John Morgenstern, Nicole Norstrud, Marc Stelmack, and Craig Skoch

Lockheed Martin Aeronautics Company, Palmdale, California

Final Report for the Advanced Concept Studies for 

Supersonic Commercial Transports Entering Service 

in the 2030 to 2035 Period, N+3 Supersonic Program

NASA/CR—2010-216796

August 2010

PMF–01623
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John Morgenstern, Nicole Norstrud, Marc Stelmack, and Craig Skoch

Lockheed Martin Aeronautics Company, Palmdale, California

Final Report for the Advanced Concept Studies for 

Supersonic Commercial Transports Entering Service 

in the 2030 to 2035 Period, N+3 Supersonic Program

NASA/CR—2010-216796

August 2010

PMF–01623

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared under Contract NNC08CA84C
Acknowledgments

This work was funded by the NASA N+3 Supersonic program administered by Peter Coen at the NASA Langley Research Center 

under prime contract (Lockheed Martin (LM)) NNC08CA84C. LM would like to thank Clayton Meyers at the NASA Glenn 

Research Center and many others at NASA for their helpful contributions, editorial support and encouragement on this program. 

LM would also like to acknowledge GE Global Research Center, Wyle Laboratories and Vic Sparrow, Dennis McLaughlin and 

Philip Morris from Penn State, LM Transportation and Security Solutions, Helen Reed and Bill Saric from Texas A&M University, 

John Hansman from MIT, and Purdue University for contributing written material, data, and ideas for the N+3 fi nal report. 

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://gltrs.grc.nasa.gov

Trade names and trademarks are used in this report for identifi cation 

only. Their usage does not constitute an offi cial endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

Level of Review: This material has been technically reviewed by NASA technical management.

This report contains preliminary fi ndings, 

subject to revision as analysis proceeds.

This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.
NASA/CR—2010-216796 iii

Executive Summary

Lockheed Martin Aeronautics Company (LM Aero), working in conjunction with seven industry and 

academia sub-contracting teammates, executed an 18 month program responsive to the NASA sponsored 

“N+3 NRA Advanced Concept Studies for Supersonic Commercial Transports Entering Service in the 

2030 to 2035 Period” contract. The key technical objective of this effort was to generate promising 

supersonic concepts for the 2030 to 2035 timeframe and to develop plans for maturing the technologies 

required to make those concepts a reality. The N+3 program is aligned with NASA’s Supersonic Project 

and is focused on providing alternative system-level solutions capable of overcoming the efficiency, 

environmental, and performance barriers to practical supersonic flight. 

In addition to FAA regulations, the Next Generation (NextGen) Air Traffic System (ATS) congestion 

levels are a concern as they are expected to increase by a factor of 1.5 to 2.5 in the 2025 timeframe. 

Understanding how supersonic aircraft affect future congestion levels requires a system of systems 

analysis that integrates vehicle design, operating environment, and economic interaction into a single 

process. LM Aero worked with a sister company, Transportation Security and Solutions (TSS), and 

Purdue University to assess the value that a supersonic transport concept vehicle brings to the NextGen 

ATS. A fast time modeling and simulation study done by TSS revealed that commercial supersonic 

vehicles will not impact future airport capacity. However, supersonic air vehicles in the 2030 timeframe 

will exert additional demand for airport operations. Purdue University simulated numerous future Civil 

Air Transport System scenarios, allocating N+3 vehicles to maximize system-wide productivity while 

also computing fleet-wide emissions and direct operating costs. These results showed that the total value 

of time saved by passengers on N+3 supersonic transports will likely exceed the added operating costs 

incurred by the aircraft. These system-level scenarios showed that supersonic transport is a viable solution 

for increased productivity and promotes the renewed viability of supersonic travel. 

Our extensive team designed a preferred supersonic configuration and developed plans for maturing 

the identified, enabling technologies required to meet the N+3 performance and environmental goals.

Working in conjunction with GE Global Research Center (GRC), John Hansman from MIT, Helen Reed 

and Bill Saric from Texas A&M, Wyle Laboratories, Purdue, and Penn State—an initial low-boom, 

supersonic configuration was used to assess potential airframe and propulsion technologies that were 

projected to meet or exceed the future supersonic boom, noise, emissions, cruise speed, range, payload, 

and fuel efficiency goals. Multi-Disciplinary Analysis and Optimization (MDAO) showed it was possible 

to achieve the N+3 boom goal with an “inverted-V”, engine-under wing configuration. Further sizing and 

quantified analysis proved that using revolutionary technologies enabled this configuration to achieve the 

range, payload, and cruise speed goals.

Based on LM provided requirements and targets, GE developed an Variable Cycle Engine (VCE) 

propulsion system and a conventional Mixed Flow Turbo Fan (MFTF) propulsion system expected to 

meet or exceed the environmental goals set by NASA, as well as an MFTF optimized solely for cruise 

efficiency. These propulsion systems take advantage of an Advanced Thermal Management System 

(ATMS) to extend the overall pressure ratio (OPR) of the engine and increase thermal efficiency. A low 

noise, high performance exhaust system takes advantage of the innovative jet noise reduction features that 

work synergistically with the VCE features to reduce the exhaust jet noise. Augmented transonic thrust 

allows the propulsion system to be favorably sized with potential take-off noise abatement. Analysis 

shows that this propulsion system, along with integrated technology sets, meets the N+3 airport noise, 

emissions, and fuel efficiency goals.

Our integrated airframe and propulsion system, along with identified/enabling technologies, is 

projected to meet or exceed all N+3 targets. Results of the environmental and performance characteristics 

of our advanced vehicle concept are summarized in Table 1. 

Through a collaboration effort, LM Aero and GE GRC identified N+1, N+2, and N+3 technologies 

critical to meet or surpass the N+3 goals. N+1 and N+2 shaping technologies were considered to be 

“endemic” or inherent to the baseline design. These configuration technologies were not included in the 

final technology roadmap, but other N+2 technologies were included to provide a comprehensive 
NASA/CR—2010-216796 iv

technology list. As a result, technology roadmaps were created for all prioritized, airframe technologies to 

demonstrate the maturation efforts required to raise each technology to a Technology Readiness Level 6 

(TRL 6). 

TABLE 1.—LM’S PREFERRED CONCEPT WITH TECHNOLOGY INPUTS MEETS OR SURPASSES ALL N+3 GOALS

NASA N+3 Efficient Multi-Mach Aircraft 

(Beyond 2030)

N+3 Goal Status

Environmental Goals

Sonic Boom 65 to 70 PLdB low boom flight

75 to 80 PLdB unrestricted flight

70 to 76 PLdB

KEY GOAL

Airport Noise 20 to 30 EPNdB

(cumulative below stage 3)

18.4

(32.2 jet only)

KEY GOAL

Cruise Emissions (g/kg fuel) <5 EINOx

Plus particular and water vapor mitigation 5 EINOx

Performance Goals

Cruise Speed Mach 1.3 to 2.0 low boom flight

Mach 1.3 to 2.0 unrestricted Mach 1.6

Range 4000 to 5500 nmi 4850 nmi

Payload 100 to 200 pax 100 pax

Fuel Efficiency 3.5 to 4.5

(pax-nmi/lb-fuel)

3.64

(pax-nmi/lb-fuel)

KEY GOAL

Recommended future work includes Phase 2 testing and Phase 3 maturation efforts to provide a 

technology set necessary to realize a vision vehicle serviceable in the 2030-2035 timeframe. 

Current N+2 efforts allow us to reasonably assume that N+2 technologies will be developed during 

those N+2 program efforts, and the developed technologies will be available for application on the N+3 

vehicle. Concentration on N+3 technologies provides a clear roadmap to achieving and surpassing the 

stated N+3 goals while providing an exciting solution to supersonic travel. Figure 1 highlights the 

comprehensive technology set for both airframe and propulsion systems.

Future work recommendations for airframe technologies include:

• Low cost, high impact tools and methodologies such as Low Boom Shaping Fidelity and CFDbased MDAO to address boom mitigation

• Distributed roughness with plasma augmentation to ensure laminar flow at supersonic conditions

• Adaptive geometry technologies including lift distribution control and inlet flow control 

technologies to address the N+3 fuel efficiency goals

Future work recommendations for propulsion technologies include:

• Continued development of VCE technologies

• Development of the Transonic Thrust Augmentation device—critical technology to meeting N+3 

goals 

• Alternate combustor/combustion concepts need to be explored, as these propulsion systems are 

developed to take advantage of the full thermal capability of the system. The currently funded 

NASA Supersonics Low Emissions combustor program will provide key validation data for high 

temperature NOx levels, and the maturation of some enabling technologies

• Predictive and design tool development in many areas need to be continued to be developed 

including:

– Aero-acoustics for both fan and jet noise 

– Combustion and emissions 
NASA/CR—2010-216796 v

Figure 1.—Technology roadmap for N+3 enabling technologies.

NASA/CR—2010-216796 vii

Contents

Executive Summary.....................................................................................................................................iii

1.0 Introduction ..........................................................................................................................................1

1.1 Subject of the Report ..................................................................................................................1

1.2 Purpose........................................................................................................................................2

1.3 Scope...........................................................................................................................................2

2.0 Work Breakdown Structure (WBS) .....................................................................................................3

3.0 Tasks and Trade Studies—Airframe Systems......................................................................................4

3.1 Advanced Vehicle Concept (WBS 3.1) ......................................................................................4

3.1.1 Concept Layout and Design (WBS 3.1.1)..................................................................... 4

3.1.2 Alternative Configurations (WBS 3.1.1)....................................................................... 8

3.2 Design Space Trade Studies With RCD (WBS 3.2).................................................................12

3.2.1 RCD Low Boom Design Approach (WBS 3.2.1)....................................................... 12

3.2.2 RCD Model Development (WBS 3.2.1) ..................................................................... 13

3.2.3 RCD Model Integration with Technology Inputs (WBS 3.2.2).................................. 17

3.2.4 Technology Benefit/Impact Assessment (WBS 3.2.4)................................................ 18

3.2.5 Quantified Analysis (WBS 3.2.3) ............................................................................... 23

3.3 Final Configuration Definition (WBS 3.1.3) ............................................................................34

3.3.1 Loft.............................................................................................................................. 34

3.3.2 Fuel Arrangement ....................................................................................................... 35

3.3.3 Aerodynamics ............................................................................................................. 35

3.3.4 Weights....................................................................................................................... 37

3.4 System of System Analysis (WBS 3.6) ....................................................................................38

3.4.1 Purdue ......................................................................................................................... 38

3.4.2 Transportation and Security Solutions........................................................................ 40

3.4.3 Results......................................................................................................................... 46

3.4.4 Conclusions................................................................................................................. 60

4.0 Tasks and Trade Studies—Propulsion Systems.................................................................................60

4.1 Technical Description ...............................................................................................................60

4.1.1 VCE Propulsion System.............................................................................................. 60

4.1.2 Conventional Mixed Flow Turbofan (MFTF) Propulsion Systems............................ 61

4.1.3 ATMS.......................................................................................................................... 61

4.1.4 Advanced Low NOx Combustor................................................................................. 61

4.2 Exhaust System.........................................................................................................................61

4.2.1 VCE Exhaust............................................................................................................... 61

4.2.2 Conventional MFTF Exhausts .................................................................................... 61

4.2.3 Transonic Thrust Augmentation ................................................................................. 61

4.2.4 Exhaust Variable Geometry Features.......................................................................... 61

4.3 Airframe Requirements.............................................................................................................62

4.3.1 NASA N+3 Goals ....................................................................................................... 62

4.3.2 Thrust Requirements................................................................................................... 62

4.3.3 Customer Installation Effects...................................................................................... 62

4.4 GE VCE Propulsion System .....................................................................................................63

4.4.1 Propulsion System Design .......................................................................................... 63

4.4.2 Noise Assessment ....................................................................................................... 65

4.5 Conventional Mixed Flow Turbofan (MFTF) Propulsion System ...........................................68

4.5.1 Propulsion System Design .......................................................................................... 68

