Optics 

Letters 9, 150-152 (1984). First means of introducing negative dispersion, and hence dramatically 

improved intra-resonator ultrashort optical pulse generation. 
NASA/TM—2012-217014 548 

(9) "Femtosecond Imaging of Melting and Evaporation at a Photoexcited Silicon Surface", M. C. 

Downer, R. L. Fork, and C. V. Shank, Journal of the Optical Society of America B2, 595-599 

(1985). First femtosecond movies. 

(10) "Compression of Optical Pulses to 6 Femtoseconds by Using Cubic Phase Compensation", R. L. 

Fork, C. H. B. Cruz, P. C. Becker, and C. V. Shank, Optics Letters 12, 483-485 (1987). World 

record for ultrashort pulses which stood for ten years. 

(11) “Harmonically mode-locked laser and applications,” R.L. Fork, K. Singh, J. Haus, R.K. Erdmann 

and S.T. Johns, SPIE Proceedings, Photonics at the Air Force Photonics Center, 2216 148-159 

(1994). 

(12) “Optical Amplifier for Space Applications”, Richard L. Fork, Spencer T. Cole, William M. Diffey, 

Lisa J. Gamble, and Andrew S. Keys. Optics Express 5, 292-301 (1999). 

(13) M.H. Smith, R.L. Fork, S.T. Cole, “Safe Delivery of optical power from space”, Optics Express 8, 

537-546 (2001). 

(14) R.L. Fork, R.L. Laycock, W.W. Walker, S.T. Cole, S.D. Moultrie, D.J. Phillips, and J.C. 

Reinhardt, “Surface High Energy Laser” (Invited Paper) Proceedings of the IEEE. 93, 1864-1873 

(2005). 

(15) Richard L. Fork, “Preventing Asteroid Earth Impacts with Laser Technology: Progress and Future 

Prospects”, (Invited paper) Proceedings of the IEEE, 95, No. 5, 847-8 (2007). 

(16) Dane J. Phillips, Rustin L. Laycock, Spencer T. Cole, Wesley W. Walker, Sean D. Moultrie, John 

C. Reinhardt, Richard L. Fork and Joe T. Howell, (Invited Paper) “Technology demonstrations and 

flight experiments validating an optical energy infrastructure for Earth–Moon space”, Acta 

Astronautica 62, 185-191 (2008). 

(17) R.L. Fork and R. L. Laycock “Solid state laser medium and laser medium heat transfer method”, 

US Patent 7,352,785, 2008. 

(18) Richard Fork, “Orbital Debris Mitigation Using Minimum Uncertainty Optical States”, 

Proceedings of the IEEE 97 951-3, June 2009. 

(19) Richard Fork, Luke Burgess, Stefani Boehme, Pat Reardon, David Pollock, Matt Wright, Nick 

Cote, Randy Gaillard, and Raymond Beach, “Surveillance and Mitigation of Orbital Debris: Laser 

Systems and Standards”, IAC-09-A6.5.9, 60th International Astronautical Congress, Daejeon, 

Republic of Korea, October 16, 2009. 

(20) Richard Fork, “Orbital Debris Mitigation Using Minimum Uncertainty Optical States”, 

Proceedings of the IEEE 97 951-3, June 2009. 

(21) Planetary Defense and an Energy Infrastructure for Near Earth Space”, Proceedings of the IEEE, 

99, 359-362 (2011). 
NASA/TM—2012-217014 549 

Dr. Jordin T. Kare 

Kare Technical Consulting 

908 15th Avenue East 

Seattle, Washington 98112 

jkare@jkare.com 206–805–9046 

Experience 

2007–Present, Founder and Chief Scientist, LaserMotive LLC 

Co-founded LaserMotive LLC (http://www.lasermotive.com) to develop and market laser power 

beaming systems for industrial and military applications. Technical lead for LaserMotive 

development of the prize-winning power beaming system and climber vehicle for the NASA 

Centennial Challenge competition for Power Beaming, November 2009. 

2007–Present, Scientific Program Manager and Senior Inventor, Intellectual Ventures LLC 

Research program director for Intellectual Ventures Laboratories, and staff inventor for Intellectual 

Ventures, an intellectual property development company based in Bellevue, WA. Information on 

Intellectual Ventures is available at http://www.intellectualventures.com. Invention activities include 

inventing over a broad range of technologies, reviewing and prioritizing invention concepts for 

potential patenting, and supporting patent drafting and application processes. Past activities (2007-

2008) included setting up Intellectual Ventures’ laboratory facilities for broad-spectrum research and 

development (http://www.intvenlab.com), hiring and managing laboratory and support personnel, and 

leading selected research projects. 

A notable project at Intellectual Ventures was the invention and initial development of the Photonic 

Fence, a system for detection, tracking, identification, and selective killing of insects, particularly 

disease-carrying mosquitoes, using digital imaging and laser technology. 

1997–Present, Consultant, Kare Technical Consulting, sole proprietor 

Consultant to aerospace companies and Federal laboratories on advanced space system architecture 

and design. Large customers include NASA (USRA), Northrop Grumman TASC, Boeing, General 

Atomics, and Lawrence Livermore National Laboratory. Small business customers include Tethers 

Unlimited, Andrews Space and Technology (superconducting cable design and evaluation for a novel 

deep-space propulsion system), ENSCO (global distributed microsensor network) and Proton Energy 

Systems (space applications of regenerative fuel cells). Particular areas of expertise include optical 

and radar imaging systems, and beamed power technology. 

6/00–10/03, Payload Engineering support for Boeing Resource21 program 

Responsible for developing requirements and technical interface to subcontractors for a $100M-class 

multispectral sensor payload. Tasking included review and analysis of subcontractor’s visible/ 

NIR/SWIR FPA design and I&T program and preliminary design of thermal IR sensor. Key member 

of the Boeing team on CONESTOGA and CONESTOGA II, unclassified NRO-sponsored projects to 

develop concepts, architectures, and technology roadmaps for U.S. Intelligence community 

information systems in the 2005–2025 timeframe. 

Two NASA Institute for Advanced Concepts (http://www.niac.usra.edu) Phase I Fellowship awards: 

modular laser launch architecture (2003–2004) and SailBeam, an innovative and technically-feasible 

propulsion concept for relativistic interstellar probes (2001). 
NASA/TM—2012-217014 550 

1997, Chief Scientist, RDL Space Corp., San Ramon, California 

 Senior Research Scientist, RDL Inc., Culver City, California 

As Sr. Research Scientist for RDL Inc., developed system architecture and sensor payload concept for 

major next-generation space system for a classified customer. Also a key participant in classified 

advanced space system concept study conducted by a major aerospace company, and in preparation of 

successful proposals for NASA LightSAR program and multiple classified programs. As Chief 

Scientist for fledgling RDL Space Corp., developed payload concepts for commercial space-based 

SAR imaging system and participated in development of business plan and DoC remote-sensing 

license application. 

1995–1996, Physicist, V-Division Space Group, Physics and Space Technology Directorate, 

Lawrence Livermore National Laboratory (LLNL), Livermore, California 

Lead designer for Tactical Imaging Constellation Architecture Studies; responsible for system 

concepts and point designs for military electro-optic and synthetic-aperture radar satellites and ground 

systems, including a 4-meter unfilled-aperture optical imaging satellite. Also developed advanced 

concepts in diverse areas including biological weapon countermeasures, laser power beaming, and 

reconnaissance systems. 

1991–1995, Physicist, Special Studies Program (O-Group), LLNL 

Developed advanced concepts and technologies for spacecraft and launch vehicles. Program leader 

for MOCKINGBIRD miniature reusable launch vehicle. LLNL interface to NASA/JPL for Pluto Fast 

Flyby mission development. Mission planner for Clementine lunar mapping mission. 

Assigned to DOE Office of Space, Washington DC in 1993–1994, as Technical Advisor to the 

Director in the area of Remote Sensing. Responsible for coordinating joint DOE/NASA projects in 

space-based remote sensing. Assisted in formulating Office of Space policy and objectives. 

1987–1991, Program Leader, LLNL/SDIO Laser Propulsion Program 

Directed multiple university and industrial research teams in developing technology for ground-to-orbit 

laser propulsion (high-volume launch of small rocket vehicles powered by large ground-based lasers). 

1985–1986, Physicist, Special Studies Program (O-Group), LLNL 

Developed laser propulsion application concept; organized 1st SDIO/DARPA Laser Propulsion 

Workshop. Also worked on adaptive optics technology for atmospheric compensation of laser beams, 

developing designs for cooled deformable mirrors. 

1984–1985, Research Associate, Lawrence Berkeley Laboratory (LBL), Berkeley, California

Education 

1978–1984 University of California, Berkeley, CA 

 Ph.D. in Astrophysics, 1984 

 Fannie and John Hertz Foundation Fellow 

1974–1978 Massachusetts Institute of Technology (MIT), Cambridge, MA 

 B.S. Physics, 1978; B.S. Electrical Engineering, 1978 

 Member Phi Beta Kappa, Sigma Xi (Physics), Tau Beta Pi (Engineering), 

Eta Kappa Nu (Electrical Engineering) 

Other Information 

 Senior member AIAA, member IEEE, SPIE. OSA, Directed Energy Professional Society 

 Approximately 40 publications; 12 issued patents 

 Inventor or co-inventor on over 350 published patent applications; currently #21 in Top 100 

Inventors, based on applications published in preceding 3 years 

 Partner in small publishing business, 1981–1987
NASA/TM—2012-217014 551 

Dr. Andrew Ketsdever 

Air Force Research Laboratory 

Propulsion Directorate 

Edwards AFB, CA 

Dr. Andrew Ketsdever is currently a Senior Research Engineer at the Air Force Research Laboratory's 

(AFRL) Propulsion Directorate at Edwards Air Force Base in California. While at AFRL, he was the 

group leader of the Advanced Concepts Group—a group responsible for identifying and developing the 

next generation of propulsion systems to enable future Air Force missions. He was also the group leader 

of the Non-Equilibrium Flows Group (AFRL/RZSA) which conducts basic research in rocket exhaust 

plume flows, signatures, and effects. AFRL/RZSA was recognized as an Air Force Office of Scientific 

Research (AFOSR) Star Team. He recently completed a Visiting Professorship in the Department of 

Astronautics at the United States Air Force Academy (2004-2007) and is also an Associate Professor at 

the University of Colorado at Colorado Springs. 

Dr. Ketsdever has worked in the areas of advanced propulsion design and testing, rarefied gas dynamics, 

microfluidics, micropropulsion and spacecraft-thruster interactions since starting at AFRL in 1992. He 

holds patents and has patents pending in the areas of micropropulsion, advanced propulsion concepts, and 

nano-ignition of propellants. Dr. Ketsdever received his Ph.D. from the University of Southern California 

(USC) in 1995 where he has taught undergraduate and graduate level courses in rarified gas dynamics, 

planetary atmospheres, microspacecraft design, and spacecraft-environment interactions. 

Dr. Ketsdever is an Associate Fellow of the American Institute of Aeronautics and Astronautics (AIAA). 

He is currently serving on the Nuclear and Future Flight Propulsion Technical Committee. Dr. Ketsdever 

has served on the AIAA Thermophysics Technical Committee (1998-2003) where he was the chairman of 

the Education Subcommittee. He has also been involved with the AIAA Working Group in Microfluidics. 

Dr. Ketsdever has authored or co-authored more than 100 technical papers. He also co-edited an AIAA 

Progress in Astronautics and Aeronautics series book entitled Micropropulsion for Small Spacecraft. Dr. 

Ketsdever is currently an associate editor for the Journal of Spacecraft and Rockets. 
NASA/TM—2012-217014 552 

Frank E. Little 

Education 

Doctor of Philosophy in Chemistry. Field: Thermodynamics and co-operative phenomena, 

University of California, Davis, 1976. 

Master of Science in Chemistry. Field: Theoretical Chemistry. California State University, 

Northridge, 1972. 

Bachelor of Science in Chemistry. California State University, Northridge, 1968. 

Fellowship 

ERDA–AWU Graduate Fellow, 1976. 

NASA Summer Faculty Fellow, Marshall Space Flight Center, 1999. 

Professional History 

TEES Distinguished Research Scientist, Space Engineering Research Center, August 2009 to 

present 

Associate Director, Center for Space Power, January 1993 to July 2009 

Associate Director, Center for Microencapsulation and Drug Delivery, May 2002 to present 

Associate Director for Commercialization, Center for Space Power, September 1990 to 

January 1993 

Business Manager, Center for Space Power, Texas A&M University, September 1989 to 

September 1990 

Assistant Director, Space Research Center, Texas A&M University, September 1986 to 

September 1990 

Project Manager, Thermodynamics, Gas Research Institute, August 1982 to September 1986. 

Staff Research Associate II, Crocker Nuclear Laboratory, University of California, Davis, 

October 1978 to August 1982. 

Lecturer in Chemistry, Department of Chemistry, University of California, Davis, October 

1976 to July 1978. 

ERDA–AWU Graduate Fellow, Crocker Nuclear Laboratory and the Department of 

Chemistry, University of California, Davis, January 1976 to September 1976. 

Research Assistant, Department of Chemistry, University of California, Davis, 1971 to 1975. 

Teaching and Research Assistant, Department of Chemistry, California State University, 

Northridge, 1968 to 1971. 

Consulting 

The Cyclotron Corporation, 1981. 

SAIC, 2002 

Technical Management Projects for Center for Space Power 

 Photovoltaic conversion, including thin film 

 Electrochemical energy storage (NiH, NiMH, Li-ion) 

 Mechanical energy storage (flywheel) 

 Power Management and Distribution (converters) 

 On-board propulsion (microwave electrothermal thruster) 
NASA/TM—2012-217014 553 

Research Activities 

 Wireless Energy Transport for Solar Power to Earth. 

 Electrochemistry 

 Closed Environment Life Support Systems for space. 

 Thermodynamics: Determination of molecular properties from measurements of fluid behavior. 

 Radioisotope production. 

 Development of applications for radioisotopes. 

 Measurement of cross-sections for nuclear reactions. 

 Synthesis of radiochemicals. 

 Chemical kinetics of gas phase reactions of atomic fluorine utilizing radiotracer methods. 

 Theoretical chemistry: Study of the liquid-vapor interface and model calculations for 

adsorption of a gas onto a solid. 

Technical Documents 

"A Statistical Mechanical Calculation of the Density Variation through a Liquid-Vapor Interface." 

Master's Thesis, California State University, Northridge, January 1972. 

"Studies of Corresponding States Potential Force Constant Scaling Relations and the Critical 

Constants for the Fluoroethanes." Doctoral dissertation, University of California, Davis, 

September 1976. 

Papers Presented at Scientific Meetings 

Little, F.E., and J.W. Root, "Critical Constants of Fluorinated Ethanes as a Basis for the Estimation of 

Molecular Sizes." Presented at the 169th National Meeting of the American Chemical Society, 

Philadelphia, 1975. 

Lagunas-Solar, M.C., F.E. Little and J.A. Jungerman, "The Radioisotope Program at U.C. Davis: 

Radionuclide Production and New Developments." Invited paper presented at the 1979 Winter 

Meeting of the American Nuclear Society, San Francisco. 

Lagunas-Solar, M.C., J.A. Jungerman, N.F. Peek, and F.E. Little, "A Remote Flow System for LargeScale High-Purity I-123 Production." Presented at the 1979 Winter Meeting of the American 

Nuclear Society, San Francisco. 

Lagunas-Solar, M.C., H.L. Thibeau, and F.E. Little, "A Remote System for Multicurie Radiochemical 

Separations." Presented at the 1979 Winter Meeting of the American Nuclear Society, San 

Francisco. 

Lagunas-Solar, M.C., F.E. Little, S.L. Waters, and J.A. Jungerman, "Cyclotron Production of Carrierfree Thallium-201 via the Pb-207 (p, 7n) Reaction." Presented at the 3rd International 

Symposium on Radiopharmaceutical Chemistry, St. Louis, Missouri, June 16-20, 1980. 

Jungerman, J.A., M.C. Lagunas-Solar, and F.E. Little, "Cyclotron Production of High-Purity Iodine123 for Medical Applications." Presented at the International Symposium on Radioiodines, 

Banff, Alberta, Canada, September 13-16, 1980. 

Little, F.E. and M. Klein, "A Rational Approach to Funding a Fluid Properties Basic Research 

Program for the Natural Gas Industry." Presented at the AIChE Spring National Meeting, 

Houston, March, 1985. 

Little, F.E. and M. Klein, "Keeping the Horse Before the Cart: The Importance of Data to Equation of 

State Development." Presented at 1st CODATA Symposium on Chemical Thermodynamic and 

Thermophysical Properties Data Bases of the CODATA International Meetings on Phase 

Equilibria and Related Property Data, Paris, September 5-13, 1985. 
NASA/TM—2012-217014 554 

Little, F.E. and M. Klein "Good Data is Forever: The Thermodynamics Program at GRI." Presented 

at 2nd CODATA Symposium on Critical Evaluation and Prediction of Phase Equilibria in 

Multi-component Systems of the CODATA International Meeting on Phase Equilibria and 

Related Property Data, Paris, September 5-13, 1985. 

Starling, K.E., M. Klein and F.E. Little, "The Thermodynamic Properties of Natural Gas- Physical 

Properties." Presented at 1st International Congress on "Gas Quality - Specifications and 

Measurement of Physical Properties of Natural Gas." Groningen, The Netherlands, April 21-25, 

1986. 

Holtzapple, Mark T., Frank E. Little, William Moses, Comer O. Patterson, and Merry Makela "A 

First Order Chemical Model for a Life Support System." Presented at the AIChE Spring 

National Meeting, New Orleans, March, 1988. 

Holtzapple, Mark, Frank E. Little, Merry Makela, C.O. Patterson and William Moses, "Analysis of an 

Algae-Based CELSS: Part I: Model Development, Part II: Options and Weight analysis." 

Presented at Controlled Environmental Life Support Systems Research Conference, Texas A&M 

University, February 22-23, 1988. 

Makela, Merry, Steve Perkins, A. Dale Whittaker, Heinz Preisig, Mark Holtzapple, and Frank Little, 

"Object-Oriented Model of a Closed Loop Life Support System.", Presented at Controlled 

Environmental Life Support Systems Research Conference, Texas A&M University, February 

22-23, 1988. 

Holtzapple, Mark T. and Frank E. Little, "Comparison of Waste Combustion and Waste Electrolysis: 

A System Approach", Presented at Life Support Systems Research Conference, Texas A&M 

University, February 14-15, 1989. 

Holtzapple, Mark T. and Frank E. Little, "Comparison of Waste Combustion and Waste Electrolysis 

Systems Analysis", Presented at Intersociety Conference on Environmental Systems, San Diego, 

July 1989. 

Preisig, H. A., Tae-Yeong Lee, Frank Little, and Bruce Wright, "On the Representation of LifeSupport System Models", Presented at Intersociety Conference on Environmental Systems, San 

Diego, July 1989. 

Preisig, Heinz. A., Tae-Yeong Lee and Frank Little "A Prototype Computer-Aided Modelling Tool 

for Life-Support System Models", Paper 901269, Presented at 20th Intersociety Conference on 

Environmental Systems, Williamsburg, VA, July 1990. 

Little, Frank E. and Mike O. Kennedy "Commercialization Strategies for Space Power Technology", 

Paper IAF--90--617 Presented at 41st International Astronautical Congress, Dresden, Germany, 

October 1990. 

Chang, Kai, Frank E. Little and Mike O. Kennedy "Microwave Power Transmission System: Space 

Flight Experiment Program", Paper IAF--90--216 Presented at 41st International Astronautical 

Congress, Dresden, Germany, October 1990. 

Chang, K., A.D. Patton, M.O. Kennedy, F.E. Little, M.A. Pollock, K.A. Hummer, J.C. McCleary, 

B.S. Wei, A.M. Brown, J.O. McSpadden “Demonstration of Microwave Power Transmission in 

Space”, Presented at SPS 91: Power from Space, Paris/Gif-sur-Yvette, August 27-30, 1991. 

Nansen, Ralph H. and Frank E. Little “A Practical Solar Power Satellite for the New Millennium”, 

Paper 95-301. Presented at the 30th Intersociety Energy Conversion Engineering Conference, 

Orlando, Florida, August 1995. 

Erb, R.B., N. Kaya, G. Maryniak, R. Leonard, M.B. Duke, F. Little, A.D. Patton, R. Nansen, J. Spies 

and W. Sadeh “International Cooperation for the Acquisition of Space-Based Energy”, Paper 

IAA-95-IAA.3.3.06. Presented at the 46th International Astronautical Congress, Oslo, Norway, 

October 1995. 
NASA/TM—2012-217014 555 

Little, Frank E. “Laser Wireless Power Transmission”, Presented at WPT’95 Second Wireless 

Power Transmission Conference, Kobe, Japan, October 16-19, 1995. 

McSpadden, James O., Frank E. Little, Michael B. Duke and Alex Ignatiev “An In-Space Wireless 

Energy Transmission Experiment”, Paper 96576. Presented at the 31st Intersociety Energy 

Conversion Engineering Conference, Washington, D.C. August 1996. 

Ehsani, M., J. Mahdavi, I. Pitel, J.E. Brandenberg, and F.E. Little “Development of an Efficient 

Power Supply for the Microwave Electrothermal Thruster”, Paper 970115 Presented at 2nd 

Conference on Commercial Development of Space, Albuquerque, New Mexico, January 1997. 

McSpadden, James O., Kai Chang, Mike Duke and Frank Little “Study of ISS Free-Flyer Power 

Beaming”, Presented at SPS ‘97 Conference, Montréal, Canada, August 24-28, 1997. 

Little, Frank .E., James O. McSpadden, Kai Chang and Nobuyuki Kaya “Toward Space Solar Power: 

Wireless Energy Transmission Experiments Past, Present and Future”, Presented at Space 

Technology and Applications International Forum (15th Symposium on Space Nuclear Power 

and Propulsion) Albuquerque, New Mexico, January 1998. 

Rakotondrainibe, Andre F., Judith A. Jeevarajan, A. John appleby and Frank E. Little, 

“Chronopotentiometric Profiles of the Anode and Cathode During the Charge/Discharge of 

Some commercial Lithium-ion Batteries”, Presented at 193rd ECS Meeting San Diego, 

California, May 3-8, 1998. 

Little, Frank E., “A wireless Power Transmission Power System for Microgravity Crystal Processing 

Satellites”, Presented at IEEE Aerospace Conference, Big Sky, Montana, March 18 – 25, 2000. 

Little, Frank E., “Solar Power Satellites: Recent Developments”, Paper 0965, Presented at the 

XXXVIIth URSI General Assembly, Maastricht, The Netherlands, August 17-24, 2002. 

Zepeda, Paola, Kai Chang and Frank Little “Optimal Antenna Taper Design for a Sandwich 

Transmitting Array in Space Solar Power Satellite” Paper IAC-02-R.2.10, Presented at the 53rd 

International Astronautical Congress, The World Space Congress, Houston, October 10-19, 

2002. 

L.H. Hsieh, B.H. Strassner, S.J. Kokel, C.T. Rodenbeck, M.Y. Li, K. Chang, F.E. Little, G.D. Arndt, 

P.H. Ngo “DEVELOPMENT OF A RETRODIRECTIVE WIRELESS MICROWAVE POWER 

TRANSMISSION SYSTEM”, Presented at 2003 IEEE International Symposium on Antennas 

and Propagation and USNC/CNC/URSI North American Radio Science Meeting, Columbus, 

Ohio, June 22-27, 2003. 

Little, Frank “Power from Space: Fulfilling the Promise” Presented at The 1st International 

Symposium on Sustainable Energy System, Kyoto, Japan, March 13-14, 2003. 

Little, F.E., S.J. Kokel, C.T. Rodenbeck, K. Chang, G.D. Arndt, P.H. Ngo, “Development of 

Retrodirective Control Transmitter for Wireless Power Transmission” Presented at 2003 Japan – 

United States Joint Workshop on Space Solar Power System (JUSPS’03) Kyoto, Japan, July 3,4 

2003. 

Little, Frank E. “Power from Space: From Promise to Reality”, Presented at URSI 2004 International 

Symposium on electromagnetic Theory, Pisa, Italy, May 23-27, 2004. 

Little, Frank and Henry Brandhorst, “An Approach for Lunar power – 24/29” Presented at Solar 

Power from Space – SPS’04/5th Wireless Power Transmission conference (WPT 5), Granada, 

Spain, June 30 – July 2, 2004. 

Rodenbeck, C., S. Kokel, K. Chang, F.E. Little, “Microwave Wireless Power Transmission with 

Retrodirective Beam Steering”, Presented at 2nd International Energy Conversion Engineering 

Conference, Providence, Rhode Island, August 16-19, 2004. 

Brandhorst, Henry W. and Frank Little, “An Approach for Continuous Lunar Power”, Presented at 

2nd International Energy Conversion Engineering Conference, Providence, Rhode Island, 

August 16-19, 2004. 

Brandhorst, Henry W. and Frank Little, “POWOW Revisited – Beamed Power for Mars Exploration”, 

Presented at 2nd International Energy Conversion Engineering Conference, Providence, Rhode 

Island, August 16-19, 2004. 
NASA/TM—2012-217014 556 

Little, Frank E. and Alan Palazzolo, “Testing of Magnetic Bearings for Flywheel Energy Storage in 

Simulated Space Conditions”, Paper No. AIAA-2005-5642, Presented at 3rd International 

Energy Conversion Engineering Conference, San Francisco, August 15-18, 2005. 

Little, Frank E., “A Wireless Power Based Space Transportation System”, Paper No. 2006-g-10, 

Presented at 25th International Space Technology and Science Conference, Kanazawa, Japan, 

June 4-11, 2006. 

Schuller, M., Lalk, T., Wiseman, L., Little, F., Godard, O., Abdel-Fattah, S., Askew, R., Klaus, D., 

Kobrick, R., Thomas, G., Rouen, M. and Conger, B, “Innovative Schematic Concept Analysis 

for a Space Suit Portable Life Support Subsystem”, Paper 2006-01-2201, Presented at the 36th 

International Conference on Environmental Systems, Norfolk, Virginia, USA, July 17-20, 2006. 

Klaus, D., Bamsey, M., Schuller, M., Godard, O., Little, F. and Askew, R., “Defining Space Suit 

Operational Requirements for Lunar and Mars Missions and Assessing Alternative 

Architectures”, Paper 2006-241 Presented at the 36th International Conference on 

Environmental Systems, Norfolk, Virginia, USA, July 17-20, 2006. 

Little, Frank, “From the Moon to the Earth: Lunar Implications for Space Solar Power”. Presented at 

the First International Symposium on Radio Systems and Space Plasma, Sofia, Bulgaria, 

September 2-5, 2007. 

Little, Frank E., “Constraints on the Design of a Lunar-to-Earth Wireless Power Transmission 

System” Paper IAC-08-C3.1.8, Presented at the 59th International Astronautical Congress, 

Glasgow, Scotland, 29 September 29 - October 3, 2008. 

Little, Frank E., Kai Chang, Rainer Fink, G. Dickey Arndt, Phong H. Ngo and Raymond F. Beach, 

“A Space to Earth Demonstration of Wireless Power Transmission” Paper No. 2009-h-27 

Presented at the 27th International Space Technology and Science Conference, Tsukuba, Japan, 

July 5-12, 2009 

Little, Frank E., “Recent US Activities in Wireless Power Transmission” Paper 2A-1b Presented at 

the 1st NSS Space Solar Power Symposium, National Space Society International Space 

Development Conference, Chicago, Illinois, May 27 – 31, 2010. 

Little, Frank E., “Opportunities and Challenges for Wireless Power Transmission” Presented at 2nd 

International Symposium on Radio Systems and Space Plasma, Sofia, Bulgaria, August 25-27, 

2010. 

Little, Frank E. “Space Experiments for Space Based Solar Power” Paper 2-10 Presented at IAA 50th 

Anniversary Celebration Symposium on Climate Change/Green Systems, Nagoya, Japan, 

August 30-31, 2010. 

Little, Frank E., Paola D. Alicea, Mindy E. Watts and Veronica N. Medrano, “Design of a Small 

Semi-autonomous Satellite for Microwave Wireless Power Transmission Demonstration”, Paper 

2011-q-04, Presented at 28th International Symposium on Space Technology and Science, 

Ginowan City, Okinawa, Japan, June 5-12, 2011. 

Little, Frank E., Veronica N. Medrano, Miguel A. Rios and Daniel J. Talamantez “Antenna System 

Design Characteristics for a Space to Earth Demonstration of Wireless Power Transmission”, 

Paper 2011-q-18, Presented at 28th International Symposium on Space Technology and Science, 

Ginowan City, Okinawa, Japan, June 5-12, 2011. 

Publications 

Mo, Siu-Hong, Edward R. Grant, Frank E. Little, Ronald G. Manning, Chester A. Mathis, Gerald S. 

Werre, and John W. Root, "Radiotracer Studies of Thermal Hydrogen Abstraction Reactions by 

Atomic Fluorine." In "Fluorine-containing Free Radicals: Kinetics and Dynamics of Reactions" 

J.W. Root, ed. ACS Symposium Series 66, pp 59-103. 

Lagunas-Solar, Manuel C., Harvey L. Thibeau, and Frank E. Little, "A Remote System for Multicurie 

Radiochemical Separations." Proceedings of 27th Conference on Remote Systems Technology. 

(1979 Winter Meeting, American Nuclear Society, November 11-15, 1979, San Francisco, 

California) pp 301- 306. 
NASA/TM—2012-217014 557 

Lagunas-Solar, Manuel C., John A. Jungerman, Neal F. Peek and Frank E. Little, "A Remote Flow 

System for Large-Scale High-Purity I-123 Production." Proceedings of 27th Conference on 

Remote Systems Technology. (1979 Winter Meeting, American Nuclear Society, November 11-

15, 1979, San Francisco, California) pp 295-300. 

Lagunas-Solar, Manuel C., Frank E. Little, and John A. Jungerman, "Proton Induced Reactions on 

Natural Pb Targets. A Potential New Cyclotron Method for Tl-201 Production." Int. J. Applied 

Radiat. Isotopes. 1981 32 (11) 817-22. 

Lagunas-Solar, M.C., F. E. Little, and H.A. Moore, Jr. "Cyclotron Productions of 128Cs (3.62 min). 

A New Positron - Emitting Radionuclide for Medical Applications." Int. J. Appl. Radiat. Isot. 

1982, 33 (8) 619-28. 

Lagunas-Solar, M.C., F.E. Little, and C.D. Goodart, "An Integrally Shielded Transportable Generator 

System for Thallium - 201 Production." Int. J. Appl. Radiat. Isot. 1982 33 (12) 1439-43. 

Little, F.E. and M.C. Lagunas-Solar, "Cyclotron Production of 67GA. Cross Sections and Thick 

Target Yields for the 67Zn (p,n) and 68Zn (p,2n) Reactions." Int. J. Appl. Radiat. Isot. 1983 34 

(3) 631-37. 

Lagunas-Solar, M.C., H.L. Thibeau, C.D. Goodart, F.E. Little, N.J. Navarro, and D.E. Hartnett, "A 

Multicurie, Transportable, Integrally Shielded 123Xe > 123I Generator and Processing System 

for High-Purity Iodine-123 Production." DOE Symp. Ser. 1985 56 (Dev. Role Short-lived 

Radionuclides Nuc. Med. Pract.) 190-202. 

Klein, M. and F.E. Little, "A Basic Research Program Aimed at Satisfying the Needs of the Gas 

Industry for Thermophysical Properties Data." in "The Role of Data in Scientific Progress" P.S. 

Glaeser ed. CODATA 1985 Elsevier Science Publishers B.V. (North Holland), Amsterdam, 

1985 pp 183-187. 

Starling, K.E., M. Klein and F.E. Little, "The Thermodynamic Properties of Natural Gas - Physical 

Properties." in "Proceedings of the Congress of "Gas Quality - Specification and Measurement 

of Physical and Chemical Properties of Natural Gas." Groningen, The Netherlands April 22-25, 

1986, J.G. van Rossum ed. Elsevier science Publishers B.V. (North Holland), Amsterdam, 1986 

pp 211-239. 

Holtzapple, Mark T., Frank E. Little, Merry E. Makela, and C.O. Patterson, Analysis of an AlgaeBased CELSS: Part I: Model Development." Acta Astronautica 19 (4), pp. 353-364, 1989. 

Holtzapple, Mark T., Frank E. Little, William Moses, and C.O. Patterson, "Analysis of an AlgaeBased CELSS: Part II: Options and Weight analysis." Acta Astronautica 19 (4), pp. 365-375, 

1989. 

Holtzapple, Mark T. and Frank E. Little, "Comparison of Waste Combustion and Waste Electrolysis 

Systems Analysis", Paper 891485. Proceedings of the 19th Intersociety Conference on 

Environmental Systems, San Diego, CA (1989). 

Preisig, H.A., Tae-Yeong Lee, Frank Little, and Bruce Wright, "On the Representation of LifeSupport System Models", Paper 891479. Proceedings of the 19th Intersociety Conference on 

Environmental Systems, San Diego, CA (1989). 

Preisig, Heinz. A., Tae-Yeong Lee and Frank Little "A Prototype Computer-Aided Modelling Tool 

for Life-Support System Models", Paper 901269. Proceedings of the 20th Intersociety 

Conference on Environmental Systems, Williamsburg, VA (1990). 

Little, Frank E. and Mike O. Kennedy "Commercialization Strategies for Space Power Technology", 

Paper IAF--90--617 Proceedings of the 41st International Astronautical Congress, Dresden, 

Germany, October 1990. 

Chang, Kai, Frank E. Little and Mike O. Kennedy "Microwave Power Transmission System: Space 

Flight Experiment Program", Paper IAF--90--216 Proceedings of the 41st International 

Astronautical Congress, Dresden, Germany, October 1990. 
NASA/TM—2012-217014 558 

Chang, K., A.D. Patton, M.O. Kennedy, F.E. Little, M.A. Pollock, K.A. Hummer, J.C. McCleary, 

B.S. Wei, A.M. Brown, J.O. McSpadden “Demonstration of Microwave Power Transmission 

in Space”, Paper B1.2 Proceedings of SPS 91: Power from Space, Paris/Gif-sur-Yvette, 

August 27-30, 1991. 

Nansen, Ralph H. and Frank E. Little “A Practical Solar Power Satellite for the New Millennium”, 

Paper 95-301. Proceedings of the 30th Intersociety Energy Conversion Engineering Conference, 

Orlando, Florida, August 1995. 

Erb, R.B., N. Kaya, G. Maryniak, R. Leonard, M.B. Duke, F. Little, A.D. Patton, R. Nansen, J. Spies 

and W. Sadeh “International Cooperation for the Acquisition of Space-Based Energy”, Paper 

IAA-95-IAA.3.3.06. Proceedings of the 46th International Astronautical Congress, Oslo, 

Norway, October 1995. 

McSpadden, James O., Frank E. Little, Michael B. Duke and Alex Ignatiev “An In-Space Wireless 

Energy Transmission Experiment”, Paper 96576. Proceedings of the 31st Intersociety Energy 

Conversion Engineering Conference, Washington, D.C. August 1996. 

Ehsani, M., J. Mahdavi, I. Pitel, J.E. Brandenberg, and F.E. Little “Development of an Efficient 

Power Supply for the Microwave Electrothermal Thruster”, Paper 970115 Space Technology 

and Aplications International Forum (2nd Conference on Commercial Development of Space) 

Albuquerque, New Mexico, January 1997, AIP Conference Proceedings 387 pp 893-899, 1997. 