4.5.2 Noise Assessment ....................................................................................................... 68

4.6 “Optimum” Cruise Conventional MFTF Propulsion System ...................................................69

4.6.1 Propulsion System Design .......................................................................................... 69
NASA/CR—2010-216796 viii

4.7 Emissions status........................................................................................................................70

4.8 Efficiency (FoM) status ............................................................................................................71

4.9 Technology Trades and Studies................................................................................................71

4.9.1 Low-Emissions Combustor......................................................................................... 71

4.9.2 Intercooler................................................................................................................... 72

4.9.3 Interturbine Combustor............................................................................................... 74

4.9.4 Constant Volume Combustor (CVC).......................................................................... 74

5.0 Technology Roadmap Development—Airframe (WBS 3.3).............................................................76

6.0 Technology Identification (WBS 3.3.1).............................................................................................76

6.1 Technology Down Selection (WBS 3.3.2)................................................................................78

6.2 Technology Prioritization (WBS 3.3.4)....................................................................................81

7.0 Technology Roadmap Development—Propulsion.............................................................................83

7.1 Technology Identification—Potential Enabling Technologies.................................................83

7.2 Technology Down Selection/Prioritization...............................................................................83

8.0 GOTCHA Analysis (WBS 3.3.3).......................................................................................................83

9.0 New Technology—Airframe (WBS 3.3.3, 3.3.4) ..............................................................................86

9.1 Tools /Methodologies—Low Boom Shaping Fidelity..............................................................86

9.1.1 Addressed N+3 Goals ................................................................................................. 86

9.1.2 Description.................................................................................................................. 86

9.1.3 Roadmap ..................................................................................................................... 86

9.2 Tools/Methodologies—CFD-Based Adjoint Shape Optimization............................................87

9.2.1 Addressed N+3 Goals ................................................................................................. 87

9.2.2 Description.................................................................................................................. 87

9.2.3 Roadmap ..................................................................................................................... 87

9.3 Laminar Flow—Distributed Roughness with Plasma Augmentation.......................................87

9.3.1 Addressed N+3 Goals ................................................................................................. 87

9.3.2 Description.................................................................................................................. 87

9.3.3 Roadmap ..................................................................................................................... 88

9.4 Pilot Situational Awareness—Low Boom Violation Cueing....................................................89

9.4.1 Addressed N+3 Goals ................................................................................................. 89

9.4.2 Description.................................................................................................................. 89

9.4.3 Roadmap ..................................................................................................................... 89

9.5 Pilot Situational Awareness—Preplanning Awareness ............................................................89

9.5.1 Addressed N+3 Goals ................................................................................................. 89

9.5.2 Description.................................................................................................................. 89

9.5.3 Roadmap ..................................................................................................................... 90

9.6 Tools/Methods—Integrated Structural Analysis.......................................................................90

9.6.1 Addressed N+3 Goals ................................................................................................. 90

9.6.2 Description.................................................................................................................. 90

9.6.3 Roadmap ..................................................................................................................... 91

9.7 Adaptive Geometry—Lift Distribution Control .......................................................................91

9.7.1 Addressed N+3 Goals ................................................................................................. 91

9.7.2 Description.................................................................................................................. 91

9.7.3 Roadmap ..................................................................................................................... 91

9.8 Plasma—Jet Exhaust Manipulation ..........................................................................................92

9.8.1 Addressed N+3 Goals ................................................................................................. 92

9.8.2 Description.................................................................................................................. 92

9.8.3 Roadmap ..................................................................................................................... 92

9.9 Adaptive Geometry—Inlet Flow Control (Choking)................................................................93

9.9.1 Addressed N+3 Goals ................................................................................................. 93

9.9.2 Roadmap ..................................................................................................................... 93
NASA/CR—2010-216796 ix

9.10 Technology Roadmap Overview ..............................................................................................94

10.0 New Technology—Propulsion...........................................................................................................94

10.1 VCE ..........................................................................................................................................94

10.1.1 Goals and Objectives................................................................................................... 94

10.1.2 Technical Description ................................................................................................. 95

10.2 Low Noise/High Performance Nozzle Exhaust ........................................................................95

10.2.1 Goals and Objectives................................................................................................... 95

10.2.2 Technical Description ................................................................................................. 95

10.2.3 Roadmap ..................................................................................................................... 95

10.3 Advanced Low NOx Combustor...............................................................................................96

10.3.1 Goals and Objectives................................................................................................... 96

10.3.2 Technical Description ................................................................................................. 96

10.3.3 Roadmap ..................................................................................................................... 96

10.4 ATMS .......................................................................................................................................97

10.4.1 Goals and Objectives................................................................................................... 97

10.4.2 Technical Description ................................................................................................. 97

10.4.3 Roadmap ..................................................................................................................... 97

10.5 Transonic Thrust Augmentation Device...................................................................................97

10.5.1 Goals and Objectives................................................................................................... 97

10.5.2 Technical Description ................................................................................................. 98

10.5.3 Roadmap ..................................................................................................................... 98

10.6 CVC Roadmap..........................................................................................................................98

10.6.1 Goals and Objectives................................................................................................... 98

10.6.2 Technical Description ................................................................................................. 99

10.6.3 Roadmap ..................................................................................................................... 99

10.7 Technology Roadmap Overview ..............................................................................................99

11.0 Conclusions......................................................................................................................................100

11.1 Tasks and Trade Studies—Airframe.......................................................................................100

11.1.1 Advanced Concept Vehicle Definition ..................................................................... 100

11.1.2 Rapid Conceptual Design Low-Boom Design Process or Design Space Trade 

Studies....................................................................................................................... 100

11.2 Tasks and Trade Studies—Propulsion ....................................................................................101

11.3 System of Systems Analysis...................................................................................................102

11.4 Technology Roadmaps—Airframe.........................................................................................102

11.5 Technology Roadmaps—Propulsion ......................................................................................103

11.6 Meeting NASA N+3 Goals.....................................................................................................104

Appendix A.—Acronym List....................................................................................................................105

Appendix B.—Enabling Technologies.........................................................................................................1

References.....................................................................................................................................................1

List of Tables

TABLE 1.—LM’S PREFERRED CONCEPT WITH TECHNOLOGY INPUTS MEETS OR 

SURPASSES ALL N+3 GOALS........................................................................................................iv

TABLE 2.—N+3 ENVIRONMENTAL AND PERFORMANCE GOALS.................................................2

TABLE 3.—PROS AND CONS OF EACH ALTERNATIVE CONFIGURATION AS 

COMPARED TO THE INITIAL CONFIGURATION...................................................................12

TABLE 4.—SIZED N+3 CONFIGURATION MISSION PERFORMANCE...........................................15

TABLE 5.—VCE VERSUS TURBOFAN SIZED CONFIGURATION PERFORMANCE ....................18

TABLE 6.—SUMMARY OF LOUDNESS, PLDB...................................................................................27
NASA/CR—2010-216796 x

TABLE 7.—LOUDNESS FOR ALL GROUND SIGNATURES, VARIOUS TURBULENCE 

REALIZATIONS ...............................................................................................................................32

TABLE 8.—BOOM MAGNITUDE ACROSS CARPET, SIG1 50 PERCENT HUMIDITY...................32

TABLE 9.—BOOM MAGNITUDE ACROSS CARPET, SIG2 50 PERCENT HUMIDITY...................33

TABLE 10.—BOOM MAGNITUDE ACROSS CARPET, SIG3 50 PERCENT HUMIDITY.................33

TABLE 11.—NONPRIORITIZED OF PUGH MATRIX TECHNOLOGIES DOWN SELECTED 

FOR FURTHER ANALYSIS ............................................................................................................81

TABLE 12.—RECOMMENDED ENABLING TECHNOLOGIES FOR FURTHER RESEARCH 

AND ANALYSIS ..............................................................................................................................81

TABLE 13.—PRIORITIZED LIST OF ENABLING TECHNOLOGIES NECESSARY TO 

ACHIEVE N+3 PERFORMANCE AND ENVIRONMENTAL GOALS ........................................82

TABLE 14.—PROPULSION SYSTEM TECHNOLOGY MATRIX .......................................................83

TABLE 15.—PRIORITIZED AIRFRAME TECHNOLOGY ROADMAP ............................................103

TABLE 16.—PRIORITIZED PROPULSION TECHNOLOGY ROADMAP ........................................104

TABLE 17.—LM’S PREFERRED CONCEPT WITH CRITICAL TECHNOLOGY MEETS OR 

SURPASSES ALL N+3 GOALS.....................................................................................................104

TABLE B.1.—N+3 SUPERSONIC TECHNOLOGY LIST—BOLT-ON TECHNOLOGIES ...................1

TABLE B.2.—N+3 SUPERSONIC ALTERNATIVE CONFIGURATION OPTIONS .............................7

TABLE B.3.—N+3 SUPERSONIC TOOLS AND METHODS LIST ......................................................10

TABLE B.4.—N+3 SUPERSONIC LIST OF MATERIALS AND GENERIC TECHNOLOGIES.........12

List of Figures

Figure 1.—Technology roadmap for N+3 enabling technologies. ............................................................... v

Figure 2.—LM WBS for N+3 Phase 1 program........................................................................................... 3

Figure 3.—Relation between vehicle length and perceived level of noise (PLdB). ..................................... 4

Figure 4.—Effect of shock separation on loudness. ..................................................................................... 5

Figure 5.—Initial sizing for reference mission. ............................................................................................ 6

Figure 6.—Initial configuration definition.................................................................................................... 7

Figure 7.—Area-ruled cabin layout. ............................................................................................................. 7

Figure 8.—Alternative configuration concepts chosen for further analysis. ................................................ 8

Figure 9.—Wing trailing edge swept to maximize compression lift and shock (and drag) 

cancellation from airfoil reflex. ....................................................................................................... 9

Figure 10.—Component and Mach 1.6 area distributions for roll angles 0° to 180°.................................. 10

Figure 11.—Lift distributions at different angles. ...................................................................................... 10

Figure 12.—Equivalent area distribution undertrack (left), and side (right) for oblique wing and 

baseline configuration.................................................................................................................... 11

Figure 13.—Undertracking ground signatures, before and after aero/boom shape optimization. .............. 13

Figure 14.—Shock strength versus L/D tradeoff points. ............................................................................ 13

Figure 15.—N+3 sizing mission................................................................................................................. 14

Figure 16.—N+3 Sizing space and results.................................................................................................. 15

Figure 17.—Range and efficiency versus MTOW. .................................................................................... 16

Figure 18.—Sizing design space with takeoff field contours. .................................................................... 16

Figure 19.—Range versus takeoff field length for MTOW = 285,000 lb................................................... 17

Figure 20.—Technology impacts to sized configuration mission range..................................................... 17

Figure 21.—NLF schematic........................................................................................................................ 18

Figure 22.—Time to Mach with and without thrust augmentation............................................................. 19

Figure 23.—Thrust augmentation (TA) impact on sizing mission time and range..................................... 19

Figure 24.—Time to altitude, with and without thrust augmentation......................................................... 20

Figure 25.—Fuel to transonic acceleration, with and without thrust augmentation. .................................. 20

Figure 26.—Sizing mission range, with and without thrust augmentation................................................. 21
NASA/CR—2010-216796 xi

Figure 27.—Cruise friction drag coefficient for 3 levels of natural laminar flow...................................... 22

Figure 28.—Cruise drag polars for 3 levels of natural laminar flow.......................................................... 22

Figure 29.—Range and efficiency [pax*nmi / lb fuel] benefits of natural laminar flow—constant 

GTOW. .......................................................................................................................................... 23

Figure 30.—Weight savings from natural laminar flow—constant range.................................................. 23

Figure 31.—F-functions for three configurations....................................................................................... 24

Figure 32.—Thin shock ground booms (no thickening)............................................................................. 25

Figure 33.—Nominal ground signatures from PC-Boom, 1/P Taylor shock structures. ............................ 26

Figure 34.—Ground Booms at 50 percent relative humidity...................................................................... 28

Figure 35.—Ground booms at 5 percent relative humidity. ....................................................................... 29

Figure 36.—Distortion of sonic booms by turbulence................................................................................ 30

Figure 37.—Sig1, 50 percent humidity, with peaking and rounding turbulence........................................ 30

Figure 38.—Signatures, peak overpressures and perceived loudness across the carpet, Sig1 

50 percent RH. ............................................................................................................................... 31

Figure 39.—Final configuration three-view. .............................................................................................. 34

Figure 40.—Fuel packaging for final configuration. .................................................................................. 35

Figure 41.—N+3 cruise drag components. ................................................................................................. 36

Figure 42.—Aerodynamic performance at Mach 1.6, 50,000 ft................................................................. 36

Figure 43.—Takeoff drag polars................................................................................................................. 37

Figure 44.—Sized N+3 configuration weights statement........................................................................... 37

Figure 45.—ATS productivity versus carbon dioxide. ............................................................................... 39

Figure 46.—DOC growth and increase attributable to SST’s..................................................................... 39

Figure 47.—SST trips, passengers, and range within ATC. ....................................................................... 40

Figure 48.—N+3 climb profile. .................................................................................................................. 41

Figure 49.—N+3 mission profile (climb phase). ........................................................................................ 42

Figure 50.—N+3 descent profile. ............................................................................................................... 42

Figure 51.—N+3 mission profile (descent phase). ..................................................................................... 43

Figure 52.—Potential N+3 route network................................................................................................... 44

Figure 53.—Sector adaptation for U.S. airspace in 2008. .......................................................................... 45

Figure 54.—Departure rate for selected U.S. airports. (Blue bars represent conventional aircraft 

and orange represent AST operations.).......................................................................................... 47

Figure 55.—Departure rate for selected U.S. airports. (Blue bars represent conventional aircraft 

and orange represent AST operations.).......................................................................................... 48

Figure 56.—Arrival rate for selected U.S. airports. (Blue bars represent conventional aircraft and 

orange represent AST operations.)................................................................................................. 49

Figure 57.—Arrival rate for selected U.S. airports. (Blue bars represent conventional aircraft and 

orange represent AST operations.)................................................................................................. 50

Figure 58.—AST route network analysis across four AST departure phases............................................. 51

Figure 59.—AST route network for the East Coast.................................................................................... 51

Figure 60.—Phase 1 analysis for selected departure sectors. ..................................................................... 53

Figure 61.—Phase 1 air traffic complexity and controller workload.......................................................... 54

Figure 62.—Phase 2 analysis for selected departure sectors. ..................................................................... 55

Figure 63.—Phase 2 air traffic complexity and controller workload.......................................................... 56

Figure 64.—Phase 3 analysis for midlevel en route sectors. ...................................................................... 57

Figure 65.—Phase 3 air traffic complexity and controller workload.......................................................... 58

Figure 66.—Phase 4 air traffic complexity and controller workload.......................................................... 59

Figure 67.—Phase 4 conflict analysis......................................................................................................... 59

Figure 68.—Phase 6 conflict analysis......................................................................................................... 60

Figure 69.—LM NASA N+3 bleed and horsepower extraction requirements ........................................... 62

Figure 70.—Results from preliminary FPR Design Space Exploration Trade Study................................. 64

Figure 71.—T4 impact on weight and SFC. ............................................................................................... 65
NASA/CR—2010-216796 xii

Figure 72.—Total jet noise level for sideline monitor location for VCE engine for different power 

setting with and without augmentation.......................................................................................... 66

Figure 73.—Typical jet noise level for jet noise reduction features versus mixed velocity scaled 

experimental data........................................................................................................................... 67

Figure 74.—Conventional MFTF mixed flow and axisymmetric plug nozzle exhaust.............................. 69

Figure 75.—Impact to range based on LM sensitivities. ............................................................................ 69

Figure 76.—GE Twin Annular Premixed Swirler (TAPS) combustor. ...................................................... 70

Figure 77.—Opposed flow flame calculated using Chemkin with Jet-A. .................................................. 71

Figure 78.—Effect of amount of mixer air on EINOx with different values of T4.................................... 71

Figure 79.—Variation of EINOx with different levels of mixedness versus T4. ....................................... 72

Figure 80.—Variation of EINOx with different levels of pilot air versus T4............................................. 72

Figure 81.—Two notional types of heat exchangers. ................................................................................. 73

Figure 82.—Schematic diagram showing implementation of PDE in NPSS model. ................................. 75

Figure 83.—Technology identification process used to construct the technology roadmaps..................... 77

Figure 84.—Technology catalogue excerpt used to define space of possible technologies, provide 

input for down selection, and disseminate ideas............................................................................ 77

Figure 85.—Technology down selection process that produced manageable list of technologies to 

further evaluate and create roadmaps............................................................................................. 78

Figure 86.—Configuration technologies assumed to be designed into the advanced vehicle 

concept........................................................................................................................................... 79

Figure 87.—Observations and insights from rankings contributed to prioritization. ................................. 80

Figure 88.—GOTCHA successfully traces prioritized technologies to N+3 environmental goals. ........... 84

Figure 89.—GOTCHA successfully traces prioritized technologies to N+3 performance goals. .............. 85

Figure 90.—Roadmap for tools/methods—low boom shaping fidelity...................................................... 86

Figure 91.—Roadmap for tools/methods—CFD-based MDAO. ............................................................... 87

Figure 92.—Laminar flow technology benefits applied to advanced vehicle concept. .............................. 88

Figure 93.—Roadmap for laminar flow technology—plasma augmented distributed roughness.............. 88

Figure 94.—Roadmap for pilot situation awareness—low boom violation cueing.................................... 89

Figure 95.—Roadmap for pilot situation awareness—preplanning awareness. ......................................... 90

Figure 96.—Roadmap for tools/methods—integrated structural analysis.................................................. 91

Figure 97.—Roadmap for adaptive geometry—lift distribution control. ................................................... 92

Figure 98.—Roadmap for plasma—jet exhaust manipulation.................................................................... 93

Figure 99.—Roadmap for adaptive geometry—inlet flow control............................................................. 93

Figure 100.—Airframe technology roadmap for enabling supersonic technologies. ................................. 94

Figure 101.—Roadmap for VCE propulsion system. ................................................................................. 95

Figure 102.—Low noise/high performance exhaust system roadmap........................................................ 96

Figure 103.—Low emission combustor roadmap....................................................................................... 97

Figure 104.—Cooled cooling air roadmap. ................................................................................................ 98

Figure 105.—Transonic thrust augmentation device roadmap................................................................... 98

Figure 106.—CVC roadmap....................................................................................................................... 99

Figure 107.—Technology roadmap overview for propulsion technologies. .............................................. 99

Figure 108.—Final configuration three-view. .......................................................................................... 100

Figure 109.—Ground signature before and after aero/boom shape optimization..................................... 101
NASA/CR—2010-216796 1

1.0 Introduction

1.1 Subject of the Report

Research in the area of Advanced Supersonic Transport (AST) has been a focus area for NASA since 

1960s, driven by maintaining U.S. leadership in the area of commercial transport. According to a 1980 

Open Travel Alliance (OTA) report (Ref. 1) on the impact of advanced air transport technology, the 

business case in favor of ASTs results from improved aircraft productivity (measured in seat miles 

generated per unit time) and its capability to transport twice the number of passengers on long distance 

flight. Higher cost of operations, concerns over environmental impact due to noise and emissions, and 

restrictions to fly supersonic on land due to sonic boom are some of the technological issues that need to 

be addressed for production and deployment of ASTs. NASA’s research efforts for the advancement of 

AST are dedicated to address these technical challenges and the AST technology is being matured under 

N+1, N+2 and N+3 projects. The goal of the N+3 project is to explore a conceptual design for multiMach aircraft in 2030 timeframe that has low sonic boom, is environmentally acceptable, fuel efficient, 

and able to fly at supersonic speed above land. Other, integrated design concerns include:

• Sonic Boom Reduction

• Cruise Efficiency

• Aero-Propulso-Servo-Elasticity

• Airport Noise

• Light Weight Structure for Airframe/Propulsion Systems

• High Altitude Emissions

A complimentary area for NASA research is the Next Generation Air Transportation System 

(NextGen) (Joint Planning and Development Office, 2009 (Ref. 2)). The U.S. Air Traffic Management 

(ATM) system is today operating at the edge of its capabilities, handling the real-time planning and 

coordination of over 50,000 flights per day. Although air traffic has seen a decline in the recent year due 

to severe economic downturn, the recent numbers suggest that traffic is currently stabilizing (Official 

Airline Guide, 2009 (Ref. 3)) however, per market forecasts by MITRE (Ref. 4) and Boeing (Ref. 5)

(2009) a strong growth in air traffic is expected in both short and long term. Additionally, Boeing’s long 

term market forecast cites that the air transportation industry is resilient and has survived many economic 

downturns in the past. It has grown at 5 percent annually and by year 2029 the number of airplanes flying 

in the National Air Space will be more than double. To address this concern, the Federal Aviation 

Administration (FAA) along with NASA and other government and industry partners are charting the 

NextGen.

One of the strategic objectives outlined in the NextGen plan is to have a system scalable enough to 

respond quickly and efficiently to increase in demand and is flexible enough to incorporate new types of 

airframe for example, Unmanned Aircraft System (UAS), Very Light Jets (VLJs), Large Civil Tiltrotor 

(LCTR), ASTs, and others. Since supersonic transports provide a step increase in passenger mobility by 

speed of travel, their incorporation within the NextGen ATS could potentially provide alternative methods 

of operation, subsonic to supersonic transition regulations, and unforeseen hazards. NASA is focused on 

providing vehicle designs and identifying enabling technologies that can meet the nation’s need for 

effective, efficient and safe air travel.

Overall, the supersonics project is designed to develop knowledge, capabilities, and identify 

innovative solutions for supersonic air vehicles. Sonic boom, environmental concerns, and NextGen ATS 

integration are major concerns for commercial supersonic travel. Revolutionary solutions are required to 

generate viable, supersonic solutions.
NASA/CR—2010-216796 2

1.2 Purpose

The purpose of this final report (PMF–01403) is to respond to the NASA sponsored program “N+3 

NRA Advanced Concept Studies for Supersonic Commercial Transports Entering Service in the 2030 to 

2035 Period.” The N+3 program is focused on generating promising supersonic concepts for the 2030 to 

2035 timeframe and to develop plans for maturing the technologies required to make those concepts a 

reality. An additional system-level focus includes understanding how a supersonic civil transport would 

integrate and operate within the 2025 NextGen ATS.

This program is committed to overcoming significant performance (cruise speed, range, payload and 

fuel efficiency) and environmental (sonic boom, airport noise, and cruise emission) challenges. The 

NASA stated N+3 goals are illustrated in Table 2.

TABLE 2.—N+3 ENVIRONMENTAL AND PERFORMANCE GOALS

N+1

Supersonic Business 

Class Aircraft

(2015)

N+1

Small Supersonic 

Airliner

(2020)

N+1

Efficient MultiMach

Aircraft

(beyond 2030)

Environmental goals 

Sonic boom 65 to 70 PLdB 65 to 70 PLdB 65 to 70 PLdB

low boom flight

75 to 80 PLdB

unrestricted flight

Airport noise 

(cum blow Stage 3)

10 EPNdB 10 to 20 EPNdB 20 to 30 EPNdB

Cruise emissions

(cruise NOx g/kg of fuel)

Equivalent to 

current subsonic

<10 <5 and particulate and water 

vapor mitigation

Performance goals

Cruise speed Mach 1.6 to 1.8 Mach 1.6 to 1.8 Mach 1.3 to 2.0

low boom flight

Mach 1.3 to 2.0

unrestricted flight

Range 

(nmi)

4000 4000 4000 to 5500

Payload 

(passengers)

6 to 20 35 to 70 100 to 200

Fuel efficiency 

(passenger-miles per lb of fuel)

1.0 3.0 3.5 to 4.5

Meeting or surpassing these goals stimulates innovation and advances the pursuit of revolutionary 

conceptual designs. System-level multi-disciplinary analysis and optimization (MDAO) and out of the 

box thinking allows for revolutionary technology identification. This fosters an environment of 

innovation and generates excitement for future supersonic travel.