Little, Frank E., James O. McSpadden, Kai Chang and Nobuyuki Kaya “Toward Space Solar Power: 

Wireless Energy Transmission Experiments Past, Present and Future”, Paper 970115 Space 

Technology and Applications International Forum (15th Symposium on Space Nuclear Power 

and Propulsion) Albuquerque, New Mexico, January 1998, AIP Conference Proceedings 420 

pp. 1225-1233, 1998. 

Little, Frank E. “A wireless Power Transmission Power System for Microgravity Crystal Processing 

Satellites”, Proceedings of the IEEE Aerospace Conference, Big Sky, Montana, March 18 – 25, 

2000. 

Wang, Chunsheng, Imran Kakwan, A.John Appleby and Frank E. Little, “In-Situ Investigation of 

Electrochemical lithium Intercalation into Graphite Powder”, Journal of Electroanalytical 

Chemistry, 489, (2000), pp. 55-67. 

Wang, Chunsheng, A. john Appleby and Frank E. Little, “Electrochemical Study on Nano-Sn, 

Li4.4Sn and AlSi0.1 Powders Used as Secondary Lithium Battery Anodes”, Journal of Power 

Sources, 93, (2001), pp. 174-185. 

Wang, Chunsheng, A. John Appleby and Frank E. Little, “Charge-discharge Stability of Graphite 

Anodes for Lithium-Ion Batteries”, Journal of Electroanalytical Chemistry, 497. (2001), 

pp. 33-46. 

Wang, Chunsheng, A. John Appleby and Frank E. Little, “Comparison of the Electrochemical 

impedance Spectroscopy Characteristics of Insertion Electrode Materials used in Secondary 

Metal Hydride and Lithium-Ion electrodes”, Journal of the Electrochemical Society, 148, (7), 

pp. A762-A767, (2001). 

Wang, Chunsheng, A. John Appleby and Frank E. Little, “Electrochemical Impedance Study of Initial 

Lithium Ion Intercalation into Graphite powders”, Electrochimica Acta, 46, (2001), pp. 1793-

1813. 

Wang Chunsheng, Andre Rakotondrianibe, A. John Appleby and Frank E. little, “Characterization of 

Metal Hydride Electrodes via Microperturbation and In Situ Intrinsic Resistance Measurement”, 

Journal of the Electrochemical Society, 147, (12), pp. 4432-4439, (2000). 

Wang, Chunsheng, A. John Appleby and Frank E. Little, “Electrochemical Study of the SnO2 

Lithium-insertion Anode using Microperturbation techniques”. Solid state Ionics, 147, (2002), 

pp. 13-22. 

Wang, Chunsheng, A. John Appleby and Frank E. Little, “Irreversible Capacities of Graphite Anode 

for Lithium-ion Batteries”, Journal of Electroanalytical Chemistry, 519, (2002), pp. 9-17. 
NASA/TM—2012-217014 559 

Zhang, Xiang-Wu, Chunsheng Wang, A. John Appleby and Frank E. Little, “Improvement in 

Electrochemical Properties of Nano-tin-polyaniline Lithium-ion Composite Anodes by Control 

of Electrode Microstructure”, Journal of Power Sources, 109, (2002), pp.136-141. 

Wang, Chunsheng, A. John Appleby and Frank E. Little, “Low-Temperature Characterization of 

Lithium-ion Carbon Anodes via Microperturbation Measurement”, Journal of the 

Electrochemical Society, 149, (6), (2002) pp. A754-A760. 

Zhang, Xiang-Wu, Chunsheng Wang, A. John Appleby and Frank E. Little, “Composite Doped 

Emeraldine-Polyethylene Oxide-Bonded Lithium-ion Nano-tin Anodes with Electronic-ionic 

Mixed Conduction”, Solid State Ionics, 150 (3,4), (2002) pp. 383-389. 

Little, Frank E., “Solar Power Satellites: Recent Developments”, Paper 0965, Proceedings of the 

XXXVIIth URSI General Assembly, Maastricht, The Netherlands, August, 2002. 

Zepeda, Paola, Kai Chang and Frank Little “Optimal Antenna Taper Design for a Sandwich 

Transmitting Array in Space Solar Power Satellite”, Paper IAC-02-R.2.10, Proceedings of the 

53rd International Astronautical Congress, The World Space Congress, Houston, October, 2002. 

Hsieh, L.H., B.H. Strassner, S.J. Kokel, C.T. Rodenbeck, M.Y. Li, K. Chang, F.E. Little, G.D. Arndt 

and P.H. Ngo, “Development of a Retrodirective Wireless Microwave Power Transmission 

System”, IEEE Antennas and Propagation Society International Symposium Digest, 2, 393-396, 

June 2003. 

Zhang, Xiang-Wu, Prashanth K. Patil, Chunsheng Wang, A. John Appleby, Frank. E. Little, David L. 

Cocke, “Electrochemical performance of lithium ion battery, nano-silicon-based, disordered 

carbon composite anodes with different microstructures”, Journal of Power Sources 125 (2004) 

pp. 206–213. 

Little, F.E., S.J. Kokel, C.T. Rodenbeck, K. Chang, G.D. Arndt, P.H. Ngo, “Development of 

Retrodirective Control Transmitter for Wireless Power Transmission”, Radio Science Bulletin 

311 (December 2004) pp. 38-46. 

Wang, Chunsheng, Prashanth Patil, A. John Appleby, Frank E. Little, Mehmet Kesmez and David L. 

Cocke, “In Situ Ionic/Electric Conductivity Measurement of La0.55Li0.35TiO3 Ceramic at 

different Li Insertion Levels”, Journal of the Electrochemical Society, 151 (8) A1196-A1201 

(2004). 

Little, Frank E. “Power from Space: From Promise to Reality”, Proceedings of the URSI 2004 

International Symposium on electromagnetic Theory, Pisa, Italy, May 23-27, 2004, pp. 311-313. 

Little, Frank and Henry Brandhorst, “An Approach for Lunar power – 24/29” Proceedings of Solar 

Power from Space – SPS’04/5th Wireless Power Transmission conference (WPT 5), Granada, 

Spain, June 30 – July 2, 2004 (in press). 

Rodenbeck, C., S. Kokel, K. Chang, F.E. Little, “Microwave Wireless Power Transmission with 

Retrodirective Beam Steering”, Paper 2004-5640, Proceedings of 2nd International Energy 

Conversion Engineering Conference, Providence, Rhode Island, August 16-19, 2004. 

Brandhorst, Henry W. and Frank Little, “An Approach for Continuous Lunar Power”, Paper 2004-

5688, Proceedings of 2nd International Energy Conversion Engineering Conference, Providence, 

Rhode Island, August 16-19, 2004. 

Brandhorst, Henry W. and Frank Little, “POWOW Revisited – Beamed Power for Mars 

Exploration”, Paper 2004-5642, Proceedings of 2nd International Energy Conversion 

Engineering Conference, Providence, Rhode Island, August 16-19, 2004. 

Little, Frank E. and Alan Palazzolo, “Testing of Magnetic Bearings for Flywheel Energy Storage in 

Simulated Space Conditions”, Paper No. AIAA-2005-5642, 3rd International Energy 

Conversion Engineering Conference, San Francisco, August 15-18, 2005. 

Little, Frank E., “A Wireless Power Based Space Transportation System”, Paper No. 2006-g-10, 

Proceedings of the 25th International Space Technology and Science Conference, Kanazawa, 

Japan, June 4-11, 2006. 
NASA/TM—2012-217014 560 

Schuller, M., Lalk, T., Wiseman, L., Little, F., Godard, O., Abdel-Fattah, S., Askew, R., Klaus, D., 

Kobrick, R., Thomas, G., Rouen, M. and Conger, B, “Innovative Schematic Concept Analysis 

for a Space Suit Portable Life Support Subsystem”, Paper 2006-01-2201, Proceedings of the 

36th International Conference on Environmental Systems, Norfolk, Virginia, USA, July 17-20, 

2006. 

Klaus, D., Bamsey, M., Schuller, M., Godard, O., Little, F. and Askew, R., “Defining Space Suit 

Operational Requirements for Lunar and Mars Missions and Assessing Alternative 

Architectures”, Paper 2006-241 Proceedings of the 36th International Conference on 

Environmental Systems, Norfolk, Virginia, USA, July 17-20, 2006. 

Little, Frank, “From the Moon to the Earth: Lunar Implications for Space Solar Power”. Proceedings 

of the First International Symposium on Radio Systems and Space Plasma, Sofia, Bulgaria, 

September 2-5, 2007. 

Little, Frank E., “Constraints on the Design of a Lunar-to-Earth Wireless Power Transmission 

System” Paper IAC-08-C3.1.8, Proceedings of the 59th International Astronautical Congress, 

Glasgow, Scotland, 29 September 29 - October 3, 2008. 

Little, Frank E., Kai Chang, Rainer Fink, G. Dickey Arndt, Phong H. Ngo and Raymond F. Beach, 

“A Space to Earth Demonstration of Wireless Power Transmission” Paper No. 2009-h-27 

Proceedings of the 27th International Space Technology and Science Conference, Tsukuba, 

Japan, July 5-12, 2009. 

Little, Frank E., “Opportunities and Challenges for Wireless Power Transmission”, Proceedings of 

the 2nd International Symposium on Radio Systems and Space Plasma, Sofia, Bulgaria, August 

25-27, 2010 pp 153-156. 

Little, Frank E., Paola D. Alicea, Mindy E. Watts and Veronica N. Medrano, “Design of a Small 

Semi-autonomous Satellite for Microwave Wireless Power Transmission Demonstration”, Paper 

2011-q-04, Proceedings of the 28th International Symposium on Space Technology and Science, 

Ginowan City, Okinawa, Japan, June 5-12, 2011. 

Little, Frank E,, Veronica N. Medrano, Miguel A. Rios and Daniel J. Talamantez “Antenna System 

Design Characteristics for a Space to Earth Demonstration of Wireless Power Transmission”, 

Paper 2011-q-18, Proceedings of the 28th International Symposium on Space Technology and 

Science, Ginowan City, Okinawa, Japan, June 5-12, 2011 

 
NASA/TM—2012-217014 561 

James McSpadden, Ph.D. 

President 

McSpadden Consulting 

1204 Padre Circle 

Allen, Texas 75013 

james@mcspadden.com 469–734–7507 

Dr. McSpadden is an expert in microwave power transmission systems and rectenna design. As a private 

consultant for over 10 years, he has lead projects dealing with the system analysis and technology 

development for various power transmission projects. He previously worked full time for The Boeing 

Company for 8 years and has been employed by Raytheon since 2005 as a Senior Principal Electrical 

Engineer. Dr. McSpadden received his B.S.E.E, M.S.E.E, and Ph.D. diplomas, all from Texas A&M 

University, in 1989, 1993, and 1998, respectively. Dr. McSpadden has over 20 published papers in 

journals, conferences, and magazines on microwave power transmission. He is a Senior Member of the 

IEEE and is currently writing a technology book titled “Microwave Power Transmission.” 

Power Beaming Publications 

(1) J.O. McSpadden, “Advances in RF Wireless Power Transmission,” presented at the International 

Symposium on Solar Energy from Space, Ontario Science Centre, Toronto, Canada, Sept. 8-10, 

2009 

(2) R.C. Hansen, J.O. McSpadden, and J.N. Benford, “A Universal Power Transfer Curve,” IEEE 

Microwave and Wireless Components Letters, vol. 15, no. 5, pp. 369-371, May 2005. 

(3) J.O. McSpadden and J.C. Mankins, “Space solar power programs and microwave wireless power 

transmission technology,” IEEE Microwave Magazine, vol. 3, no. 4, pp. 46-57, Dec. 2002. 

(4) J.O. McSpadden, “The current capabilities of microwave WPT technology” in 2001 Asia-Pacific 

Radio Science Conference AP-RASC’01 Conference Digest, Aug. 1-4, 2001, Chuo University, 

Tokyo, Japan, pp. 14-17. 

(5) J.O. McSpadden, L. Fan, and K. Chang, “Design and experiments of a high-conversion-efficiency 

5.8-GHz rectenna,” IEEE Trans. Microwave Theory Techn., vol. 46, no. 12, pp. 2053-2060, Dec. 

1998. 

(6) J.O. McSpadden, R. M. Dickinson, L. Fan and K. Chang, “A novel oscillating rectenna for wireless 

microwave power transmission,” in 1998 IEEE MTT-S International Microwave Symposium 

Digest, 1998, pp. 1161-1164. 

(7) J.O. McSpadden, K. Chang, F. Little, and M. Duke, “Study of ISS free-flyer power beaming,” in 

Proceedings of SPS '97 Conference, Montreal, Canada, Aug. 24-28, 1997, pp. 169-174. 

(8) J.O. McSpadden, L. Fan and K. Chang, “A high conversion efficiency 5.8 GHz rectenna,” in 1997 

IEEE MTT-S International Microwave Symposium Digest, 1997, pp. 547-550. 

(9) J. McSpadden, K. Chang, and A.D. Patton, “Microwave power transmission research at Texas 

A&M University,” Space Energy and Transportation, vol. 1, no. 4, pp. 368-393, 1996. 

(10) J.O. McSpadden, F.E. Little, M.B. Duke, and A. Ignatiev, “An in-space wireless energy 

transmission experiment,” in Proceedings of the 31st Intersociety Energy Conversion Engineering 

Conference, 1996, pp. 468-473. 

(11) J. McSpadden and K. Chang, “Review of the solar power satellite,” in Proceedings of the Fifth 

International Conference on Space ‘96, 1996, pp. 254-259. 

(12) J.O. McSpadden, A.M. Brown, K. Chang and N. Kaya, “A receiving rectifying antenna for the 

international space year - microwave energy transmission in space (ISY-METS) rocket 

experiment,” IEEE Aerospace and Electronic Systems Magazine, vol. 9, no. 11, pp. 36-41, Nov. 

1994. 
NASA/TM—2012-217014 562 

(13) J.O. McSpadden, N. Kaya, A.M. Brown and K. Chang, “A receiving rectifying antenna for the 

international space year - microwave energy transmission in space (ISY-METS) rocket 

experiment,” in 29th Intersociety Energy Conversion Engineering Conference, Monterey, CA, 

1994, pp. 723-727. 

(14) J.O. McSpadden and K. Chang, “A dual polarized circular patch rectifying antenna at 2.45 GHz for 

microwave power conversion and detection,” in 1994 IEEE MTT-S International Microwave 

Symposium Digest, 1994, pp. 1749-1752. 

(15) J.O. McSpadden, “Construction and testing of a space ready rectenna, final report,” Center for 

Space Power, Texas Engineering Experiment Station, The Texas A&M University System, College 

Station, TX, NASA Lewis Research Center Grant NAG 3-1368, June 1993. 

(16) “Wireless space power experiment (WISPER), final report,” NASA/URSA Advanced Design 

Program, Dept. of Electrical Engineering, University of Alaska Fairbanks, Fairbanks, AK, June 

1993. 

(17) J.O. McSpadden, T. Yoo, and K. Chang, “Theoretical and experimental investigation of a rectenna 

element for microwave power transmission,” IEEE Trans. Microwave Theory Techn., vol. 40, no. 

12, pp. 2359-2366, Dec. 1992. 

(18) J.O. Mc Spadden and K. Chang, “Suppression of rectenna harmonic radiation by a frequency 

selective surface,” presented at 1992 IEEE-APS/URSI/NEM Joint Symposia, Chicago, 1992. 

(19) T. Yoo, J.O. McSpadden, and K. Chang, “35 GHz rectenna implemented with a patch and a 

microstrip dipole antenna,” in 1992 IEEE MTT-S International Microwave Symposium Digest, 

1992, pp. 345-348. 

(20) J.O. McSpadden, T. Yoo, and K. Chang, “Diode characterization in a microstrip measurement 

system for high power microwave power transmission,” in 1992 IEEE MTT-S International 

Microwave Symposium Digest, 1992, pp. 1015-1018. 

(21) K. Chang, A.D. Patton, M.O. Kennedy, F.E. Little, M.A. Pollock, K.A. Hummer, J.C. McCleary, 

B.S. Wei, J.O. McSpadden, and A.M. Brown, “Demonstration of microwave power transmission in 

space,” in SPS 91 Power From Space, Paris, 1991, pp. 343-347. 

 
NASA/TM—2012-217014 563 

Leik N. Myrabo, Ph.D. 

Lightcraft Technologies, Inc. 

1914 Walloomsac Road 

Bennington, Vermont 05201 

802–447–6275 (w) 802–989–2874 (c) 

Professional History 

Professor Leik N. Myrabo is recognized as an aerospace pioneer by the Smithsonian National Air 

and Space Museum for his pioneering work on laser propulsion. His company, Lightcraft 

Technologies, Inc. (founded in 1999), holds the world altitude record for laser launched vehicles. 

Myrabo has authored and co-authored more than 210 journal, symposium, and conference articles in 

addition to two books: The Future of Flight (1985), and The Lightcraft Flight Handbook: LTI-20 

(2009). 

He received his B.S. degree in Aerospace Engineering from Iowa State University in 1968, and Ph.D. 

in Engineering Physics from the University of California at San Diego in 1976, and, Thereafter, he 

spent a total of 7 years at Physical Sciences, Inc., W.J. Schafer Associates, and the BDM Corporation, 

pursuing “Star Wars” research as a staff scientist and consultant on directed energy applications, 

aerospace systems, space prime power, and advanced propulsion. He joined the Rensselaer 

Polytechnic Institute (RPI) faculty in 1983, retiring in August 2010 to devote his full-time attention 

to commercializing beamed energy propulsion (BEP) under Lightcraft Technologies, Inc.

For nearly three decades at RPI, Myrabo’s burning desire has been to create and demonstrate viable 

concepts for non-chemical propulsion of future flight vehicles for an era beyond oil. Prior to June 

2005, when he was awarded a $2.25M MURI grant by AFOSR, he had previously attracted $3.5M 

(in 2005 dollars) into RPI, all focused on BEP research for powering future aircraft and spacecraft. 

Myrabo’s MURI combined experimental/experimental research investigated laser propulsion 

physics for affordable, rapid access to space: i.e., launching pico/nano/micro-satellites (i.e., 0.1 to 

100 kg). 

Research & Development Interests 

Advanced space propulsion & power, aerospace systems, directed energy, hypersonic gas dynamics, 

energy conversion, space technology, alternative energy. Myrabo’s experimental and theoretical 

studies at RPI were focused on alternative (non-chemical) energy sources for aeronautical, aerospace, 

and space flight propulsion of the 21st century and beyond. This advanced energetics research took a 

long term, high risk approach to identifying areas of potential technological breakthroughs. His 

primary R&D emphasis was on the application of Beamed Energy Propulsion (BEP) with laser, 

microwave, or millimeter wave sources, in developing future air-breathing and rocket shuttlecraft for 

a variety of hypersonic boost-glide, space launch, and orbit-raising missions. 

Recent International Collaboration 

The Lightcraft Project at the Umea Institute of Design (UID) was a collaboration between three MA 

Programmes (Transportation Design, Advanced Product Design, and Interaction Design) and 

Lightcraft Technologies, Inc. conducted in an intensive ten-week period during the Fall’09 semester 

in Umea, Sweden. Thirty-three 2nd year design UID students in eleven teams (with one student from 

each programme) aggressively pursued the research and development of concepts, ideas and scenarios 

to design the ideal lightcraft travel experience—requiring the intimate integration of advanced 

technology with the human element. Their final presentations disclosed novel Lightport (equivalent to 

today’s airports) designs, tickets with Lightport navigation/ departure & arrival/and passport info, 

unique vehicle boarding methods, elaborate and functional interior designs with reconfigurable seats, 

g-suits and restraints, holographic in-flight displays, and so much more. The UID Lightcraft Project

exploited a sophisticated laser powered multi-cycle engine/vehicle concept that was evolved over a 

10+ year period at RPI. This new global flight transportation system is based on beaming power from 
NASA/TM—2012-217014 564 

satellite solar power stations—envisioned to replace the commercial jumbojet airline system by 

2025–for an era beyond oil.

Innovative Work of Potential Long-Term Significance 

(R&D milestones for Myrabo and his investigative teams; Reference numbers “[xx]” refer to Selected 

Publications/References list—see below)

 First demonstration of a hypersonic laser-scramjet at Mach 6-10 in the IEAv hypersonic shock 

tunnel in Sao Jose dos Campos, Brazil, with I.I. Salvador, M.A.S. Minucci, P.G.P. Toro [1-2]. 

 Created 1 km diameter, rotating LEO solar power station concept (engineering analysis, and 

orbital mechanics study) exploiting rechargeable superconducting magnetic energy storage 

system [3-4], and a compatible concept for a 20-m microwave lightcraft designed to transport 6-

12 person crews [5]. Invented concept for new ultra-high power density superconductor (with 

D.A. Gross): metal matrix superconductor composites [4]. 

 First laser-boosted light sail demonstration, and vacuum photonic thrust measurements 

(pendulum mount) of candidate ultralight light sail materials (5 cm diameter molybdenum-coated, 

carbon micro-truss fabrics) using a high power CW infrared laser called LHMEL II [6]. First 

vertical wire-guided flight attempts (i.e., 1 g acceleration) of laser-boosted sails using advanced 

flight-weight materials [7]. 

 First laser-boosted, spin-stabilized lightcraft flown in vertical free-flights propelled by 

airbreathing pulsejet engine at White Sands Missile Range [8]. Holds a 30.2 m world altitude 

record for airbreathing laser-propelled lightcraft, demonstrating infinite specific impulse [9]. 

 First laser-propelled, spin-stabilized rocket lightcraft in free-flight [10]. Myrabo holds the 72m 

world altitude record for laser-boosted rocket lightcraft, with longest boost duration and greatest 

“air time” flown on the PLVTS pulsed CO2 laser at WSMR, NM [11]. 

 First measurements of beam-riding characteristics of a laser-propelled lightcraft engine [12], and 

first investigation and simulation of lightcraft flight dynamics [13-14]. 

 First high power pulsed laser experiments examining the survivability of refractory metal 

coatings on graphite bell engines and ceramic matrix composite materials for laser pulsejet 

engines [15]. 

 Established theoretical foundations for the Directed Energy AirSpike (DEAS) with Yuri 

Raizer[16]. First electric-arc AirSpike experiments in Mach 8-10 hypersonic flow with and 

without a blunt body [17-19]; electric-arc AirSpike experiments in Mach 3 flow [20]; Laserinduced AirSpike experiments in hypersonic flow with M.A.S Minucci and P.G.P. Toro [21]. 

 First magnetic-augmented, airbreathing laser pulsejet experiments—using the terawatt 1 μm Ndglass Pharos III laser [22-23]. 

 Created the Lightcraft Technology Demonstrator (LTD) concept for laser-launching of 

microsatellites (i.e., 10-100 kg payloads), under support of the SDIO Laser Propulsion Program 

[24]. (The laser lightcraft design flown at WSMR evolved directly from this advanced 

engine/optics/vehicle concept.)

Selected Publications/References 

(1) I.I. Salvador, I.I., Myrabo, L.., et.al., 2-D Air-Breathing Lightcraft Engine Experiments in 

Hypersonic Conditions,” in 7th International Symposium on Beamed Energy Propulsion, edited by 

H.-A. Eckel and S. Scharring, AIP Conference Proceedings, American Institute of Physics, 

Melville, NY, 2011. 

(2) I.I. Salvador, I.I., Myrabo, L.., et.al., 2-D Airbreathing Lightcraft Engine Experiments in 

Quiescent Conditions,” in 7th International Symposium on Beamed Energy Propulsion, edited by 

H.-A. Eckel and S. Scharring, AIP Conference Proceedings, American Institute of Physics, 

Melville, NY, 2011. 
NASA/TM—2012-217014 565 

(3) Myrabo, L.N., “Microwave Power-Beaming Infrastructure for Manned Lightcraft Operations,” in 

4th International Symposium on Beamed Energy Propulsion, edited by K. Komurasaki, T. Yabe, S. 

Uchida, and A. Sasoh, AIP Conference Proceedings 830, American Institute of Physics, Melville, 

NY, 651-662 (2006). 

(4) Gross, D.A., and Myrabo, L.N., “Metal Matrix Superconductor Composites for SMES-Driven, 

Ultra High Power BEP Applications: Parts 1 & 2,” in 4th International Symposium on Beamed 

Energy Propulsion, edited by K. Komurasaki, T. Yabe, S. Uchida, and A. Sasoh, AIP Conference 

Proceedings 830, American Institute of Physics, Melville, NY, 411-436 (2006). 

(5) Myrabo, L.N., “Horizon Mission: 2025 Space Command’s Ultra-Energetic Lightcraft with SuperPressure Airship Structure,” in 3rd International Symposium on Beamed Energy Propulsion, edited 

by A.V. Pakhomov and L.N. Myrabo, AIP Conference Proceedings 766, NY: American Institute 

of Physics, 86-99 (2005). 

(6) Myrabo, L.N., Knowles, T.R., Bagford, J.O., and Siebert, D.B., “Experimental Investigation of 

Laser-Pushed Light Sails in a Vacuum,” in High-Power Laser Ablation III, edited by Claude R. 

Phipps, Proceedings of SPIE Vol. 4065 (2000) pp. 521-532. 

(7) Myrabo, L.N., Knowles, T.R., Bagford, J.O., Seibert, D.B., and Harris, H.M., “Laser-Boosted 

Light Sail Experiments with the 150 kW LHMEL II CO2 Laser,” in High-Power Laser Ablation 

IV, edited by Claude R. Phipps, Proceedings of SPIE Vol. 4760 (2002) pp. 774-798. 

(8) Myrabo, L.N., Messitt, D.G., and Mead Jr., F.B., "Ground and Flight Tests of a Laser Propelled 

Vehicle," AIAA Paper 1998-1001, 36th AIAA Aerospace Sciences Meeting & Exhibit, Reno, NV, 

12-15 Jan. 1998. 

(9) Mead Jr., F.B., Myrabo, L.N., and Messitt, D.G., "Flight and Ground Tests of a Laser-Boosted 

Vehicle," AIAA Paper 1998-3735, 34th AIAA Joint Propulsion Conf. & Exhibit, Cleveland, OH, 

13-15 July 1998. 

(10) Messitt, D.G., Myrabo, L.N., and Mead Jr., F.B., "Laser Initiated Blast Wave for Launch Vehicle 

Propulsion," AIAA Paper 2000-3848, 36th AIAA Joint Propulsion Conference, 16-19 July 2000, 

Huntsville, AL. 

(11) Myrabo, L.N., “Brief History of the Lightcraft Technology Demonstrator (LTD) Project,” Review 

of Laser Engineering, 34, 423-428 (2006). 

(12) Libeau, M., and Myrabo, L., “Off-Axis and Angular Impulse Measurements on a Lightcraft 

Engine,” in 3rd International Symposium on Beamed Energy Propulsion, edited by A.V. Pakhomov 

and L.N. Myrabo, AIP Conference Proceedings 766, American Institute of Physics, Melville, NY, 

166-177 (2005). 

(13) Libeau, M.A., Myrabo, L.N., and Filippelli, M., “Combined Experimental & Theoretical Flight 

Dynamics Investigation of a Laser-Propelled Vehicle,” in 1st International Symposium on Beamed 

Energy Propulsion, edited by A.V. Pakhomov, AIP Conference Proceedings 664, American 

Institute of Physics, Melville, NY, 125-137 (2003). 

(14) Ballard, C.G., Anderson, K.S., and Myrabo, L.N., “Flight Dynamics Simulation of Lightcraft 

Propelled by Laser Ablation,” Paper 6261-74, in High-Power Laser Ablation VI, edited by Claude 

R. Phipps, Proceedings of SPIE, Vol. 6261, 2006. 

(15) Myrabo, L.N., Libeau, M.A., Meloney, E.D., Bracken, R., and Knowles, T.B., “Survivability of 

Thin Metal Mirror Coatings on Graphite Pulsed Laser Propulsion Engines,” AIAA Paper 2002-

3783. 

(16) Myrabo, L.N., and Raizer, Y.P, “Laser-Induced Air Spike for Advanced Transatmospheric 

Vehicles,” AIAA Paper 1994-2451. 

(17) Marsh, J.J., Myrabo L.N., Messitt, D.G., Nagamatsu, H.T., and Raizer, Yu.P., "Experimental 

Investigation of the AirSpike Hypersonic Inlet at Mach 10," AIAA Paper 1996-0721. 

(18) Myrabo, L.N., Raizer, Yu.P., and Shneider, M.N., “The Calculation and Similarity Theory of 

Experiment Simulating the Air-Spike Effect in Hypersonic Aerodynamics,” High Temperature, 

36, n.2, 287-292 (1998). 
NASA/TM—2012-217014 566 

(19) Myrabo, L.N., Raizer, Yu.P., Shneider, M.N., and Bracken, R.M., “Drag and Total Power 

Reduction for Artificial Heat Input in Front of Hypersonic Blunt Bodies,” High Temperature, 42,

n. 6, 1-9 (2004). 

(20) Misiewicz, C., Myrabo, L.N., Shneyder, M. and Raizer, Y.P., “Combined Experimental and 

Numerical Investigation of Electric-Arc Airspikes for Blunt Body at Mach 3,” in 3rd International 

Symposium on Beamed Energy Propulsion, edited by A.V. Pakhomov and L.N. Myrabo, AIP 

Conference Proceedings 766, American Institute of Physics, Melville, NY, 528-550 (2005). 

(21) Minucci, M.A.S, Toro, P.G.P., Oliveira, A.C., Chanes Junior, J.B., Ramos, A.G., Pereira, A.L., 

Nagamatsu, H.T., and Myrabo, L.N., “Investigation of a Laser-Supported Directed Energy Air 

Spike in Hypersonic Flow,” Journal of Spacecraft and Rockets, 40, n. 1, 133-136 (2003). 

(22) Lyons, P.W., Myrabo, L.N., Jones, R.A., Nagamatsu, H.T., and Manka, C., “Experimental 

Investigation of a Unique Airbreathing Pulsed Laser Propulsion Concept,” AIAA Paper 1991-

1922. 

(23) Myrabo, L.N., Lyons, P.W., Jones, R.A., Liu, S., and Manka, C., “Airbreathing Laser Propulsion 

Experiments with 1 μm Terawatt Pharos III Laser – Parts 1 & 2,” in 7th International Symposium 

on Beamed Energy Propulsion, edited by H.-A. Eckel and S. Scharring, AIP Conference 

Proceedings, American Institute of Physics, Melville, NY, 2011. 

(24) Myrabo, L.N., et al, “Transatmospheric Laser Propulsion,” Final Technical Report, prepared under 

Contract No. 2073803 for Lawrence Livermore National Laboratory and the SDIO Laser 

Propulsion Program, 1989. 
NASA/TM—2012-217014 567 

Dr. Kevin L.G. Parkin 

Dr. Kevin L.G Parkin is a Research Faculty member at the Silicon Valley campus of Carnegie Mellon 

University. He holds an M.Phys. in Physics with Space Science and Technology from the University of 

Leicester, and an M.S. in Aeronautics and Ph.D. in Aeronautics with minor in Electrical Engineering from 

Caltech. As an Undergraduate, Dr. Parkin undertook research in non-linear time series analysis, CatSat 

satellite battery performance, low thrust trajectory optimization, and finite element structural modeling 

tools for conceptual design. As a first year Graduate Student, Dr. Parkin developed ICEMaker, an Excelbased system for team collaboration on complex conceptual designs. ICEMaker was subsequently used by 

satellite design teams at NASA JPL, NASA GRC, Northrop Grumman, Boeing, UTRC, Raytheon, 

Caltech, MIT, Stanford, and The Naval War College. Subsequently, he undertook research on Pulse 

Detonation engines and AJAX hypersonic airbreathing propulsion. In 2002, he invented the Microwave 

Thermal Rocket, and subsequently it became the topic of his Ph.D. thesis. In 2005, he demonstrated the 

operating principle of microwave thermal propulsion in the laboratory for the first time. In recognition of 

his accomplishments in this area, Dr. Parkin was awarded the Korolev Medal by the Russian Federation 

of Cosmonautics. In 2006, he founded the Mission Design Center at NASA ARC, developing a software 

infrastructure to supersede the ICEMaker-era approach to conceptual design. Dr. Parkin served as deputy 

director of the Mission Design Center until 2008, subsequently returning to his present work on 

microwave thermal rocketry. Most recently, Dr. Parkin invented the Engineering Inference Engine (EIE). 

EIE is a forward-chaining inference engine and corresponding language used for system analysis of the 

microwave thermal rocket. Dr. Parkin is a Member of the Institute of Physics (IOP) and the American 

Institute of Aeronautics and Astronautics (AIAA). 

NASA/TM—2012-217014 569 

Appendix C.—Thrust Augmentation of an In-Space Propulsion System 

Andrew Ketsdever 

Air Force Research Laboratory 

Propulsion Directorate 

Edwards AFB, CA 93524 

C.1 Concept 

This study investigates a concept by which the thrust (and specific impulse) of an in-space propulsion 

system is augmented using beamed energy. A 1 Mega-Watt (MW) level space asset is assumed for much 

of the study. The 1MW power level is assumed to be the beamed source power at the space asset (i.e., not 

necessarily at the receiving vehicle). The space asset is initially assumed to be either a CW laser, pulsed 

laser, or microwave based energy beaming system. 

A method of augmenting thrust in an in-space propulsion system through the use of off-satellite 

resources is to beam electromagnetic energy to the space vehicle. The beamed energy can be added to a 

proposed propulsion system in a variety of ways. First, the beamed energy can be collected by the 

receiving spacecraft and converted into electrical energy that can be used to power an electric propulsion 

system. A concept of this nature requires a beam collector on the receiving spacecraft and a system to 

convert the photon energy (or resulting thermal energy) into electrical power. Second, the energy can be 

absorbed directly in the combustion (or stagnation) chamber of the propulsion system. This concept 

would be very similar to a solar thermal propulsion system where photon energy is used to heat the 

propellant gas (directly or indirectly through the use of a thermal mass). This concept can also be 

envisioned as a way to augment the performance of chemical thrusters by adding energy directly into the 

combustion chamber. Third, the beamed energy can be absorbed in the expanding section of a 

converging-diverging nozzle to add energy after the combustion chamber as shown in Reference 1. 

Finally, a laser ablation thruster can be considered where the laser beam is used to directly ablate and 

subsequently heat a propellant from a solid surface. There are some benefits and drawbacks to each of 

these approaches as shown in Table C.1. 

C.2 Proposed Mission 

A space asset with a 1 MW source being beamed to a receiving spacecraft is assumed. From an initial 

orbital mechanics look, this much power only makes sense from the standpoint of an orbital maneuver 

that requires a large change in semi-major axis and/or inclination (i.e., high delta-V). Furthermore, this 

much power is enabling for responsive missions that need to be done quickly. Although several orbital 

scenarios can be envisioned, a nominal LEO–GEO co-planar transfer will be investigated in this study. 

For this study, the space asset will be assumed to be in MEO with a transferring vehicle going between 

LEO and GEO. The MEO location for the space asset allows the total distance of the traveling beam to be 

minimized (i.e., the maximum distance of travel can be half the distance between LEO and GEO). 
NASA/TM—2012-217014 570 

Table C.1.—Advantages and disadvantages of in-space beamed energy concepts. 