Overall, the N+3 effort is driven by the need for alternative solutions capable of overcoming the 

efficiency, environmental, and performance barriers to practical supersonic flight. Results from these 

studies aid in upcoming research efforts and provides a roadmap for future supersonic funding. 

1.3 Scope

LM Aeronautics conducted research, testing, trade studies and sensitivity analysis in support of the 

NASA’s N+3 Supersonic Vehicle effort. A combination of advanced design and an integrated system 

analysis was taken to define a conceptual vehicle capable of meeting the environmental and performance 

goals. Viable technology development paths were produced by the design, engineering, and test 

capabilities of our team. In addition, core technology trades were performed to provide estimates of the 

advanced vehicle concept’s noise, emissions and performance characteristics. LM was also responsible 

for the coordination and management of all subcontractors and resulting work. LM is committed to 
NASA/CR—2010-216796 3

helping NASA successfully achieve their goals of first understanding what is necessary in 2030 to 2035, 

generating a suite of enabling concepts and technologies to meet those needs, and socializing that vision 

with the broadest possible audience.

2.0 Work Breakdown Structure (WBS)

LM Aeronautics was responsible for the overall design, development, and technology identification 

necessary to realize a visionary vehicle capable of achieving the supersonic N+3 environmental and 

performance goals. A combination of advanced design and an integrated system approach was required to 

define an advanced concept vehicle serviceable in the 2030 to 2035 timeframe (Task 3.1). Design of the 

vehicle included configuration layout, design, analysis, and definition to produce a concept tightly 

integrated with airframe and propulsion technologies. Using a system-level design space, LM Aeronautics 

was also tasked to perform various trade and sensitivity studies to understand how a future Next 

Generation (NextGen) scenario with supersonic transports drove design requirements (range, noise, 

emissions, boom, fuel, and mobility). The interplay of design constraints was modeled and analyzed in 

physics based multi-disciplinary analysis and optimization (MDAO) process using Rapid Conceptual 

Design (RCD). Task 3.2 included RCD model development, integration with technology inputs, 

quantified analysis, and technology benefit/impact assessments. After multiple design iterations and 

system-level analysis of the preferred configuration, LM Aeronautics was responsible for developing a 

technology roadmap of enabling technologies for the N+3 vehicle. This roadmap includes a list of key 

technologies, definition of roles and quantification of impacts on the concept vehicle, traceability to N+3 

goals, baseline TRLs, proposed TRL maturations schemes for future N+3 phases, and prioritization. 

Overall, LM Aeronautics was ultimately responsible to optimize complex multi-variable combinations of 

airframe and propulsion technologies while iterating, maturing, identifying, and ultimately downselecting critical technologies required to realize an N+3 vehicle. Figure 2 illustrates the overall WBS of 

the tasks and duties required for the program.

Figure 2.—LM WBS for N+3 Phase 1 program.
NASA/CR—2010-216796 4

Our efforts focused on four major tasks: Advanced Concept Vehicle Definition (Task 3.1), Design 

Space Trade Studies (Task 3.2), Technology Roadmap Development (Task 3.3), and ATS System-ofSystems Analysis (Task 3.6). LM was also responsible for the management and coordination of seven 

subcontractors to provide subject-matter data and expertise to the program. Collaboration included

teaming with GE Global Research Center (GRC) with GE Aviation for advanced propulsion concepts as 

well as fuel efficiency and emissions, Penn State for jet noise reduction, Purdue for system of system 

analysis, MIT for green initiatives, Wyle labs for real-world loudness effects and boom guidance, LM 

Transportation and Security Solutions for air traffic analysis, and Helen Reed and Bill Saric for laminar 

flow analysis. All required tasks include subsequent subtasks that align with the main task. The WBS 

encompasses all work necessary to oversee and direct the execution of the N+3 Phase 1 Program. 

3.0 Tasks and Trade Studies—Airframe Systems

3.1 Advanced Vehicle Concept (WBS 3.1)

3.1.1 Concept Layout and Design (WBS 3.1.1)

3.1.1.1 Description

Before laying out a configuration, we looked at the N+3 goals and addressed design methods and 

strategies necessary to meet those challenges. Based on our past history designing and analyzing 

supersonic configurations, we first focused our energy on the sonic boom requirement. The N+3 sonic 

boom goal of 65 to 70 PLdB is significantly lower than the state of the art 107 PLdB of the (408,000 lb, 

100 passenger) Concorde with a shock strength of 2 psf, or the 102 PLdB of the (12,000 lb—33 times 

lighter than Concorde) F-5 with a shock strength of 1.3 psf. Meeting the sonic boom goal requires a 

minimum shock (ramp signature) shock strength of 0.12 to 0.17 psf. One way of meeting this goal is 

increasing the fuselage length used by SEEB to calculate the minimum shock signature, as shown in 

Figure 3. In order to reduce the length required, it is anticipated that the perceived level of noise on the 

ground can be reduced through shock blending, as shown in Figure 4, and through taking into account 

real world absorption and turbulence. Results from Wyle’s analysis on Effects of Atmospheric 

Propagation on Low-Boom Shaped Signatures can be seen in Section 3.2.5.

Figure 3.—Relation between vehicle length and perceived level of noise (PLdB).
NASA/CR—2010-216796 5

The other noise challenge was meeting the airport noise goal of 20 to 30 EPNdB cumulative below 

FAR36 Stage 3 limits. Current subsonic airplanes, like the Boeing 777–200 with GE 90–85B and the 

Airbus A380 with RR Trent 970, already meet this goal at 23 EPNdB and 26 EPNdB cum below Stage 3 

respectively. However, it is more of a challenge for supersonic aircraft. Using the Concorde as a state-ofthe-art (SOA) comparison, its supersonic transport status is 45 EPNdB cumulative above Stage 3. Our 

strategy for meeting the noise goal was to first require GE to meet sideline –3 EPNdB at 90 percent power 

also known as PLR (programmed lapse rate), use the GE VCE, and optimize takeoff procedures. Second, 

reduce approach noise with a low-noise fan design, inlet liners and inlet flow choking. Third, investigate 

other promising advanced technologies.

Figure 4.—Effect of shock separation on loudness.

Signature Variations for Loudness vs. Shock Separation

-0.8

-0.6

-0.4

-0.2

0

0.2

0.4

0.6

0.8

0 20 40 60 80 100 120 140 160 180 200

Time, msec

Ground Overpressure, psf

15 msec Separation

6 msec Separation same Duration

6 msec Separation same Expansion

Signature Variations for Loudness vs. Shock Separation

-0.8

-0.6

-0.4

-0.2

0

0.2

0.4

0.6

0.8

0 20 40 60 80 100 120 140 160 180 200

Time, msec

Ground Overpressure, psf

15 msec Separation

6 msec Separation same Duration

6 msec Separation same Expansion
NASA/CR—2010-216796 6

Figure 5.—Initial sizing for reference mission.

As part of the iterative design process, we looked at a number of different vehicle concepts that would 

integrate features necessary to achieve the N+3 mission requirements and performance goals. Desirable 

configuration features included items that would provide low boom, low drag, low weight, and good 

aeroelasticity performance for cruise and off-cruise conditions. Drawing on previous Quiet Supersonic 

Transport (QSST) experience, our process started with applying the desirable configuration features to a 

modified inverted-V, “QSST-like” concept. The four-engine inverted V-tail concept was proposed to 

better capture advantages of the inverted tail concept—particularly greater wing bending moment relief.

Preliminary vehicle sizing with QSST and historical data established the weight breakdown necessary 

to determine engine thrust, wing sizing, and fuselage length for boom. A slight improvement was 

assumed, giving an L/D of 10 and an SFC of 0.95 lb fuel/lb thrust/hr. These assumptions were applied to 

the reference mission of 100 passengers, 4000 nmi range, and Mach 1.6 cruise. This resulted in a Max 

Take-off Gross Weight (MTOW) of just over 300,000 lb, with an efficiency of 3.07 pax-nmi/lb fuel, as 

shown in Figure 5. However, this did not meet the requirement of efficiency between 3.5 to 4.5 paxnmi/lb fuel. It was calculated that the efficiency could be raised to 3.97 pax-nmi/lb fuel if the L/D 

increased to 11, the SFC improved to 0.90 lb fuel/lb thrust/hr, and empty weight reduced by 5 percent.

This quantified the N+3 vehicle improvement values to achieve NASA’s desired performance goals. 

These values were status indicators as opposed to targets. N+3 technologies were sought to maximize 

performance as much as possible and potentially go beyond these goals.

3.1.1.2 Results

The initial configuration was sized with an assumed MTOW approximately equal to 300,000 lb, 

resulting in a wing area approximately equal to 3,000 ft2

, and a take-off thrust approximately equal to 

100,000 lb The benefits of this low-boom configuration include stretched boom signature due the inverted 

V-tail and nose droop, favorable aerodynamic interference and compression lift for aft-under-wing 

mounted engines, efficient propulsion integration due to the planform trailing edge sweep and airfoil 

reflex, aerodynamic efficiency for wing planform design, reduced wing gull roll penalties due to wing tip 

and inverted V-tail anhedral, and structural flexibility suppression due to inverted V-tail wing bracing.

Once designed, these specific elements were considered endemic to the configuration and always a part of 

the initial configuration technology set. The design was used as the “yardstick” to compare other potential 

configurations. Figure 6 highlights the overall initial configuration definition and design features that 

were modeled within CATIA V5.

Once the initial concept was defined, an initial inner mold-line (IML) cabin volume constraint was 

determined to insert passengers within the loft. The initial configuration held 101 passengers including 

future projected economy seat sizing comfort improvements relative to the Concorde and other regional 

jets plus the provision for 10 percent first class seats. The cabin layout included one galley, two 

lavatories, one supplemental space, and three emergency constraints. The boom constraints on the 

fuselage outer mold line (OML) forced cabin camber and cross section pinching on each end. This 

limitation required one first class seat to be removed from the forward section of the cabin, and a 

unification of the next set of seats. Nine rows in the aft section of the cabin changed from four across to 

three across while the cabin was lengthened. Cambered cabin slopes less than 5 percent have to be 

reconciled in a future design phase. Figure 7 demonstrates a realistic cabin layout that establishes fuselage 

IML constraints for the initial configuration.
NASA/CR—2010-216796 7

Figure 6.—Initial configuration definition.

Figure 7.—Area-ruled cabin layout.
NASA/CR—2010-216796 8

Figure 8.—Alternative configuration concepts chosen for further analysis.

3.1.2 Alternative Configurations (WBS 3.1.1)

3.1.2.1 Description

The N+3 concept vehicle definition also included exploration of alternative concepts, both 

conventional and unconventional, to investigate all potential configuration solutions. Figure 8 highlights 

the various configurations that were studied starting with the family of inverted-v tail configurations and 

branching off to an oblique wing, a twin-fuselage concept, and a variety of brainstorming concepts. 

3.1.2.1.1 “Blue Sky” Configurations

After the initial ideas listed above were considered, further brainstorming sessions, called “blue sky,” 

were conducted with leading experts from outside the program, to identify more revolutionary concepts.

However, no further configurations were discovered that could reasonably outperform those concepts 

already being considered.

3.1.2.1.2 Engines-Over-Wing Configuration

The engine-over-wing configuration was considered to address potential 

structural benefits (shorter landing gear) and noise level reductions possible 

with engine placements above the wing. When the engines are placed over the 

wing, engine spillage shocks are blocked from the ground by the wing.

However, this results in higher pressure on the upper surface of the wing 

predicted to reduce L/D by 2 points.

In order to assess the need for noise reduction with the engines over wing configuration, it needed to be 

determined how low the noise could be for the engine under wing configuration. This was done through a 

wing configuration study to address propulsion/airframe integration (PAI) issues of a low-boom design.

Figure 9 exhibits the trailing edge design study used for favorable interference drag. The wing trailing 

edge was swept to capture maximum nacelle shock compression lift and airfoil reflex for shock (and drag) 

cancellation. The nacelle shock was substantially countered; it met a 65 to 70 PLdB equivalent area target 

as easily as above the wing engine placements. The high pressure caused by the nacelle shock on the 

lower surface of the wing resulted in higher efficiency (lower angle-of-attack) through an increased L/D.

Since it was possible to meet the sonic boom requirement with the higher efficiency of the engines-underwing concept, further development of the engines-over-wing configuration was discontinued.
NASA/CR—2010-216796 9

Wing planform

Figure 9.—Wing trailing edge swept to maximize compression lift and shock (and drag) cancellation from airfoil reflex.

3.1.2.1.3 T-Tail Configuration

The T-tail configuration was considered in order to raise the aft lift and stretch 

the boom signature. The configuration reduces interference drag and eliminates 

the inverted-V structure. A set of coarse geometry trades was performed to reduce 

the T-tail’s sonic boom level while minimizing adverse impacts to drag. These 

trades followed the same approach used on the inverted V-tail, which is described 

in Section 3.2.1. Results for both configurations are provided in Table 4 in that 

section. This configuration was not considered further, however, because it also entailed an increase in 

wing weight (because the T-tail does not help support the wing as the inverted V-tail does), possible 

flutter issues, and a larger fin.

3.1.2.1.4 Oblique Wing Configuration

Oblique wings have been studied since the 1970s and have been proven to 

provide good aerodynamic performance at supersonic and subsonic speeds. The 

variable sweep allows better aerodynamic optimization at a variety of Mach 

numbers. If the takeoff constraint is driving the wing size, this would allow for a 

smaller wing with good low-speed, transonic, and supersonic performance. The 

smaller wing would require less thrust, possibly requiring fewer engines.

However, an oblique wing design complicates the boom design and introduces a weight penalty with the 

wing pivot mechanism. Analysis of the area distributions was performed to determine whether the 

complications to the boom design could be overcome.

Figure 10 shows area distributions at Mach 1 for the components and area distributions for the 

combined vehicle at Mach 1.6 for roll angles from 0° to 180°. The Mach 1 area distribution shows one of 

the benefits of the oblique wing, in that the area of the wing is spread over a greater length of the vehicle 

than a traditional wing, which can help reduce wave drag. The Mach 1.6 area distributions show that there 

are some smile angles that are better than others. For a smile angle that causes the Mach angle to align 

more closely with the wing, the wing appears quickly in the area distribution, increasing that angle’s 

contribution to wave drag.

Figure 11 shows the lift distributions for different angles. At 49°, which is opposite the angle of the 

wing, the lift distribution is spread out over a greater length of the vehicle at about 1250 lb/ft. For the 

opposite –49° angle, the lift is concentrated over a much shorter length, at almost 3000 lb/ft. This results

0

10

20

30

40

50

100 120 140 160 180 200 220

BL FEET

FS FEET (nose at 8.333) Rear Spar

0

10

20

30

40

50

100 120 140 160 180 200 220

BL FEET

FS FEET (nose at 8.333) Rear Spar
NASA/CR—2010-216796 10

Mach 1 Area Distribution of Components Mach 1.6 Area Distributions

Figure 10.—Component and Mach 1.6 area distributions for roll angles 0° to 180°.

Figure 11.—Lift distributions at different angles.

in an asymmetric sonic boom signature, where the vehicle is quieter on one side and louder on the other.

In Figure 12, the equivalent area distribution is shown compared to the baseline configuration. On the left 

plot, the undertrack distribution for the oblique wing is better than for the baseline, since it does not have 

the overshoot. The right plot shows the equivalent area distributions off to the sides compared to the 

baseline. At +49°, this again shows that the oblique wing has a better distribution than baseline, but on the 

other side at –49°, the area distribution is significantly worse.

The oblique wing configuration allows for improved signature in some areas at the expense of others.

The efficiency can be higher at off-design Mach numbers through rotations; however, since the vehicle 

spends the majority of the time at the design cruise Mach number, this has a limited benefit. Unsweeping 

the wing makes it easier to meet takeoff requirements. However, takeoff is not predominantly limiting the 

baseline design, so this benefit is likely to be attenuated using the engine cycle matched to the baseline. In 

addition, the weight penalty of the rotation mechanism counters the off-design benefits. In conclusion, the 

supersonic transport mission spends more than 85 percent around the cruise condition, so cruise 

efficiency while maintaining a low sonic boom disturbance is the predominant characteristic for the best 

supersonic transport design. While the boom constrained cruise efficiency analysis is incomplete, it does

show both advantages and disadvantages. At this point, the asymmetry inherent to this design makes it 

more challenging and resource intensive without a clear benefit indicated to justify the extra work. The
NASA/CR—2010-216796 11

Figure 12.—Equivalent area distribution undertrack (left), and side (right) for oblique wing and baseline configuration.

show oblique wing appears to be a viable design alternative at this depth of analysis; however, it was not 

chosen as a baseline design.

3.1.2.1.5 Twin Fuselage Configuration

The twin fuselage configuration is another proposed N+3 configuration but 

presented fuel efficiency and boom design complications. The benefit of the twin 

fuselage configuration is the favorable interference between the fuselages that 

reduce drag. The fuselage weight is also split on the wing, reducing bending 

moment and wing structural weight. With the separation distance between the two 

fuselages, the far-field method would not accurately predict wave drag. Using an 

appropriate near-field method, the twin fuselage’s favorable interference at 

optimal separation was small (less than 2 percent drag reduction) while the minimum practical cabin 

cross-section created a large fineness ratio penalty for the 100 to 200 passenger size. The twin fuselages’ 

increased surface area skin friction and increased cross-section wave drag overwhelmed the possible wing 

structural benefit, prompting the elimination of this configuration.

3.1.2.1.6 Mother/Daughter Configuration

The mother/daughter concept involves two vehicles: a “daughter” optimized for cruise (high wing 

loading, thrust sized for cruise) and not subject to any takeoff field constraints, and a “mother” optimized 

for takeoff and landing performance (low wing loading, thrust sized for takeoff) without prioritizing 

cruise efficiency. Operations would involve the two vehicles taking off in tandem, relying partly on the 

mother’s thrust for takeoff and climb. The pair then separates upon attaining cruise altitude, at which 

point the daughter cruises efficiently and the mother returns to its home airport, most likely to serve 

several daughters and perhaps returning unmanned. The daughter could similarly rendezvous with another 

mother at its destination airport if additional power is needed for landing. Comparison of the baseline 

configuration and a daughter (nontakeoff constrained) version of it, developed for this purpose, showed 

that takeoff was not a strongly limiting constraint. In other words, the takeoff and cruise thrust 

requirements of the baseline were closely balanced. Entirely new airframe and engine designs that did not 

already mitigate takeoff requirements would be required to understand the full benefits of no takeoff

requirements. Since the mother/daughter configurations would require different propulsion systems, 

aerodynamics, etc., a judgment was made that the added complexity of a second vehicle did not warrant 

further investigation.

3.1.2.2 Results

Analysis and consideration of each alternative configuration resulted in each one’s elimination. 

Table 3 summarizes the pros and cons of each alternative configuration as it was compared and analyzed 
NASA/CR—2010-216796 12

in regards to the initial configuration design. Within the family of configurations, the T-tail configuration 

was eliminated due to a lack of substantial benefits, and the engine over wing design was eliminated due 

to a two point reduction in L/D. The oblique wing configuration promised high aerodynamic 

performance, but the design was ultimately eliminated due to the weight penalty of the rotation 

mechanism and the sonic boom design difficulties. The twin fuselage configuration provided potential 

weight savings, but these benefits were not enough to outweigh the boom design complications to achieve 

the N+3 sonic boom goal. And finally, the mother/daughter configuration was eliminated due to 

complexity. As a result, the “as-drawn” initial configuration was used to perform additional sizing and 

mission analysis and lofted in CATIA V5 as the final configuration as seen in Section 3.3.

TABLE 3.—PROS AND CONS OF EACH ALTERNATIVE CONFIGURATION

AS COMPARED TO THE INITIAL CONFIGURATION

3.2 Design Space Trade Studies With RCD (WBS 3.2)

3.2.1 RCD Low Boom Design Approach (WBS 3.2.1)

The first set of trade studies defined the inverted V-tail in terms of its longitudinal location, span, 

area, dihedral and leading edge sweep angles, and thickness-to-chord (t/c) ratio. An analogous set of 

trades was performed to define the geometry of the T-tail configuration as mentioned in Section 3.1.2.

The approach was to identify the configuration that minimized drag, then adjust each geometry parameter 

individually to reduce sonic boom. By focusing on both drag and sonic boom, these trades constituted the 

first step in reducing the sonic boom associated with the inverted V-tail and T-tail configurations.

This initial set of trades transitioned the inverted V-tail configuration from a purely minimum-drag design 

to one with reduced sonic boom. It defined the planform and provided the starting point for the subsequent 

shape optimization process, which further reduced boom through finer adjustments in the shape of the lifting 

surfaces and ultimately enabled the inverted V-tail configuration to meet the N+3 noise goal.
NASA/CR—2010-216796 13

3.2.2 RCD Model Development (WBS 3.2.1)

3.2.2.1 Aero/Boom Shape Optimization

3.2.2.1.1 Description

Rapid Conceptual Design (RCD) was used to modify the lift distribution of the as-drawn 

configuration such that its boom signature conformed to the prescribed target. This model included 

induced and wave drag, ground signature, and equivalent area analyses to enable shaping to account for 

both drag and boom. This shape optimization was performed with the planform fixed in accordance with 

the results of the tail-location trade described above. Its goal was to ensure that the configuration could be 

shaped to meet the N+3 sonic boom goal of 70 PLdB.

3.2.2.1.2 Results

The resulting improvement in the as-drawn configuration’s sonic boom signature is illustrated in 

Figure 13. The left-hand plot corresponds to the sonic boom signature of the configuration prior to 

optimization (which resulted from coarse geometry trades performed previously). It was characterized by 

the large, 1.38-psf shock (increase in overpressure) occurring at 200 msec. The sonic boom signature 

corresponding to the resultant configuration, in which the maximum shock strength has been reduced to 

0.29 psf, is shown in the right-hand plot in Figure 13.

The consequence of this improvement in sonic boom signature was a reduction in L/D from 10.1 to 

8.0 (at beginning of cruise). This decline incurred because the initial shape did not meet the constraints 

required for sonic boom. The aero/boom shape optimization process yielded the five Pareto points in 

Figure 14, which represent the trade space between sonic boom and L/D.

Figure 13.—Undertracking ground signatures, before and after aero/boom shape optimization.

Figure 14.—Shock strength versus L/D tradeoff points.

Opt. Start

7

8

9

10

11

0 0.4 0.8 1.2 1.6

Max Shock (psf)

L/D

Max. Shock versus L/D 

Pareto (tradeoff) Points

Opt. Finish

-2

-1

0

1

2

0 50 100 150 200 250

Time (msec)

Overpressure (psf)

Initial Signature:

max shock = 1.38 psf

L/D = 10.1

-2

-1

0

1

2

0 50 100 150 200 250

Time (msec)

Overpressure (psf)

Optimized Signature:

max shock = 0.29 psf

L/D = 8.0
NASA/CR—2010-216796 14

This aero/boom shape optimization yielded two important results regarding the ability of the N+3 

vehicle to meet its goals. First, it verified that the aircraft’s lifting surfaces can be shaped to meet the 

70-PLdB boom target. Second, it quantified the aerodynamic cost of doing so in terms of L/D. The latter 

result, shown to be a 21 percent reduction (from 10.1 to 8.0), can be offset by aerodynamic technologies 

such as Natural Laminar Flow (NLF) (to be described in Section 3.2.4.2) to enable a low-boom 

configuration to continue to meet N+3 range goals of over 4000 nmi.