Concept Advantages Disadvantages 

I. Beamed energy power 

conversion 

 Converted power can be used for any 

propulsion system 

 Existing, Off-the-shelf propulsion 

systems can be considered 

 High power solar arrays with energy 

concentration are increasing in efficiency 

 Relatively low overall system 

efficiency 

 New generation of power conversion 

systems need to be demonstrated for 

high power level 

II. Stagnation chamber energy 

absorption 

 Energy addition to combustion chamber 

may also increase combustion efficiency 

 High temperature operation 

throughout entire system 

 Completely re-designed systems 

necessary to handle higher 

temperatures 

 Requires both primary propulsion and 

beamed energy to work (reliability) 

III. Diverging nozzle energy 

addition 

 Lower overall temperatures throughout 

entire system 

 Slight modification to existing systems 

 Energy absorbers (aluminum oxide, 

Carbon nanotubes, etc) can be added to 

almost any existing system 



 Direct energy absorption from 

expanding gas is extremely low 

 Fuel additives (absorbers) required 

for many systems which might 

adversely affect intrinsic performance 

 Requires both primary propulsion and 

beamed energy to work (reliability) 



IV. Laser ablation  Direct conversion of energy to propulsion 

 Solid propellants can be used 

 High power density required for 

ablation 

 Not demonstrated at this level 

C.3 Laser Versus Microwave In-Space Asset 

Because of the large distance of travel of the beamed energy, a laser asset is considered optimal. 

Although the power level near 1 MW is more indicative of current microwave technology, the laser 

system will be advantageous from the standpoint of beam divergence (diffraction limited optics favor 

lower wavelength), beam optics (i.e., focusing, collimating, etc.), and beam steering. 

For a laser with an initial 5-m diameter laser beam in MEO at a wavelength of 500 nm, the minimum 

spot size in LEO or GEO would be 8.8 m. A 10-m diameter collector on the spacecraft will be assumed 

such that the entire 1-MW beam is collected along the entire spacecraft trajectory. To make a 5-m 

diameter beam in MEO, a significant effort is required in low mass, high efficiency, and high power 

optics. Something similar to the USAF Academy’s photon sieve might be ideal (Ref. 2). To collect an 

equivalent amount of solar power, the collector would need to be 30.6 m, assuming a solar flux of 

1358 W/m2

 at Earth. Therefore, the beamed energy would significantly add to the overall system power 

over what the sun by itself could supply for the same size solar collector. 

C.4 Continuous-Wave Mode Versus Pulsed 

Pulsed operation of the proposed laser system is advantageous from the standpoint of the spacecraft 

power required to produce the laser beam. Stored energy source can be used to operate the laser so that 

the host spacecraft does not have to produce multi-MW of power. An orbital dynamics simulation needs 

to be done to assess whether a continuous source in MEO could access a transferring vehicle through its 

entire trajectory or where periods of eclipse might occur. 

C.5 Thruster Concepts Investigated 

A proposed thruster from each of the four concepts listed in Table C.1 will be evaluated. The 

following thruster concepts were investigated for thrust augmentation (the value in the parentheses 

indicates the concept from Table C.1 that the thruster could be considered for). 
NASA/TM—2012-217014 571 

1. Arcjet (I, II) 

2. Resistojet (I, II) 

3. MPD (I) 

4. Solar Thermal Thruster (II) 

5. Liquid Monopropellant (II, III) 

6. LOx/LH2 Bipropellant (II, III) 

7. Aluminum Solid Propellant (II, III) 

8. Laser Ablative Thruster (IV) 

Only the LOx/LH2, aluminum solid, and MPD systems have been demonstrated at the MW level to 

date. However, there are no known physical reasons why the other systems could not be demonstrated, 

although their efficiencies at the MW level would have to be investigated. Table C.2 gives some relevant 

performance parameters ( augmented) for the thruster systems listed above. 

For a 1MW energy addition to make sense, the augmented thruster system should have an energy 

level near 1MW, or the thruster should operate completely with the 1 MW of beamed energy. If the 

augmented thruster jet power is significantly less than 1MW, the energy addition to the propulsion system 

will essentially act as the only means of thrust. This is technically plausible; however, in this case, the 

beamed-energy system might as well act as the only propulsion means thereby reducing the overall 

complexity of the system on the transferring spacecraft side. If the augmented thruster power is much 

larger than 1MW, the energy added to the vehicle from the beamed-energy source will not significantly 

impact the thruster performance. Alternatively, concepts that utilize the 1MW beamed energy source as 

its only means of propulsion can also be considered. 

TABLE C.2.— AUGMENTED THRUSTER SYSTEM PERFORMANCE 

Thruster Isp (sec) Thrust (N) Mass Flow (kg/s) Jet Power (kW) Propellants 

Monopropellant 220-240 1-600 4e-4 – 0.3 1-710 Hydrazine, Hydrogen 

Peroxide 

LOx/LH2 (a) 420-450 

(444) 

10,000-70,000 

(66,700) 

2.3-17 

(15.3) 

20,600 – 155,000 

(145,260) 

LOx/LH2 

Aluminum Solid(b) 250-300 

(289.5) 

27,000-88,000 

(27,000) 

9.2-36 

(9.5) 

33,000 - 130,000 

(38,340) 

Aluminum/Aluminum 

Oxide 

Arcjet(c) 500-1,000 

(800) 

0.2-2 (2) (0.24) 1.8-30 

(26) 

Ammonia, Hydrazine 

Resistojet 400-1,000 0.2-6 2e-5 – 1.5e-3 3-30 Hydrogen, Ammonia, 

Hydrazine 

MPD(d) 2,000-5,000 

(4,500) 

40-100 (22.6) 1e-4 – 5e-3 

(5e-4) 

1,000 

(500) 

Hydrogen, Lithium, 

Argon 

Laser Ablation(e) 2,500 81.5 3.3e-3 1,000 Solid 

Solar Thermal 500-1,000 20-220 2e-3 – 4.5e-2 50-1,000 Hydrogen, Water, 

Ammonia a

 After the RL-10 in () 

b After STAR 27 Apogee Kick Motor in () 

c

 After ESEX Arcjet in () 

d

 After Energia (Russian) Li-MPD in () 

e

 Assumes no losses. Power set to 1 MW and Isp limited to 2500 s (arb) 

TABLE C.3.—THRUSTER SYSTEM PARAMETERS 

Parameter Thrusters Notes 

Demonstrated MW power operation 2, 3, 6 At least lab level demonstration 

Plasma naturally available for absorption 4, 6 High efficiency laser photon absorption 

Particles/surfaces naturally available for absorption 3, 7 Relatively high efficiency laser photon absorption 

Will not require major redesign for power augmentation 7, 8 Heat addition issues—thermal redesign 

System benefits including high density propellants 1, 2, 3, 7 Solid and liquid propellants 
NASA/TM—2012-217014 572 

C.5.1 Concept I 

Using the beamed energy as a power source to operate an electric thruster can be considered (i.e., 

concept I). However, the power conversion efficiency will be relatively low (~40%). Therefore, concept I 

systems will not be considered in this study. 

C.5.2 Concept II 

From the argument that the augmented thruster power should be equivalent to the beamed power, 

electric propulsion systems are not viable for concept II implementation since they will require 

approximately 1MW of on-board spacecraft power. A high on-board power requirement will make the 

transferring vehicle power system as complex as the laser beaming spacecraft asset. Concepts that utilize 

thermochemistry with energy addition to the combustion chamber are viable; however, their hindered by 

the complexity of an increased thermal loading throughout the propulsion system while maintaining 

efficient combustion. An augmented solar thermal propulsion system has several benefits in that its 

complexity is essentially the same whether using a solar source or a beamed energy source. Also, solar 

thermal propulsion systems are relatively efficient and the waste heat from the propulsion system could 

be converted to electrical power that the receiving spacecraft could use for other processes. 

C.5.3 Concept III 

Neutral gases do not readily absorb photon radiation. This is true of most gases over a wide range of 

wavelengths. However, high temperature (>1500K) materials and plasmas generally are highly efficient 

absorbers of both visible and microwave energy. Therefore, fuel additives or plasma formation will be 

necessary for concept III. From a chemical reaction perspective, adding energy absorbers into a fuel will 

affect the combustion process for both the bipropellant and monopropellant systems. Fuel additives such 

as aluminum oxide and carbon nanotubes could be considered. A full redesign of turbomachinery, 

injectors, mixers, and combustion chambers may be necessary. Solid thrusters contain alumina as a 

product in the combustion process. Although these particles (both liquid and solid) will readily absorb 

microwave energy at high temperature, they are not good absorbers of visible or near-IR energy. 

Therefore, an additive would be necessary to absorb in typical laser wavelengths. From a complexity 

standpoint, a monopropellant thruster is beneficial over a LOx/LH2 system and will, therefore, be used as 

a representative liquid propellant system. 

C.5.4 Concept IV 

Laser ablative thrusters using solid propellants will be considered. 

C.6 Mission Analysis 

A notional 2,500 kg (total initial mass including propellant) is transferred from a nominal LEO orbit 

with an altitude of 200 km to GEO. The delta-V required to perform this mission is approximately 

4.18 km/s via a Hohmann transfer. The monopropellant system will be assumed to be an impulsive 

maneuver whereas the solar thermal and laser ablation thrusters will be assumed to be nearly constant 

thrust, spiral transfers. For the spiral transfers, a delta-V of 6 km/s is assumed. For the augmented cases, 

the laser energy addition is assumed to occur for the entire burn time of the thruster. 

For the monopropellant thruster, the thrust augmentation is assumed to add to the already existing 

chemical power in the expanding section of the nozzle. The diverging section of the nozzle could be 

envisioned as the collector of the incident energy as well, although providing a nozzle exit diameter of 

10 m might be impractical. The monopropellant system would need to be seeded with an absorbing 

particle (alumina or CNTs). Previous studies suggest that upwards of 50% of the beam energy could be 

converted to useful thrust (Ref. 1). The mass flow of the system is considered constant so that the 

augmentation energy goes into increasing the propellant exit velocity. The monopropellant thruster is 
NASA/TM—2012-217014 573 

assumed to be a scaled engine that could operate at a nominal 1MW of chemically provided jet power. 

The augmented Isp is assumed to be 230 s. 

The augmented solar thermal propulsion system is assumed to operate on hydrogen propellant with a 

specific impulse of 800 s. This specific impulse was chosen so that a reasonable thrust value could be 

obtained for responsive maneuvering from LEO–GEO while still weighing the importance of efficiency. 

The augmented (solely solar collection) and augmented (solely beamed energy collection) thrusters 

include a 10-m diameter collector at the vehicle. The efficiency of the augmented solar thermal thruster is 

assumed to be 80%. 

For the laser ablation thruster, only the 1MW power augmentation from the LEO asset is considered. 

An aluminum propellant is assumed. A specific impulse of 1,500 s is assumed as a compromise between 

efficiency and responsiveness. The laser ablation thruster includes a 10-m diameter propellant bed. It is 

assumed that the thruster is 75% efficient. 

TABLE C.4.—MISSION PARAMETERS FOR AUGMENTED AND AUGMENTED THRUSTER SYSTEMS 

Thruster ΔV

(km/s) 

Power 

(MW) 

Thrust 

(N) 

Isp 

(sec) 

Propellant 

mass (kg) 

Burn time 

(sec) 

Resulting 

payload 

mass to 

GEO (kg) 

Difference versus 

augmented 

thruster, 

percent 

Monopropellant 

(Unaugmented) 

4.2 1.0 886 230 2,111 5,377 389 ----- 

Monopropellant 

(Augmented) 

4.2 1.5 1,085 282 1,953 4,973 547 40.6 

Solar Thermal 

(Unaugmented) 

6 0.11 28 800 1,336 374,058 1,164 ----- 

Solar Thermal 

(Augmented) 

6 0.80 76 2,157 617 156,757 1,883 61.8 

Laser Ablation 6 0.75 102 1,500 837 120,844 1,663 N/Ai

i

Indicates that an augmented thruster would not be able to perform the mission. 

C.7 Conclusions 

Augmented systems, albeit highly conceptual, lead to significant improvements in the metrics used to 

assess their viability for in-space propulsion missions. An augmented monopropellant system could 

potentially increase the payload to GEO by 40% over an augmented system, and an augmented solar 

thermal system could increase the payload mass by 60%. It is also important to note that these missions 

will be responsive with the longest maneuver time. Although similar performance in terms of payload 

mass to orbit could potentially be achieved by traditional electric propulsion systems, the time of transfer 

would be significantly larger. This performance is also being achieved through thrust augmentation via 

beamed energy without a significant penalty to mass of the transferring spacecraft (the most obvious 

advantage to beamed energy propulsion). 

C.8 References for Appendix C

1. Cornella, B., Ketsdever, A., Gimelshein, N., and Gimelshein, S., “Thrust Augmentation of Solid 

Rocket Motors Using Beamed Microwave Energy,” J. Prop. and Power, Vol. 26, No. 5, pp. 1016-

1024, 2010. 

2. Andersen, G.,” Large optical photon sieve,” Opt. Lett., Vol. 30, pp. 2976–2978, 2005.

NASA/TM—2012-217014 575 

Appendix D.—Microwave Energy Issues 

Design Reference Mission 1–C Report 

High Power Microwave Antennas

D.1 Introduction 

This appendix describes a specific type of ground transmitter, implemented with an array of antennas, 

which tracks and provides a high power millimeter-wave beam to launch a microwave thermal rocket 

(MTR) into low Earth orbit (LEO). The transmitter requirements and possible configurations are specified 

along with the background, theory of operation, and design examples. 

This implementation of a high power transmitter is to spatially combine the power radiated from 

many small antennas elements into a single coherent power beam. These antenna elements must be 

properly phased together to deliver the maximum amount of power to a target. Thus, this aperture type is 

termed a phased array antenna. By equally distributing an RF signal generated from a single source to all 

of the antenna elements and phase correcting the signal for each antenna element, a coherent beam is 

radiated into free space. High power radiation is achieved by amplifying the RF signal at each element 

either by solid-state or tube amplifiers. 

D.2 Power Beaming Scenario 

The baseline MTR power beaming scenario is shown graphically in Figure D.1 where the launch 

platform is located 100 lm from the ground transmitter. The rocket ascent trajectory consists of two 

segments. From ground to an altitude of 45 km, the rocket is launched vertically using conventional 

chemical fuel. During this initial phase an onboard MTR beacon signal is broadcast to the transmitter 

ground station. A ground receiver at the transmitter acquires and initiates the tracking of the antenna array 

towards the MTR. At an altitude of 45 km, the ground transmitter begins power beaming at an elevation 

angle of 25° from the horizon. This second phase of the ascent is then powered by the microwave beam 

where the MTR levels off and accelerates horizontally to fly overhead and past the ground transmitter. 

The rocket reaches a maximum altitude of 90 km where the power beaming concludes at an elevation 

angle of 52° from the horizon. 

Figure D.1.—Baseline power beaming configuration and MTR details.

120 km

-100 km 0 km 68 km

Cutoff at 20 g’s 

15 km drop 

launch at 2 g’s

Ground transmitter 

offset from launch 

site by 100 km Power beaming begins 

at 45 km altitude

Power beaming stops 

at 90 km altitude

Frequencies of interest:

95, 140, 170 & 220 GHz

3.3m x 6.7m heat 

exchanger target
NASA/TM—2012-217014 576 

For the transmitter design, the most important parameter is the distance or range to the MTR. The 

range S is calculated with a known trajectory by 

( ) 2 ( )cos( ) 2 2 S  Re  Re  h  Re Re  h  (D1)

where 

Re Radius of the Earth, 6378.14 km 

h Altitude of spacecraft, km 

 Earth central angle from transmitter to spacecraft target, degrees 

The MTR’s slew angle  as viewed from the transmitter is determined by the trajectory and is related 

to the Earth’s central angle  by 





 







     

R h

R

e

e sin sin( ) 1 (D2) 

Figure D.2 shows the baseline trajectory parameters used in this study. Rocket altitudes, range to the 

MTR from the transmitter, and elevation angles as viewed by the transmitter vary from power beaming 

commencement to cut-off where the maximum range is balanced between the two points. 

Figure D.2.—MTR altitude, range, and elevation angle with respect to the transmitter.

Table D.1 provides the key requirements and design variables for the ground transmitter. The most 

important requirement is to achieve a power density of 30 MW/m2 across the MTR’s heat exchanger 

surface throughout the launch trajectory. The variables in this system are the power beaming frequency, 

location of the transmitter, and transmitter size. Because the power density deceases with the square of the 

distance to the MTR, the launch trajectory and location of the transmitter are carefully chosen to provide 

the required power density at the extreme ends of the trajectory path from launch to power beaming cutoff. Additionally, the launch and transmitter locations are ideally located in areas with arid climates and 

high altitudes to minimize the gaseous attenuation through the atmosphere. To this regard, two ground 

transmitter locations identified for this study: (1) China Lake, CA (Latitude 35.94°, Longitude E 242.55°, 

Elevation 1,679m) and (2) the Llano de Chajnantor Observatory in the Atacama Desert in Chile (Latitude 

-23.02°, Longitude W -67.75°, Elevation 5058.7m). 

0

10

20

30

40

50

60

70

80

90

30

40

50

60

70

80

90

100

110

120

-100 -80 -60 -40 -20 0 20 40 60 80

Elevation Angle (deg)

Altitude (km), Range (km)

Downrange Displacement (km)

Altitude

Range

Elevation Angle

Power

Beaming

Stop

Transmitter

Location

Power

Beaming

Start

Slew 

Angle 

Target

Transmitter

Altitude h



Elevation 

Angle 
NASA/TM—2012-217014 577 

TABLE D.1.—POWER BEAMING REQUIREMENTS, VARIABLES, AND ASSUMPTIONS 

Requirements Variables Assumptions 

 30 MW/m2 minimum power density 

at spacecraft’s heat exchanger 

 ≤ 0.5dB loss in power density at the 

edges of the 3.3m x 6.7m heat 

exchanger 

 Spacecraft trajectory with respect to 

the transmitter shown in Figure D.2 

 Power beaming frequency

95, 140, 170, and 220 GHz 

 Transmitter location

China Lake, CA and Atacama Desert, Chile

 Transmitter size

 Output power

 Power beaming occurs under 

clear sky conditions to avoid 

atmospheric attenuation due to 

rain and clouds

Figure D.3 compares the specific atmospheric attenuation in dB/km at the two ground station 

locations (Ref. 1). Although both sites have low specific attenuations, the Atacama Desert is lower due to 

its higher elevation. The attenuation peaks at 60 GHz and 119 GHz are caused by oxygen resonances, and 

the peaks at 22 GHz and 183 GHz are caused by water vapor resonances. Obviously, these frequencies 

are avoided for this power beaming application. 

Figure D.3.—Clear sky atmospheric specific attenuation at China Lake, CA and the Llano de 

Chajnantor Observatory in the Atacama Desert, Chile.

Using the MTR trajectory information in Figure D.2 and specific attenuations in Figure D.3, Figure 

D.4 shows the atmospheric attenuation comparison at the four frequencies of interest for the two 

transmitter locations. China Lake has higher atmospheric loss at the higher frequencies due to its lower 

elevation. In both locations, the highest loss occurs at a downrange displacement of -100 km the power 

due to the lower elevation look angles from the transmitter (see Figure D.2). 

Figure D.4.—Clear sky atmospheric attenuation at China Lake and Atacama.

0.001

0.01

0.1

1

10

100

0 25 50 75 100 125 150 175 200 225 250

Attenuation (dB/Km)

Frequency (GHz)

China Lake

Atacama

0

0.5

1

1.5

2

2.5

3

3.5

4

-100 -80 -60 -40 -20 0 20 40 60 80

Attenuation (dB)

Downrange Displacement (km)

China Lake

95 GHz

140 GHz

170 GHz

220 GHz

0

0.5

1

1.5

2

2.5

3

3.5

4

-100 -80 -60 -40 -20 0 20 40 60 80

Attenuation (dB)

Downrange Displacement (km)

Atacama

95 GHz

140 GHz

170 GHz

220 GHz

China Lake

Water Vapor Density: 2.68 g/m3 

Temperature: 13.4°C 

Barometric Pressure: 1029.5 hPa 

Atacama

Water Vapor Density: 0.52 g/m3 

Temperature: -1.3°C 

Barometric Pressure: 505.2 hPa
NASA/TM—2012-217014 578 

Again using the MTR trajectory information in Figure D.2 and specific attenuations in Figure D.3, the 

radiated power densities at the MTR are shown in Figure D.5 at the four power beaming frequencies and 

at the two transmitter locations. Power density is calculated using the far-field equation 

2 4 R

EIRP Pdens

  (D3)

where EIRP is the transmitter’s effective isotropic radiated power (i.e., the RF power multiplied by the 

antenna gain) and R is the range to the target. 

At 95 GHz and 140 GHz, the power densities are relatively balanced at power beaming start and stop 

points. At 170 GHz and 220 GHz, there is a noticeable difference in the power density curves attributed to 

the different atmospheric losses between the two ground locations. In all cases, the power density peaks 

around the transmitter’s zenith due to the shortest range and lower atmospheric attenuation. 

(a) (b)

 (c) (d)

Figure D.5.—Power densities at the MTR from power beaming turn-on to turn-off at (a) 95 GHz, (b) 140 GHz, 

(c) 170 GHz, and (d) 220 GHz. 

The derived EIRPs used to create these power density curves are given in Table D.2. They range from 

187 dBW to 190 dBW to provide a minimum of 30 MW/m2 over the trajectory range. 

20

30

40

50

60

70

80

90

100

-100 -80 -60 -40 -20 0 20 40 60 80

Power Density (MW/m2)

Downrange Displacement (km)

95 GHz

China Lake

Atacama

20

30

40

50

60

70

80

90

100

-100 -80 -60 -40 -20 0 20 40 60 80

Power Density (MW/m2)

Downrange Displacement (km)

140 GHz

China Lake

Atacama

20

30

40

50

60

70

80

90

100

-100 -80 -60 -40 -20 0 20 40 60 80

Power Density (MW/m2)

Downrange Displacement (km)

170 GHz

China Lake

Atacama

20

30

40

50

60

70

80

90

100

-100 -80 -60 -40 -20 0 20 40 60 80

Power Density (MW/m2)

Downrange Displacement (km)

220 GHz

China Lake

Atacama
NASA/TM—2012-217014 579 

TABLE D.2.—REQUIRED TRANSMITTER EIRPS TO ACHIEVE 

A MINIMUM OF 30 MW/M2 AT THE MTR 

Frequency (GHz) EIRP (dBW) 

China Lake Atacama 

95 187.5 187.2 

140 188.1 187.2 

170 190.4 187.6 

220 190.2 187.6 

D.3 Transmit Phased Array Antennas 

A transmit phased array antenna consists of multiple antenna elements that spatially combines the RF 

power from each radiating element. This property is due to the phase coherency of the RF signal through 

the antenna’s beam forming network. When the phases of each antenna element are exactly aligned and 

each antenna radiates equal levels of RF power, the transmitted effective isotropic radiated power (EIRP) 

is N2

GP where N is the number of antenna elements, G is the gain of a single antenna, and P is the 

radiated RF power from a single element. If the antenna element phases are random (i.e., non-coherent), 

the anticipated EIRP is NGP but it can vary from zero to N2

GP. To achieve the N2 effect on EIRP, the 

antenna elements must be properly phase calibrated. 

Phased array antennas have long been used for radar, communication, and radio-astronomy 

applications. Advantages and disadvantages of phased arrays are given in Table D.3. 

TABLE D.3.—PHASED ARRAY ANTENNA ADVANTAGES AND DISADVANTAGES 

Advantages All Electronic Array Disadvantages All Mechanical Array Disadvantages 

 Fast Beam Agility—Electronically beam 

steering can be accomplished in microseconds 

 Beamforming Complexity  Beamforming Complexity 

 Graceful Degradation  Scan Loss  Tracking Speed 

 Scalability  Cost  Mechanical Wear 

 Multiple Independent Beams 

To meet the high power density demands of the MTR, an all mechanical steered phased array of dish 

antennas is chosen to provide a high EIRP. To achieve phase coherency throughout the array, a single RF 

source is distributed to each dish antenna where phase corrections are applied to correct for spatial 

differences, Doppler shifts, electrical errors, and temperature effects. This relatively low power RF signal 

at each antenna is then amplified using high power amplifying tube such as a gyroklystron. Because the 

dish antennas steer mechanically, the antennas are spaced to avoid blocking the view to the MTR from 

each other. Because the antennas are electrically large (i.e., larger than a wavelength), large grating lobes 

will form if the antennas are spaced in a uniform pattern. These grating lobes reduce the total RF power in 

the main beam and therefore reduces its power density at the MTR. Thus, the relative location of each 

antenna with respect to each other should be placed in a manner to suppress these undesirable grating 

lobes. 

D.3.1 Uplink Array Background 

The first demonstration of arraying large dish antennas occurred at the JPL Goldstone facility in 1993 

at X-band (Figure D.6) (Refs. 2 and 3). Using a two-element array of 34 m diameter dishes, signal 

combining efficiencies of 98% at 7.19 GHz were achieved while tracking low-earth orbiting (LEO) debris 

using time-overlapped radar pulses. Both antennas tracked debris from about 10° elevation at signal rise 

to 4° elevation at signal set under varying weather conditions (e.g., hail falling on one antenna). JPL 

demonstrated this technique using 5 kW peak pulses at 100s pulse widths and 50 pulses/sec. Both 

antennas were fed from a common oscillator source, and a common receiver was used to measure in real 

time the phase-path differences from the antennas from signals reflected by orbital debris. Conjugate 
NASA/TM—2012-217014 580 

phases incorporating the predicted geometry path lengths were then applied through a phase locked loop 

for the uplink signal. Thus, tracking was accomplished retrodirectively on the orbiting targets. 

The phased locked loop schematic and receive pulse structure are shown in Figure D.7. One of the 

antennas was designated as the master whose phase was set to the system phase reference. The second 

antenna was designated as the slave whose phase was adjusted by phase control software so that the 

signal at the target was coherent with the master. A master oscillator is used to drive the master antenna 

and the numerically controlled oscillator (NCO) of the slave transmitter. Doppler compensation is applied 

to the common receiver local oscillator. A pulse operation table (POT) controlled the pulse separation, 

pulse width, transmit trigger, receiver data inputs to the phase loop, application of the output of the phase 

loop to the phase of the next pulse, and offsetting both the timing and phase by predetermined values. 

The master and slave pulses were typically offset by 50 s to form a pulse 150 s total length with the 

first 50 s being from the master, the middle composed of both master and slave, and the final 50 s from 

the slave. This receive pulse structure was measured in actual experiments when the pulses were in phase 

on the target. 

Figure D.6.—JPL’s 1993 uplink phased array demonstration (Ref. 2).

Figure D.7.—Uplink array phased locked loop schematic and receive signal structure (Ref. 2).

Master 

Slave 

204 m
NASA/TM—2012-217014 581 

JPL later followed with a second uplink demonstration on June 27, 2008, also at the Apollo antenna 

cluster at Goldstone (Refs. 4 to 7). Using an array of three 34m diameter antennas, a 7.18 GHz signal was 

coherently combined to the Extrasolar Planetary Observation and Characterization/Deep Impact Extended 

Investigation (EPOXI) spacecraft (Figure D.8). The received signal power level at the spacecraft was 

recorded at the theoretical limit of 9 times over the power level from a single 34m antenna. Using righthand circular polarized signals, EPOXI measured power gains of 6dB and 9.5dB over a single antenna for 

both 2-element and 3-element arrays. Figure D.8 shows the 3-element array configuration where antenna 

Deep Space Station (DSS)-25 is the reference antenna and the other two antennas are slaved to its 

reference phase. Although the radiated power from each 34m antenna could be as high as 20kW, the 

power was reduced to 2kW to avoid saturating EPOXI’s receiver. 

Figure D.8.—JPL’s 2008 uplink phased array demonstration (Ref. 4).

Figure D.9 shows the measured received power over time as recorded on EPOXI’s carrier lock 

accumulator (CLA). At SCET (spacecraft event time) 13:36, the received power increased 9.2 dB with 3 

transmitting antennas which is very close to the theoretically 9.5 dB limit (i.e., 20 log(3) =9.5 dB). 

Figure D.9.—EPOXI’s measured receive power from the uplink phased array demonstration (Ref. 5).

DSS-25 

reference 

DSS-24 

DSS-26 

258 m 

500 m 

301 m 

9.2 dB 
NASA/TM—2012-217014 582 

The arrayed power distribution in the far-field is the product of a single antenna with the array factor 

of all three antennas. As seen in Figure D.10, these widely separated antennas create an interference 

pattern within its main beam from two or three antenna point sources located at each antenna phase center 

(Ref. 8). The narrowest fringe space, in pattern (b), corresponds to the widest antenna separation of 500m 

between antennas DSS 24 and DSS 26. For this power beaming application where all the power needs to 

be concentrated on the launch vehicle single point, these grating lobes are an undesirable trait. The key 

attribute from this EPOXI experiment is the coherent power combining from an array of 3 antennas while 

tracking the spacecraft. 

Figure D.10.—Simulated far-field radiation patterns of two and three 34 m diameter antennas at the distance of the 

moon: (a) antennas DSS 25 and DSS 26, (b) antennas DSS 24 and DSS 26, (c) antennas DSS 24 DSS 25, and 

(d) three-antenna far-field intensity pattern (Ref. 8).

To track the satellite, DSS-25 transmitted a carrier tone that was received and acquired by EPOXI to 

ensure that’s onboard high gain antenna was pointed towards Earth. EPOXI then transmitted an X-band 

downlink carrier that was acquired by DSS-25’s receiver for recording telemetry from the spacecraft. The 

phase of this downlink signal was carefully monitored and distributed to the two other transmitters. 

Slowly varying phase drifts in the ground system were monitored and controlled by the Phase Comparator 

and Control Assembly (PCCA) that is located in the Signal Processing Center (SPC)-10 located 16 km 

from the Apollo Complex antennas. The PCCA contained a signal distribution assembly (SDA), two 

phase comparator assemblies (PCAs) to measure round-trip and cross-phase from all three Apollo 

antennas, and a phase modulation assembly (PMA) that can be used to add correction phases to either the 

DSS‐24 or the DSS‐26 carriers (Figure D.11). 
NASA/TM—2012-217014 583 

The ground system consists of the X-band exciters at SPC-10, X-band couplers and a “round-trip” 

PCA at SPC-10, optical fibers for signal distribution to the transmitting antennas, X-band couplers at the 

output of the power amplifiers at each antenna, and additional optical fibers to return the coupled signal 

samples to SPC-10 for comparison. The two-way optical fiber distribution network to and from the 

antennas is located in the same bundle for most of the 16 km distance from SPC-10 to the Apollo cluster, 

ensuring similar thermal behavior for the outgoing and returning signals. At the Apollo Station, the 

individual fibers are broken out from the common bundle and routed to their respective antennas, 

typically a distance of a few hundred meters, over which the fibers may experience independent thermal 

environments. Following power amplification at the antenna pedestal room, a small fraction (–54 dB) of 

the amplified X-band signal is coupled off the waveguide and routed back to SPC-10, where the phase of 

this sampled signal is compared to the transmitted phase using the real-time PCA. The inputs to the PCA 

are the outgoing (reference) and return (sample) signals from the Apollo antennas. The PCA outputs 

consist of complex samples of equal magnitude representing the phase difference between the reference 

and sample of each antenna (round-trip phase) that is used to measure changes in total path length or the 

difference phase between the reference outputs of two different antennas at SPC-10 (cross-reference 

phase). 

Figure D.11.—PCCA block diagram for controlling the phase of the distributed uplink signal (Ref. 4).

D.3.2 Array Pattern Theory 

The pattern of a phased array antenna is composed of the element pattern multiplied by the array 

factor. To model the element pattern of a dish antenna, the Hansen one-parameter distribution for a 

circular aperture is selected (Ref. 9). This distribution is chosen based on its efficient taper and selectable 

sidelobe levels (SLLs). For example, a 1.86-m diameter dish antenna with a 25 dB SLL operating at 

95 GHz has an element pattern given in Figure D.12. A Hansen parameter of 0.89 provides this sidelobe 

level with an 87% taper efficiency. 
NASA/TM—2012-217014 584 

Figure D.12.—Element antenna pattern of a 1.86m dish antenna at 95 GHz. 

The radiated field pattern of a square array of M antennas is given by 

 (D4)

where xi and yi denote the locations of the antenna elements,  is the angle from the aperture boresight, 

and  is the azimuth angle (Ref. 10). The antennas are spaced on a 2 m square grid, and the array factor 

pattern is shown in Figure D.13. Many grating lobes are created with this spacing. 

 

Figure D.13.—Array of 9 antenna elements separated by 2m and array factor pattern at 95 GHz. 

0 0.1 0.2 0.3 0.4 0.5

0

10

20

30

40

50

60

70

Angle (degrees)

Element Power Pattern (dBi)

Main Beam:

1.86m Diameter

64.7 dBi Gain

0.114° HPBW

25 dB SLL

s( ) φ θ

0

M

i 0

M

j

e

j k sin( ) θ xicos( ) φ yj    sin( ) φ











 2 0 2

 2

0

2

X-Direction (m)

Y-Direction (m)

0 0.1 0.2 0.3 0.4 0.5  10

 5

0

5

10

Angle (degrees)

Array Factor Power Pattern (dB)
NASA/TM—2012-217014 585 

When the element pattern and array factor are multiplied with each other, the resulting power patterns 

at three pattern slices are shown in Figure D.14. As seen, the grating lobes are suppressed, but at the 0° 

azimuth pattern cut reveals a sidelobe that is only 4 dB below the main beam. 

Figure D.14.—Power patterns of 9-element array of 1.86 m dish antennas at 95 GHz. 

One method of grating lobe suppression is to proper antenna element spacing and amplitude tapering. 

The array factor should have one complete cancellation null and it should fall between the first null of the 

element factor. However, the antenna spacing is dictated by the maximum steering angle from zenith, 

and this criteria may not be met. 

Another issue that surfaces with an electrical large aperture operating at millimeter-wave frequencies 

is the far-field patterns. If the target, in this case the MTR, is in the near-field of the array, the on-axis 

beam power density could be much less than predicted by the far-field equations. As Figure D.15 

indicates, the power radiated from an aperture is contained in a corrugated tube whose mean diameter is 

the same as the aperture (Ref. 11). This tube extends to roughly twice the hyperfocal distance of L2

/2

where L is the aperture’s largest dimension. At L2

/ the far-field pattern forms and simple closed-form 

equations for calculating power densities (e.g., Equation (DD3)) or half-power beamwidth (HPBW). 

Figure D.15.—Power flow diagram from an aperture antenna (Ref. 11).

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09 0.1

50

55

60

65

70

75

80

0 degrees

45 degrees

60 degrees

Angle (degrees)

Array Power Pattern (dBi)

Main Beam:

74.3 dBi Gain

0.037° HPBW

4 dB SLL

R = L2 / 

HPBW

L

Aperture
NASA/TM—2012-217014 586 

Because the phases of each antenna can be controlled, the beam can be focused within the radiating 

near-field region. The radiating near-field region from a large aperture is defined as the distance  to the 

far-field distance 2L2

/ from the aperture where L is the aperture’s largest dimension (Figure D.16) 

compares the calculated on-axis power densities at distances from 0.02L2

/ to 2L2/ for a square 

aperture. For an unfocused beam, the on-axis power density undulates due to the beam broadening 

occurring as the beam is being formed. However, the focused beam on-axis power densities follow the 

predicted power densities using the far-field Equation (DD3). Thus, it is possible to use the far-field 

equations for power densities and HPBWs with the assumption that the radiated pattern is focused on the 

target in the near-field. 

Figure D.16.—Comparison between on-axis power densities in the near-field for a square aperture with a focused 

beam, an unfocused beam, and a comparison to the far-field power density Equation (DD3) (Refs. 11 and 12).