3.2.2.2 Sizing and Mission Performance

LM’s Rapid Conceptual Design (RCD) process was also used to size the as-drawn configuration to 

achieve maximum range subject to fuel volume and takeoff performance constraints. The integrated, 

multidisciplinary sizing model included multiple aerodynamic components (profile, induced, and wave 

drag), a parametric weights buildup, and mission performance and balanced field length (BFL) analyses. 

In addition to sizing, it facilitated assessments of N+3 technology benefits at the system (vehicle) level.

The sizing mission for the N+3 vehicle is illustrated in Figure 15 and consisted of the following 

segments:

1. Warm-up and takeoff—3 min at idle + 1 min at max power

2. Subsonic climb to 10,000 ft at 250 kn calibrated air speed (KCAS) (with afterburning)

3. Level acceleration at 10,000 ft: Mach 0.45 to Mach 0.66 (365 kn equivalent air speed (KEAS)) (with afterburning)

4. Subsonic climb and acceleration (maximum Mach 0.9) to 24,700 ft (with afterburning)

5. Level transonic acceleration to 450 KEAS (constant q=685 psf) (with afterburning)

6. Supersonic climb and acceleration at 450 KEAS until Mach 1.6 (with afterburning)

7. Cruise at Mach 1.6

8. Level deceleration to Mach 0.9

9. Descend at 4000 ft/min to 5,000 ft (constant 250 KEAS from 10,000 to 5,000 ft)

10. 5 min loiter awaiting clearance (no distance credit)

11. Descend to land (no distance credit)

12. Climb to 5000 ft (250 KCAS) after wave-off (no distance credit)

13. 5 min Loiter awaiting clearance (no distance credit)

14. Subsonic climb to 35,000 ft diversion altitude (250 KCAS to 10,000 ft, then 250 KEAS)

15. Diversion at best cruise mach

16. Descend at 4000 ft/min to 5,000 ft (250 KEAS above 10,000 ft, then 250 KCAS)

17. Descend to land (no distance credit)

18. Land with fuel reserves to loiter 30 min at 5,000 ft

Figure 15.—N+3 sizing mission.

1

2

3

4

5

6

7

9

10

11

10,000 ft

365 KEAS (Max Mach 0.9)

24,700 ft

450 KEAS

Seg’s 2 to 6 with 

afterburning

Cruise Mission Range

(no distance credit below 5000 ft on approach/landing)

8
NASA/CR—2010-216796 15

Figure 16.—N+3 Sizing space and results.

At a gross takeoff weight (GTOW) of 285,000 lb, the sized N+3 configuration achieved a range of 

4640 nmi using conventional fuel tanks (i.e., without expanding fuel capacity through Biologically 

Inspired Morphing Structure or other technologies). Its thrust-to-weight ratio (T/W) and wing loading 

(W/S) were 0.34 and 92 psf, respectively. This range lies well within the N+3 goal bounds of 4000 to 

5500 nmi. The two-dimensional sizing design space (T/W, W/S) is shown graphically in Figure 16, on 

which the orange and yellow circles correspond to the as-drawn and sized configurations, respectively. As 

indicated on the graph, the sized configuration is fuel volume constrained; its internal fuel volume, when 

full, corresponds to exactly 285,000 lb. GTOW. Configurations in the shaded blue area of the graph have 

insufficient fuel volume to take off at this weight. Configurations in the red shaded area of the graph in 

Figure 24 had balanced field lengths (BFL) exceeding 11,000 ft; the sized configuration’s BFL was 

9,770 ft. Its all-engines-operating (AEO) and FAR 25 takeoff field lengths were 9040 and 10,400 ft, 

respectively. Therefore, while sizing based on FAR25 field length would reduce margin relative to the 

11,000 ft limit, doing so would yield the same (sized) configuration in terms of T/W and W/S. Table 4

details the sized configuration’s mission performance according to individual segments.

TABLE 4.—SIZED N+3 CONFIGURATION MISSION PERFORMANCE

Phase Initial alt 

(ft)

Initial 

Mach

Initial wt. 

(lb)

Phase dist 

(nmi)

Phase time 

(min)

Phase fuel 

(lb)

Avg. L/D Avg. SFC 

[(lb/hr)/lb]

1. Warm-up and takeoff 0 0 285,000 ------- ----- 851 ------ ------

2. Climb to 10,000 0 0.381 284,149 8 2 2,913 10.6 1.08

3. Level accel 10,000 0.453 281,236 4 1 951 10.5 1.12

4. Climb and accel 10,000 0.665 280,285 24 3 3,892 10.5 1.20

5. Transonic accel 24,700 0.900 276,393 10 1 1,239 10.3 1.26

6. Supersonic accel 24,700 1.11 275,154 40 3 4,083 11.5 1.28

7. Supersonic cruise 40,500 1.60 271,116 4,447 293 104,127 8.71 0.881

8. Level decel 53,600 1.60 166,989 40 3 147 9.24 0.105

9. Descend and decel 53,600 0.900 166,842 77 12 921 10.8 1.72

10. 5-min loiter 5,000 0.343 165,921 ------- 5 902 10.9 0.707

11. Descend to land 5,000 0.342 165,019 ------- 4 440 11.2 1.21

12. to 18. Diversion and reserves 0 ------- 164,579 200 68 12,849 ------ ------

Mission totals ------- ------- ---------- 4,850 395 133,270 Range (nmi)

4000

4100

4200

4300

4400

4500

4600

4700

4800

4900

100

95

90

85

80

0.325

0.300

0.350

0.375

0.400

W/S 

(psf)

T/W

BFL = 11,000 ft

Fuel Volume 

Constrained

Sized Vehicle

As-Drawn Vehicle
NASA/CR—2010-216796 16

The sensitivity of the sized N+3 vehicle’s range to changes in maximum takeoff weight (MTOW) is 

illustrated in Figure 17. The blue line on the figure illustrates the nearly linear relationship between 

MTOW and range. Each 1000 lb. of additional weight yields approximately 11.2 nmi of added range; this 

translates to 89.3 lb per mile. The red line and right-hand axis on Figure 17 denote passenger-miles flown 

per lb. fuel on each mission, an efficiency metric which decreases with additional weight. The 

configuration was re-sized in accordance with the fuel volume and takeoff field constraints to produce the 

results in Figure 17. Extrapolating the red (efficiency) line on the plot suggests that ranges up to 

5000 nmi, which would correspond to a MTOW of approximately 320,000 lb, are possible while meeting 

or exceeding N+3’s minimum efficiency goal of 3.5 pax*nmi per lb. fuel.

Reducing the maximum allowable takeoff field, which was originally 11,000 ft during sizing, has no 

impact until the requirement falls below the sized configuration’s takeoff field length of 9.970 ft. As the 

field performance gets more restrictive, it requires increasing the thrust-to-weight ratio and decreasing 

wing loading, as shown in Figure 18. Both of these lead to inefficient cruise, decreasing the achievable 

mission range as illustrated in Figure 19. This analysis was performed with a constant MTOW of 

285,000 lb.

Figure 17.—Range and efficiency versus MTOW.

Figure 18.—Sizing design space with takeoff field contours.

Range (nmi)

4000

4100

4200

4300

4400

4500

4600

4700

4800

4900

10

9

9

85

8

0.30

0.32

0.35 0.37 0.40

W/S

(psf)

T/W

BFL = 11,000 ft

10,000 

9,000 

8,000 

Fuel Vol.

4300

4400

4500

4600

4700

4800

4900

5000

260 270 280 290 300 310

MTOW (x1000 lb)

Range (nmi)

3.5

3.6

3.7

3.8

3.9

4.0

4.1

4.2

Pax * nmi / (lb fuel)

Range

Efficiency
NASA/CR—2010-216796 17

Figure 19.—Range versus takeoff field length for MTOW = 285,000 lb.

Figure 20.—Technology impacts to sized configuration mission range.

Finally, the sensitivities of the sized N+3 configuration’s range to changes in empty weight and 

specific fuel consumption (SFC) were evaluated by perturbing each of those quantities by 5 percent and 

resizing the vehicle to the same constraints. The results of this study are summarized in Figure 20. A 

5 percent increase or decrease in SFC resulted in a range of 4370 or 4940 nmi, respectively. The resultant 

range due to a 5 percent increase or decrease in empty weight was 4440 or 4880 nmi, respectively. The 

corresponding, local sensitivities are thus –57 nmi per percent increase in SFC, and –44 nmi per percent 

increase in empty weight.

3.2.3 RCD Model Integration with Technology Inputs (WBS 3.2.2)

3.2.3.1 Description

The same N+3 vehicle model used for sizing also incorporated analysis components specific to some 

of the technologies that will enhance the vehicle’s performance. Namely, GE’s VCE and the aerodynamic 

benefits of NLF were represented in the sizing RCD model itself.

0

500

1,000

1,500

2,000

2,500

3,000

3,500

4,000

4,500

5,000

Baseline +5% SFC -5% SFC +5% We -5% We

Range (N Mi)

4,35

4,40

4,45

4,50

4,55

4,60

4,65

8,00 9,000 10,000 11,000

BFL (ft)

Range (nmi)
NASA/CR—2010-216796 18

Figure 21.—NLF schematic.

RCD integration of the FLF4-L1 engine was achieved by including a cycle deck provided by GE 

directly in the mission performance component of the model. Also, takeoff thrust was evaluated based on 

that deck and accounted for in the takeoff field performance analysis within the sizing model. The effects 

of NLF were captured using a friction drag analysis that supported both laminar and turbulent flow 

assumptions on all surfaces, top and bottom. Transition points, illustrated schematically in Figure 21, 

were varied parametrically and the aircraft’s friction drag based on appropriately weighted sums of 

laminar and turbulent skin friction coefficients.

3.2.3.2 Results

Results of this task are described in the following Section 3.2.4. 

3.2.4 Technology Benefit/Impact Assessment (WBS 3.2.4)

3.2.4.1 GE VCE

3.2.4.1.1 Description

Performances of both the initial and sized configurations were evaluated using GE’s VCE, which 

helps the N+3 vehicle meet its goals in two ways. First, it is lighter than a similarly-powered turbofan. 

This reduces the aircraft’s empty weight fraction, thereby permitting an equal increase in fuel fraction and 

longer range, compared to a vehicle designed for a conventional engine. 

3.2.4.1.2 Results

GE’s VCE impact on range performance was investigated by sizing an alternate configuration, based 

on a GE turbofan engine, to the same fuel volume and takeoff constraints. A comparison of the sized N+3 

configuration, with its VCE, and the turbofan-based alternate is summarized in Table 5.

TABLE 5.—VCE VERSUS TURBOFAN SIZED CONFIGURATION PERFORMANCE

VCE Alternate Turbofan 

Range 4640 nmi 4400 nmi

T/W (total thrust) 0.34 (98,000 lb) 0.38 (108,000 lb)

W/S (wing area) 92 (3090 ft2) 94 (3040 ft2)

Empty weight 129,600 lb 131,500 lb

Mean cruise SFC (cruise fuel) 2.32% more than AT 2.27% less than VCE

BFL 9,770 ft 8,720 ft

Time to M0.9 / M1.6 (with AB) 5.8 min / 10.1 min 4.9 min / 8.1 min

As shown in the table, the VCE yields greater range compared to the turbofan, 4640 nmi versus 4400, 

primarily due to the increased weight of the turbofan. Even though the VCE does cruise at a slightly higher 

SFC, fixing the GTOW (285,000 lb) resulted in more fuel available for the corresponding configuration. In resizing for the turbofan, the consequent weight gain was partially offset by shrinking the wing slightly (W/S 

increased from 92 to 94 psf), which also improves cruise efficiency slightly. However, further increases in 

W/S result in a net loss of range due to reduced fuel volume. The turbofan's power does reduce time to Mach 

0.9 and 1.6 by 1 and 2 min, respectively, and reduce BFL by approximately 1000 ft, but the VCE more 

effectively helps the N+3 vehicle meet its range performance goal.

Transition Laminar 

Flow

Turbulent 

Flow
NASA/CR—2010-216796 19

Figure 22.—Time to Mach with and without thrust augmentation.

Figure 23.—Thrust augmentation (TA) impact on sizing mission time and range.

Mission performance was evaluated using the thrust augmentation technology during the climb and 

acceleration segments of the sizing mission (Figure 15, segments 2 to 6). Comparing this performance 

with that attained using max dry power showed that thrust augmentation reduced the time from takeoff to 

beginning of cruise by over 50 percent. As shown in Figure 25, thrust augmentation enables the N+3 

vehicle to reach Mach 1.6 10.1 min after takeoff. Doing so via the same climb profile would take 

22.6 min using max dry power. Similarly, the aircraft reaches Mach 0.9 (the beginning of its transonic 

acceleration phase) in 5.8 min using the thrust augmentation and 11.2 min without them.

The impact of employing thrust augmentation during climb and acceleration on overall range and 

mission fuel is illustrated in Figure 23. An analogous mission flown entirely on max dry power takes 16 

min longer (340 versus 324 min.) and reaches 172 miles farther (4812 versus 4640 nmi) than the sizing 

mission as flown (i.e., with thrust augmentation). The 3.7 percent increase in range is attributable to fuel 

savings achieved by limiting operations to max dry power.

In terms of operational flexibility, the more important benefit of thrust augmentation technology is 

that it allows the level transonic acceleration to be moved above commercial jet traffic if necessary. This 

is exhibited in Figure 24, which shows the time necessary to reach altitudes up to 40,000 ft by following 

340 324

0

60

12

18

24

30

36

w/o TA w/ TA

Total Mission Time (min.)

Excluding Diversion

4812 464

0

100

200

300

400

500

w/o TA w/ TA

Range (nmi)

Excluding Diversion

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

0 5 10 15 20 25

Time (min)

Mach

with Thrust Aug.

without Thrust 

• 5.8 min to M0.9

• 10.1 min to M1.6

• 11.2 min to M0.9

• 22.6 min to M1.6
NASA/CR—2010-216796 20

Figure 24.—Time to altitude, with and without thrust augmentation.

Figure 25.—Fuel to transonic acceleration, with and without thrust 

augmentation.

the sizing mission profile to 10,000 ft and climbing at Mach 0.9 thereafter. As indicated by the green line 

on the graph, the N+3 vehicle with thrust augmentation activated reaches 40,000 ft in only 22 min. 

Conversely, absent the thrust augmentation, the vehicle is effectively forced to cruise-climb starting from 

near 32,000 ft. A direct consequence of this is that, as the transonic acceleration is delayed to 32,000 ft or 

higher, the climb phase accounts for a substantially larger fraction of mission time and available fuel, as 

illustrated in Figure 25. The plot on that figure depicts the total fuel consumption, from takeoff through 

subsonic climb, as a function of the altitude at which acceleration takes place. As the time and fuel spent 

during the subsonic climb grow to dominate the mission, the ultimate effect is to compromise total 

mission range. This undesirable result is shown on the plot of mission range versus transonic acceleration 

altitude in Figure 26.

These results demonstrate that, if supersonic flight were sometimes restricted by ATC (air traffic 

control) limitations until above most subsonic traffic, thrust augmentation is necessary to do so without 

compromising the N+3 vehicle’s speed and range. This is particularly important in light of the airport and 

airspace analysis performed by LM Transportation and Security Solutions (to be described in Section 

3.4.2), which showed that the greatest potential for conflicts between N+3 vehicles and existing subsonic

0

4

8

12

16

20

24 28 32 36 40

Transonic Acceleration Altitude (x1000 ft)

Fuel to Trans Accel (x1000 lb)

w/ Thrust Aug

w/o Thrust Aug

0

10

20

30

40

0 5 10 15 20 25 30

Time (min)

Altitude (x1000 ft)

w/ Thrust Aug

w/o Thrust Aug
NASA/CR—2010-216796 21

Figure 26.—Sizing mission range, with and without thrust augmentation.

traffic exists during transonic acceleration and supersonic climb. The VCE can minimize this risk by 

moving these segments above existing traffic while still enabling the N+3 vehicle to meet its range goal.

3.2.4.2 NLF

3.2.4.2.1 Description

Experimentation by Bill Saric and Helen Reed from Texas A&M showed that the use of Distributed 

Roughness Elements (DREs) can delay flow transition to produce significant regions of swept wing

laminar flow. The impact of this technology and unswept NLF on the sized N+3 configuration was 

investigated by parametrically varying the transition point, in terms of percent of chord, in analyzing the 

aircraft’s friction drag. For each aircraft component (wing, fuselage, etc.), skin friction coefficients were 

computed for both laminar and turbulent flow. Each component’s total skin friction and drag coefficients 

were calculated using a sum weighted according to the assumed transition point.

3.2.4.2.2 Results

The effects of laminar flow on the N+3 vehicle’s friction drag coefficient and total drag polars is 

visible in Figure 27 and Figure 28, respectively. Both were evaluated at a cruise condition of Mach 1.6 

and 50,000 ft altitude. Of particular interest was the improvement in lift-to-drag ratio (L/D) attained 

through NLF. Without this technology, the N+3’s cruise CL at 50,000 ft during its sizing mission is 

0.145. As highlighted on Figure 28(b), inducing laminar flow over 40 percent of chord length leads to an 

L/D increase of 2, which would offset the drop in lift-to-drag ratio (L/D) observed as a result of shaping 

for sonic boom (described previously in Section 3.2.2.1). Therefore, this technology is critical in ensuring 

that a vehicle shaped to comply with the N+3 noise goal of 70 PLdB can also achieve the sized 

configuration’s range of over 4600 nmi.

The performance benefits of friction drag reduction achieved through NLF are summarized in Figure 

29 and Figure 30. The first figure is a graph of mission range and passenger efficiency, again defined as 

passenger-miles per lb. Fuel, as a function of the flow transition point with GTOW fixed at 285,000 lb

The solid blue line shows the increase in mission range if the sized configuration’s T/W and W/S are 

maintained at 0.34 and 0.92, respectively, while the dotted blue line depicts the additional range benefit 

that would result from re-sizing the configuration to the same constraints. In the latter case, the range 

benefit of additional fuel volume outweighs the detriment of more wetted area, which is lessened by 

maintaining laminar flow. This is evidenced by the slightly increased wing size (reduced W/S). Finally, 

the solid and dashed red lines show the increase in passenger efficiency, again corresponding to 

maintaining and resizing T/W and W/S, respectively.

4200

4300

4400

4500

4600

4700

4800

4900

24 28 32 36 40

Transonic Acceleration Altitude (x1000 ft)

Range (nmi)

w/ Thrust Aug

w/o Thrust Aug
NASA/CR—2010-216796 22

NLF also has the effect of reducing the fuel, and consequently the gross takeoff weight, of a fixedlength mission, as shown in Figure 30. The dark and light blue lines on the plot correspond to the sized 

configuration’s GTOW and fuel weight, respectively, for its 4640-nmi mission. The red line depicts the 

consequent increase in efficiency, which is attributable entirely to the decrease in fuel required.

Figure 27.—Cruise friction drag coefficient for 3

levels of natural laminar flow.

Figure 28.—Cruise drag polars for 3 levels of natural laminar flow

-0.1

0.0

0.1

0.2

0.3

0.00 0.01 0.02 0.03 CD

CL

0% Chord

40% Chord

80% Chord

0

4

8

12

16

0.0 0.1 0.2 0.3 CL

L/D

0% Chord

40% Chord

80% Chord

(a) Drag Polars (b) Lift-to-Drag Ratio

Δ(L/D) = 2

CD0 @ Mach 1.6 / 50,000 ft

0.000

0.002

0.004

0.006

0.008

0.010

0% 40% 80%

Flow Transition (% Chord)
NASA/CR—2010-216796 23

Figure 29.—Range and efficiency [pax*nmi / lb fuel] benefits of natural 

laminar flow—constant GTOW.

Figure 30.—Weight savings from natural laminar flow—constant range.

3.2.5 Quantified Analysis (WBS 3.2.3)

3.2.5.1 Description

The sonic boom ground signatures of the N+3 aircraft were originally computed by modeling shock 

waves as either thin shocks or as simple “1/P Taylor structure” shocks. In an additional study by Wyle 

Laboratories and Penn State University, the boom signatures were completed by calculation of actual 

shock structures accounting for molecular relaxation effects, and the variations associated with 

atmospheric turbulence. The effect of lateral offset of the boom, i.e., propagation off track as well as 

under track, was examined. The key result of this analysis was perceived loudness, PLdB, of the booms 

accounting for those real world effects.

Three near-field signatures provided by LM to Wyle and PSU and served as the starting point for this 

analysis. Specifically, they were target (desired) signatures derived from SEEB corresponding to loudness 

levels of 80, 75, and 70 PLdB in accordance with N+3 goals. The first two were undertrack signatures 

while the last one was derived off-track at 40° azimuth. Figure 31 shows those signatures, normalized as 

F-functions. The shorter names noted in the figure captions (“denoted Sig1,” etc.) are to refer to these 

signatures in this report. Flight conditions for all three are Mach 1.6 at 48,000 ft in the standard 

atmosphere (Ref. 6).

0

5

100

15

20

25

300

0 20 40 60 80

Flow Transition (% Chord)

Weight (x1000 lb)

0

1

2

3

4

5

6

Efficiency (Pax*Nmi) / (lb fuel)

GTOW

Mission Fuel

Efficiency

GTOW for 4640 N. 

0

1000

2000

3000

4000

5000

6000

7000

8000

0% 20% 40% 60% 80%

Flow Transition (% Chord)

Range (N mi)

0

1

2

3

4

5

6

7

8

Efficiency (Pax*Nmi)/(lb fuel)

Range (T/W, W/S constant)

Range (Resized)

Efficiency (T/W, W/S const.)

Efficiency (Resized)

Range @ 285 klb

T/W=0.34

W/S=92 

T/W=0.30

W/S=86 
NASA/CR—2010-216796 24

Figure 31.—F-functions for three configurations.

Figure 32 and Figure 33 show nominal ground booms. These were computed using PCBoom3, 4 for 

those flight conditions. Figure 32 shows booms with thin shocks, while Figure 33 incorporates 1/P Taylor 

shock structures. Both include a ground reflection factor of 1.9. Note that the individual shocks in the Ffunctions are still distinct: they have not coalesced. Thickening (Figure 33) smoothes them so that the 

ground booms have the appearance of ramp signatures. The Perceived Loudness levels should be 

regarded as qualitative, since the frequency content of a Taylor shock is not the same as that of a 

relaxation shock.
NASA/CR—2010-216796 25

Figure 32.—Thin shock ground booms (no thickening).
NASA/CR—2010-216796 26

Figure 33.—Nominal ground signatures from PC-Boom, 

1/P Taylor shock structures.
NASA/CR—2010-216796 27

Three propagation effects were considered and are addressed in Sections 3.2.5.2, 3.2.5.2.2 and 

3.2.5.2.3:

• Shock structures due to molecular relaxation atmospheric absorption

• The effects of turbulence

• The effects of off-track propagation.

3.2.5.2 Results—Effects of Molecular Relaxation

3.2.5.2.1 Effects of Molecular Relaxation

Loudness of sonic booms (Ref. 7) is determined primarily by the structure of the shock waves, which 

contain the high frequency audible components. The structure is governed by a balance between nonlinear 

steepening and molecular absorption, quantified by the Burgers equation (Ref. 8). PCBoom6 (Ref. 9)

contains a Burgers solver, which uses a mixed time and frequency domain algorithm similar to that 

devised by Anderson (Ref. 10). Molecular absorption is defined by the current ANSI/ASA standard

(Ref. 11).