D.3.3 Tracking 

One technique for the phased array to accurately tracking the MTR is to employ retrodirective beam 

pointing. Retrodirectivity was demonstrated in the 1993 JPL experiment discussed in Section D.3.1. A 

retrodirective array is an array of widely-separated (separations much greater than a wavelength) antennas 

that each receive and detect the phase of a beacon signal from the target to be tracked. As shown in 

Figure D.17, each antenna conjugates, or reverses the sign of the beacon signal, amplifies it, and 

re-transmits it towards the spacecraft so that the signal retraces its path and converges on and around the 

beacon. 

0

5

10

15

20

25

30

35

40

0.01 0.1 1

Power Density (dB)

R/(2L2/)

Focused

Unfocused

Far-Field Extrapolation
NASA/TM—2012-217014 587 

Figure D.17.—Retrodirective tracking scheme.

The beacon broadcasts a signal at frequency f of the form 

    2 0 VB t  Acos ft  (D5)

that is received by all array elements, where A is the signal amplitude and 0 is an arbitrary phase. As 

illustrated each array antenna is equipped with separate receive and transmit antennas in order to maintain 

high isolation between transmit and receive signal paths. The signal received by the k

th element of the 

array at time t is proportional to that radiated by the beacon at an earlier time t – rk /c, where rk is the 

distance from the beacon to the receive antenna and c is the velocity of light 

        k t

k

R V t  Bcos 2 f t r /c 0  Bcos 2 ft  (D6)

where ft = -2frk /c + 0 is the accumulated phase of the received signal. The received signal is processed 

by a phase conjugator, which simply reverses the sign of the phase, so that after amplification by the highpower microwave amplifier, the signal transmitted by the k

th element of the array is of the form 

        0 V t Ccos 2 ft t Ccos 2 f t  rk /c  k

T (D7)

The high-power signal incident on the MTR is transmitted at an earlier time t – rk/c by the k

th element 

of the transmitter array so that 

     

  0 cos 2

cos 2 /

   

    

D ft

V t D f t r c k t

k

HP (D8)

f1 

Phase 

Conjugator

RF Beacon 

  0 Acos 2ft 

    0 Bcos 2f t  rk / c 

    0 C cos 2f t  rk / c 

  0 Dcos 2ft 

Transmitting (High Gain) Antenna 

Receiving Antenna 
NASA/TM—2012-217014 588 

that is the same for all array elements. The radiated power from each array element arrives at the MTR 

with the same phase. The received power adds vectorially to yield a power density exceeding that from a 

single element by a factor of N2

 where N is the number of array antennas 

D.3.4 Configurations 

Phased array antennas have long been used for radio-astronomy applications for receiving images of 

distant targets. Figure D.18 shows an example of a 84-element radio telescope array operating at 17 GHz 

and 34 GHz. Unlike uplink transmit phased arrays, signals can be coherently combined in these receive 

phased arrays by post processing the phases of the received signals. For large transmit phased arrays, the 

individual antenna elements must be phased calibrated to single reference antenna as described earlier. 

Figure D.18.—84-element Nobeyama Radioheliograph array for 17 GHz and 34 GHz reception. Each antenna is 

80cm in diameter (Ref. 13).

For the array to avoid blockage from adjacent antennas while steering towards the rocket, the 

minimum antenna separation is dependent on antenna diameter and minimum elevation angle. From the 

baseline trajectory of Figure D.2, the minimum elevation angle of 25° occurs at the downlink range of 

–100 km at power beam turn-on. The minimum separation is given by 

sin( )

1.1

  D Smin (D9) 

where the 1.1 factor avoids scattering effects from the adjacent antenna when pointed at the minimum 

elevation angle, D is the antenna diameter and  is the elevation angle with respect to the horizon. Driven 

by this minimum spacing requirement, grating lobes will be very high with respect to the main beam for a 

uniform grid of antenna separations. Stadter et al., studied the effect of randomizing the locations of a 

large array of dish antennas. Figure D.19 shows one result of an array of 217 dish antennas separated by 

21.7 m and operating at 7.2 GHz. The individual locations of each 3.8 m diameter antenna was varied 

from one half-wavelength up to 500 half-wavelengths, and the array pattern was monitored for 

suppressing the array factor grating lobes. When the locations were randomized up to 500 halfwavelengths, the grating lobes were substantially suppressed. This technique would be recommended to 

suppress grating lobes for this application. 
NASA/TM—2012-217014 589 

Figure D.19.—Suppression of array grating lobes by randomizing the location of each antenna (Ref. 14). Each 3.8m 

antenna location is varied from one half-wavelength to 500 half-wavelengths. Element patterns of the 3.8 m antenna 

are shown in blue, and the array patterns are shown in red.

With the assumptions that grating lobes can be suppressed from antenna element location 

randomization and the power beam can be focused in the near-field, a comparison of the different 

frequencies can be evaluated. Table D.4 shows assumptions made in this evaluation to determine the 

array sizes to meet the EIRPs provided in Table D.2. 

TABLE D.4.—PHASED ARRAY ANTENNA INPUT PARAMETERS 

Parameter Value Comments 

Taper Efficiency 87% 25 dB SLL One-Parameter Hansen Taper 

Dish Surface Efficiency 80% Ruze Equation (Ref. 15)

Array HPBW 0.001 degrees Provides a 41m HPBW at the MTR at zenith 

Antenna Output Power 500 kW Radiated power from each dish antenna 

Figure D.20 shows the calculated results to size the array for the very large EIRP requirements. 

Although the array diameter reduces as the frequency increases, the number of dish antennas increases 

with frequency. The antennas are spaced using Equation (DD9), and the required surface accuracy is also 

determined using the Ruze equation. 
NASA/TM—2012-217014 590 

Figure D.20.—Antenna array summary for meeting the 30 MW/m2 power density requirement at the MTR.

D.4 Conclusions 

The two overriding technical challenges to a phased array antenna approach are the following: 

1. Phasing possibly thousands of dish antennas to a single reference that tracks the MTR in the 

near field. Dynamic focusing would be required to keep the beam focused on the moving 

MTR at approximately 4 m/second. 

2. Developing a 500 kW amplifying tube to generate the needed RF power output of each 

antenna. Current millimeter-wave tubes are produced with 10kW of output power, and an 

order of magnitude higher power is needed in this application. 

These barriers could be overcome with future technology development and demonstrations. The 

optimum array configuration also depends on the cost of its components, and future studies should 

consider the cost impact of output power with individual antenna size. 

 

Frequency (GHz): 95 140 170 220

Wavelength (m): 0.0032 0.0021 0.0018 0.0014

Antenna Spacing (m): 4 2.5 1.7 1.3

Antenna Spacing (): 1268 1167 964 954

Antenna Diameter (m): 1.86 1.16 0.79 0.60

Required RMS Surface Accuracy (m): 119 80 66 51

Antenna HPBW (deg.): 0.114 0.123 0.150 0.151

Antenna Gain (dBi): 63.8 63.1 61.4 61.3

Antenna EIRP (dBW): 120.8 120.1 118.4 118.3

Array Diameter (m): 211 143 118 91

No. of Array Antennas: 2190 2580 3790 3870

Array Gain (dBi): 97.2 97.2 97.2 97.2

Array Pout (dBW): 90.4 91.1 92.8 92.9

Array EIRP (dBW): 187.6 188.3 190.0 190.1
NASA/TM—2012-217014 591 

D.5 References for Appendix D 

1. “Attenuation by atmospheric gases,” Recommendation ITU-R P.675.5, 2001. 

2. R.M. Dickinson, D.L. Losh, R.D. Barber, and J.K. Dempsey, “A phase-control approach for a largeelement coherent microwave power uplink system,” IEEE Trans. Antennas and Propagation, vol. 47, 

no. 3, pp. 487-495, March 1999. 

3. W.A. Imbriale, “Large Antennas of the Deep Space Network,” Monograph 4, Jet Propulsion 

Laboratory, California Institute of Technology, Chapter 9, Feb. 2002, 

http://descanso.jpl.nasa.gov/Monograph/series4_chapter.cfm. 

4. V.A. Vilnrotter, P.C. Tsao, D.K. Lee, T.P. Cornish, L. Paal, and V. Jamnejad, “EPOXI uplink array 

experiment of June 27, 2008,” IPN Progress Report 42-174, pp. 1-25, Aug. 15, 2008. 

5. V. Vilnrotter, D. Lee, T. Cornish, P. Tsao, L. Paal, and V. Jamnejad, “Uplink array concept 

demonstration with the EPOXI spacecraft,” IEEE A&E Systems Magazine, pp. 29-35, May 2010. 

6. P. Tsao, V. Vilnrotter, and V. Jamnejad, “Pointing-vector and velocity based frequency predicts for 

deep-space uplink array applications,” in 2009 IEEE Aerospace Conference, March 7-14, 2009, pp. 

1-5. 

7. F. Davarian, “Uplink arrays for the deep space network,” Proceedings of the IEEE, vol. 95, no. 10, 

pp. 1923-1930, Oct. 2007. 

8. V. Vilnrotter, D. Lee, R. Mukai, T. Cornish, and P. Tsao, “Three-antenna doppler-delay imaging of 

the crater tycho for uplink array calibration applications,” IPN Progress Report 42-169, pp. 1-17, 

May 15, 2007. 

9. R.C. Hansen, “Antennas,” in Reference Data for Engineers: Radio, Electronics, Computer, and 

Communications, M.E. Van Valkenburg and W.M. Middleton, Eds., 9th Edition, Boston: Newnes, 

2002, p. 32-53. 

10. E. Brookner, Practical Phased Array Antenna Systems, Lexington, MA: Lex Book, 1997, p. 2-20. 

11. R.C. Hansen, “Focal region characteristics of focused array antennas,” IEEE Trans. Antennas and 

Propagation, vol. 33, no. 12, pp. 1328-1337, 1985. 

12. R.C. Hansen, Microwave Scanning Antennas. Vol. 1 Apertures, Boston: Academic Press, 1964, pp. 

33-46. 

13. http://solar.nro.nao.ac.jp/norh/html/introduction.html

14. P.A. Stadter, B.L. Kantsiper, D.G. Jablonski, A.R. Golshan, and J. Costrell, “Uplink arraying 

analysis for NASA’s deep space network,” in 2010 IEEE Aerospace Conference Proceedings, Big 

Sky, Montana, March 6-13, 2010, pp. 1 - 6. 

15. J. Ruze, “Antenna Tolerance Theory – A Review,” Proceedings of the IEEE, vol. 54, no. 4, pp. 633-

640, 1966. 

NASA/TM—2012-217014 593 

Appendix E.—ABL Experiment 

White Paper 

Beamed Energy Propulsion Demonstrations 

Using The Airborne Laser Test Bed 

E.1 Background 

This paper discusses using the Airborne Laser Test Bed (ALTB) as the energy source for a series of 

low-cost experimental demonstration vehicles that use Beamed Energy Propulsion (BEP) engines. It will 

address at a top level how the ALTB operates, its capabilities and its constraints as well as basic 

demonstration configurations. 

The BEP demonstration vehicles will explore key BEP technologies with the purpose of expanding 

the knowledge base and envelope of understanding. 

E.2 Benefits of Demonstration 

Beamed Energy Propulsion uses a beam of energy, either laser or microwave to add energy in the 

form of heat to a working fluid. The heated working fluid is then expelled and creates thrust to propel the 

vehicle. Most rocket engines use both a fuel and an oxidizer, which are burned to generate the energy 

needed for thrust. A benefit of using BEP is the fuel can be a single fluid and be as benign as water. The 

fuel stills needs to be heated up, but instead of using a combustion process to add energy to the system, 

the energy is beamed to the vehicle from an external source. The external source of the beamed energy 

can be amortized over a large number of uses therefore reducing the total cost per use. 

E.3 History of BEP 

Konstantin Tsiolkovsky published the original idea for BEP in 1924, he pointed out that energy could 

be delivered to a space rocket by means of tight light beams (first laser wasn’t built until 1960). The 

modern history of BEP, started in 1972, when Arthur Kantrowitz first popularized the idea of using lasers 

for propulsion. 

E.4 History of the ALTB 

The Airborne Laser (ABL) is technology demonstration that grew out of the first Gulf War SCUD 

hunts. Air Force pilots would report seeing SCUDs launch but had no way of engaging them. The ABL 

program contract was signed in 1996 and in Feb 2010 that aircraft shot down its first boosting threatrepresentative missile. With the declaration of Defense Secertary Gates that the ABL was not moving 

forward to an acquisition program but to remain a science and technology research platform, aircraft was 

renamed the Airborne Laser Test Bed (ALTB). The new designation was to reflect that it was a national 

asset that was to be used to learn and understand high-energy lasers, their effects and uses. The Missile 

Defense Agency’s Advanced Technologies Directorate (MDA/DVL) currently manages the ALTB. 

E.4.1 How the ALTB Works 

The ALTB is comprised of four main subsystems; the mega-watt class high-energy laser (HEL) 

which uses chemical oxygen iodine laser (COIL) technology, the Beam Control/Fire Control (BC/FC) 

system, the Battle Management system and the 747 aircraft. These subsystems are integrating into a 

seamless capability that can detect, track, target and kill boosting ballistic missiles. 
NASA/TM—2012-217014 594 

The ALTB method of engagement is as follows: The Battle Management’s infrared search and track 

(IRST) sensors detect the plume of a launching missile. The BC/FC slews the turret to optically find the 

target plume. Once the system locks onto the target optically the BC/FC fires a kilowatt-class tracking 

illuminator laser (TILL) to determine the targets range. With the range determined, the TILL now 

“walks” up the missile body to find the nose and then continues to track the target from the nose. The 

BC/FC now fires a second kilowatt-class laser that measures the atmospheric turbulence. This laser, the 

beacon illuminator laser (BILL) is offset behind the TILL but just forward of where the HEL will hit the 

target. This offset is to allow the system to measure the turbulence along the path of were the HEL will be 

when it is fired. As soon as the system as determined that it has a solid lock on the target and the 

turbulence information the HEL fires until the target is destroyed. The whole process takes only seconds. 

E.4.2 Capabilities 

The ALTB is capable of tracking boosting missiles and aircraft that have a sufficient heat signature 

such as the modified Gulfstream use as an Airborne Diagnostics Target (ADT). The ALTB can then put 

and hold an atmospherically compensated HEL beam on that target. A surrogate high-energy laser 

(SHEL) is used on the ADT. The SHEL has the same wavelength as the HEL but is only about 10 watts. 

The ALTB specific technical capabilities required to design BEP demonstration vehicles are either 

“For Official Use Only” or classified but can be obtained through proper channels from the Missile 

Defense Agency’s program office MDA/DVL. 

E.4.3 Constraints 

In order for the demonstration vehicles to be lased by the ALTB they will need to meet certain 

physical criteria and dimensions. The software in the ALTB can be adjusted to fine tune the exact aim 

point of the HEL on the demonstration vehicle. Below are some of the design constraints. 

 Demonstration vehicle needs enough (TBR) infrared signature to emulate a boosting ballistic 

missile. 

 Distance x ± y (TBR) between IR source and leading point (nose cone). 

 Distance x ± y (TBR) between leading point and where the HEL will hit. 

In addition, there are airspace consideration and constraints that must be taken into account when 

designing the vehicle and the engagement geometry. The best engagement field of view is a cone ~xxo 

from the forward centerline. Based on the ALTB operating environment and the FAA, the engagements 

need to take place above 40,000 ft. Depending on which range the tests are conducted on there is also an 

altitude at which the beam must be above when it leaves the controlled airspace. 

Engagement windows and times must be coordinated with US Strategic Command’s Laser 

Clearinghouse, the FAA and the host range. This can limit the amount of desired testing available and 

needs to be worked continuously. 

 As part of the Laser Clearinghouse coordination for engagement windows, an analysis of the 

reflected and scattered laser energy must be completed for each vehicle and geometry. This process can 

take up to six months to complete and may require physical samples of the vehicle to measure its 

reflectivity. 

E.5 Basic Demonstration Vehicle Configurations 

The demonstration vehicles will most likely be launched using a booster rocket or air launched to get 

it into the desired engagement geometry. Depending on which BEP engine is being used on the 

demonstration vehicle it will either have an orthogonal or an axial incidence angle. 
NASA/TM—2012-217014 595 

E.5.1 Orthogonal Incidence 

Engines that have the laser come at the side of the vehicle will be categorized as Orthogonal. 

Depending on the sensitivity of the engine to the change in incidence angles, the vehicle flight path or the 

laser receiver will have to be adjusted to maintain the required incidence angle. 

E.5.2 Axial Incidence 

Engines that have the laser come at the back of the vehicle will be categorized as Axial. Depending 

on the sensitivity of the engine to the change in incidence angles, the vehicle flight path or the laser 

receiver will have to be adjusted to maintain the required incidence angle. 

E.6 Demonstration Risks 

E.6.1 Technical 

One of the key risks that will need to be addressed early in the design of the demonstration vehicle 

will be maintaining the laser incidence angle throughout the engagement. 

The ALTB is still a technology demonstrator and as such is not 100% reliable. Specific GO/NO GO 

criteria will need to be developed to ensure the highest probability of success. 

If an air launched solution is defined for the demonstration vehicle, then the demo vehicle will need 

to be able to “fly” on it own until the carrier craft clears the engagement air space. 

E.6.2 Programmatic 

The ALTB funding is to be reviewed in 2013 to determine if it should continue or be terminated. 

NASA and DARPA will need to work closely with the Missile Defense Agency to ensure that the ALTB 

will be available when the demonstration vehicles are ready for test. 

NASA/TM—2012-217014 597 

Appendix F.—Environmental Considerations for a Ground-Based BEP Station

Daniel E. Raible, Scott Darpel, and Maciej Zborowski 

F.1 Introduction 

Beamed Energy Propulsion (BEP) is a technology that typically employs heat exchangers or optical 

plasma engines to provide thrust to a vehicle. By reducing or eliminating the onboard fuel requirements, 

vehicles may be developed with very high specific impulse (ISP). By providing the propulsive energy 

from a remote location where power is abundant, the mass of the vehicle may be dramatically reduced. 

To date several scaled prototypes of BEP vehicles have been successfully flown, but a full size 

demonstration has yet to be realized. The implementation of the BEP concept to provide launch-to-orbit 

capabilities will require a shift in facility approaches from that of conventional rocket technologies, since 

the fundamental system and safety implications differ greatly. 

Figure F.1 depicts an illustration of a proposed laser-based launch complex, in which multiple laser 

cavities are beam combined through a multi-faceted mirror and delivered to the receiving spacecraft 

through a beam director. 

Figure F.1.—Proposed European Space Agency laser launch facility (reproduced from source 1). 

Traditionally, locating launch facilities has been done with one eye towards orbit, and the other on 

safety. When examining the following figure, depicting the location of all current launch facilities, it is 

easy to see that the primary attribute is having either a large body of water or largely uninhabited land 

available for the departing vehicle’s trajectory. These exclusions zones provide a margin of safety from 

launch mishaps, such as explosions, or even simple exposure to launch site gases and deployables. 
NASA/TM—2012-217014 598 

Figure F.2.—Existing launch sites around the world. 

(1–Vandenberg, 2–Edwards, 3–Wallops Island, 4–Cape Canaveral, 5–Kourou, 6–Alcantara, 7–Hammaguir, 

8–Torrejon, 9–Andoya, 10–Plesetsk, 11–Kapustin Yar, 12–Palmachim, 13–San Marco, 14–Baikonur, 15–Sriharikota, 

16–Jiuquan, 17–Xichang, 18–Taiyuan, 19–Svobodny, 20–Kagoshima, 21–Tanegashima, 22–Woomera) 

While safety will always be a primary concern, the challenges for locating a Beamed Energy Launch 

(BEL) facility will be unique and required more considerations. It is true that a BEL facility will eliminate 

the risk of explosion and inhalation associated with chemical propellants, but the nature of high-energy 

power beams presents its own exposure risks. Beam reflection and refraction can expose personnel and 

bystanders to vision and skin damage. Additionally, the effect of launch site atmospheric conditions on 

the propagation of concentrated energy beams rises to a key consideration. Humidity and atmospheric 

particulate levels can affect the ability to maintain beam integrity. From a utilities perspective, access to 

sufficient power, with multiple ties into a nearby grid is desirable to reduce the need for on-site 

generation, but is of smaller concern. 

Do any of the current launch locations meet all the attributes for a safe, effective and affordable BEL 

Facility? Or, would a completely new set of locations, perhaps some not even fixed, be considered? This 

paper presents the needs for each of these attributes, as well as a discussion of a selection of locations. 

F.2 Weather and Potential Launch Site Suitability 

There are many components of “weather” that affect the suitability of a location, or set of locations to 

be home to a BEL facility. Aerosol, precipitation and cloud obscuration all contribute to meeting a new, 

rapid launch paradigm. An initial survey, based on a collection of weather charts can provide some insight 

into this suitability. A brief note regarding political implications: by constraining the list of potential BEL 

sites to the US and related territories, other ramifications such as international agreements and 

transportation logistics are avoided thus lowering the overall launch costs. Since the future of BEP for 

launch lies primarily on its ability to offer low cost and rapid turnaround capability to orbit, the location 

of the launch facilities must maintain BEP as a competitor to conventional launch technologies. 

Launch site availability is an important metric, since it will impact the site’s ability to repeatedly 

launch on schedule. One primary detrimental impact on launching has always been cloud cover. 

Historically this has been centered on avoiding potential static charge buildup related to more mature 

cloud formations, but BEL introduces an additional dimension since cloud obscuration will have a 

deleterious effect on beam propagation. A first look may be made into the average annual number of 

cloudy days for locations within the continental U.S. (CONUS), as shown in Figure F.2. Here it can be 
NASA/TM—2012-217014 599 

seen that a region exists in the Southwestern area of the country that supports clear day availability with 

obscurations present less than 89.5 days per year. 

 

Figure F.3.—Average annual number of cloud obscuration days across the U.S. (reproduced from 2). 

As expected, locations to the Southwestern U.S., through the dry western desert present a positive 

case for cloud obscuration. This would indicate a high rate of availability for launches where beamed 

columns would not have to navigate clouds, and suffer the associated loss of power. 

Aerosol content in the atmosphere also contributes to the attenuation effects of the beam due to 

absorption and scattering. This is wavelength dependant related to the size of the aerosols, and the type of 

attenuation may be parsed out between humidity and direct rainfall contributions. The average rainfall for 

locations in the CONUS points to some of the same general regions in Figure F.4. 

Figure F.4.—Mean maximum daily precipitation across the U.S. (reproduced from 2). 
NASA/TM—2012-217014 600 

Aerosol content is another contributor to beam attenuation. The capacity for the atmosphere to 

contain moisture is related to the temperature of the air, which can be seen across the U.S. in Figure F.5. 

 

Figure F.5.—Annual mean daily average temperature (reproduced from 2). 

For a more accurate depiction of the atmospheric conditions at a potential location, a focused site 

survey should be conducted at selected candidate positions. Considering a combination of favorable 

attributes including atmospheric availability, remoteness, obtainable power, existing infrastructure, 

accessible trajectories and trajectory several potential sites in the Southwestern region of the U.S. may be 

identified to support BEL. These may include Vandenberg Air Force Base (currently supports 

conventional launch operations), Goldstone, White Sands and China Lake (all government installations 

containing some degree of existing infrastructure), Mauna Kea (high altitude offers a circumvented 

approach to atmospheric losses) and finally a mobile shipboard location that would offer diversity with 

launch inclinations. These sites, shown in Figure F.6, could potentially serve as primary launch location, 

or as a secondary auxiliary downrange beaming station. 

Figure F.6.—Potential BEP ground complex locations 
NASA/TM—2012-217014 601 

F.3 Range Safety Zones 

At a BEP launch complex, a range safety keep-away zone (KOZ) must be established and maintained 

around and beyond an ascending craft, just as in traditional chemical propulsion launch sites. With BEP 

however, these KOZ’s will serve to protect against exposure to reflection, refraction and scattering 

components of the beam exceeding levels considered above the acceptable standards. These KOZ’s will 

be based upon maximum permissible exposure for humans, and damage thresholds for ground, airborne 

and space-based assets within proximity of the BEP activity. These thresholds may be calculated utilizing 

resources established by the American National Standard for Safe Use of Lasers. The nature and type of 

hazard that may be encountered by exposure to the radiated beam will vary as a function of the 

wavelength of the energy used for the launch system. Typical skin photochemical effects (burns) are 

generally limited to tissue absorption of wavelengths shorter than 600 nm (UV and some visible). This 

range is typically less than what would be practically considered for BEP source technologies due to high 

coefficients for atmospheric extinction. For wavelengths between 400-1400 nm (visible and nearinfrared), the ocular absorption hazards would mainly concern retinal burns. At wavelengths between 

1400 nm to 1 mm (mid to far-infrared, the cornea would be the most at risk for damage. Both of the 

regions are favorable for BEP, and so this places eye safety very prominently as a factor to consider when 

implementing the system. 

Considering a laser hazard evaluation for outdoor operation, the components of intrabeam exposure, 

specular and diffused reflections are all key. Additionally, issues such as the stability and operational 

limitations of the laser platform, likelihood of people being within the range during operation, people 

viewing the operation with optical aids such as telescopes and binoculars, and operations through 

airspaces traversed by aircraft must be considered. To perform these calculations, it is necessary to define 

a metric known as the Maximum Permissible Exposure (MPE). This is the maximum level of laser 

radiation to which an unprotected person may be exposed without adverse biological changes in the eye 

or skin. The MPE is generally set with a safety factor of 1/10th the level where there exists a 50% chance 

of injury, based on an empirically derived model. The MPE is dependent on the type of laser 

(wavelength), potential exposure duration, whether eye or skin exposure is anticipated and pulse 

repetition frequency if the laser is operated in pulsed mode. Table F.1 contains the worst-case 

recommended accidental limiting exposure durations for CW and repetitive-pulse MPE calculations as a 

function of wavelength (Ref. 3). From these starting point acceptable durations, the maximum level of 

energy exposed may be calculated. 

TABLE F.1.—EXPOSURE LIMITS BY WAVELENGTH AND EXPOSURE TYPE (REPRODUCED FROM SOURCE 3)
NASA/TM—2012-217014 602 

The Nominal Hazard Zone (NOHZ) for all possible beam paths needs to be evaluated for all cases of 

outdoor operation. The NOHZ can be thought of as an area within which the level of direct, reflected or 

scattered laser radiation during normal operation exceeds the applicable MPE. Generally the potential 

exposure (irradiance) decreases with distance, until a point where the irradiance is less than the applicable 

MPE, and this boundary defines the NOHZ. The most serious potential exposure condition would be 

intrabeam ocular viewing, followed by ocular exposure of specular reflections from the incident beam 

upon the vehicle surface. Although the human eye is most sensitive in the visible region of light centered 

on 550 nm, exposure to non visible optical radiance may also be harmful. Consider source wavelengths 

calculated at both 1.064 and 2.0 μm. For the intrabeam case where the eye is in the direct path of the 

laser beam, both the near infra-red (NIR) and far infra-red (FIR) limiting exposure durations are given as 

10 seconds from Table F.1. The ANSI expressions for establishing the exposure value thresholds at these 

frequencies are given by: 

5.0*CC*10-3 W*cm-2 (NIR 1.050 to 1.400 um) 

0.1 W*cm-2 (FIR 1.800 to 2.600 um) 

Note that the NIR case involves the correction factor CC, which increases the MPE values for ocular 

exposure because of pre-retinal absorption of radiant energy in the spectral region between 1050 and 

1400 nm. 

For BEP it is necessary to dynamically focus the optical energy on the ascending craft, and for the 

purposes of calculating the intrabeam Nominal Hazard Distance (NHD) we consider the ‘Lens-on-Laser’ 

case, which includes properties of focal length and emergent beam diameter: 

Specular (or mirror-like) reflections can occur when the size of the surface irregularities (or 

roughness) is less than the wavelength of the incident radiation, resulting in an angle of reflection equal to 

the angle of incidence. In this scenario the NHZ for an eye in the path of a specularly reflected beam due 

to a mirror-like surface is almost as serious as the direct exposure condition. 
NASA/TM—2012-217014 603 

Figure F.7.—Illustration of specular reflection (reproduced from source 4). 

Diffuse (Lambertian) reflections from randomly oriented irregularities that are greater than the 

wavelength of the incident radiation create a diffused region of reflection. In this case, a diffusely 

reflected portion of the laser beam enters the eye, and this represents the least serious potential exposure 

condition. 

Figure F.8.—Illustration of diffuse reflection (reproduced from source 4). 

The nature of the diffused reflections may act as either a ‘point’ or ‘extended’ source based on the 

geometry of the system, and each case will have a different procedure for calculating the diffused NOHZ. 

The angular size (subtense) of the apparent source as viewed by the observer will determine the type of 

reflection, where anything appearing smaller than α = 1.5 mrad will be considered a point source and 

everything larger will be treated as an extended source. The diffuse reflection NOHZ may be 

characterized by the expression in Figure F.9: 
NASA/TM—2012-217014 604 

Figure F.9.—Diffuse reflection nominal hazard zone 

The diffuse NOHZ will follow the trajectory of the ascending vehicle, and may be traced back to the 

initial point of departure from the ground to define in-part the KOZ. The definition of these regions can 

be used to determine the overall KOZ around the launch area. Safety must be considered for each of the 

following types of observers: 

 Ground observer 

o Unaided eye 

o “Binocular safe” (50 mm aperture) – largest optics for casual observer 

o “Celestron safe” (200 mm aperture) 

 Civil aircraft (low altitude, relatively close) 

 Commercial aircraft (30-40,000 ft altitude, 250 m/s) 

 Spacecraft 

Generally need to consider several possible exposures: 

 Direct exposure to main beam – presumably only an issue for spacecraft 

 “Stray beam” – direct exposure to beam that misses vehicle, by accident or due to sidelobe of 

beam 

 Specular reflection (“Glint”) from vehicle 

o Known (e.g., specular component of reflection from heat exchanger surface) -- 

predictable area, reflectivity, and divergence 

o Accidental -- reflection due to surface damage or other problem that exposes a shiny 

surface 

 Diffuse reflection from vehicle 

o “Lambertian” wide angle scattering 

o “Hot spot”, with significant peaking in a particular direction (usually either near the 

specular direction, or retroreflection toward beam source. 

For extended exposure (which we expect for diffuse reflection – viewers are likely to watch the 

vehicle ascend) and an extended source: 

0.25

2

2 ( / ) 1.8  MPE mw cm  CACET

2( 0.7) 10  CA  for 0.7 – 1.05 μm; CA = 5 for 1.05 – 1.15 μm 

CE accounts for the angular size of the source. For a source which is reasonably uniform (i.e., no 

individual “hot spot” is more hazardous than the overall source) and smaller than 100 milliradians in 
NASA/TM—2012-217014 605 

apparent width (i.e., observer is farther than ~50 meters from the vehicle, which we assume will always 

be the case) CE = max(a/1.5 milliradians, 1) where a is the apparent angular width. 

For an apparently rectangular source of length l and width w, viewed from a distance D, a is defined 

as l w D For typical heat exchanger vehicles, the physical source width is 3-5 meters. So for 

distances greater than a few kilometers, the vehicle is always an effective point source (to the unaided 

eye), but at moderate range, or through a telescope, it may be an extended source. 

T2 is the “integration time” for continuous viewing. The integration time is defined as 10 seconds for 

near IR point sources (so that T2^-0.25 = 1/1.8) but in the 2007 revision of Z136.1, it’s made dependent 

on the angular size of the source, so that the MPE is (slightly) reduced for large extended sources. For a

between 1.5 and 100 mrad, T2 = 10 * 10^(a-1.5 / 98.5), so it varies from 10 to 100 seconds. In practice, 

the observing time at potentially hazardous range will never be 100 seconds for any observer, so the 

calculated T2 overestimates the hazard slightly (makes the MPE too low) in some cases. However, the 

effect is never more than a factor of 1.8 in MPE, and generally much less. 

The actual flux scattered from the vehicle, to which the MPE must be compared, is just 

2  Ptotal cos() D

where α is the albedo (“whiteness”) of the surface, Ptotal is the total laser power hitting the vehicle, and θ

is the viewing angle relative to the normal to the (nominally flat) heat exchanger. The worst case will 

generally be θ = 0 (looking at the heat exchanger face-on) so we’ll just assume cos(θ) = 1. 

The actual heat exchanger will have an albedo as close to zero as we can make it, to absorb as much 

laser energy as possible. However, it’s not possible to get a perfectly black surface. Also, parts of the 

heat exchanger (e.g., where sections join) may be deliberately made white to reduce the local heat load, 

and some fraction of the beam may spill over onto, or accidentally hit, white areas. 

For the purposes of this rough analysis, we assume an effective α of 0.1 – 10% of the incident laser 

power is scattered from the vehicle surface. (Note that it’s easy to monitor the scattered power, so the 

laser can certainly shut down if the scattered power is much larger than expected.) 

See the Figure F.10 for the results. This plot is for 

 P = 100 MW 

 Wavelength = 808 nm (laser diode) 

 l = w = 3 m 

 alpha = 0.1 
NASA/TM—2012-217014 606 

Figure F.10(a).—MPE compared with 3 m BEP exposure as a function of distance. 

So for this case the eye-safe distance is just over 100 meters (111 m, to be exact). The “kink” in MPE 

is at 30 m, where the heat exchanger apparent width is 100 mrad. 

The MPE is lower if the source area is smaller. If the beams were concentrated into a smaller 1 m 

spot, the plot becomes: 
NASA/TM—2012-217014 607 

Figure F.10(b).—MPE compared with 1 m BEP exposure as a function of distance 

In this case the eye-safe viewing distance is ~300 meters, which is still a very reasonable result given 

the large amount of radiated energy. 

For 1.06 um, the MPE is about a factor of 5 higher than at 808 nm. The eye-safe range for a 3 m spot 

and 10% albedo is 45 meters. However, even at 1.06, if the albedo is higher and the spot size smaller, the 

hazardous range can be substantial; this might be the case for the Lightcraft if it tumbles and the beam 

strikes the forebody. For a 1 m spot and 100% albedo (100 MW scattered power) the eye-safe range is 

~800 m. 

F.4 Aided Viewing: Telescopes and Binoculars 

In general, the safe viewing distance for an observer with a telescope of aperture A is less than or 

equal to the naked-eye safe distance, multiplied by A/7 mm = M (for magnification). This is because the 

reference pupil diameter for IR laser safety is 7 mm. A larger aperture increases the collection area by

M2

; increasing the distance to the source by a factor of M decreases the flux by the same factor M2. 

This turns out to be true for an extended source as well as a point source, although the explanation is a 

bit complex; essentially, any optical magnification >M makes the apparent angular extent of the source 

larger; any magnification <M makes the source smaller (reduces the MPE) but also results in the 

telescope exit pupil diameter being bigger than 7 mm, so some light from the source doesn’t enter the eye. 

Therefore, the safe viewing distance for an observer with binoculars (50 mm aperture) is 

approximately 7.1 times the distance for naked-eye viewing, or ~800 meters for the baseline case (3 meter 

source size). 
NASA/TM—2012-217014 608 

The safe viewing distance for an observer with a portable telescope (200 mm aperture) is another 

factor of 4 larger, or about 3.2 km. (In practice, this is conservative, as amateur telescopes have 

substantial obscuration and, unless specially coated, have significantly less than 100% transmission in the 

infrared.) 

F.5 Hot Spots and Non-Lambertian Scatter 

While it’s not hard to ensure that scattered light is reasonably diffuse, ensuring a true uniform 

distribution is difficult. Conservatively, therefore, we would expect to design for some much higher-flux 

reflections. These are likely to be in more or less known directions, most likely aimed at least generally 

back towards the laser source. 

If we assume the peak flux is 10-fold higher than the nominal Lambertian flux, then the naked-eye 

hazard range increases to ~300 meters at 1.06 microns, 550 m at 940 nm, and 965 m at 808 nm – almost 

10-fold more than the nominal ranges. Binocular and telescope safe distances would be similarly 

increased. 

Ten-fold higher than Lambertian flux implies that the entire scattered energy is spread over 

approximately 0.3 steradians, which is very roughly a 30 degree full-angle cone, or a 5 by 180 degree fan. 

This is clearly achievable by shaping surfaces appropriately if necessary; e.g., a very modest degree of 

diffusion in the vertical direction plus curvature of the heat exchanger in the horizontal direction. 

F.6 Extension to KOZ 

For diffuse, approximately Lambertian scattering, and an assumed scattered power of 10 MW from a 

3-meter heat exchanger, the naked-eye hazard range for the laser-launched vehicle is only of order 

100 meters, and even an observer with a sizeable telescope is safe beyond a range of roughly 3 km. 

However, since it’s difficult to ensure that scattering will be uniform, especially from the heat 

exchanger, a conservative assumption pending more detailed analysis is that the keep-out zone should be 

at least 1 km for all observers, and 3 to 7 km for users of binoculars or cameras with through-the-lens 

viewing, depending on the laser wavelength. This is also the safe range for aircraft close to the vehicle, 

since pilots and passengers might use binoculars, but are unlikely to use any larger optics. 

Larger telescopes may pose some hazard out to ranges as much as 25-30 km (for a 200 mm aperture, 

808 nm), although that would require a very unlikely set of circumstances, including very good tracking 

(following the accelerating vehicle without significant jitter for many seconds, using a telescope with high 

infrared transmission, and having the vehicle follow a trajectory and orientation that keeps the hot spot 

centered on the observer, also for many seconds.) 5-10 km is more likely to be the actual “Celestron 

safe” range. 

Beyond ~30 km there is essentially no hazard even to observers with telescopes, so once the vehicle 

reaches 30 km altitude it can be viewed from anywhere. 

An order-of-magnitude comparison may be made between the BEP parameters and the existing range 

safety zone existing at the NASA Kennedy Spaceflight Center’s pad 39A, which until recently was the 

sole launch facility for the remainder of the Shuttle Transportation System (STS) missions. The STS 

safety zone is established at an 8,500 ft radius from the launch pad, so determined by the amount of 

distance required to keep the airborne particulate level of hydrogen chloride at a safe level for inhalation. 

This area has been illustrated by the green circle in Figure F.11. Overlaid onto pad 39B are two red 

boundaries representing the approximate area of a ground-based field array source encircled by a 1 km 

KOZ for unaided viewing. Minus the appropriate safety factor measures, this initial graphical comparison 

demonstrates that the BEP approach presents a KOZ within the same order-of-magnitude of conventional 

launch systems. 
NASA/TM—2012-217014 609 

Figure F.11.—Initial safety zone comparison with conventional established facilities. 

Beyond the immediate launch area, the KOZ needs to also reflect the trajectory of the vehicle during 

ascent. For the STS there are multiple zones to cover the field of uncertainty for the flightpath of the 

Shuttle, the splashdown zones for the solid rocket boosters (SRBs) and the external tank (ET) impact 

zone. These areas are shown at the bottom of Figure F.11, and the total length extends downrange 

approximately 11,000 km. 
NASA/TM—2012-217014 610 

Figure F.12.—Geometric safety zone comparison with conventional established launch facilities. 

During initial BEP trajectory analysis, it is determined that a laser sustainer station would likely be 

placed in the vicinity of 400 km downrange from the launch booster facility. One reason for this is to 

limit the pointing angles of the beams to 50-60 degrees from the vertical, and to maximize the amount of 

power on target by limiting the amount of slant range propagation. Since both laser stations and their 

KOZs occupy a 4 km footprint, a 4 by 400 km rectangular area may be used to define the overall KOZ, as 

shown in the top part of Figure F.12. 

Beyond the sustainer laser system, a KOZ field of uncertainty may be added, but the merits of BEP 

may be further exploited to unconstrain the solution. Since the typical BEP launch vehicles are minimal 

mass and do not contain volatile propellants, in the case of a mid-trajectory ‘failure’ (such as a loss of 

beam track) the beam would shut down, and the vehicle could execute a tumbled deceleration and 

eventual parachute deployment, to result in a relatively uneventful recovery on land or sea. This type of 

abort would eliminate the need for an extended-field KOZ.

F.7 Emitter Interception 

The ability for the BEP system to provide high launch availability and repeatability is based on the 

number of launch window opportunities presented when there exists a beam path sufficiently clear of 

existing orbital assets. This may be analyzed in Satellite Tool Kit (STK) by simulating a repetition of 

BEP launch profiles and overlaying existing satellites to examine beam interception events. 

The beam divergence may be estimated by assuming a conical beam with a half angle calculated to 

coincide with the width of the (120km) focused beam at GEO (~36,000km). An additional 30% was 

added to the width of the beam as a margin, to account for the shortened focus during the vertical ascent. 

The intensity was estimated by assuming the 30MW/m2

 average intensity at the focus to be spread 

normally across the diverging beam, with no atmospheric losses. 
NASA/TM—2012-217014 611 

F.7.1 CASE 1: Equatorial Launch 

Launch into 36° inclination from China Lake. Launches were simulated every 4 minutes for 8 days. 

Each intersection between the beam and background satellites is recorded. On average there were 5.25 

intersections per day, each for 0.2 seconds. 

Figure F.13.—Plot of launch time against range of intersecting satellite. Each color represents a single day of 

launches (i.e., the 8 days are overlaid). The NORAD sat ID is listed along with intersection duration and range. 

Figure F.14.—Plot of launch time against 1/R2 of intersecting satellite. Each color represents a single day of launches 

(i.e., the 8 days are overlaid). 
NASA/TM—2012-217014 612 

Figure F.15.—Probability of an intersection happening in each half hour period, over the 8 days. 

Figure F.16.—Histogram of clear launch windows—periods that have no intersections. Over the 8 days there are 

29 periods with more than 1 hour of clearance. 
NASA/TM—2012-217014 613 

Figure F.17.—Histogram showing the distribution of intersection maximum intensity during the 8 days—note that 

most have Imax less than 100 W/m2

 (i.e., 8% of solar const.). 

Figure F.18.—Histogram of the max energy incident on a satellite per square meter during an intersection—this is 

calculated using Imax x intersection duration.
NASA/TM—2012-217014 614 

F.7.2 CASE 2: Sun Synchronous Orbit 

Direct launch into 98° sun synch inclination from China Lake. Launches were simulated every 

4 minutes for 8 days. Each intersection between the microwave beam and background satellites is 

recorded. On average there were 12.75 intersections per day, each for 0.03 seconds. 

Figure F.19.—Plot of launch time against range of intersecting satellite. Each color represents a single day of 

launches (i.e., the 8 days are overlaid). The NORAD sat ID is listed along with intersection duration and range. 

Figure F.20.—Plot of launch time against range of intersecting satellite. Each color represents a single day of 

launches (i.e., the 8 days are overlaid). The NORAD sat ID is listed along with intersection duration and range. 
NASA/TM—2012-217014 615 

Figure F.21.—Probability of an intersection happening in each half hour period, over the 8 days. 

Figure F.22.—Histogram of clear launch windows—periods that have no intersections. Over the 8 days there are 

41 periods with more than 1 hour of clearance. 
NASA/TM—2012-217014 616 

Figure F.23.—Histogram showing the distribution of intersection maximum intensity during the 8 days—note that 

most have Imax less than 200 W/m2

 (i.e., 15% of solar const.). 

Figure F.24.—Histogram of the max energy incident on a satellite per square meter during an intersection—this is 

calculated using Imax x t, where t is the intersection duration.

F.8 Propellant Injection into the Atmosphere 

In the case of propellant ablation or heat exchangers for launch phase mission, there would exist byproducts that are released into the atmosphere. Currently discussed propellants have included LH2, LN2, 

H2O, H2, LOX/H2, peroxide/H2, H2O2 and He pressurant (for pressure-fed propellants). Water would be the 

most benign to the environment, and this should be considered when trading systems parameters to 

optimize performance. Upon selection of a propellant, the expected flow rates should be compared with 

the trajectory and velocity profile to determine the residual concentrations in the atmosphere, and any 

outstanding levels noted. 
NASA/TM—2012-217014 617 

F.9 Spectrum 

A primary challenge in the radio frequency (RF) based communications industry is spectrum 

allocation. Both RF and laser sources have been considered for BEP applications. In the RF case a 

proposed micro/mm-wave BEP system would have to work within the spectrum constructs and 

constraints of the NTIA/FCC system. This may cause a significant roadblock due to the level of radiated 

power and interference with currently allocated bands, along with adding a level of programmatic 

complexity in obtaining the proper allocation permits for operation. Likewise, although there are not 

current spectrum allocations for free-space optical communications, over time it may be required to 

obtain a similar wavelength dependant license. In a similar operational scenario to existing high power 

sources, the Laser Clearing House would also need to be involved with the scheduling of the BEP launch 

facility. 

F.10 Range Security 

Similar security measures would be needed for BEP as with conventional launch systems, including 

patrol of the KOZ when active and personnel security at the launch control complex. Furthermore, given 

the high energy capability of the BEP sources, proper homeland security measures would have to be 

established to safeguard the facilities against terrorist attack or overthrow. 

F.11 Operations 

The conduction of a BEP launch would include a set of weather criterion established from historical 

NOAA data and realtime local monitoring from ground stations and balloons. Local radar would be used 

for weather characterization, as well as monitoring air traffic in the area and tracking the launch vehicle’s 

trajectory. These measures could be integrated into the fire control system as a way to provide safety 

interlocking against airspace breach events or out-of-bound flightpath deviations. 

The impact on bird populations could be mitigated by understanding local migratory patterns when 

establishing the location of the complex (Ref. 5). In general these are well known and predictable, as 

shown in Figure F.25. In areas of known bird concentrations, low level visible light laser emitters could 

be temporarily activated to discourage activity in the vicinity of the main beam, in a similar manner to 

laser crop protection systems. 
NASA/TM—2012-217014 618 

Figure F.25.—Migratory bird patterns (reproduced from source 6). 

F.12 References for Appendix F 

1. “Laser Propulsion for ESA Missions: Ground to Orbit Launch,” Instituto Superior Técnico, 

Solscientia Lda., Contract No. 17048/03/NL/PA Final Report, 8 November 2004. 

2. National Oceanic and Atmospheric Administration’s (NOAA) National Weather Service (NWS) 

Climate Services, http://www.weather.gov/climate/

3. American National Standard for Safe Use of Lasers, ANSI Z136.1-2007 

4. Molecular Expressions, Interactive Java Tutorials, Specular and Diffuse Reflection 

http://micro.magnet.fsu.edu/primer/java/scienceopticsu/reflection/specular/index.html

5. Robert H. Diehl, Ronald P. Larkin, John E. Black, “Radar Observations of Bird Migration over the 

Great Lakes,” The Auk, 2003. 

6. North American Migration Flyways 

http://alaska.usgs.gov/science/biology/avian_influenza/images/NA_migration_flyways.gif
NASA/TM—2012-217014 619 

Appendix G.—Effects of Beaming Energy Through the Atmosphere 

G.1 Executive Summary 

Power beaming the large amounts of energy through the atmosphere that is required for propulsion 

either within the atmosphere or in transferring payloads from low Earth orbit to geosynchronous orbit 

(LEO to GEO) will be met with the extremely deleterious effects of the thermal nonlinearities induced by 

atmospheric heating from beam power absorption. The resulting phenomena, collectively called thermal 

blooming, will be the source of two major effects: (1) beam steering away from the intended target 

vehicle and (2) beam broadening, which will make the beam larger at the target than intended. Both of 

these effects have been modeled and are shown here to be remedied by the use of high-order (10th-order 

Hermite) phase compensation at the transmitter. Although it has been shown, in principle, that these 

perturbations can be mitigated, there may remain many engineering obstacles that must be overcome, 

such as the elimination of the mechanical jitter of the transmitter platform and other problems during 

beam propagation. However, there is nothing in the prevailing physics of the situation that would 

preclude power beaming through the atmosphere as discussed in this report. 

It is important to begin to capture the prevailing effects and overall system operation using a scaled 

atmospheric experiment that would simulate the realistic environment in which an adaptive optics system 

must operate, from the variable wind velocity up to the mechanical jitter of the transmitter platform. The 

authors recommend that, due to the immediate availability of high-power sources, scaled atmospheric 

experiments be implemented (tailored after the Scaled Atmospheric Blooming Experiments (SABLE) 

Project by Lincoln Laboratory in the early 1990s) in which the operation of a closed-loop millimeterwave adaptive-optics algorithm is assessed in the presence of a moving extended target. In addition, this 

experimental scenario could be used to address the possibility of air ionization, and subsequent 

breakdown, across the apertures of the combined millimeter-wave gyrotron sources as discussed in the 

DRM 1–C section. 

Finally, there is the need to look at various general beam wave profiles such as hypergaussian waves 

and fractional charge (in the topological sense) Laguerre-gaussian beam waves, which show great 

promise in their ability to be robust with respect to atmospheric nonlinearities. However, it is important to 

keep the modeling effort to the level of yielding analytical results, rather than requiring numerical 

evaluation, so that all the nuances of the physics involved are captured, and at the same time, to provide a 

tool for overall system evaluation as well as the design of adaptive optics algorithms. Finally, the model 

should be incorporated into a trajectory analysis program so that a cadre of launch geometries can be 

evaluated from the point of view of atmospheric thermal nonlinearities. 

G.2 Introduction 

When large amounts of power are being delivered through the Earth’s atmosphere via millimeter or 

infrared ‘beams’ (i.e., laser beams or beams formed at the output of a millimeter-wave antenna system), 

many propagation mechanisms must be addressed that may be potentially deleterious to such power 

transmission. The most obvious problematic mechanism is the ever-present random variation of the 

atmospheric refractive index due to local temperature variations known as turbulence. This naturally 

occurring phenomena is driven by thermal convection of heat from the Earth’s surface; once the resulting 

air motion exceeds a critical value of velocity, laminar flow essentially evolves into turbulent flow and 

fluctuations in the temperature distribution become statistically random (Ref. 1). These temperature 

fluctuations then act directly on the prevailing refractive index, thus rendering the refractive index a 

random quantity. These refractive index variations randomly focus and defocus the intervening 

electromagnetic wave field. Thus, the atmosphere can be considered to be composed of “lenses” of 

random focusing and defocusing characteristics that, because of the gross atmospheric motion due to 

wind, move across the beam. This gives rise to many beam quality variations: the major ones being beam 

broadening and beam steering. The statistical analysis and modeling of this type of atmospheric 
NASA/TM—2012-217014 620 

propagation has a long and rich history and has resulted in analytical descriptions for the effect of 

turbulence on the operation of systems relying on such beam propagation. Many models and descriptions 

exist for the engineering analysis of the operation of transmission systems that rely on the propagation of 

electromagnetic beams in the atmosphere. (For a good recent treatment of this topic, see Ref. 2 and the 

references therein.) 

This scenario may be considered as passive electromagnetic wave propagation; that is, the wave field 

moves through an atmosphere whose refractive index is determined by other sources, not by the field 

itself. However, as the energy density of the beam increases, absorption of the beam energy by 

atmospheric gas components results in local heating of the atmosphere, which does indeed act directly on 

the refractive index and cause it to decrease in value. The possibility of this situation was first advanced 

in 1966 (Ref. 3). This thermal change of the refractive index field then acts on the electromagnetic wave 

field causing it to also change, and so on. The propagation scenario now becomes an active one, whereby 

the propagating field modifies the very medium it which it exists. This heating process is called thermal 

blooming, and it substantially differs from that of the passive propagation discussed earlier (see Ref. 4 

and Ref. 5, which contains a very comprehensive review of the work and references that existed up to 

1990). Here, a “thermal lens” is created within the atmosphere by the heating due to the energy density of 

the beam. This self-action of the beam will not only bend the beam into regions of higher refractive index 

(beam steering), but convection within the atmospheric fluid will also arise, which is the source of the 

self-induced turbulent flow of the medium. The situation is further complicated when one includes the 

effects of atmospheric wind and aerosols and the abovementioned passive propagation effects. 

Defocusing and other associated nonlinear thermal-blooming distortions of the beam cross section 

will then result. In extreme cases of very large energy densities, the propagating beam will essentially 

break up into smaller beams, or filaments, which severely constrains the amount of energy density that the 

beam will be able to possess as it travels through the atmosphere. Unlike in passive propagation, the 

thermal-blooming mechanism introduces nonlinearities into the analysis of the phenomena that 

substantially complicate a complete mathematical description. Complete analyses of these types of 

propagation scenarios can only be done numerically, which was a major activity within the United States 

and Russia in the late 1980s. Other than the usual order-of-magnitude estimates using the equations of 

fluid mechanics and wave propagation, only numerical modeling of the effects of atmospheric thermal 

nonlinearities abound in the literature. Analytical treatments appropriate for an engineering analysis of 

atmospheric propagation systems encountering thermal blooming have been lacking, especially those that 

endeavor to describe the result of adaptive correction of such nonlinear effects. This situation makes a 

comparative assessment of the operation of through-the-atmosphere power transmission difficult. 

The propagation environment of high-energy electromagnetic wave transmission within the Earth’s 

atmosphere is presented to assess the atmospheric effects on the various beamed-energy propulsion 

scenarios considered in this report. Section G.3 presents a brief review of the major physical mechanisms 

that prevail in the atmosphere that will deleteriously affect energy transfer via electromagnetic beam 

waves. The critical power thresholds for the wavelengths of 2.0 m (infrared) and 2.0 mm (millimeter 

wave) are derived for the various transmitter output aperture sizes at which these propagation mechanisms 

arise and which must be addressed. Next, Section G.4 presents a propagation model that is later used to 

calculate, for a variety of aperture sizes and output powers, the thermally induced beam broadening and 

steering that will occur for a 2.0-m laser beam transmitted from the Earth’s surface to a target 800 to 

35 200 km distant. The adaptive correction to mitigate the thermal nonlinearity effects on the beam will 

then be considered. Of the various performance parameters that an adaptive optics system can be designed 

to optimize at the target, the minimization of the beam radius at target is used as the optimization 

parameter. Section G.5 dwells on the dynamics that an adaptive optics system must satisfy, particularly 

with regard to the control delays inherent with the propagation distances that are involved. Finally, 

Section G.6 highlights some scaled ground experiments that can be performed using available high-power 

millimeter-wave sources. In particular, assessing the effects of using an extended target with a closedloop adaptive optics approach and studying the possible electrical breakdown of moist air in the vicinity 

of the combined gyrotron outputs. To aid the reader, Section G.8 gives the definitions of the symbols and 
NASA/TM—2012-217014 621 

Section G.9 derives the nonlinear propagation equations for high-energy transmission through the 

atmosphere. 

G.3 Deleterious Atmospheric Propagation Mechanisms at High Energies 

At the power densities that are required for beamed-energy propulsion, several aspects of propagation 

within the atmosphere must be addressed that can potentially perturb the beam wave and render it 

unreliable for energy transfer to a small target. Once these aspects have been identified, physically 

understood, and mathematically modeled, appropriate mitigation procedures can be specified and applied 

to the problem to optimize energy transfer. These propagation features and the critical power levels at 

which they will be seen are given next. 

G.3.1 Ionization and Electrical Breakdown 

When the transmission of high-energy electromagnetic waves through the atmosphere is being 

evaluated, the first phenomenon that is considered is the ionization and subsequent electrical breakdown 

of air (Ref. 6, which contains a good introduction and the original work). Within the atmosphere, this 

occurs via the process of cascade ionization, whereby a free electron is created by multiquantum 

absorption by the atmospheric gas and through the inverse bremsstrahlung process, accelerates, and 

subsequently collides with an atom. The collision produces another electron, and both accelerate, collide 

with other atoms, and so on. This cascading process terminates in the release of light known as electrical 

breakdown. At a wavelength of  = 10.6 m (i.e., the wavelength of high-power CO2 lasers), electrical 

breakdown occurs at a power density of 109

 W/cm2 at atmospheric pressures and densities typical of 

those at sea level. This intensity is reduced by 2 orders of magnitude by the formation of the shock front 

due to the explosive detonation of atmospheric aerosols. Thus, for purposes of comparison, one can 

choose a power density of 107

 W/cm2 as the threshold for electrical breakdown within the atmosphere. 

At millimeter wavelengths in the 100- to 200-GHz frequency range, electrical breakdown occurs at a 

smaller power density of 1010 W/m2

. One can then assume the worst case of a twofold decrease in this 

value due to atmospheric aerosols, thus giving a threshold power density of 108

 W/m2. 

G.3.2 Induced Molecular Polarization—The Kerr Effect 

The next process that must be considered is the Kerr effect: that is, the process whereby the 

permittivity of the atmosphere, which is a function of the polarizability of the constituent gases, varies 

because of the intense electric field of the laser beam acting to induce molecular orientation (Refs. 7 

and 8). As discussed in Appendix G.9, the Kerr process can be described by a single differential equation, 

making its analysis straightforward in comparison to the thermally induced permittivity variations 

discussed in Section G.3.3. The associated relaxation times for molecular polarizability within the 

atmosphere are ~1011 s, so in the case of pulsed laser propagation, the medium can be considered to be in 

steady state for pulse lengths larger than this value. As the brief analysis in the Appendix G.9 shows, PcrK

is the threshold value of beam power at which the Kerr effect is only a function of the Kerr constant and 

the wavelength (see Eq. (G47)); it becomes an issue as a perturbing propagation mechanism at 

PcrK 1.5  109

 W for the wavelength  = 2.0 m and at PcrK 1.5  1016 W for  = 2.0 mm. The Kerr 

effect increases the permittivity in the location of the beam and thus results in what has become to be 

known as self-focusing; that is, the beam tends to move into regions of larger permittivity and thus acts as 

if it has encountered a lens. The lens is essentially induced by the beam through the Kerr effect. This selffocusing phenomena results in a spatial instability within the beam that causes the beam to break into 

individual filaments, where each filament takes on the size that corresponds to PcrK and the associated 

electric field strength. 
NASA/TM—2012-217014 622 

G.3.3 Induced Heating of the Atmosphere—Thermal Nonlinearities and Thermal Blooming 

The absorption of electromagnetic radiation by atmospheric gases becomes a source for the 

generation of heat. At large field intensities, the resulting temperature increase also changes the 

atmospheric permittivity. However, unlike the Kerr effect, the permittivity decreases and the beam tends 

to defocus (Refs. 4, 5, and 9 to 12). I n addition, one must also admit the description of the atmospheric 

fluid dynamic processes that are elicited by local heating and thus incorporate the Navier-Stokes and heattransfer equations for a moving medium; the medium is in motion not only because of the Archimedean 

forces that appear but also because of wind, which removes the heat that is generated. The entire problem 

is severely complicated and is usually relegated to numerical analysis applied to specific cases. However, 

within certain approximations, analytical results can be obtained. As shown in Appendix G.9, one can 

obtain an expression for the critical beam power PcrT above which thermally induced propagation issues 

can become prominent. Because of the interplay of the several physical mechanisms that prevail in this 

propagation process, one must specify the prevailing wind velocity V as well as the effective radius reff of 

the output aperture of the beam source and the wavelength-dependent absorption constant . 

2

p

crT

eff T

C V

P

k r

   

 (G1) 

As given by Equation (G1), one has for V = 4.4 m/s (10 mi/hr), reff = 10 m and  = 2.0 m, for 

which   410–6 m–1 and PcrT = 16.7 W, a surprisingly small power. This is due mainly to the very large 

radius subtended by the beam through which the field-induced temperature increase must dissipate 

because of wind. For a more typical radius, reff = 0.1 m, one has PcrT  1.7 kW. At the much larger 

wavelength  = 2.0 mm, one has   2.210–4 m–1, giving for reff = 10 m, PcrT = 3.5105

 W. 

When the radiation source is pulsed with a time duration of 1 s, the threshold power becomes much 

higher because the atmosphere does not have sufficient time to heat as the pulse moves through it. 

2

p

crP

T p

C

P

k t

  

 (G2) 

From Equation (G2), the critical power at  = 2.0 m is PcrP = 3.7107

 W and the critical power at 

 = 2.0 mm is PcrP = 7.71011 W. 

G.3.4 Atmospheric Aerosols 

Aerosols in the form of fog and clouds are very efficient absorbers of electromagnetic radiation 

(Refs.13 and 14). At the energy densities that prevail for beamed-energy propulsion, aerosols will 

essentially explosively detonate, thus diminishing their absorption and scattering abilities. The dynamics 

of this complex process has been well studied, and it has been established that, once a steady state has 

been established, propagation channels will appear in the medium through which the beam can propagate 

(Refs. 15 and 16). 

G.3.5 Atmospheric Turbulence 

The effect of random variations of the atmospheric permittivity field due to heating of the atmosphere 

(the sources of which are solar absorption in the atmosphere and on the Earth’s surface) is independent of 

the beam power level for the lowest critical power thresholds derived in Section G.3.2 for the thermalblooming case. This ever-present deleterious mechanism has the most heritage in terms of study and 

understanding. The effects of this phenomenon on the propagation of a beam wave show themselves in 

terms of loss of spatial coherence, beam steering, and beam broadening and are attendant with thermal 

blooming (Ref. 17). Of course, these characteristics are the same as those seen with thermal blooming, but 
NASA/TM—2012-217014 623 

the adaptive compensation in the case of turbulence is straightforward in that the effect being corrected is 

not the result of a nonlinear process. 

TABLE G.1.—THRESHOLD POWER LEVELS BEYOND WHICH PROPAGATION EFFECTS OCCUR AT 

INFRARED AND MILLIMETER WAVELENGTHS FOR VARIOUS OUTPUT APERTURE RADII, REFF

Propagation mechanism Threshold power level, W General relation for critical power Pcr and 

Wavelength,  power density pcr

2.0 m 2.0 mm

Electrical breakdown

 Effective radius of the beam, reff

 1 m 10 m 30 m 

3.11011

3.11013

2.81014

3.1108

3.11010

2.81011

No general relation 

Kerr effect 1.5109 1.51016 Critical power and power density for the 

Kerr effect 

2

8 crK

K

c P

k



 2 2 8 crK

eff K

c p

k r



 

Thermal blooming (continuous wave, CW)

 Effective radius of the beam, reff

 1 m 10 m 30 m

Thermal blooming (pulsed, 1 s)

166.7 

16.7 

5.6 

3.7107

3.5106 

3.5105

1.2105

7.71011

Critical power and power density for 

thermally induced perturbations

2

p

crT

eff T

C V

P

k r

 

  2 3

p

crT

eff T

C V

p

k r

 

 

Critical power and power density for the 

pulsed case 

2

p

crP

T p

C

P

k t

 

 2 2

p

crP

eff T p

C

p

kr t

 



Turbulence >0 >0

The results given in Sections G.3.1 to G.3.5 are condensed in Table G.47 and stated in terms of the 

critical power for the appropriate propagation mechanism and output aperture sizes (where they matter) of 

1.0, 10.0, and 30 m. Also shown, for reference, are the general equations for the critical power and 

associated critical power density. As can be seen, discussions of the use of beam powers up to 100 MW at 

infrared wavelengths easily exclude the effects of electrical breakdown and the atmospheric Kerr effect, 

leaving thermally induced nonlinearities and, of course, turbulence to be dealt with. For millimeter 

wavelengths, the thresholds for electrical breakdown are 3 orders of magnitude smaller than for the 

infrared region. Hence, aperture sizes larger than 10 m must be used for 100-MW power levels at these 

wavelengths using contiguous apertures. 

G.4 Modeling the Effects of Atmospheric Thermal Nonlinearities on Beamed Energy 

Propulsion and the Required Level of Their Mitigation 

Although it is not in the purview of this work to physically model the thermal nonlinearities attendant 

with the propagation of high-energy radiation through the Earth’s atmosphere, it has been found 

necessary to introduce a simple analysis employing the peculiarities of the propagation scenarios of 

beamed-energy propulsion in order to analytically assess and compare the overall affects of the 

atmosphere in these various scenarios as well as to evaluate the level of the corrective adaptive optics that 

will be required to make feasible the goals of beamed-energy propulsion. Two such power beaming cases 

are considered here: (1) orbit transfer from a low Earth orbit (LEO) to a geosynchronous orbit (GEO) 

from a ground-based laser transmitter operating at a wavelength of 2.0 m in the infrared spectrum and 

(2) beaming to a vehicle being launched within the Earth’s atmosphere using a ground-based array of 

gyrotrons operating at the wavelength of 2.0 mm within the far infrared spectrum. Because orbital transfer 
NASA/TM—2012-217014 624 

from LEO to GEO presents the worst-case situation in these two beaming cases because of the distances 

to be traversed, it is used as an example to form an atmospheric propagation model upon which the 

overall operation can be assessed. It is then shown how the model can be used to analyze the millimeter 

power-beaming case for which the results are also presented. 

The major property of the use of such propulsion to transfer payloads from LEO to GEO is, of course, 

the ratio of the diffraction length Ld associated with the beam wavelength and radius, to the focal length F

and the thickness H of the atmosphere. Here, the diffraction length 2 Ld eff  kr (where k is the wave number 

of the radiation and reff is the effective radius of the transmitting aperture; see Appendix G.9) and F is 

such that Ld >> H/Ld. 

This circumstance allows for the natural diffraction of the beam to be neglected within the region in 

which the beam is perturbed by induced nonlinear effects. This, along with the very large values of 

parameters describing the nonlinear interaction of the electromagnetic radiation with the fluid dynamics 

of the atmosphere (see Appendix G.9), makes for the creation of an easily implementable model to 

describe the major beam parameters that are of interest: the radius of the beam at the target and the 

displacement of the beam from the target when no adaptive correction is applied. Once these have been 

secured, the corrections can be applied to the various orders of aberrations of the initial phase of the beam 

and their effect at the target can be assessed. 

It is important to note here that only the effects of the thermal nonlinearities of the atmosphere are 

considered. The additional effects of the ever-present turbulent fluctuations of the atmosphere are not 

included here because this problem has much heritage, it is well understood, and techniques for its 

mitigation are established. However, because of the very large power that the beams will possess that are 

considered here, it is expected that the induced convective velocity within the atmosphere in the beam 

path will swamp any turbulent velocity fluctuations. Thus, the turbulent mechanism affecting beam wave 

propagation at these power levels will be rendered negligible. The goal of the present analysis is to isolate 

the effects of the much more deleterious thermal nonlinearities on very long-range power beaming and to 

estimate the required level of adaptive compensation to make the orbital transfer mechanisms discussed 

elsewhere in this report realistic and feasible. 

G.4.6 Quick Overview of the Model and a Description of Beam Behavior at Targets in Low Earth 

Orbit and Geosynchronous Orbit Due to Atmospheric Thermal Nonlinearities 

At the outset, the basis of the propagation model will be reviewed, leaving the details of the 

derivation to a forthcoming NASA TM. The initial equations are 

2

0

Cr V p eff T T E

P xx

                 (G3) 

  2 2 2 0 exp 2 V d

E T i E RTE E E L z E

z x 

   

                   (G4) 

(also in Appendix G.9), where the dimensionless nonlinearity parameter RV is given by 

2

eff T 0

V

p

kr P

R

C V

   

 (G5) 

connecting the total incident power P0 in the beam with the effective radius reff of the transmitted beam 

and the prevailing wave number k  2/ of the radiation of wavelength . The parameter E is the electric 

field of the beam at the target, z is a coordinate,  is the atmospheric absorption coefficient at the 
NASA/TM—2012-217014 625 

particular wavelength, T  |/T| is the variation of the permittivity with respect to the associated 

temperature variation T,  is the density of the atmosphere, Cp is the specific heat of the atmosphere at 

constant pressure, and V is the total relative velocity of the beam slewing across the atmosphere and that 

of atmospheric wind. (The normalization of the variables, indicated by the prime signs, is discussed in 

Section G.9.) Using the parameter values discussed in Section G.9, one has, for a 1-MW beam (i.e., P0 = 

1.0×106 W), RV = 6.0×104 for  = 2.0 m. In this instance, the second term in the first relation of 

Equation (G4) can be neglected relative to the third term: that is, the evolution of the beam within the 

atmosphere due to natural diffraction is negligible compared with the phase perturbation due to the 

thermal nonlinearity acting on the permittivity function. Hence, the model equations become 

  2

2 0 exp 2 V d

E T i RTE E E L z E

z x

   

               (G6) 

The solution of these relations must now be augmented with the prevailing boundary condition for the 

initial radiation profile of the beam at the output aperture of the transmitter at z  0. Using 

  2

0 ,0 exp 1 2

Ld EA i

F

     

           



 (G7)

which is derived in Section G.9 (where A0 is the initial amplitude of the field), and a gaussian beam 

profile, one has 

  2

222

0 ,0 exp 1 2

Ld E A ix y F

                            

 (G8) 

However, in most power-beaming applications, the focal length F of the beam will be set to the distance L

to the target: F = L. In addition, the full normalized representation for a gaussian beam wave must be 

employed in the last relation of Equation (G6): 

  0 1 2 , exp 1 21

A E z

iz iz

                          

 (G9) 

where  is the control, or update, coefficient. 

Finally, one must consider the effective thickness H of the atmosphere. Although there are several 

models that can be used to represent the absorption constant  with respect to the height within the 

atmosphere (the most popular is an exponential variation with height), it proves expedient here to use a 

single value for H and a corresponding effective value for . 

The solution of Equations (G6) and (G8) can now be considered for Ld/F >> H/L. To be sure, one has 

H  10 km, F < 35 200 km, and for reff  30 m, Ld  2.8×109

 m at infrared wavelengths. 

G.4.7 Application of the Foregoing to the Calculation of Beam Spread and Deflection Due to 

Atmospheric Thermal Nonlinearities for Power Beaming to Low Earth Orbit and 

Geosynchronous Orbit 

Again, leaving the details concerning the derivations to a forthcoming NASA TM, one can employ 

Equations (G6) to (G9) to obtain expressions for two very important performance parameters for power 

beaming to a target: that is, (1) the amount that the beam radius is widened because of the induced 

thermal defocusing (beam spread) and (2) the amount that the entire beam is steered away from the 
NASA/TM—2012-217014 626 

intended target because the propagation path of the beam tends to favor regions of larger permittivity 

(i.e., steers away from regions of higher atmospheric temperature caused by absorption (beam 

displacement)). It is important to note that, as discussed in Section G.9, the wind direction is taken to 

occur along the x-axis of a coordinate system whose origin is situated at the transmitter aperture with 

beam propagation occurring along the z-axis. Also, the radius of the transmitter aperture at z = 0 is given 

by an effective radius reff (0) that is related to the corresponding beam waist radius W0 by reff (0) = W0 /2. 

Because of the asymmetry introduced into the problem by the wind velocity V moving along the 

x-axis of the originally circular beam, there are two effective radii, reff,x (L) and reff,y (L), that characterize 

the beam in the x- and y-axis, respectively, at the distance L to the target vehicle. Again, this circumstance 

is brought about by the removal of heat from the beam channel along the x-axis. However, heat is not 

convectively removed from the beam along the y-axis. Instead, it diffuses in the y direction, which is a 

much slower process than convection. The result is that the defocusing in the y direction is much greater 

than in the x direction, and the beam becomes elliptical as it leaves the atmosphere and continues to 

enlarge until it interacts with the target in space. 

The results of the calculation for the effective radii and the target are given by 

    2

, 0

1 1

3 eff x eff NLV r LRL P   (G10) 

    2

, 00 1 0.866 1.612 d

eff y eff NLV NLV

L r LRL P P

L

    (G11) 

where 

   

0

0 4 0

T

NLV

p eff

k HP P

C Vr

   



 (G12) 

is yet another nonlinear parameter that enters the problem and characterizes, within the thin phase screen 

approximation, the level of the thermal nonlinearity effect on a beam wave of initial power P0 through an 

atmosphere of thickness H. The quantity Reff(L)  L/(kreff(0) is what the radius of a beam focused on the 

target at a distance L would be if the atmosphere were not present: that is, if transmission occurred 

entirely within a vacuum. (The reader is asked to excuse the rather excessive notations related to the 

various beam radii; the initial radius of the beam at the transmitter, denoted here as reff(0), is usually 

stated in terms of the waist radius W0, which is a factor of 2 smaller than the former. However, it is 

desired here to show how the effective radii “evolve” throughout the propagation process.) Thus, the 

quantities given by the radicals in Equations (G11) and (G12) are the factors by which the beam radii 

increase at a space-borne target because of the induced thermal nonlinearities within the atmosphere. 

In addition, as discussed earlier in this chapter, there is a deflection (L) of the beam axis into the 

direction opposite to that of the wind. This is given by 

  L 2RL P eff NLV    0  (G13) 

The implications of these results to the behavior of a beam wave to targets at LEO (L = 800 km) as 

well as GEO (L = 35 200 km) are given next. The operating wavelength of the continuous wave case 

considered here is taken to be  = 2.0 m, and the atmospheric wind velocity along the x-axis is taken to 

be Vx = 4.4 m/s  10 mi/hr. Three transmitter aperture radii—reff(0) = 5 m, 10 m, and 30 m—are 

considered. Table G.48 displays the beam radii that would occur at targets in LEO and GEO if the 

atmosphere were not present. 
NASA/TM—2012-217014 627 

TABLE G.2.—BEAM RADII AT TARGETS IN LOW EARTH 

ORBIT (LEO) AND GEOSYNCHRONOUS ORBIT (GEO) 

FOR VARIOUS OUTPUT APERTURE SIZES. 

[Wavelength  = 2.0 m.] 

Orbit of 

target 

Distance to 

target, 

L, 

km 

Effective beam output aperture radius, 

reff(0), m 

5 10 30 

Beam radius at target, m

LEO 800 0.1 0.051 0.017 

GEO 35 200 4.48 2.24 0.75 

Only a subset of these possibilities are considered in what is to follow. To keep things realistic, 

reff(0) = 5 m at GEO and reff(0) = 30 m at LEO are not subject to further analysis. 

Figures G.1 to G.4 display plots of reff,x(L), reff,y(L), and (L) versus P0 for transmission to LEO and 

GEO for the various initial beam radii at the output aperture using Equations (G10) to (G13). In addition 

to values quoted in Section G.9, the nominal value of H = 10 km was used in conjunction with the 

average absorption coefficient   4×106

 m1. As can easily be seen in all the examples, the thermal 

action of the atmosphere has a severe impact on the integrity of the beam wave at the target locations. 

Comparing the radii that result at the targets in LEO and GEO, because of a beam that has traversed 

thermal nonlinearities within the atmosphere, to those quoted in Table G.48 shows the rather 

unsatisfactory, but not unexpected, results that cause researchers to deem beaming power to such 

locations as unfeasible without the proper compensation to the beam applied at the transmitter. The 

induced deflection of the beam can easily be removed by a tracker, but the aberrations that remain, which 

result in the beam broadening, can only be removed by appropriate phase compensation applied at the 

transmitter. 

G.4.8 Phase Compensation of the Effects of Atmospheric Thermal Nonlinearities 

The first research into the compensation of thermal-blooming effects considered the application of a 

Zernike polynomial expansion of the transmitted phase front (Refs. 18 and 19). This technique provides 

for an optimal representation of the phase front of the beam wave so long as the cross section of the beam 

remains circular; after all, Zernike polynomials are, by design, orthogonal on a unit circle. The problem 

encountered in the Section G.4.7 development is different: a beam asymmetry arises because of the 

differences in convective versus diffusion cooling of the atmospheric channel in which the beam 

propagates. The model briefly reviewed in Section G.4.6 allows for the analytical assessment of phase 

compensation of the deleterious effects shown in Section G.4.7. 

The different behavior of the beam along the transverse axes suggests that a more general set of 

orthogonal polynomials should be used than those of Zernike. Here, since the initial form of the beam 

wave is given by a gaussian function, the transmitted phase front S(x,y) at the output aperture of the 

transmitter can be expanded in a set of orthogonal polynomials associated with a gaussian weight 

function: that is, the Hermite polynomials Hn(x) and Hm(y) (Ref. 20). Thus, 

    ,

M N

nm n m

n m

S xy a H xH y  (G14) 

The expansion coefficients anm are determined by applying one of several performance metrics. For 

example, for power-beaming applications considered in this study, it is desired to shape the phase front so 

as to minimize the radius of the beam at the target so the beam does not interfere with the adjacent 

structures of the vehicle. Hence, the performance metric to be minimized through the selection of anm is 

related to the electric field E(,L) of the beam at the target vehicle by 
NASA/TM—2012-217014 628 

 

 

 

 

2 3

0 2 2

2

0

, d

, d

E L

L xy

E L





     

      

     





 (G15) 

in which E(,L) results from the application of the boundary condition 

    2

0 ,0 exp , 1 2

Ld E A iS x y i

L

                         

 (G16) 

to the solution of Equation (G6). Ideally, the number of terms of the expansion given in Equation (G14) is 

infinite: N, M . The linear terms for which n + m = 1 determine the inclination correction of the beam 

wave. The quadratic terms for which n + m = 2 determine the focusing correction, and the higher order 

terms for which n + m  3 give the higher order aberration corrections. However, in realistic applications, 

the number of aberrations N and M is finite, and in fact, it is desired to find the smallest number of 

aberration corrections that can be used to represent the compensated phase. The specific expressions for 

the coefficients anm are found by equating to zero the derivative of (L) with respect to these coefficients. 

The propagation problem defined here, as well as the associated minimization problem, can be solved 

analytically and will also appear in a forthcoming NASA TM. The results applied to the calculation of the 

phase-corrected beam radii are as follows. The elliptical shape of the beam, which is characterized by two 

radii at the target vehicle, facilitates the calculations if a single effective radius reff(L) defined by 

    2 2

, ,

2

eff x eff y

eff

r Lr L

r L

   (G17) 

is considered. Applying the modeling procedure outlined in Equations (G14) to (G16), one finds that, for 

the radius reff,corr(L) of a focused beam wave corrected for the first M phase aberrations represented by the 

model of Equation (G14), 

r L RL eff,corr      eff  

        2 2

2 2 1

0 0 3( )/2 3 /2

10 0

00 0

1 2.95 4 2 ! ! 2 !

2 !! 2 ! 4

MM M nm n nm n

NLV NLV nm n

nm n

HH H

P P n mnm nn

nm n

 



 

                               

 

 (G18) 

Figures G.5 to G.8 show the results of this relation applied to the various LEO and GEO beaming cases 

considered in Section G.4.7. 

These plots display the fact that, for initial beam powers of 10 MW or less, phase correction using 

M-order aberrations returns the radius of the beam to its desired value at the target vehicle. Beyond this 

power level, only for the case shown in Figure G.8, which uses a beam radius of 30 m at the transmitter, 

can a 10th-order phase correction be effective up to about 70 MW. 

One assumption that has been prevailing in all these calculations is that the wind velocity V is a 

constant along the entire propagation path. In reality, this is certainly not the case. In fact, there may be 

regions in which the V  0 (i.e., a stagnation zone), and the convective heating scenario assumed here will 

transform into a diffusive heating scenario characterized by much lower critical power thresholds. This 

case will not be studied here since it is the subject of a recent study (Ref. 21), but it must be kept in mind 
NASA/TM—2012-217014 629 

that a statistical description of the wind profile along the propagation path will be required in a more 

careful examination. 

Finally, there is the whole issue of the effect that the beam profile will have on the level of thermal 

nonlinearity perturbations. In fact, very high energy laser beams are characterized by initial beam profiles 

that are tubular: that is, have a local minimum of intensity in the center that increases toward the outer 

portions of the beam. Such beams, which can be modeled by hypergaussian profiles, tend to be 

characterized by higher critical power thresholds. This is due to the fact that the center of the beam, with a 

lower intensity, is thermally cooler than the periphery and thus tends to bend toward the center and 

remain stable. This effect should be studied for the particular high-energy sources that will be used as the 

design of the power-beaming scenarios evolve. 

G.4.9 Using the Propagation Model for the Analysis of Millimeter-Wave Beaming Within the 

Earth’s Atmosphere 

The methodology developed in this chapter for the infrared power-beaming case for LEO-to-GEO 

orbital transfer can be applied to the millimeter-wave case for power beaming to vehicles launched from 

within the Earth’s atmosphere. The wavelength of concern here is  = 2.0×103

 m = 2.0 mm with an 

output aperture radius reff = 50 m. This gives Ld = 7.9×106

. The propagation distances involved range 

from L = 20 to 120 km. Although the beam will not be directly focused on the target vehicle (the 

dimensions being envisioned will require the beam to have a focal length just beyond the vehicle), for the 

purposes of this discussion, F  L. The effective distance within the atmosphere responsible for most of 

the absorption at these wavelengths is H  1.8 km. Hence, the condition Ld/F >> H/Ld is easily met. In 

addition, there is a characteristic length LT that is a measure of the distance at which thermally induced 

diffraction effects will occur. This characteristic length can be defined using the length Ld as well as the 

nonlinear distortion parameter RV given by Equation (G19). 

2

2 2 eff T 0 0

V eff T V

p crT

kr P P R kr T

CV P

 

  



 (G19) 

A detailed analysis similar to that given in Section G.9 gives for the thermal diffraction distance 

3

0

eff p

T

T

r CV

L

P

    

 (G20) 

At the wavelength considered in this section, the associated absorption coefficient is   2.210–4 m–1, 

which gives LT  73 km for the total transmitter power of P0 = 800 MW (which is considered later in this 

section). Hence, thermal diffraction effects will certainly not prevail within the region H = 1.8 km or, for 

that matter, within the entire troposphere, and the model constructed earlier in this chapter can be adopted 

here. (This circumstance is solely due to the very large output aperture that is being used.) 

 
NASA/TM—2012-217014 630 

Equations (G10) to (G13) can now be used to consider four scenarios for a typical power-beaming 

launch: 

Horizontal beaming ( = 0) at L = 20 km 

Beaming at  = 30 and L = 50 km 

Beaming at  = 50 and L = 100 km 

Beaming at  = 90 and L = 120 km 

In each case, the target is to be illuminated in an area with a radius of Reff(L)  1.8 m. Thus, a 

dynamic focus is to be employed; the focal length is determined by 

  2

where 1 1 2 2

d eff

d

L L kL R L

F

L L

   

 (G21) 

and where the “-” sign is used in front of the radical for a beam focused behind the target. Figures G.9 to 

G.12 display the results of applying the propagation model to these cases. 

As can be seen from these results, the worst case is, of course, the horizontal beaming at an elevation 

angle  = 0, as in Figure G.9. Here, the beam traverses 20 km of atmosphere. The beam severely 

broadens along the y-axis, perpendicular to the wind velocity, since the heat that is generated is not 

convectively removed in this direction as it is along the x-axis. Related to this phenomenon is the fact that 

the beam is deflected in the x-direction up to 12 m from the target at a power level of 800 MW. Of course, 

as the elevation angle of the beam path increases, the distance within the atmosphere decreases up to the 

best case, in which  = 90, as depicted in Figure G.12. It must be cautioned, however, that the model 

used here, as mentioned earlier, does not include viscous and Archimedean forces that occur during the 

thermally induced motion of the atmosphere within the vertical column in which the beam propagates. 

For situations in which large elevation angles are realized, these effects must be included in a future 

extension of the analysis. Finally, it must be noted that in order to follow the relatively close target 

vehicle, the beam will have a discernable slewing velocity in comparison to that of the LEO-to-GEO 

transfer case. This slewing velocity may (depending on the relative wind velocity) tend to lessen the 

thermal effects. This will certainly be the case for the upward movement of the beam. As for the 

horizontal component of the slewing velocity, it can either detract or worsen the thermal effects. Such 

aspects of the problem can only be assessed via a simulation of the particular launch environment. 

Just as with the infrared beaming case, the adaptive correction of these deleterious effects is 

accomplished through the use of phase compensation at the transmitter output aperture. Again, because of 

the asymmetry induced within the structure of the beam, it is advantageous to represent the corrected 

phase front as an expansion of Hermite polynomials. Figures G.13 to G.16 show the results of applying 

various orders of Hermite correction to the millimeter-wave beam. The goal is, of course, to return the 

beam radius to the desired value of 1.8 m. As expected, the  = 0 case requires the most compensation; 

correction up to 10th-order aberrations is needed after the removal of the beam deflection via a tracker. 

The other cases only need correction up to the 5th order. 

The additional advantage that this millimeter-wave power-beaming scenario has over the LEO/GEO 

orbital transfer case is that the target vehicle will have a much larger apparent velocity in both the traverse 

and longitudinal directions than an object executing a LEO-to-GEO transfer. This circumstance brings to 

the fore the conditions of stability of the control algorithms, which compose a closed-loop adaptive optics 

system. The beam also will have a significant slewing velocity that, in some respects, eases the effects of 

thermal nonlinearities, but in others, places a burden on the operation of the adaptive optics control 

system. This is discussed in Section G.5. 

Finally, it should be mentioned that for the power levels considered in this millimeter-wave case, one 

should consider the mechanism of explosive detonation (Ref. 16) of small atmospheric aerosols that may 
NASA/TM—2012-217014 631 

traverse the beam, especially in the horizontal beaming case. This will give rise to a time-dependent 

absorption coefficient that may indeed lessen the long-term heating effects within the atmosphere. 

G.4.10 Tool for Quickly Assessing Nonlinear Thermal Effects on Power Beaming 

Section G.9 introduces several expressions for threshold power levels at which thermal effects will 

become prevalent. Although these derivations are based essentially on normalization and dimensional 

analysis, they can be given credence by using the results obtained in Section G.4.8. However, these 

critical power levels need to be related to a specific property or characteristic of a beam wave that will 

impact the performance of the power-beaming system. Because there are many such characteristics and 

because there will be as many corresponding critical power levels, there is no single universal quantity 

that can be stated to convey the specific impact that thermal blooming will have on the entire performance 

of a beamed-energy system. However, the most important characteristic in power-beaming situations is 

the variation of the beam radius due to the defocusing that occurs when the atmosphere is heated. As 

shown by Equation (G11) and, for example, in Figures G.1 to G.4, the radius of a beam in an atmosphere 

with convection is severely perturbed in the direction perpendicular to the wind direction. To attempt to 

secure an analytical estimate at which the radius of such a beam will begin to increase because of thermal 

blooming, one can expand Equation (G11) to obtain 

, 0     0.866 1

2

d

eff y eff NLV

L r LRL P

L

          (G22) 

and require the value for P0 at which 

    2

0

0

0.866 0

0.1 1

2

d eff T

NLV

p

L k r PH

P

L C VL

 

   

 (G23) 

Setting L = H, one finally obtains for the critical power at which the beam radius begins to expand 

  2 10

0

p

crR

eff T

C V

P

k r

   

 (G24) 

which is just 10PcrT. One can now use this relation for millimeter wavelengths and compare it to the 

infrared case. Figures G.17 and G.18 show plots of Equation (G24), as well as the associated power 

density   2 0 , P r crR eff  versus aperture radius for both cases. Here, the nominal wind velocity is taken to be 

V = 0.5 m/s. Of course, Equation (G24) is simple enough to apply a statistical model describing the value 

for V. 

Evaluating Equation (G24) using the numerical values of the parameters that do not depend on the 

wavelength, one has for this critical power 

2

crR 200 MW

eff

P

r

  

 (G25) 

where PcrR is in megawatts and  and reff are in meters. (The units of  are inverse meters.) These results 

clearly show that propagation within the atmosphere can sustain much larger power levels at millimeter 

wavelengths than at infrared wavelengths. 
NASA/TM—2012-217014 632 

G.5 Dynamics of Adaptive Phase Compensation for Power-Beaming Applications 

There are two major approaches that can be used to implement phase compensation to mitigate 

atmospheric effects: (1) wave front “reversal,” or conjugation, and (2) wave front control employing the 

optimization of a specific system metric for performance. The first presents a fundamental problem for 

the mitigation of thermal nonlinearities in that it relies on a principle of reciprocity, a basic tenant of 

which is violated in thermal-blooming applications. The use of such a principle essentially requires, 

through a wave front conjugation condition, that the amplitude and phase be corrected. However, it is 

much easier to correct the phase perturbations and not the associated amplitude perturbations. This is fine 

if amplitude variations are small (e.g., astronomical imaging). In thermal-blooming scenarios, amplitude 

variations cannot be neglected. Using this approach is not warranted in power beaming. What is 

recommended here is the use of a method using the actual radiant intensity at the target. What used to be 

called aperture optimization, or tagging (Ref. 22), has come to be known as target-in-the-loop, Its 

implementation is known as gradient descent optimization (GDO) wave front control (Refs. 23 to 27) and 

is suggested for the power-beaming applications addressed in this report. Here, the control rule is based 

on the direct optimization of an easily measured system performance metric, such as the radiant intensity 

at the target. 

The model governing the operation of such an adaptive optics system is given by first identifying a 

performance metric   S a     , as defined by Equation (G14), which depends on the phase given by 

Equation (G14) and the array a a  nm   of the associated expansion coefficients (Ref. 24). The 

coefficients in Equation (G14) were selected to minimize the beam radius on the target vehicle. Here, the 

optimal selection of the values of anm during the actual operation of the system is given by the control rule 

    nm

nm

a

nm

at t

t a

 

     (G26) 

in which nm a  are the time constants and  is the control (or update) coefficient. Here, J is taken to be the 

size of the focal spot at the target vehicle, and t is the time. The intensity can be inferred by recording the 

reflected radiation from the target at a point separated from the transmitter so that a slightly different and 

unperturbed propagation path is used. It must now be established that such a control structure will 

dynamically operate in the two extreme cases of power beaming considered here: (1) infrared beaming to 

a vehicle in GEO and (2) millimeter-wave beaming to a near Earth vehicle. 

Consider first power-beaming to GEO. For a brief analysis in which the required temporal 

characteristics of the adaptive optics system (and thus the temporal stability) are to be derived, one can 

consider the isolated case for which n + m = 2: that is, according to the discussion in Section G.4.8, the 

beam wave is focused dynamically. Thus, one has from Equation (G14) for the associated (now timedependent) aberration coefficient 

20   ( )

Ld a t

F t  (G27) 

where, for example, H2(x) ~ x

2

. For the case in which the focus is to be placed at the target at a distance L

(i.e., F(t) = L), one has from Equation (G26) 

    20 20 0 exp d d

a

L L t

at a

L L t

   

          

 (G28) 
NASA/TM—2012-217014 633 

where   2 2

a d 20 t LL   is the adaptation time of the adaptive optics system. This very important time 

constant is, although proportional to 20, inversely proportional to the distance to the target. This wellknown property is that the adaptive system will possess a faster response the farther the target is from the 

transmitter. If this were all that the description of the adaptation system required, then the state of 

focusing on the target would monotonically approach the desired result Ld/L. However, there is a limiting 

factor placed on this circumstance by the delay inherent in overall system response due to the propagation 

time to and from GEO. When the propagation delay is allowed to enter into the control rule, 

Equation (G26) becomes 

    nm

nm d

a

nm

a t tt

t a

  

     (G29) 

where td is the delay time (in this case, light time to and from the target). Expanding the right side of this 

equation in a Taylor series in td and keeping the first few terms yields 

        2 2 2

2 2 nm

nm d

a d

nm nm d nm d

at t t t t

t

t a at a t

       

             

 (G30) 

The corresponding characteristic equation  for the resulting differential relationship for a20(t) = exp(–t)

in the example considered in Equation (G30) is given by 

2 2

2

2 a d a ad d

d

t t t tt t

t 

  

  (G31) 

which shows that the system converges to a stable state so long as ta > td; that is, the adaptation period of 

the system is longer than the delay time. For a target in LEO, td  0.005 s, and for GEO, td  0.24 s (for 

the round trip time). If on the other hand, td > ta, the system does not converge to the stable state and 

  20 limt

a t . Thus, the adaptive system must be continuously “tuned” (by adjusting  in response to 

changes in td as well as the evolving values for L in the case of a moving target) so that the condition 

ta > td is satisfied. 

To these considerations must also be added the temporal delay in the overall system response. This 

will not be discussed here because these considerations can be found in the literature. Suffice it to say that 

the delay that is met with in beaming to GEO can be easily factored into the operation of the adaptive 

system. It must be remembered that the distance to the target is always increasing in time for a beamedpower scenario and that this, above and beyond that of the propagation delay, must be properly treated in 

the design and specification of the adaptive optics control system. 

Work on the actual implementation of the closed-loop control of the gradient descent algorithm of 

Equation (G26) has rapidly progressed. The technology is now mature and forms the basis for reliable 

adaptive control. The only possible exception to its application is for power beaming to locations within 

the Earth’s atmosphere for surface or near-surface launching. Here, the target vehicle, unlike in the LEOto-GEO transfer case, is rapidly moving and is considered to be an extended target; that is, its spatial 

extent cannot be relegated to a point object, and its possible rotation and orientation become an issue for 

use as a reflector of radiation back to the adaptive optics receiver. Here, the randomly rough surface of the 

target reflector creates a speckle field characterized by bright and dark regions of intensity that have been 

found to complicate the implementation of a closed-loop adaptive optics scenario. This can lead to the 

slowing of the convergence speed of the system and, ultimately, may compromise the efficiency of its 
NASA/TM—2012-217014 634 

operation in terms of the optimization of the focal spot at the target. Methods using the precompensation 

of the beam show promise (Ref. 28). 

Although it was shown that the basic physics of the situation does not preclude beaming energy to a 

prescribed area on a target vehicle in GEO, the engineering implementation will offer some major 

obstacles to be negotiated. In particular, beaming into a GEO location will require 30-nanoradian pointing 

accuracy on the part of the transmitter optics. The presence of mechanical jitter of the transmitter platform 

will have to be continuously mitigated. In addition, on the other side of the spectrum, launching from 

within the atmosphere will challenge the application of the adaptive optics at millimeter wavelengths. The 

fact that the beam is rapidly slewing because of the relatively rapidly moving object makes it difficult for 

the closed-loop adaptive optics system to converge to a stationary value since the medium within the 

column of atmosphere in which the millimeter-wave beam exists is constantly being exchanged. Here, 

instead of employing the target as the beacon source, the use of an artificial beacon placed ahead of the 

moving vehicle may help in characterizing the atmosphere ahead of the beam. This will assess the phase 

perturbations of the nonheated atmosphere, but the problem still remains concerning just how the 

atmosphere will respond to the heating from the beam as it arrives at that particular column. This suggests 

that, because of the availability of high-power sources at millimeter wavelengths, one could configure an 

in situ experiment to assess the operation of both closed-loop and artificial beacon adaptive optics 

approaches. This will be discussed further Section G.6. 

G.6 Experiments Assessing the Compensation of Thermal Nonlinearities of High-Power 

Beams in the Atmosphere 

Experiments dealing with induced effects due to thermal nonlinearities elicited by high-power 

propagation are usually performed in a laboratory using liquids placed in cells in which the thermal 

nonlinearity thresholds are much smaller than in air. In these scaled laboratory experiments, where 

atmospheric turbulence effects are simulated by transmission phase screens, various adaptive optics 

algorithms have been tested and evaluated (Refs. 29 and 30). Some experiments have been performed in 

the open atmosphere along horizontal paths in a program called the Scaled Atmospheric Blooming 

Experiments (SABLE) directed by Lincoln Laboratory (Ref. 31). Work also is continuing along these 

lines in other countries (Ref. 32). All these experiments endeavor to evaluate the in situ operation of 

adaptive optics systems on a scaled basis. The same must be recommended for the operation of the 

beamed power scenarios discussed in this chapter. Because of the maturity and availability of high-power 

sources in the millimeter range, the authors recommend that a scaled atmospheric experiment be 

performed on a moving target to assess the operation of various adaptive optics algorithms. Such an 

experiment should be modeled after the SABLE project. It is important to note that currently no database 

exists that addresses the power-beaming cases considered here. It is important to begin to capture the 

prevailing effects and system operation using a scaled atmospheric experiment that would simulate the 

realistic environment in which an adaptive optics system must operate—from the variable wind velocity 

to the mechanical jitter of the transmitter platform. In addition, this experimental scenario could be used 

to address the possibility of the ionization and subsequent breakdown of air across the apertures of the 

combined millimeter-wave gyrotron sources as discussed in the DRM 1–C section. 

G.7 Conclusions and Recommendations 

The various deleterious propagation mechanisms associated with high-power electromagnetic wave 

propagation through the atmosphere have been discussed. In addition to turbulence, thermal nonlinearities 

associated with the absorption of radiation by atmospheric gases will contribute to the major effects of 

beam wave propagation for beamed-energy propulsion at the power levels considered here. The 

simplified propagation model advanced here showed that the beam radius and its deflection are severely 

affected by the phenomena of thermal blooming. However, the model also showed that appropriate phase 

compensation at the transmitter output aperture can mitigate these effects and return the propagation 
NASA/TM—2012-217014 635 

situation to one that is acceptable for the power transfer requirements that must be satisfied for beamed 

propulsion. For infrared transmission out of the atmosphere for low Earth orbit to geosynchronous orbit 

(LEO-to-GEO) beaming, up to 10th-order (once the tilt has been removed) aberration correction will be 

needed to maintain a minimal focal spot at a GEO location using a 30-m-diameter transmitter aperture. 

For millimeter-wave beaming within the atmosphere, 5th-order aberration correction will suffice, with the 

exception of the horizontal case, in which, once again, a 10th-order correction will be required. In 

principle, these corrections will allow the beaming system to operate within the prevailing specifications. 

In practice, however, some challenges remain in the implementation. 

The adaptive optics approach recommended here for the LEO-to-GEO launch case is a closed-loop 

system that uses the specular reflection from the target vehicle as a beacon source (i.e., a target-in-theloop system). The target vehicle is seen by the adaptive optics system as a simple point reflector. The 

round-trip propagation time delay inherent in this scenario can be tolerated as long as the adaptation time 

of the adaptive optics system is properly set to a slightly larger time than that of the delay to ensure proper 

convergence. 

In the case of millimeter-wave beaming to a moving vehicle within the Earth’s atmosphere, the object 

is considered as an extended target with a rotation and orientation that can complicate the application of a 

closed-loop system. The reflection from the vehicle will have a speckle structure that could severely 

impact the wave front sensor used by the adaptive optics system. Here, it may be that the simpler artificial 

beacon method could be used, but there is an additional complication: namely, the rapid movement of the 

beam across the atmosphere. The changes induced in the beam column through the atmosphere will not 

all be those due to the adjustment by the adaptive optics, and the system will not be able to properly 

adapt. This situation presents itself for a scaled atmospheric high-power millimeter-wave transmission 

experiment in which both closed-loop and artificial beacon-based adaptive optics systems are tested and 

evaluated. 

The propagation modeling for the various beamed propulsion scenarios presented in this report should 

be extended beyond that employed here in Section G.4 by incorporating additional fluid mechanical 

descriptions of the atmosphere as well as more general beam wave profiles such as hypergaussian as well 

as fractional charge (in the topological sense) Laguerre-gaussian beam waves which show great promise 

in their ability to be robust with respect to atmospheric nonlinearities [Ref. 33]. However, it is important 

that the modeling effort be kept to the level of yielding analytical results, rather than requiring numerical 

evaluation, so as to capture all the nuances of the physics involved and, at the same time, provide a tool 

for overall system evaluation as well as the design of adaptive optics algorithms. Finally, the model 

should be incorporated into a trajectory analysis program so that a cadre of launch geometries can be 

evaluated from the point of view of atmospheric thermal nonlinearities. 
NASA/TM—2012-217014 636 

G.8 Symbols 

A0 initial amplitude of the field (Eq. (G7)) 

anm expansion coefficients 

a20(t) time-dependent aberration coefficient 

a



 array of expansion coefficients

Cp specific heat of the atmosphere at constant pressure 

c velocity of light 

E electric field of the beam at the target 

E0 defined by equation after Equation (G50) 

F focal length

H thickness of the atmosphere 

Hm(x) Hermite polynomial 

Hn(x) Hermite polynomial 

I intensity of radiation 

I∥ longitudinal size of the spatial variation of the variable permittivity 

J size of the focal spot at the target vehicle 

k wave number of the radiation; 

L distance to target 

(L) deflection of the beam source 

Ld diffraction length 

LT characteristic distance at which thermally induced diffraction effects occur 

M number of expansion terms or phase aberrations (Eq. (G14)) 

N number of expansion terms or phase aberrations (Eq. (G14)) 

m, n subscripts in Eq. (G14)

Pcr critical power 

PcrK threshold value of beam power at which the Kerr effect is only a function of the Kerr constant 

and the wavelength 

PcrP critical power 

PcrR critical power at which the beam radius is affected because of thermal focusing (Figs. G.17 

and G.18) 

PcrT critical beam power above which thermally induced propagation issues can become 

prominent 

P0 total incident power in the beam 
NASA/TM—2012-217014 637 

pcr critical power density 

pcrR critical power density at which the beam radius begins to expand (Figs. G.17 and G.18) 

Reff(L) radius of a beam focused on a target at a distance L if there were no atmosphere 

RP prevailing nonlinear parameter (defined by Eq. (G57)) 

RV dimensionless nonlinearity parameter, or distortion parameter, for the thermal nonlinearity 

R distortion parameter for Kerr nonlinearity 

reff effective radius of the output aperture of the beam source 

reff,x(L) effective radius of the beam in the x-axis 

reff,y(L) effective radius of the beam in the y-axis 

reff(0) transmitter aperture radius 

r'eff(L) particular effective radius 

r'eff,corr radius of a focused beam curve corrected for the first M phase aberration 

S(x,y) transmitted phase front at the output aperture of the transmitter 

T temperature variation 

t time 

ta adaptation time of the adaptive optics system 

td delay time 

tp duration of radiation pulse; also Table G.1 

TV characteristic temperature 

V prevailing wind velocity; total relative velocity of the beam slewing across the atmosphere 

(after Eq. (G5) 

Vx atmospheric wind velocity along the x-axis 

V



 total relative velocity of the motion (slewing) of the beam across the atmosphere and that of 

the atmospheric wind 

W0 waist radius of the beam at the exit of the output aperture 

x' coordinate (Eqs. (G4) and (G51)) 

xˆ unit vector (after Eq. (G49)) 

y' coordinate (Eq. (G8)) 

z' initial radiation profile of beam at output aperture of transmitter (Eq. (G4)) 

 wavelength-dependent absorption coefficient 

 control, or update, coefficient (Eq. (G9)) 

 variable permittivity 

 variation of permittivity with respect to temperature variation 
NASA/TM—2012-217014 638 

 Kerr effect perturbation to the permittivity 

0 nominal permittivity 

 Eq. (G31) 

 Peclet number 

NLV(P0) parameter that characterizes level of thermal nonlinearity effect on a beam wave of initial 

power through an atmosphere of thickness H 

 elevation angle

 thermal conductivity of the atmosphere 

 wavelength 

 velocity of the wave field in the medium 

(L) performance parameter defined by Eq. (G15) 

 density of the atmosphere 

' dimensionless transverse coordinate; also Eqs. (G4) and (G5) 





 position vector in the plane transverse to the direction of propagation along the z-axis 

2 

 tranverse Laplacian 

 temporal variation of variable permittivity

anm  time constants 

 phase front 

 thermal diffusivity 

2T diffusivity 

 
NASA/TM—2012-217014 639 

G.9 Derivation of the Fundamental Nonlinear Propagation Equations for High-Energy 

Transmission Through the Atmosphere and for the Associated Critical Powers 

For the propagation of scalar electromagnetic waves through a medium characterized by a variable 

permittivity , one has for the equation describing the evolution of the electric field strength 

2 2

0

1 2 0 E E ik E k E

z t 

                     

 (G32) 

where 



 is a position vector in the plane transverse to the direction of propagation along the z-axis, 

k  2/ is the wave number of the radiation field of wavelength ,  is the velocity of the wave field in 

the medium ( 0   c, where c is the velocity of light), and 0 is the nominal permittivity (for the 

atmosphere, 0  1). As will be discussed later in this appendix, the function describing  is, in general, 

dependent on the value of E: that is, , thus making the propagation problem a nonlinear one. 

Only in the case of weak fields (the typical situation for atmospheric optics) is  independent of E. 

To simplify the analysis, one can, to a very good approximation, assume that ∥/, where l∥ is the 

longitudinal size of the spatial variation of and is its temporal variation. This will preclude the 

description of the propagation pulses of very short temporal duration, where dispersion will become 

important, but will suffice for the purposes of this presentation. Thus, Equation (G32) reduces to 

2 2

0

2 0 E ik E k E

z 

   

        

 (G33) 

which is known in the literature as the quasi-optical approximation. The first two terms of Equation (G33) 

describe the propagation (within the quasi-optical approximation) of a beam wave (see Eq. (G35)) in a 

vacuum; the last term is the perturbation added to the evolution of the beam wave through a medium 

described by the factor  It is through this factor that the various effects of the atmosphere on 

propagation come into play. 

For the power ranges of interest in high-energy laser propagation, the total permittivity function can 

be written as 

    2 2

0 , E Ti E T K T k

        (G34) 

in which  is the wavelength-dependent absorption coefficient. Absorption within the propagation 

medium gives rise to the increase in the associated temperature T from the nominal value, which, in turn, 

modifies the value of the permittivity by the amount TT, where T  T is the variation of the 

permittivity with respect to T. (Strictly speaking, the temperature increase T is also a function of |E|

2

, it is 

not a direct function as is the Kerr effect.) This gives rise to thermal nonlinearities and the associated 

thermal blooming. Also, because of the large value of E in cases of very high energy propagation, induced 

orientation of anisotropically polarized molecules of the atmosphere occurs due to the interaction of the 

induced dipole moment of the molecule due to the intense electric field. This contributes another 

perturbation  to the permittivity through the quantity |E|

2

. This phenomenon is known as the Kerr 

effect. The functional relations of both |E|

2

 and T to the quantity |E|

2

 also are given through the 

prescription of a differential equation. In this appendix, each of these contributions is discussed separately 

and the critical field powers at which they will occur within the atmosphere are derived. Before this is 
NASA/TM—2012-217014 640 

done, however, it will prove to be convenient to make the following digression concerning the description 

of a beam wave. 

The consideration of the general solution for Equation (G33) makes it necessary to have an 

expression for the associated boundary condition. This brings to the fore the normalization of spatial 

coordinates. The boundary condition is given by the expression for the general form of a gaussian beam 

wave: 

  2

0 , exp 1 21

A k E z

iz iz

     

           

 (G35) 

where A0 is the initial amplitude of the field and 

12 1 2 2

0

2 1

i

kW F        (G36) 

where W0 is the waist radius of the beam at the exit of the output aperture and F is the focal length. The 

initial condition for the field at the output aperture is thus given by 

  2

2 2

0 0 2

0

1 2 ,0 exp exp 2 2k k EA A i

W F

                                

 (G37) 

At this point, it is convenient to normalize the spatial coordinates with respect to the parameters of the 

problem. For example, if one defines the effective radius of the beam wave at the exit of the output 

aperture as 0 2 eff r W (this definition is consistent with the definition for the power distribution in the 

transverse plane) one can define the dimensionless transverse coordinate  /reff and write 

Equation (G37) as 

  2

2 2 0

0

1 1 ,0 exp 2 2

kW EA i

F

                            

 (G38) 

One can now apply the same coordinate normalization to Equation (G33). That is, since 

    2 2

2

0

2 E , , z Ez

W             (G39) 

Equation (G33) can be written as 

2 22

0 0 2 2 0

2 2

kW k W E i EE

z 

 

 

  (G40) 

(letting 0 = 1 for atmospheric scenarios). Now, normalizing the longitudinal coordinate z with respect to 

the quantity defined by 2

0 2 L kW d  (i.e., z = z/Ld), and noting that 
NASA/TM—2012-217014 641 

    2

0

E , , z Ez 2

z kW z

      

 

 (G41) 

causes Equations (B7) and (B9) to become 

  2

0 ,0 exp 1 2

Ld EA i

F

     

           

 (G7) 

and 

  2 2 22 2 ,0 eff

E i E kr E T E

z 



    

   (G42) 

This form of the quasi-optical equation facilitates the developments that follow. The quantity Ld is called 

the diffraction length and, as seen from Equation (G7), it is the length that the focus must achieve for the 

associated diffraction to become appreciable. 

G.9.1 Critical Power for the Atmospheric The Kerr Effect 

To isolate the effect of the Kerr nonlinearity from the thermal effects due to absorption, one simply 

lets |E|

2

 in Equation (G42) and obtains 

  2 2 22 2 0 eff K

E i E kr E E

z 



    

   (G43) 

The differential equation that governs the evolution of is given by 

2 K

K K E

t

     

 (G44) 

where  is the relaxation constant for the Kerr process and is the Kerr constant for the atmosphere; 

typically,  10–11 s. Relative to the other processes to be considered (and the fact that very short pulse 

lengths are not being considered here), one can ignore the first term on the left of Equation (G44). In this 

case, |E|

2

and Equation (G43) becomes 

2 2 22 2 0 eff K

E i E kr E E

z 



   

   (G45) 

In the literature, Equation (G45) is sometimes stated to describe a cubic medium because of the 

occurrence of E in the last term. If |E|

2

 is connected to the associated intensity I of the beam using the 

relation |E|

2 = (8/c)I and then to the corresponding beam power P0 using 2

0 , P rI   eff Equation (G45) 

becomes 

2 2 0 K

E i E RE

z 



  

   (G46) 
NASA/TM—2012-217014 642 

where RK  P0/PcrK is the nonlinear parameter (also called the distortion parameter) for the Kerr 

nonlinearity defined in terms of the critical power PcrK associated with the Kerr nonlinearity, 

2 8 crK

K

c P

k  

 (G47) 

For Earth’s atmosphere at sea level, K  2.510–16 CGSE units9

 (Ref. 8). Thus, for a laser wavelength of 

= 2.0 m, the Kerr effect has a threshold at PcrK  1.5109

 W. This value is less than 1 order of 

magnitude from the powers that are being considered (~70106

 W, giving RK  0.05) for power-beaming 

applications. Hence, the nonlinearities associated with the Kerr effect may become a minor factor in 

power-beaming applications (and can be easily removed via adaptive optics), but as will be shown in the 

next section, RK is several orders of magnitude below the corresponding parameter for thermal 

nonlinearities, which dominate the beaming process in the atmosphere. 

In the case of millimeter-wave propagation at  = 2.0 mm, one obtains PcrK = 1.51016 W, and at 

P0 = 70106

 W, one obtains RK  4.610–9, thus alleviating any concerns about the Kerr effect at these 

wavelengths. 

G.9.2 Critical Power for Atmospheric Thermal Effects 

G.9.2.1 Continuous Wave Source 

Using Equation (G34) in Equation (G42) and neglecting the Kerr contribution gives 

2 22 2 0 eff T

E i E kr i T E

z k 

             

 (G48) 

The temperature variation from the nominal atmospheric temperature is given by the application of the 

conservation of energy to the atmospheric heat budget, given in its entirety by 

2

p

T C VT T I

t

              

  (G49) 

where  is the density of the atmosphere, Cp is its specific heat at constant pressure, is its thermal 

conductivity, V



is the total relative velocity of the motion (slewing) of the beam across the atmosphere 

and that of the atmospheric wind, and  is the absorption coefficient for the radiation of intensity 

I  (c/8)|E|

2

. Strictly speaking, one should also include the set of Navier-Stokes equations in the 

Boussinesq approximation to account for the viscous and Archimedean forces that occur during the 

thermally induced motion of the atmosphere. However, for the purposes of this discussion and subsequent 

analysis, a description at this level is not required. 

Equation (G49) can be simplified straight away by first specializing the slewing and wind velocity to 

be along the x axis of the coordinate system, V Vx  ˆ  , where xˆ is the unit vector. Furthermore, the 

diffusivity 2

T will be taken to have contributions only in the direction transverse to the laser beam 

propagation: that is, 2 2   T T

 . Finally, only the stationary heating case will be considered for this 

particular discussion whereby the time derivative can be dropped. (That is, the continuous wave (CW) 

radiation is taken to be acting long enough for the steady-state case to be achieved. This is not the case for 

 

9

Electrical units of the centimeter-gram-seconds system of units. 
NASA/TM—2012-217014 643 

pulsed radiation, which is discussed in Section G.9.2.2.) Applying these considerations to Equation (G49) 

allows it to be written as 

2 2

8 p

T c VTE

x C 

        (G50) 

where   /Cp is the associated thermal diffusivity. As was done earlier with Equation (G33), one can 

normalize the transverse coordinates with respect to reff. In addition, the field value E can be normalized 

with respect to E = E/E0 where 2

0 0 8 8. E Ic P r c   eff (This is what was essentially done to 

Equation (G45) to obtain Equation (G46)). Equation (G50) then becomes 

2

2 2

2

0

eff

p

eff eff

r V T C TE

P r xr 

                     

 (G51) 

Now, the two heat-transfer mechanisms that are described here can be isolated. Consider the ratio of 

the coefficients of the two terms within the parentheses of Equation (G51):   Vreff/. This 

dimensionless ratio is known in fluid mechanics as the Peclet number. Taking the nominal wind velocity 

V = 4.4 m/s (10 mi/hr), reff = 10 m, and using the documented value for the thermal diffusivity of the 

atmosphere  = 2.1210–5 m2

/s, one has that  = 2.1106 >> 1, allowing one to neglect the second term in 

the brackets. In this instance, one can write 

2

0

Cr V p eff T T E

P xx

                

 (G3) 

where the temperature has been normalized with respect to the characteristic temperature 

TV  P0/CpreffV. Finally, applying this prescription to Equation (G48) gives 

2 22 2 0 d eff T V

E i E iL E k r T T E

z 

 

            (G52) 

As done with Equation (G45), one can define the distortion parameter for the thermal nonlinearity 

2

2 2 eff T 0 0

V eff T V

p crT

kr P P R kr T

CV P

 

  



 (G19) 

where 

2

p

crT

eff T

C V

P

k r

   

 (G1) 

is the critical power associated with the thermal nonlinearity in the case of large Peclet numbers. Using 

Equation (G19) and noting that the third term in Equation (G52) can be transformed away, one has 
NASA/TM—2012-217014 644 

  2 2 0 exp 2 V d

E i E RTE E E L z

z 

 

               (G53) 

The problem of propagation through a thermal nonlinearity thus reduces, to within the 

approximations adopted, to the solution of Equations (G3) and (G53). For Earth’s atmosphere,  = 1.2 

kg/m3

, Cp = 1.0103 J/K-kg, and T  –2.310–6 K–1. That is, the atmospheric permittivity decreases with 

an increase in temperature. This indicates that the laser beam will steer toward regions with larger 

permittivity: that is, cooler regions. In the calculations, the absolute value of this quantity is used. At 

 = 2.0 m,   410–6 m–1. If one uses the earlier stated values for the other quantities, one finds that 

PcrT = 16.7 W. Thus, for P0 = 70106

 W, RV  4.2106. The very small value for the critical power occurs 

because of the rather large beam radius. The temperature must diffuse over the large cross section of a 

cylindrical column. Of course, larger values of V will increase PcrT. For a smaller, more typical value for a 

beam radius, reff = 0.1 m, one finds that PcrT = 1667 W and RV  4.2104

 for P0 = 70 MW—still a rather 

large value for the nonlinearity parameter. As a comparison, a more typical laser power of P0 = 100 kW

corresponds to RV  60. It is thus seen that nonlinearities due to thermal effects at infrared wavelengths 

within the atmosphere will dominate those of the Kerr effect. 

The small values for PcrT at the large output aperture radii and the corresponding very large values for 

RV do indeed have severe ramifications for the long distance propagation of beam waves at P0  70 MW 

as shown earlier. However, these values facilitate the use of approximation procedures to be used in 

solving Equations (G3) and (G53) from which specifications for the adaptive correction of the various 

aberrations are derived, and it is shown that complete mitigation is possible. 

The situation for millimeter-wave propagation is better. For  = 2.0 mm,  = 2.210–4 m–1 and 

PcrT = 3.5105 W, and for P0 = 70 MW, RV = 200.8. Although the absorption coefficient  is 2 orders of 

magnitude larger than that for infrared wavelengths, the overall optical effects due to the thermally 

induced refractive index variations are much smaller because of the larger wavelength: the “thermal lens” 

that is induced has weaker focusing capabilities at millimeter wavelengths. 

G.9.2.2 Pulsed Source 

In the case of a source that produces radiation pulses of temporal duration tp, the prevailing fluid 

dynamics of the situation differs from that of a CW source. Here, for tp < reff/V, the convection/diffusion 

equation, Equation (G49), is approximated as 

p

T C I

t



  

 (G54) 

Coupling this equation with the propagation equation, Equation (G48), and following the same 

normalization procedure as outlined in Section G.9.2.1 for the CW case and normalizing the time to tp, 

one obtains in place of Equation (G53), 

  2 2 0 exp 2 P d

E i E RTE E E L z

z 

 

               (G55) 

where 

2

2

0

p eff

p

C r T T E

Pt t t

                

 (G56) 
NASA/TM—2012-217014 645 

and the prevailing nonlinear parameter is now 

2

T p 0 0

P

p crP

k tP P R

C P

    

 (G57) 

in which the corresponding critical power is 

2

p

crP

T p

C

P

k t

  

 (G2) 

which is independent of reff. 

Taking a nominal pulse width of tp  1 s, one has for  = 2.0 m, PcrP = 3.7107

, and for 

 = 2.0 mm, PcrP = 7.71011. Thus, temporally modulating the source does not elucidate the convective 

thermal effects that are inherent in the CW case and allows for much larger critical values of peak power. 

Pulse lengths smaller than 1 s require the use of Equation (G32), which takes into account pulse 

dispersion through the medium. This effect is especially important when atmospheric turbulence is taken 

into account. However, for the purposes of the present analysis, which is to obtain the critical powers at 

which various nonlinear propagation effects show themselves, these effects are not considered. 

The issue of the repetition rate of a pulse train within the atmosphere in conjunction with scanning 

(slewing) the beam through a region of the atmosphere to aid in heat dispersion (for high pulse repetition 

rates) is characterized by yet another critical power threshold, which will be larger than that given by 

Equation (G2). 

G.9.3 Summary 

The parameters RK, RV, and RP provide for a comparative assessment of the nonlinear propagation 

mechanisms that are induced because of the large field power densities. Thus, at the power levels 

considered here, the effect of Kerr nonlinearities are almost negligible relative to that of thermal effects, 

and it is the later that must be completely understood so as to be effectively mitigated via properly 

designed adaptive optics algorithms. They do not, however, give an assessment of the behavior of specific 

laser beam properties, such as the beam steering and beam broadening that accompany propagation. This 

can only be done through modeling using the specific propagation geometries that are being considered. 

The very large values for RV in the case of the power levels considered for beamed propulsion are a 

benefit in that they allow approximation procedures to be applied to analytically evaluate propagation and 

adaptive optics models. 
NASA/TM—2012-217014 646 

G.10 References for Appendix G 

1. Manning, R.M.: Stochastic Electromagnetic Image Propagation and Adaptive Compensation. Ch. 2, 

McGraw-Hill Book Co., New York, NY, 1993. 

2. Wheelon, A. D.: Electromagnetic Scintillation: Vol. 2, Weak Scattering. Cambridge University Press, 

2003. 

3. Litvak, A.G.: Self Focusing of Powerful Light Beams by Thermal Effects. Sov. Phys. JETP Letters, 

vol. 4, no. 9, 1966, pp. 230232. 

4. Smith, D.C.: High Power Laser Propagation—Thermal Blooming. Proc. IEEE, vol. 65, no. 12, 1977, 

pp. 16791714. 

5. Vorob’ev, V.V.: Thermal Blooming of Laser Beams in the Atmosphere. Prog. Quantum Electron., 

vol. 15, 1991, pp. 1152. 

6. Keldysh, L.V.: Ionization in the Field of a Strong Electromagnetic Wave. Sov. Phys. JETP, vol. 20, 

1965, pp. 13071314. 

7. Boyd, R.W.: Nonlinear Optics. Ch. 4, Academic Press, San Diego, CA, 1992. 

8. Sutherland, R.L.: Handbook of Nonlinear Optics. Marcel Dekker, Monticello, NY, 1996. 

9. Reichert, J.D.; Wagner, W.G.; and Chen, W.Y.: Instabilities of Intense Laser Beams in Air. J. Appl. 

Phys., vol. 44, no. 8, 1973, pp. 36413646. 

10. Reichert, J.D.; Wagner, W.G.; and Chen, W.Y.: Propagation of an Intense Gaussian Laser Pulse in 

Air. J. Appl. Phys., vol. 44, no. 8, 1973, pp. 36473658. 

11. Fleck, J.A.; Morris, J.R.; and Feit, M.D.: Time-Dependent Propagation of High Energy Laser Beams 

Through the Atmosphere. Appl. Phys., vol. 10, 1976, pp.129160. 

12. Vedenov, A.A.; and Markin, O.A.: Propagation of Intense Laser Radiation in an Absorbing Medium. 

Sov. Phys. JETP, vol. 49, no. 4, 1979, pp. 608615. 

13. 13. Sprangle, P.; Penano, J.; and Hafizi, B.: Propagation of High Energy Laser Beams in Various 

Environments. Report prepared for High-Energy-Laser Joint Technology Office (NRL Report 

NRL/MR/6790079032), May 2007. 14. Fischer, R., et al.: Absorption and Scattering of 1.06m 

Laser Radiation From Oceanic Aerosols. Appl. Opt., vol. 48, no. 36, 2009, pp. 69906999. 

14. Glickler, S.L.: Propagation of a 10.6 Laser Through Cloud Including Droplet Vaporization. Appl. 

Opt., vol. 10, no. 3, 1971, pp. 644650. 

15. Zuev, V.E.; and Kuzikovskii, A.V.: Thermal Bleaching of Aqueous Aerosols by Laser Radiation. 

Russ. Phys. J., vol. 20, no. 11, 1977, pp. 14801500. 

16. Agrovskii, B.S., et al.: Thermal Blooming of Laser Beams in Turbulent Media. J. Opt. Soc. Am., vol. 

A2, no. 12, 1985, pp. 23042312. 

17. Bradley, L.C.; and Herrmann, J.: Phase Compensation for Thermal Blooming. Appl. Opt., vol. 13, no. 

2, 1974, pp. 331334. 

18. Novoseller, D.E.: Zernike-Ordered Adaptive-Optics Correction of Thermal Blooming. J. Opt. Soc. 

Am., vol. A5, no. 11, 1988, pp. 19371942. 

19. NIST Handbook of Mathematical Functions. Ch. 18, Cambridge University Press, 2010. 

20. Penano, J.R.; Sprangle, P.; and Hafizi, B.: Propagation of High Energy Laser Beams Through 

Atmospheric Stagnation Zones. NRL Report NRL/MR/6790068925, 2006. 

21. Hardy, J.W.: Active Optics: A New Technology for the Control of Light. Proc. IEEE, vol. 66, 1978, 

pp. 651697. 

22. Vorontsov, M.A., et al.: Adaptive Imaging System for Phase-Distorted Extended Source/Multiple 

Distance Objects. Appl. Opt., vol. 36, 1997, pp. 33193328. 

23. Vorontsov, M.A., et al.: Adaptive Optics Based on Analog Parallel Stochastic Optimization: Analysis 

and Experimental Demonstration. J. Opt. Soc. Am., vol. A17, no. 8, 2000, pp. 14401453. 
NASA/TM—2012-217014 647 

24. Vorontsov, M.A.: Decoupled Stochastic Parallel Gradient Descent Optimization for Adaptive Optics: 

Integrated Approach for Wave-Front Sensor Information Fusion. J. Opt. Soc. Am., vol. A19, no. 2, 

2002, pp. 356368. 

25. Vorontsov, M.A.; and Kolosov, V.: Target-in-the-Loop Beam Control: Basic Considerations for 

Analysis and Wave-Front Sensing. J. Opt. Soc. Am., vol. A22, no. 1, 2005, pp. 126141. 

26. Piatrou, P.; and Roggemann, M.: Beaconless Stochastic Parallel Gradient Descent Laser Beam 

Control: Numerical Experiments. Appl. Opt., vol. 46, no. 27, 2007, pp. 68316842. 

27. Vorontsov, M.A.; Kolosov, V.V.; and Kohnle, A.: Adaptive Laser Beam Projection on an Extended 

Target: Phase- and Field-Conjugate Precompensation. J. Opt. Soc. Am., vol. A24, no. 7, 2007, 

pp. 19751993. 

28. Primmerman, C.A.; and Fouche, D.G.: Thermal-Blooming Compensation: Experimental 

Observations Using a Deformable-Mirror System. Appl. Opt., vol. 15, no. 4, 1976, pp. 990995. 

29. Primmerman, C.A.; Johnson, F.B.; and Wigdor, I.: Thermal Blooming Compensation Using the 

CLASP System. Appl. Opt., vol. 17, no. 18, 1978, pp. 29092912. 

30. Fouche, D.G.; Higgs, C.; and Pearson, C.F.: Scaled Atmospheric Blooming Experiments (SABLE). 

Lincoln Lab. J., vol. 5, no. 2, 1992, pp. 273293. 

31. Qiao, C.-H., et al.: Simulation Experiment of High Energy Laser Propagation in the Atmosphere. 

Proc. SPIE, vol. 7382, Art.# 73822I, 2009. 

32. Molchan, M.A.; Doktorov, E.V.; and Vlasov, R.A.: Propagation of Vector Fractional Charge 

Laguerre-Gaussian Light Beams in the Thermally Nonlinear Moving Atmosphere. Opt. Lett., vol. 35, 

no. 5, 2010, pp. 670672. 

NASA/TM—2012-217014 649 

Appendix H.—Summary Presentation 

Beamed Energy Propulsion (BEP) Study

Final Stakeholder Review

Raymond F. Beach

Patrick J. George

Briefing prepared for:

NASA, Office of Chief Technologist

DARPA, Tactical Technology Office

NASA, Space Operations Mission Directorate

NASA, Office of Chief Engineer

NASA HQ

14 April 2011

Final Report 1

Review outline

2

• Executive summary

• Original objectives / direction from mid-term review

• Assessment of feasibility, potentials, challenges

• Recommended developments

• Summary

• Details of findings

• LEO to GEO

• Deep space

• Launch

• A path forward

• Technology investments

• BEP development program

• Summary

Final Report
NASA/TM—2012-217014 650 

Direction from the mid-term review:

This memorandum is a confirmation of the direction provided at the 

Beamed Energy Propulsion midterm review on 16 December 2010. The 

overall study direction received is to reduce the current scope and 

focus on areas which have the highest payoff. This includes:

1. Focus the Millimeter wave thermal rocket (DRM 1-C') as an 

expendable design making an easier direct comparison with the 

laser design (DRM 1-B). (pgs. 43-50)

2. Modify the in-space missions (DRMs 2 and 3) to include in-space 

power generation such as a solar array powering an orbiting laser 

beaming station. (pgs. 25 & 34)

3. Drop any further development of the In-Space GEO Servicing 

mission (DRM 2-B) and the Horizontal Launch mission (DRM4)

4. Do not consider the effects of launch rates for this study

5. Do not consider the amortization of non-recurring vehicle and 

ground system costs

6. Develop ROM non-recurring costs for vehicles and ground 

systems (pgs. 26, 36 & 50)

7. Develop ROM vehicle and ground facility costs for recurring 

launches 

8. Investigate and analyze the effects of thermal blooming on laser 

and millimeter wave propagation through the atmosphere. (pg. 

28)

9. In-space missions will be reported to the same depth as the 

launch missions (pgs.22-36)

10. Emphasize concepts utilizing and extending from high TRL 

commercial lasers like diode-pumped lasers. De-scope concepts 

that require low TRL lasers. (pg. 45)

11. The primary product of this study needs to be a determination of 

TRL1s for near term concepts of:

a) beam energy technology (pg 58)

b) rocket vehicle technology (pg 58)

Summary of BEP study original task 

and mid-term redirection

3

What makes BEP attractive ?

The attraction of beamed energy propulsion is 

the potential for high-specific impulse while 

removing the power-generation mass from 

the spacecraft.

Why look at BEP now ?

The rapid advancements in high-energy 

beamed power systems and optics over the 

past 20 years warrant a fresh look at the 

technology and potential game-changing 

applications.

Goals & objectives:

Review and analyze the state-of-art in 

beamed energy propulsion and: 

(a) identify potential game-changing 

applications (pgs. 13 &14)

(b) formulate a roadmap of technology 

development (pg. 59)

(c) identify key near-term technology 

demonstrations to rapidly advance 

elements of BEP technology to TRL 6. 

(pgs. 41 & 56)

Final Report

Launch:

To earth orbit

LEO to GEO:

Low earth to geosynchronous 

orbit transfer module

Deep space:

Transport to Jupiter/Saturn

4

Beamed energy propulsion applications

Final Report
NASA/TM—2012-217014 651 

Technically Feasible Game Changing Economically Viable

• Closed design solution –

trajectory, beam power, 

thrust/weight, mass etc.

• For low TRL - defined basis for 

technology and technology 

improvements

• Allowable margins – mass, engine 

performance, thermal, structural

• Reusability

• Payload mass fraction 

• Improved mission capability –

increase operations time, etc.

• Reduced Transfer Time (2X)

• “...Technology that is innovative,

unique and promises to enable 

revolutionary improvements to 

the efficiency and effectiveness of 

our country’s space capability.” 

[OCT BAA]

• Lower vehicle costs ($/kg)

• Lower launch facility operation 

costs – range safety, vehicle 

prep/handling

• Infrastructure costs

LEO to GEO YES

Low TRL’s

YES/NO, but might be 

combined with other 

applications

NO, high cost of 

beaming stations

Deep Space YES

Low TRL’s YES, with investments NO, high cost of beaming stations

Launch YES

Low TRL’s YES, with investments NO, high cost of beaming stations

Feasibility 

Ranking

YES

Green – All

YES/NO

Yellow - >50%

NO

Red - <50%

Beamed Energy Propulsion feasibility study results

• Beamed Energy Propulsion is feasible, but very challenging for high power, with large upfront costs

• Study findings indicated two areas of great interest:

• Revolutionary propulsion system for launch and space

• Additional finding – Remote operation of small spacecraft via laser beam

Final Report 5

Technically Feasible Game Changing Economically Viable

• Closed design solution –

trajectory, beam power, 

thrust/weight, mass etc.

• For low TRL - defined basis for 

technology and technology 

improvements

• Allowable margins – mass, engine 

performance, thermal, structural

• Reusability

• Payload mass fraction 

• Improved mission capability –

increase operations time, etc.

• Reduced Transfer Time (2X)

• “...Technology that is innovative,

unique and promises to enable 

revolutionary improvements to 

the efficiency and effectiveness of 

our country’s space capability.” 

[OCT BAA]

• Lower vehicle costs ($/kg)

• Lower launch facility operation 

costs – range safety, vehicle 

prep/handling

• Infrastructure costs

LEO to GEO YES

Low TRL’s

YES/NO, but might be 

combined with other 

applications

NO, high cost of 

beaming stations

Deep Space YES

Low TRL’s YES, with investments NO, high cost of beaming stations

Launch YES

Low TRL’s YES, with investments NO, high cost of beaming stations

Feasibility 

Ranking

YES

Green – All

YES/NO

Yellow - >50%

NO

Red - <50%

LEO to GEO mission evaluation

Final Report 6
NASA/TM—2012-217014 652 

Technically Feasible Game Changing Economically Viable

Potentials

• Scalable for various size spacecraft

• Primary and secondary optics 

reduces pointing requirements

• Can Be Time Phased with In-Space 

BEP Development to provide 

Optimum Program Development 

Scenario

• Synergy with Manned GEO 

Servicing mission

• Environmentally friendly water 

based propellant and simplified fuel 

depot requirements

• Reusable

• Increased payload mass fraction

• High thrust, high Isp propulsion 

engine utilizing water for propellant

• LEO to GEO Vehicle Synergistic 

with Deep Space Transfer Stage

• Reusable design for lower life cycle 

cost

Challenges

• Beaming 80 MW through 

atmosphere. Max current is Air 

Borne Laser Test Bed MW class

• Operation of GEO based remote 

complex, high-energy beaming 

station. 

• 188 nanoradians pointing and 

tracking (TRL 2-3). Best proven is 

Hubble Space Telescope 50 nr. 

Web Telescope 24 nr

• Required technologies at low TRL

• Plasma engine (TRL 2)

• Reflective mirror: 4m x 6m; 

99.99% reflectivity, 

micrometeoroid tolerant (TRL 2)

• Multi MW space/ground 

operational laser (TRL 2-3)

• High cost to establish beaming 

station on ground & in GEO

RCS Pod

Solar Array

Secondary

Reflector

Patch Antennas

(every 90 deg around

bus)

Propellant

(Water) Tank

Engine

Shielding for Engine

RCS Tank

Pressurant

Tank

Primary Reflector

LEO to GEO mission evaluation:

Potentials and challenges

Final Report 7

Technically Feasible Game Changing Economically Viable

• Closed design solution –

trajectory, beam power, 

thrust/weight, mass etc.

• For low TRL - defined basis for 

technology and technology 

improvements

• Allowable margins – mass, engine 

performance, thermal, structural

• Reusability

• Payload mass fraction 

• Improved mission capability –

increase operations time, etc.

• Reduced Transfer Time (2X)

• “...Technology that is innovative,

unique and promises to enable 

revolutionary improvements to 

the efficiency and effectiveness of 

our country’s space capability.” 

[OCT BAA]

• Lower vehicle costs ($/kg)

• Lower launch facility operation 

costs – range safety, vehicle 

prep/handling

• Infrastructure costs

LEO to GEO YES

Low TRL’s

YES/NO, but might be 

combined with other 

applications

NO, high cost of 

beaming stations

Deep Space YES

Low TRL’s YES, with investments NO, high cost of beaming stations

Launch YES

Low TRL’s YES, with investments NO, high cost of beaming stations

Feasibility 

Ranking

YES

Green – All

YES/NO

Yellow - >50%

NO

Red - <50%

Deep Space mission evaluation

Final Report 8
NASA/TM—2012-217014 653 

Technically Feasible Game Changing Economically Viable

Potentials

• Scalable for various size 

spacecraft

• Primary and secondary optics 

reduces pointing requirements

• Can be time phased with InSpace BEP Development to 

provide optimum program 

Development Scenario

• Synergy with Manned GEO 

Servicing mission

• Faster transit to outer planets 

(2X)

• Environmentally friendly water 

based propellant and simplified 

fuel depot requirements

• 118 kg of instrument payload 

vs. 82 kg (Enceladus Mission 

Study) 

• Deep Space Transfer Stage 

Vehicle Synergistic with LEO to 

GEO

• Can Be Time Phased With Inspace BEP Development To 

Provide Optimum Program 

Development Scenario

Challenges

• Required technologies at low 

TRL

• Plasma engine (TRL 2)

• Reflective mirror: 4m x 6m; 

99.99% reflectivity, 

micrometeoroid tolerant 

(TRL 2)

• Multi MW space/ground 

operational laser (TRL 2-3)

• 2.3 Nanoradians pointing and 

tracking (TRL 2-3). Best 

proven is Hubble Space 

Telescope 50 nr

• Beaming 80 MW through 

atmosphere. Max current is 

Airborne Laser Test Bed MW 

class

• High cost to establish beaming 

station in GEO or on ground

• High operation cost of GEO 

based remote complex, highenergy beaming station. 

Deep Space mission evaluation:

Potentials and challenges

Final Report 9

Launch mission evaluation

Technically Feasible Game Changing Economically Viable

• Closed design solution –

trajectory, beam power, 

thrust/weight, mass etc.

• For low TRL - defined basis for 

technology and technology 

improvements

• Allowable margins – mass, engine 

performance, thermal, structural

• Reusability

• Payload mass fraction

• Improved mission capability –

increase operations time, etc.

• Reduced Transfer Time (2X)

• “...Technology that is innovative,

unique and promises to enable 

revolutionary improvements to 

the efficiency and effectiveness of 

our country’s space capability.” 

[OCT BAA]

• Lower vehicle costs ($/kg)

• Lower launch facility operation 

costs – range safety, vehicle 

prep/handling

• Infrastructure costs

LEO to GEO YES

Low TRL’s

YES/NO, but might be 

combined with other 

applications

NO, high cost of 

beaming stations

Deep Space YES

Low TRL’s YES, with investments NO, high cost of beaming stations

Launch YES

Low TRL’s YES, with investments NO, high cost of beaming stations

Feasibility 

Ranking

YES

Green – All

YES/NO

Yellow - >50%

NO

Red - <50%

Final Report 10
NASA/TM—2012-217014 654 

Technically Feasible Game Changing Economically Viable

Potentials

• Removal of the launch powergeneration mass 

• High-specific impulse engines

• Established trajectory for each DRM 

and identified power levels needed at 

vehicle and ground source

• Identified mass and performance 

margins required 

• Identified all technology performance 

requirements

• Use of less toxic launch propellants

• Multiple automated launches per day 

from same facility

• Launch processing, vehicle and 

operational safety

• Low cost launch alternative for 

small payloads (1-100kg) 

• Use of existing laser / microwave 

technology with reduction of facility 

costs as development occurs

• Automated processing and launch

• Identified electrical energy and 

storage costs to allow buffer from 

utility power systems 

Challenges

• Large number of low TRL 

technologies need to be developed 

(synchronized lasers; large dish 

antenna; heat exchanger; plasma 

engine; atmospheric adaptive optics)

• Microradian pointing and tracking 

(TRL 2-3). Best proven is Hubble 

Space Telescope 50 nanoradians

• Beaming up to 600 MW through the 

atmosphere. Max current is Airborne 

Laser Test Bed low single MW class

• High cost to establish beaming 

station on ground

• All concepts did not meet advertised 

level of cost to orbit

• Large launch range safety zone, 

greater than Shuttle

11

Launch mission evaluation:

Potentials and challenges 

Final Report

Technically Feasible Game Changing Economically Viable

LEO to GEO YES

Low TRL’s

YES/NO, but might be 

combined with other 

applications

NO, high cost of 

beaming stations

Deep Space YES

Low TRL’s YES, with investments NO, high cost of beaming stations

Launch YES

Low TRL’s YES, with investments NO, high cost of beaming stations

Answered feasibility questions:

• LEO to GEO and Deep Space:

• Near, mid, far term applications

• Low, medium, high power/flux

• Launch:

• Far term application

• High power/flux

• Synergistic with solar thermal propulsion

• BEP can enable significant increase in payload 

capability for LEO to GEO transfer, and faster 

transit times to planets

• Costs of solar GEO based beaming station(s) 

and reflector(s) is high

• Did not meet $/kg launch costs. Projections 

indicate investment could result in high payoff 

• Costs of multiple ground beaming station(s) or 

on-orbit reflector(s) is high

Summary of study results

Final Report 12
NASA/TM—2012-217014 655 

Remote control of small spacecraft 

through the integration of propulsion, 

power, communications, and 

navigation/attitude control Thermal plasma engine, using water 

as propellant, provides high thrust 

and high Isp

Beamed Energy Propulsion feasibility study results:

Unexpected findings

Final Report 13

Benefits

• Remote operation – coordinated maneuver with 

source and receiving spacecraft without 

mechanical connection

• Reduce complexity, mass and cost of the 

receiving spacecraft

• Applicable for variety of space and aero vehicles 

with different propulsion systems

• Possible to control beam receiving “Shepherd” 

spacecraft from “Mother” spacecraft to perform 

specific maneuvers - similar to berthing

Remote control of small spacecraft 

through the integration of propulsion, 

power, communications, and 

navigational/attitude control through a 

laser beam 

Concept

• Think small spacecraft

• Think integrated systems

• Think space

Unexpected findings:

Remote spacecraft positional control

14

If successfully developed, could lead to a new class of spacecraft 

with enhanced operational capability, and lower mass.

Final Report
NASA/TM—2012-217014 656 

Benefits

• Able to utilize diverse new propellant types (water, 

ammonia, etc.)

• Eliminate toxic chemicals and associated safety issues both 

in space and on the ground

• Environmentally friendly water propellant potentially 

available on the Moon and Mars. 

• Can utilize solar, laser, or mm wave beamed energy

• Potential for highest performance – eliminates oxidizer or 

energy source on spacecraft

• Optical-to-optical high efficiency energy transfer 

• Both high thrust, and high Isp

Concept

• Think water as propellant

• Think launch and space

• Think combined beamed energy 

and thermal plasma engine

Potential for highest propulsion 

engine performance – eliminates 

oxidizer / energy source on 

spacecraft

Unexpected findings:

Beamed Energy Propulsion engine

15

Currently at very low TRL. If successfully developed, could lead to 

higher thrust/mass performance than current propulsion systems 

and the resultant lower mass.

Final Report

Implications from study:

Beamed energy technology directions

Development path

• Develop standard / protocol for Beamed Energy 

transmission:

• Types, quality, regulations, organizations

• Ground, aircraft, UAV, space

• Develop and demonstrate integrated power, 

propulsion, control/communications:

• Ground, aircraft, UAV, space

• NASA/USAF last/solar electric propulsion demonstration

• DARPA Exo-SPHERES

• Develop and demonstrate engine propulsion 

technologies:

• Airborne Laser, X prize, NASA Centennial Challenge

Development synergy

• Scalable for both space and launch 

applications

• Will provide benefits for near, mid, and far 

term systems

• Builds on early solar thermal propulsion 

work:

• Air Force/NASA integrated solar upper state

• DARPA HiDVE

• NASA inflatable concentrator/secondary lens, 

optical engine cavity

Manned GEO Servicing Remote

Positional Control

Fractionated Satellites

F6 Fuel Depots

2015 2025 2020

Small

Launch

ETO

STP

Low flux, low power

2035

BEP

High flux, low power

BEP ETO

High flux, power

Beamed Energy

Protocols/Standards

LEO to Geo

Tug Thermal Plasma Engine

Final Report 16
NASA/TM—2012-217014 657 

• Executive summary

• Original objectives / direction from mid-term review

• Assessment of feasibility, potentials, challenges

• Recommended developments

• Summary

• Details of findings

• LEO to GEO

• Deep space

• Launch

• A path forward

• Technology investments

• BEP development program

• Summary

Review outline

Final Report 17

Review outline

Mid‐term direction:

2. Modify the in‐space missions

(DRMs 2 and 3) to include in‐

space power generation such as

a solar array powering an

orbiting laser beaming station.

(pgs. 25  & 33)

3. Drop any further development of

the Space GEO Servicing mission

(DRM 2‐B) and the Horizontal

Launch mission (DRM4)

• Executive summary

• Original objectives / direction from mid-term review

• Assessment of feasibility, potentials, challenges

• Recommended developments

• Summary

• Details of findings

• LEO to GEO

• Deep space

• Launch

• A path forward

• Technology investments

• BEP development program

• Summary

Final Report 18
NASA/TM—2012-217014 658 

19

• All BEP mission defined by study team included consultants:

• Utilized outside experts/advocates, Air Force, academia to help in mission concept 

development

• All Missions are unique to BEP study

• Both GRC and GSFC mission design centers used as independent analyst to evaluate 

BEP missions

• LEO to GEO and Deep Space mission concept development:

• Initial evaluations considered efficiency (optics, energy conversions, regulators, 

cables/connectors/switchgear, thruster), mass, and response for SEP/beamed 

electric, STP/beamed thermal, and solar sails/beamed sail Mass

• Decision made to base space missions on thermal propulsion

• Synergistic with beamed thermal for ETO

• High ISP and thrust as FOM referenced in BEP study plan

Space study envelope

Consultants may not agree with all study findings.

Final Report

20

BEP study comparison of propulsion methods

• Thermal propulsion energy technology:

• Synergistic with solar thermal propulsion 

technology

• Technology demonstrated at system and 

component levels:

• ISUS, HiDVE

• RCS

• Significant performance enhancement with 

optical engine:

• Lower TRL

Beamed electric system efficiency – Space 41%

S/A Array

Regulator

Cables

Connectors DDCU Switchgear PPU Thruster CablesConnectors

55% 97% 98% 92% 99% 98% 92% 96%

Beamed thermal system efficiency - launch and space

Primary

Mirror

Secondary

Lens

Optical

Cavity Thruster

89% 98% 98% 80%

Heat

Exchanger Thruster

81% 80%

Heat exchanger engine 65%

Optical engine 68%

Final Report
NASA/TM—2012-217014 659 

Comparison of propulsion methods

11.7 m

3 m

• Centaur:

• 2000 kg dry mass

• 21000 kg propellant (LOX/H2)

• 2,000 kg payload

• RL10 110,000 Newtons thrust

• 464 sec. Isp

• 5 hours to GTO

• LEO to GEO tug:

• 500 kg dry mass

• 4900 kg water propellant

• Two 2000 kg satellites transferred GEO

• 5000 Newtons thrust

• 813 sec. Isp

• Variable transfer time to GEO 

dependent on number of ground 

stations:

• 5 Stations, 32 hours

• 2 Stations, 11 days

• 1 Station, 20 days

• Reusable

Final Report 21

National Aeronautics and Space Administration

www.nasa.gov

Low Earth to Geosynchronous Orbit Transfer Mission

Ground Based Beaming Station

LEO

R

C

S

P

o

d

P

r

o

p

e

l

l

a

n

t

GEO (W

a

t

e

r

)

T

a

n

k

E

n

g

i

n

e S

h

i

e

l

d

i

n

g

f

o

r

E

n

g

i

n

e

R

C

S

T

a

n

k

• Designed for 10 transits, 20 

GEO satellites delivered

• 20 days total transit time vs. 

7 days with chemical system

LEO-GEO “Tug” with Laser Propulsion

• Launch of two 2000 kg GEO satellites and 

propellant on Falcon 9 from Kwajalein Atoll

• 400 km circular orbit, 9o inc.

• Autonomous rendezvous with tug

• Spiral transfer

Vehicle 

• Thermal plasma engine

• H2O water propellant

• 6000 K, 50 psia chamber press, 813 s Isp

• 550 kg dry mass, 4900 kg propellant

REUSABILITY ENABLES REDUCTION IN LAUNCH & TRANSPORTATION COSTS

BUT LARGE INITIAL INFRASTRUCTURE COSTS

Beaming Station Located on Ground

• Single 78 MW emitter, 195MW input

• 30m dia. Dish

• Located in low moisture environment 

close to Equator - Hawaii 

Final Report 22
NASA/TM—2012-217014 660 

Final Report 23

LEO to GEO ground and space facilities

Ship Based

80MW Output

3 m Diameter Aperture

80 MW Output

30 m Diameter

1150 to 2900 x 

120 kW modules; 

Diode-Pumped 

Solid State Lasers

0.5 MW Output

Diode-Pumped Solid 

State Lasers

1 MW Space Platform 

With or Without 

Steering Mirrors

Potential synergy with launch ground system and facility control:

• Dual use for ground system

• Number of ground and ship based facilities dependant on performance requirements

• Moveable shipped based system can reduce facility numbers without performance reduction

LEO to GEO

Apogee Facility Perigee Facility

LEO to GEO

24

• 500 kg dry mass tug:

• 50 N thermal plasma 

engine

• 813 s Isp

• Highly reflective 

shielding for tug and 

satellites

• Gimbaled primary 

reflector to track laser

• Single Falcon 9 launch

• Two 2000 kg GEO 

satellites transferred:

• 4900 kg water propellant

• Single Falcon 9 launch

RCS Pod

Solar

Array

Secondary

Reflector

Patch Antennas

(every 90 deg

around bus)

Propellant

(Water)

Tank

Engine

Shielding for

Engine

RCS Tank

Pressuran

t Tank

Primary

Reflector

GEO

Satellite

Primary Reflector

4.0 x 5.6 meter elliptical

Propulsion Stage

(Tug)

Single Launch

5.78 m 2.06 m

Propellant

Stage

GEO

Satellite

Single Launch

1.60 m

2.46 m

LEO to GEO laser transport:

Vehicle description

Final Report
NASA/TM—2012-217014 661 

National Aeronautics and Space Administration

www.nasa.gov

Low Earth to Geosynchronous Orbit Transfer Mission

Space Based Beaming Station

Beaming Station

• Single 0.5 MW in-space GEO emitter

• 1.5 MW input required 

• 70 “FAST” arrays equivalent – ISS sized 

spacecraft

• Ground controlled

• Radiate 0.5 MW waste heat

LEO

R

C

S

P

o

d

P

r

o

p

e

l

l

a

n

t

GEO (W

a

t

e

r

)

T

a

n

k

E

n

g

i

n

e S

h

i

e

l

d

i

n

g

f

o

r

E

n

g

i

n

e

R

C

S

T

a

n

k

• Designed for 10 transits, 20 

GEO satellites delivered

• 20 days total transit time vs. 7 

days with chemical system

LEO-GEO “Tug” with Laser Propulsion

• Launch of two 2000 kg GEO satellites and 

propellant on Falcon 9 from Kwajalein Atoll

• 400 km circular orbit, 9o inc.

• Autonomous rendezvous with tug

• Spiral transfer

Vehicle 

• Thermal plasma engine

• H2O water propellant

• 6000 K, 50 psia chamber press, 813 s Isp

• 550 kg dry mass, 6400 kg propellant

REUSABILITY ENABLES REDUCTION IN LAUNCH & TRANSPORTATION COSTS

BUT LARGE INITIAL INFRASTRUCTURE COSTS

Final Report 25

Executive summary of study findings:

LEO to GEO mission 

* Extrapolation based on mission analysis of high efficiency space power 

systems project performed by COMPASS; 300kw/$1B

** Based on Launch mission COMPASS analysis

Mid-term direction:

1. Do not consider the effects of launch rates for this study

2. Do not consider the amortization of non-recurring vehicle and ground system costs

3. Develop ROM non-recurring costs for vehicles and ground systems (pgs. 26, 36 & 50)

4. Develop ROM vehicle and ground facility costs for recurring launches 

Space vehicle cost assumptions:

• Prime contractor costs (no fee)

• No technology development costs

• Protoflight development approach

• Before reserves, insight/oversight, launch services
NASA/TM—2012-217014 662 

Technically Feasible Game Changing Economically Viable

Potentials

• Scalable for various size spacecraft

• Primary and secondary reduces 

pointing capability

• Can Be Time Phased with In-Space 

BEP Development to provide 

Optimum Program Development 

Scenario

• Synergy with Manned GEO 

Servicing mission

• Environmentally friendly water 

based propellant and simplified fuel 

depot requirements

• Reusable

• Increased payload

• High thrust, high Isp propulsion 

engine utilizing water for propellant

• LEO to GEO Vehicle Synergistic 

with Deep Space Transfer Stage

• Reusable design for lower life cycle 

cost

Challenges

• Beaming 80 MW through 

atmosphere. Max current is 

Airborne Laser Test Bed MW class

• Operation of GEO based remote 

complex, high-energy beaming 

station. 

• 188 nanoradians pointing and 

tracking (TRL 2-3). Best proven is 

Hubble Space Telescope 50 nr

• Required technologies at low TRL

• Plasma engine (TRL 2)

• Reflective mirror: 4m x 6m; 

99.99% reflectivity, 

micrometeoroid tolerant (TRL 2)

• Multi MW space/ground 

operational laser (TRL 2-3)

• High cost to establish beaming 

station on ground and/or in GEO

27

LEO to GEO mission evaluation:

Potentials and challenges

Final Report

28

Summary: “The simplified propagation model advanced here shows that the beam

radius as well as its deflection are severely affected by the phenomena of thermal

blooming. However, the model also showed that appropriate phase compensation at

the transmitter output aperture can mitigate these effects and return the

propagation situation to one that is acceptable for the power transfer requirements

that are needed to be satisfied for beamed propulsion.” “...nothing in the

prevailing physics that would preclude power beaming [at these

power levels] through the atmosphere.”

Analysis reviewed by :

J E. Lowder, Chief Scientist of Cobham Analytic Solutions, formerly 

known as SPARTA Inc. Missile Defense Sector. (Recommendation 

from Denise Podolski NASA OCT) “I have read the paper and was 

impressed with the depth of the theoretical analysis. The results 

look quite reasonable as far as I can tell with out running some 

numbers myself.”

Daniel v Murphy; MIT Lincoln Lab Directed Energy Group: We

would have to spend much more time to reach an independent

conclusion. But, work in the late 80's on the GBFEL TIE program

showed that thermal blooming was not a show stopper for missile

defense, with system parameters that are not far from what you

envisage for power beaming.

Energy costs for launch are low

• Maximum wallplug electricity needed for launch is 900 MW for the 

Boost station and 1.8GW for the Sustain station

• Boost station on for 100 secs, Sustain station on for 200 secs.

• Average Commercial cost for electricity is $0.10 per kW-hr (U.S Energy 

Information Administration 2009)

• For 900MW @ 100 secs; 900,000kW X (100/3600 = 0.0278) = 25000 

kW-hrs

• 25000 kW-hrs X 0.10 $/kW-hr = $2,500

• For 1.8GW @ 200 secs = $10,000

• Total electricity cost for launch = $12,500

• Similar in electrical demand to NASA GRC’s Wind 

Tunnels

Atmospheric effects do not preclude power beaming

Report Released: Effects of Beaming Energy Through the 

Atmosphere by Robert Manning, NASA GRC; Investigated and 

analyzed the effects of thermal blooming on infrared and 

millimeter wave propagation through the atmosphere for power 

beaming applications at the megawatt level. A. Ionization and Electrical Breakdown

B. Induced Molecular Polarization 

– The Kerr Effect

C. Induced Heating of the Atmosphere 

– Thermal Non-Linearities and “Thermal Blooming”

D. Atmospheric Aerosols

E. Atmospheric Turbulence

Pointing requirements are challenging

If a 1.5 meter diameter engine is used, one would require a 60 meter 

diameter transmitting aperture focused at the GEO distance of 32,500 

km. For a 5% spatial variation of the beam at this location (i.e., a 0.075 

meter beam wander from its intended axis), one requires a 

0.075/(32,500*10^(3))=2.3*10(-9) radian accuracy. For an initial 

target position in LEO at 400 km, this accuracy backs-off to 

0.075/(400*10^(3))=188*10^(-9) radians. 

Therefore, given the assumptions made:

The accuracy in going from LEO to GEO essentially goes 

as 188 nanoradians down to 2.3 nanoradians. 

The Hubble Space Telescope has a proven accuracy of 

50 nanoradians

Answers to concerns from mid-term review:

Atmospheric effects, energy costs, & pointing requirements

Final Report
NASA/TM—2012-217014 663 

29

Present and future Gyrotron & laser costs

Future VCSEL 

laser diodes

Current costs

• Gyrotron projections based on a 1.2 

MW unit sold for $2M in 2010 

(Fedbizopps contract award number NNA10DF73P)

• Fiber laser projections based on 1 

kW unit sold for $150k in 2003

Impact of future technology

• CW laser diode cost represents the 

lower limit of fiber laser cost, and 

could be lowered with the 

development of VCSEL diodes

• Gyrotron costs could be lowered by 

the development of ‘super power’ 

tubes which is currently underway

Impact of volume production

• LLNL Mercury Laser diodes 

demonstrated a learning curve of 

60% from 1994-2001 (~20 

cumulative MW)

• Magnetrons demonstrated 

comparable learning curve from 

1945-present (~200,000 cumulative 

MW)

Final Report

30

Source: “Solid‐State Laser Weapon Systems, Bridging the

Gap — or Bridge Too Far?”, by Andrew Krepinevich, Tom

Ehrhard, and Barry Watts, Center for Strategic & Budgetary

Assessments (CSBA), May 20, 2009.

Source: Kevin Parkin NASA ARC 

Exponential growth of continuous working solid 

state lasers and Gyrotrons

Final Report
NASA/TM—2012-217014 664 

Review outline

Mid‐term direction:

2. Modify the in‐space missions

(DRMs 2 and 3) to include in‐

space power generation such as

a solar array powering an

orbiting laser beaming station.

(pgs. 25  & 33)

3. Drop any further development of

the Space GEO Servicing mission

(DRM 2‐B) and the Horizontal

Launch mission (DRM4)

• Executive summary

• Original objectives / direction from mid-term review

• Assessment of feasibility, potentials, challenges

• Recommended developments

• Summary

• Details of findings

• LEO to GEO

• Deep space

• Launch

• A path forward

• Technology investments

• BEP development program

• Summary

Final Report 31

Deep Space mission overview

DRM 3 BEP booster showing major vehicle systems.

Laser BEPS orbit raising and Jupiter 

Transfer Transition Insertion 

Booster and payload launch 

separately and dock on orbit

Booster and payload separate: 

payload coasts to Jupiter

Laser BEPS allows direct Jupiter 

transfer orbit for Galileo-like spacecraft

Booster and payload 

configuration for operational 

stages of DRM 3 

100 minute 

Jupiter Transfer 

Trajectory 

Insertion burn 

(red) from earth 

orbit (green) to 

Jupiter direct 

trajectory 

(white). Jupiter 

direct trajectory 

available every 

400 days

Final Report 32
NASA/TM—2012-217014 665 

National Aeronautics and Space Administration

www.nasa.gov

GEO

• Jupiter – 2.5 yrs total transit time vs. 

6 yrs with chemical system 

Vehicle 

• Advanced Sterling Radioisotope Generators 

• 5000 N thermal plasma engine

• H2O steam propellant

• 6000 K, 50 psia chamber press, 813 s Isp

• 550 kg dry mass, 4900 kg propellant

• 118 kg of instruments Deep Space Configuration 

(90% Similar to LEO to GEO)

Deep Space Mission

Transport to Outer Planets

Ground Based Beaming Station

Beaming Station Located on Ground

• Single 78 MW emitter, 200 MW input

• 30m dia. Dish

• Located in low moisture environment 

close to Equator - Hawaii 

MISSION TRANSIT TIME TO JUPITER CUT IN HALF; 

LAUNCH OPPORTUNITIES BECOME MUCH MORE FREQUENT, MORE PAYLOAD MASS

Final Report 33

National Aeronautics and Space Administration

www.nasa.gov

Beaming Station Located in GEO

• Single 50 MW emitter, 150 MW input 

required 

• Equivalent to 100 ISS sized spacecraft 

each with 75 “FAST” arrays 

• Ground controlled

• Radiate 50MW waste heat

GEO

• Jupiter – 2.5 yrs total transit time vs. 

6 yrs with chemical system 

Vehicle 

• Advanced Sterling Radioisotope Generators 

• 5000 N thermal plasma engine

• H2O steam propellant

• 6000 K, 50 psia chamber press, 813 s Isp

• 550 kg dry mass, 4900 kg propellant

• 118 kg of instruments

Deep Space Configuration 

(90% Similar to LEO to GEO)

Deep Space Mission

Transport to Outer Planets

Space Based Beaming Station

MISSION TRANSIT TIME TO JUPITER CUT IN HALF; 

LAUNCH OPPORTUNITIES BECOME MUCH MORE FREQUENT, MORE PAYLOAD MASS

Final Report 34
NASA/TM—2012-217014 666 

Deep Space mission evaluation:

Potential and challenges

Potentials

• Faster transit to outer planets (nearly 2X)

• Environmentally friendly water based propellant and 

simplified Fuel Depot requirements

• Deep Space Transfer Stage Vehicle Synergistic with LEO to 

GEO

• Synergy with Manned GEO Servicing mission

• 118 kg of instrument payload vs. 82 kg (Enceladus Mission 

Study) 

• Can Be Time Phased With In-space BEP Development To 

Provide Optimum Program Development Scenario

Challenges

• High cost to establish beaming station in GEO or on ground

• Required technologies at low TRL

• Plasma engine (TRL 2)

• Reflective mirror: 4m x 6m; 99.99% reflectivity, micrometeoroid tolerant 

(TRL 2)

• Multi MW space/ground operational laser (TRL 2-3)

• 2.3 nanoradians pointing and tracking (TRL 2-3). Best proven is 

Hubble Space Telescope 50 nr. Webb Telescope designed for 24 nr 

• High cost operation of GEO based remote complex, high-energy 

beaming station . Not pointing to Earth

• Beaming 80 MW from ground through atmosphere. Max current is 

Airborne Laser Test Bed MW class

DRM 3 BEP booster showing major vehicle 

systems.

Final Report 35

Summary of study findings:

Deep Space mission

* Estimated based on LEO to GEO cost analysis

** Based on Enceladus mission study by GSFC

*** Nuclear system would be a better technology

Prometheus 200 kW $3B

^ Based on Launch mission analysis

Mid-Term Direction

1. Do not consider the effects of launch rates for this study

2. Do not consider the amortization of non-recurring vehicle and ground system costs

3. Develop ROM non-recurring costs for vehicles and ground systems (pgs. 26, 36 & 50)

4. Develop ROM vehicle and ground facility costs for recurring launches 

Space vehicle cost assumptions:

• Prime contractor costs (no fee)

• No technology development costs

• Protoflight development approach

• Before reserves, insight/oversight, launch services

Final Report 36
NASA/TM—2012-217014 667 

Synergistic space mission benefits

• Additional missions enabled by in-space 

beaming facility:

• Debris mitigation

• Laser communications

• Space protection

• Beaming to Earth

• Disaster relief

• Powering highflying UAV's

• Additional applications of power beaming:

• Additional power to old or damaged satellites 

requiring a short-term boost to complete, extend, 

or recover their mission

• Manned GEO servicing missions

Final Report 37

Unexpected findings

Remote spacecraft positional control 

Final Report 38
NASA/TM—2012-217014 668 

RCS solar engine

Remote spacecraft positional control concept:

Some components have been tested

• Notional concept based on synergistic 

previous propulsion, power, and 

communications technology developments:

• Solar propulsion:

• ISUS, HiDVE, optical cavity, inflatable concentrator, 

secondary lens, integrated RCS

• Can utilize diverse new propellant types (water, 

ammonia, etc.)

• Solar power:

• High intensity, high voltage solar cell

• Communications:

• Laser

• Can utilize solar, laser, or mm wave source

GRC/MSFC optical cavity 

and secondary lens

Hi Voltage/Hi 

intensity laser 

for power and 

communications

Integrated Power, Communications, Propulsion 

Final Report 39

Remote spacecraft positional control concept

systems approach

Integrate propulsion, power, 

communications, and 

navigational/attitude control on the 

same laser beam

• Remote operation – coordinated maneuver 

with source and receiving spacecraft 

without mechanical connection

• Reduce complexity, mass and cost of the 

receiving spacecraft

• Applicable for variety of space and aero 

vehicles with different propulsion systems

• Possible to control receiving “Shepherd” 

spacecraft from “Mother” spacecraft to 

perform specific maneuvers - similar to 

berthing

40

If successfully developed, could lead to a new class of spacecraft 

with enhanced operational capability, and lower mass.

Final Report
NASA/TM—2012-217014 669 

Review outline

Mid‐term direction:

1. Focus the Millimeter wave thermal rocket

(DRM 1‐C') as an expendable design making

an easier direct comparison with the laser

design (DRM 1‐B)..

10. Emphasize concepts utilizing and extending

from high TRL commercial lasers like diode‐

pumped lasers. De‐scope concepts that

require low TRL lasers.

• Executive summary

• Original objectives / direction from mid-term review

• Assessment of feasibility, potentials, challenges

• Recommended developments

• Summary

• Details of findings

• LEO-GEO

• Deep space

• Launch

• A path forward

• Technology investments

• BEP development program

• Summary

Final Report 42
NASA/TM—2012-217014 670 

Launch mission technology envelope

All three are founded on concepts from 

advocate community, which we analyzed 

and further developed – note: variations 

could be studied endlessly

• Laser powered

• Thermal engine

• LH2 & H2O propellant

• 10-100 kg Payload

• Launch mass 5337 kg

• Microwave powered

• Thermal engine

• LH2 & solid propellant

• 10 – 100 kg payload

• Launch mass 2153 kg

• Continuous/pulsed laser powered

• Optical plasma engine

• Air & H2O propellant

• 1-100 kg payload (e.g. 6 CubeSats

plus support h/w is 40 kg)

• Total mass launched 334 kg

Final Report 43

44

Thermal 

Microwave

FOM

VALUE

Payload mass 40kg

Wet mass 2153 kg

ISP 760 s

Thrust/mass 13

Max launch 

Power (output)

Millimeter 

Wave 600 

MW

Stage to orbit 2

Propellant LH2/Solid

Optical 

Laser

FOM

VALUE

Payload mass 40kg

Wet mass 332kg

ISP 985 s

Thrust/mass 233

Max launch 

power (output)

Laser 350 MW 

Cont./1 GW 

Pulsed 

Stage to orbit 1

Propellant Air/H2O

Launch vehicle comparison

(pictorial representations to scale)

Thermal 

Laser

FOM

VALUE

Payload mass 80kg

Wet mass 5415 kg

ISP 785 s

Thrust/mass 29

Max launch 

Power (output)

Laser 240 

MW

Stage to orbit 1

Propellant LH2/H2O

Pegasus XL: 

• Payload 420 kg 

• Air launch at 40,000 ft

• Reduced range safety

• Consistent weather

• Three stage to orbit

16.6m

Final Report
NASA/TM—2012-217014 671 

45

20 km 400 

km

750 MW Station 420 MW Station

Launch Control

Laser 

optical

Millimeter 

wave 

thermal

Laser 

thermal

100 km

900 MW Gyrotron station 1.8 GW Gyrotron station

120m dia. 

antenna

active surface 

green bank 

class

45m dia.

antenna 

15 km

420 MW continuous or 

1.0 GW pulsed station

400 km

3 m 

diameter 

Telescope

140 MW output

1150 to 2900 x 

120 kW modules; 

diode-pumped 

solid state lasers 

(repped pulse)

Ground based

600 X 1 MW 

Gyrotrons

incoherently 

combined 

Ground based

140 MW output laser power

• 2300 X 60 kW beam modules

• 6 x 10 kW fiber lasers (COTS)

• Shared aperture (not phase 

locked)

• 1.2 m telescope, adaptive optic

5 urad divergence

Ground based

240 MW output laser power

• 1200 X 200kW modules 

• Incoherent laser diodes 

arrays. 60% efficient; $2/watt 

• “Celestron-class” optics

• 80 urad beam divergence

Ground Based

300 X 1 MW 

Gyrotrons

Incoherently 

Combined 

Ground Based

350 MW Output

3000 to 7250 x 120 kW 

modules; Diode-Pumped 

Solid State Lasers 

1 GW Continuous or 

2.4 GW Pulsed Station

Relay satellite 

station

3.5m x 5m

“Monocle”

Mirror in 400 km orbit. 

OR

Launch ground facilities comparison

Final Report

46

Launch application ROM costs (FY11$)

Vehicle

NonRecurring 

Cost ($M)

Recurring

Cost (First 

Unit) $M

Commercial

(First Unit)

$M

Laser Optical

DRM 1-A 63 15 5

Laser Thermal 

Rocket DRM 1-B 126 35 13*

MM Wave Thermal

Rocket DRM 1-C’ 110 28 11*

Vehicle

Ground Facility

Non-Recurring 

Cost ($B)

Ground Operations 

Yearly Cost

360 

Launches/yr($M)

Laser Optical

DRM 1-A 5 240

Laser Thermal

Rocket DRM 1-B 6.7 330

MM Wave Thermal

Rocket DRM 1-C 5.7 285

Mid-term direction - costs

4. Do not consider the effects of 

launch rates for this study

5. Do not consider the amortization 

of non-recurring vehicle and 

ground system costs

6. Develop ROM non-recurring costs 

for vehicles and ground systems

7. Develop ROM vehicle and ground 

facility costs for recurring 

launches 

Number of launch vehicles produced

The differences between the curve sets indicates the 

opportunity to reduce the per unit cost of going from 

NASA “government” to commercial requirements and 

production methods.

* Extrapolated

NASA vs. Commercial Costs ($M)

Final Report
NASA/TM—2012-217014 672 

• Option to explore the use of the Lightcraft concept to 

launch Cubesats

• Ground ruled 

• keeping the same trajectory/power levels stations – kept 

everything needed for launch and insertion

• Added a dual pea-pod cube sat launcher – launches six (1 

kg) cubesats once on-orbit

• P-Pod device seems to fit

• Issues of high spin rate will require redesign of P-Pod and 

cubesats

47

Laser lightcraft derivatives

DRM description

• Nano-sat launch capability to LEO (400 km)

• Multiple launches per day (1.33 kg payload)

• Ground-based power-beaming facility; launch from high 

desert (e.g., HELSTF, NM) or 3 km mtn.

Power source description

• 12 MW pulsed 1.053 μm solid state diode-pumped laser 

design with multiple 120 kW beamlets, each pulsing at 60 

Hz.

• 96 beam-lines, 2 kJ @ 5760 Hz, 50-100 ns pulses

• >10% wall-plug efficiency with long pulse

• Beam combining w/ rotating mirror (3600 rpm) “Gatling 

gun” thru 3 m diameter telescope

Vehicle description

• Separable, de-spun 1.33 kg CubeSat

• 35 cm diameter vehicle; 8 kg launch mass (wet)

• Combined-cycle air-breathing / rocket engine

• Water used as rocket propellant

Technical challenges

• High power/ reliability pulsed laser exploiting existing 100-

150 kW solid state, diode-pumped laser technology (R)

• Rotating mirror beam combining system w/ targeting 

accuracy of 2 nrad (R)

• High reflectivity (~99.99%) engine optics in hot plasma 

environment (R)

• Retro-directive control for pointing (R)

(R) = Technology Required for Success

35 cm laser lightcraft 1.2 m laser lightcraft w/pod 

Final Report

48

Safety zone comparison

Burnout (400km altitude, 

1200km down range)

BEP range safety corridor

(4km width)

Sustainer site 

(400km) Launch site

Booster site 

(110km)

400 km

4 km 

~285km

Total area:

~1,600 km2

Total area:

~910,000 km2 64km

~90

32km

*Not to scale

SRB

ET

Impact 

Zone

KSC

Restr. 

Area

2100km

400km

Launch site

SRB impact zone 

(~250km)

ET impact zone 

(~11000km) *Not to scale

*Not to scale

Final Report
NASA/TM—2012-217014 673 

49

Launch mission evaluation:

Potentials and challenges

Technical risk and difficulties increase for Earth to LEO

• Potentials

• Strong synergy with in-space thermal plasma engine 

propulsion

• Lightcraft mass and vehicle technology

• Study identified technology needs and directions to 

allow development

• Can be time phased with in-space BEP development 

to provide optimum program development scenario

• Challenges

• Significant facility investment requiring large 

amortized costs

• Higher than expected amortized vehicle R&D costs

• Higher recurring vehicle costs based on current 

launch vehicle experience

• Atmospheric and weapons and sensitivity

Shown to scale

16.6m

Final Report

Summary of Study Findings

Launch Missions

* 1.3 kg CubeSat payload appears to scale favorably. Needs further investigation.

Launch vehicle cost assumptions:

• Prime contractor costs (no fee)

• No technology development costs

• Protoflight development approach

• Before reserves, insight/oversight, launch services

Mid-term direction:

1. Do not consider the effects of launch rates for this study

2. Do not consider the amortization of non-recurring vehicle and ground system costs

3. Develop ROM non-recurring costs for vehicles and ground systems (pgs. 26, 36 & 50)

4. Develop ROM vehicle and ground facility costs for recurring launches 

Final Report 50
NASA/TM—2012-217014 674 

Unexpected findings

Beamed Energy Propulsion engine

Final Report 51

14/04/2011 NASA/DARPA 52 Pre‐Decisional – for BEPS Team Use Only

Optical plasma engine operational concept
NASA/TM—2012-217014 675 

53

Introduction of optical thermal plasma engine

• BEP study found significant potential performance

increase from thermal plasma engine

• Utilization of not-traditional dense propellants

• Optical thermal plasma engine:

• Synergistic with previous solar thermal propulsion 

technology:

• Inflatable concentrator, optical cavity engine, 

secondary lens, integrated RCS

• Increased performance due to elimination of heat 

exchanger

• Potential relaxation of optical pointing requirements 

using primary and secondary optics

Isp T/W

Heat

Exchanger 785,760 12.9,29

Plasma 985,813 233,1

Final Report

54

Beamed Energy Propulsion engine components

• Optics

• Utilize primary mirror and secondary lens to reduce surface and pointing 

requirements and enable use of inflatable:

• Significant pointing relaxation compared with FAST concentrator 

• Use of elastomeric adaptive optics has significant potential to improve primary 

capability with inflatable:

• Synergy with ISRU

• Engine

• Plasma stabilization and control of critical importance to prevent thermal 

degradation of mirror:

• Potential control methods include propellant flow and electro-magnetic plasma manipulation

• System

• Potential utilization of steering mirrors to reduce space or ground infrastructure:

• Synergistic with development of primary optic and utilization of elastomeric/adaptive optics

Final Report
NASA/TM—2012-217014 676 

55

Game changing technology: Engine propulsion

Unproven technology (TRL 1-3) Demonstrated technology (TRL 4-6) Operational systems (TRL 7-9)

10-4

102

103

104

Vehicle acceleration or T/W ratio (g’s)

Specific impulse (seconds)

10-5 10-3 10-2 10-1 1 10 102

SAILS

106

105

107

CHEMICAL ROCKETS

ELECTROSTATIC

ELECTROTHERMAL

ELECTROMAGNETIC

PULSED 

FISSION

THERMAL

FISSION

SOLAR 

THERMAL * * * LASER/SOLAR 

THERMAL *

Isp T/W

Heat

Exchanger

Plasma

785,760 12.9,29

985,813 233,1

Final Report
NASA/TM—2012-217014 677 

Review outline

Mid-term direction:

11. The primary product of this 

study needs to be a 

determination of TRL’s for near 

term concepts of:

A. Beam energy technology 

(BET)

B. Rocket vehicle technology 

(RVT)

57

• Executive summary

• Original objectives / direction from mid-term review

• Assessment of feasibility, potentials, challenges

• Recommended developments

• Summary

• Details of findings

• LEO-GEO

• Deep space

• Launch

• A path forward

• Technology investments

• BEP development program

• Summary

Final Report

58

Prioritized list of technology investments:

Enabling & enhancing 

Engine/Beam 

Technology

Technology Current TRL

(S of A)

Comments

Estimated 

Time to 

Achieve 

TRL 6

Enabling

BT

Satellite integrated systems to 

allow remote operation via laser 

beam

2 Systems 

development 2 yrs

BT Laser power beam combining 

communications

2

Laser comm.

currently at 

low power 

levels

3 yrs

ET Engine and controls 3

Concept 

engine 

produced, 

limited testing 

3 yrs

Enhancing

ET Optical engine 3 Some testing 2 yrs

ET Highly reflective mirror with ability 

to withstand micrometeoroid debris 1 Basic research 4 yrs

ET Large inflatable mirror structure 2 Some testing 3 yrs

ET Adaptive elastomeric optic 2 Basic research 3 yrs

Final Report
NASA/TM—2012-217014 678 

High power laser/microwave 

development 

BEPS

NASA/Air Force

20 years

Near term

• In-space focused

• Synergistic with solar 

thermal propulsion

Develop energy beaming transfer protocol/standards

Technology development

BEPS

NASA/DARPA

Demonstrations leading to applications

Mid term

• Fractionated satellites

• Satellite constellations

Far term

• Launch ETO

• Synergistic with in-space

Industry

Applications

Low power 

applications

High power 

applications Space based

SPACE Mission

Ground 

based

Orbital Low power 

applications Sub-orbital Ground 

based

LAUNCH Missions

59

Potential beamed energy development program

DoD (Joint)

Air Force

Navy

Final Report

60

Overall study summary

• The study concluded using beamed energy for launching vehicles into space is technically feasible 

assuming commitment to developing new technologies and large investments can be made over 

long periods of time. These include specific technologies such as high power lasers and microwave 

sources as well as launch facility infrastructure. However, from a competitive standpoint, a 

potential advantage for beamed energy for ETO rests with smaller, frequently launched payloads to 

achieve economic viability. The major issue remains high infrastructure costs

• The study concluded using beamed energy for space propulsion is definitely feasible and showed 

distinct advantages and greater potential over current propulsion technologies. However, this 

conclusion assumes mission commitments and investments in critical technologies are made over 

long periods of time. BEP energy source requirements are much less than for ETO. Lower 

propulsion costs with shorter transit times for LEO to GEO servicing missions and faster science 

missions to the outer planets are the major benefits. The major issue remains high infrastructure 

costs.

• The study findings indicated two areas of great interest, potential and worthy of consideration for 

investment. Beamed energy could provide the key to combining power, propulsion and 

communications thus creating a new class of small spacecraft with minimal systems, reducing the 

complexity, cost and mass. This novel capability may allow spacecraft off-loading of some systems 

previously thought of as a necessity. Beamed energy to a thermal optical plasma engine could 

provide a highly efficient technology for high thrust and high Isp space propulsion. 

Recommend proceeding with concurrent integrated laser, 

communications and propulsion flight demonstration and the thermal 

plasma engine technology development project.

Final Report
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14. ABSTRACT

The scope of this study was to (1) review and analyze the state-of-art in beamed-energy propulsion (BEP) by identifying potential game-changing applications, (2) formulate a 

roadmap of technology development, and (3) identify key near-term technology demonstrations to rapidly advance elements of BEP technology to Technology Readiness 

Level (TRL) 6. The two major areas of interest were launching payloads and space propulsion. More generally, the study was requested and structured to address basic mission 

feasibility. The attraction of beamed-energy propulsion (BEP) is the potential for high specific impulse while removing the power-generation mass. The rapid advancements in 

high-energy beamed-power systems and optics over the past 20 years warranted a fresh look at the technology. For launching payloads, the study concluded that using BEP to 

propel vehicles into space is technically feasible if a commitment to develop new technologies and large investments can be made over long periods of time. From a 

commercial competitive standpoint, if an advantage of beamed energy for Earth-to-orbit (ETO) is to be found, it will rest with smaller, frequently launched payloads. For space 

propulsion, the study concluded that using beamed energy to propel vehicles from low Earth orbit to geosynchronous Earth orbit (LEO-GEO) and into deep space is definitely 

feasible and showed distinct advantages and greater potential over current propulsion technologies. However, this conclusion also assumes that upfront infrastructure 

investments and commitments to critical technologies will be made over long periods of time. The chief issue, similar to that for payloads, is high infrastructure costs.

15. SUBJECT TERMS

Beamed energy; Laser propulsion; Laser launch 

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