The three signatures have been processed by PCBoom6's Burgers solver to obtain ground signatures 

for two humidity conditions: 50 percent, representing a typical low-absorption condition that occurs most 

of the time, and 5 percent, representing dry, high-humidity conditions. The analysis procedure is to first 

compute the thin-shock solution at 47,000 ft (1000 ft below the flight altitude), then begin the Burgers 

calculation. The reason for beginning the calculation a short distance away from the flight path is to avoid 

anomalies associated with the Anderson-like algorithm in regions of very high wave pressure. 

Figure 34 and Figure 35 illustrate the booms for the 50 and 5 percent humidity cases. Note that their 

appearance is similar to the Taylor-thickened booms: the same shape as the thin shock solutions (Figure 

33) but with the shocks smoothed. Table 6 summarizes the loudness of the three booms and three types of 

thickening. Booms under dry (high absorption) conditions are considerably less loud, by about 8 dB for 

these examples, than under moist (low absorption) conditions. The nominal 1/P Taylor structure booms 

fall about midway between the extremes.

TABLE 6.—SUMMARY OF LOUDNESS, PLdB

Shock Type Sig1 Sig2 Sig3

Taylor 89.4 82.9 76.3

50% RH 94.1 87.6 82.7

5% RH 85.8 79.7 74.6
NASA/CR—2010-216796 28

Figure 34.—Ground Booms at 50 percent relative humidity.
NASA/CR—2010-216796 29

Figure 35.—Ground booms at 5 percent relative humidity.
NASA/CR—2010-216796 30

3.2.5.2.2 Effects of Atmospheric Turbulence

Atmospheric turbulence is known to distort sonic booms. Figure 36 is a classic NASA measurement 

result, showing booms measured under calm and turbulent conditions (Ref. 12). There is a 

considerable body of literature and analysis on the specific effects of turbulence on booms. Early 

predictive models tend to be statistical in nature, which makes it difficult to the estimate of the effects on 

loudness.

Empirical filters have, however, recently been developed for the effect of individual realizations of 

turbulence (Refs. 13 to 15), the loudness of which can be computed. Ten such filters have been applied to 

each of the relaxation-thickened booms in Figure 37 and Figure 38. Figure 37 shows two of these 

applications, a “peaking” and a "rounding" instance applied to Sig 1, 50 percent humidity. The original 

boom, from Figure 37(a), is drawn in blue and the turbulent-distorted boom is drawn in black. 

Table 7 shows the Perceived Loudness for all six ground signatures (Sig1, Sig2 and Sig3, each at 

50 percent and 5 percent relative humidity) and the ten realizations. The minimum, maximum, 

average and standard deviation (sigma) are shown for the ten turbulence realizations. Note that 

the average of each of the turbulent-distorted booms is always less than that of the corresponding 

nondistorted boom. The range of loudness from minimum to maximum for each boom is up to 10 dB. It 

should be noted that the turbulence filters are based on flight test data that, in general, had less distortion 

than seen in other tests such as that shown in Figure 36. While it is expected that the average loudness of 

booms would follow the trend of small reduction (or possibly no reduction), variations of individual 

booms may be greater than computed here.

Figure 36.—Distortion of sonic booms by turbulence.

Figure 37.—Sig1, 50 percent humidity, with peaking and rounding turbulence.
NASA/CR—2010-216796 31

Figure 38.—Signatures, peak overpressures and perceived loudness across the carpet, Sig1 

50 percent RH.
NASA/CR—2010-216796 32

TABLE 7.—LOUDNESS FOR ALL GROUND SIGNATURES, VARIOUS TURBULENCE REALIZATIONS

Signature and relative humidity

Turbulence Sig1, 50% Sig1, 5% Sig2, 50% Sig2, 5% Sig3, 50% Sig3, 5%

None 94.10 85.81 87.56 79.72 82.70 74.64

1 92.28 85.50 85.77 79.58 81.13 73.37

2 87.16 81.01 80.99 75.60 74.85 67.91

3 95.01 88.50 89.19 82.90 84.76 77.62

4 87.79 79.83 81.43 74.19 76.78 68.26

5 92.40 83.43 84.80 76.65 81.04 72.62

6 92.99 86.00 85.20 78.99 81.87 74.48

7 94.17 86.46 86.60 79.95 83.33 75.55

8 94.43 86.34 86.36 79.61 83.49 75.71

9 94.55 86.99 87.11 80.25 83.88 75.91

10 93.73 85.71 85.80 78.88 82.79 74.96

Minimum 87.16 81.01 80.99 74.19 74.85 67.91

Maximum 95.01 88.50 89.19 82.90 84.76 77.62

Average 92.40 84.95 85.29 78.61 81.33 73.60

Sigma 2.65 2.59 2.36 2.41 3.03 3.07

3.2.5.2.3 Off-Track Booms

Off-track boom propagation was predicted by using the under-track F-functions at lateral azimuths, in 

10° increments, out to lateral cutoff of 49.8°. Physically, roll angles off-track generally have a substantial 

reduction in impulse due to lift reducing with the cosine of the roll angle degrees from under-track. 

Therefore actual off-track sonic boom generally decreases in impulse; however, shaped booms (non Nwave) can decrease or increase in loudness depending on how well their shaping is maintained at other 

roll angles (a very configuration-specific dependency). The following indicates just one difficulty of 

maintaining good off-track shaping: an F-function that works under-track may work (Sig3) or may not 

work off-track (Sig1 and Sig2). Figure 41 shows the ground signatures for Sig1 50 percent humidity. The 

peak overpressure and perceived loudness of each boom is presented in Table 8. Note that the peak 

pressure progressively decreases across the carpet (due to the longer propagation distance), as expected, 

but that loudness increases. The reason for this is apparent from the signatures in Figure 38: the off-track 

positions, with their longer propagation distances, are beyond the design point for the under-track ground 

signature. The individual steps, which did not coalesce under track, progressively coalesce off-track. At 

the carpet edge most of the steps at the bow of Sig1 and all of the steps at the tail have coalesced into 

single shocks. The same coalescence occurs for Sig2, with peak pressure and perceived loudness shown 

in Table 9. On the other hand, Sig3 was intentionally designed with a lower slope in its ramp because it 

was shaped for the roll angle furthest off-track, which also results in good shaping at lower coalescence 

everywhere else. Table 10 shows the peak pressure and perceived loudness of Sig3 across the track. Both 

decrease toward the carpet edge. Note that this is a coalescence issue, independent of the humidity and 

shock structure. 

The off-track calculation is approximate because a single under-track F-function was used at all 

azimuths, while a physical configuration’s F-function changes off-track both from the cosine drop-off in 

overall lift and from asymmetric changes in the distribution of lift and volume at each roll angle. It is 

important that optimal designs be developed off-track as well as under-track. 

TABLE 8.—BOOM MAGNITUDE ACROSS CARPET, Sig1 50 PERCENT HUMIDITY

φ

(°)

Pmax

(psf)

Loudness

(PLdB)

0 1.56 94.10

10 1.54 95.16

20 1.50 95.63

30 1.43 96.84

40 1.31 98.51

49.8 1.03 99.48
NASA/CR—2010-216796 33

TABLE 9.—BOOM MAGNITUDE ACROSS CARPET, Sig2 50 PERCENT HUMIDITY

φ

(°)

Pmax

(psf)

Loudness

(PLdB)

0 1.53 85.91

10 1.52 85.89

20 1.48 85.69

30 1.41 86.68

40 1.29 90.44

49.8 1.02 96.03

TABLE 10.—BOOM MAGNITUDE ACROSS CARPET, Sig3 50 PERCENT HUMIDITY

φ

(°)

Pmax

(psf)

Loudness

(PLdB)

0 0.97 82.70

10 0.96 81.11

20 0.93 80.25

30 0.88 78.65

40 0.81 76.79

49.8 0.64 76.50

3.2.5.3 Conclusions

Ground booms accounting for molecular relaxation shock structure, at 50 and 5 percent relative 

humidity, have been computed for three nearfield signatures. Thin shock and 1/P Taylor shock signatures 

were also presented for comparison. Loudness results are presented and discussed. Humidity has a clear 

effect on the loudness of booms. Under dry conditions (5 percent RH) the sample booms are about 8 dB 

less loud than under moist/typical (50 percent RH) conditions. Loudness of the nominal 1/P Taylor shock 

booms is about midway between moist and dry relaxation shock booms.

The effect of turbulence was evaluated by applying filters to the under-track booms for ten 

realizations of turbulence. This turbulence set yielded a slight decrease in average loudness, but with 

individual variations of up to plus or minus 5 dB. The sample turbulence set is modest in amplitude, and 

real-world turbulent variations may be bigger. The turbulence model used (Ref. 13 to 15) should be 

expanded to cover a wider range of atmospheric conditions.

Off-track propagation was examined, assuming the under-track F-functions applied at all azimuths. 

For booms Sig1 and Sig2 the longer propagation distances off-track resulted in partial coalescence of 

shocks, losing the shaping effect and yielding higher perceived loudness than under track. 

Coalescence did not occur for the lower amplitude Sig3, so its loudness did decrease off-track across 

the carpet. These off-track trends used to same F-function for all roll angles, but a real configuration’s Ffunction changes considerably since lift drops off with cosine of roll angle and the distribution of volume 

and lift changes at every roll angle. Still, it illustrates that low boom shaping must be developed off-track 

as well as under-track. Finally with regard to low boom, variations in humidity, and mild turbulence were 

investigated. Variations of +2 PLdb (turbulence peak) to –15 PLdB (turbulence and humidity rounding) 

were seen relative to 50 percent humidity Burgers analysis. Strong turbulence has changed N-wave 

average loudness as much as –6 PLdB along with greater scatter in loudness. While humidity variations 

are on average close to the 50 percent humidity level, there are other uncertainties (atmospheric winds, 

greater rounding/over-prediction of N-waves, ground reflection arrival time rounding) that seem to all 

push in the quieter direction. Since this result comes to us at the end of the program, hereafter we will 

carry a –4 to –16 PLdB uncertainty in “real” loudness relative to molecular relaxation loudness 

calculations (or more approximately +2 to –11 PLdB uncertainty from 1/dp Taylor loudness).
NASA/CR—2010-216796 34

3.3 Final Configuration Definition (WBS 3.1.3)

3.3.1 Loft

3.3.1.1 Description

Sizing and mission performance from Section 3.2.2.2 drove the final configuration loft. The drooped 

nose in the initial configuration reduced the vehicle length, but this was not enough to overcome the 

added complexity of the shape. A straighter nose may cause difficulties with pilot visibility, but this will 

be solved with the use of Synthetic Vision as the design moves closer to reality. As a result, the final 

configuration is shown with a reduction in the drooped nose.

3.3.1.2 Results

The RCD process was also used to size the final configuration to meet the mission, as detailed in 

Section 3.2.2.2. The final configuration was sized to have a gross takeoff weight was 285,000 lb, with a

thrust to weight ratio of 0.34 and a wing loading of 92 psf. The final configuration with all included 

changes is shown in Figure 39.

The final wing, canard and tails were optimized to reduce the sonic boom with minimal aerodynamic 

efficiency reduction. The Rapid Conceptual Design (RCD) model and results for this boom optimization 

are detailed in Section 3.2.2.1. This optimization resulted in a reduction of the 1.38 psf max shock to 

0.29 psf, and a reduction in L/D from 10.1 to 8.0.

Figure 39.—Final configuration three-view.
NASA/CR—2010-216796 35

3.3.2 Fuel Arrangement

3.3.2.1 Description

Initially, the fuel arrangement was performed by only placing fuel within the final configuration’s 

wings. However, this fuel placement was ultimately inadequate and fuel had to be placed in the forward 

fuselage and V-tail. The internal fuel assessment was performed in CATIA V5.

3.3.2.2 Results

The fuel volume was found by examining space for fuel tanks in the vehicle, as shown in Figure 40.

The fuel tanks included a small tank in the fuselage to allow for some trimming of the CG through fuel 

placement. The total fuel available is 133,900 lb.

3.3.3 Aerodynamics

3.3.3.1 Description

Cruise drag of the sized N+3 configuration was consisted of 3 induced, profile, and wave drag. These 

items are plotted in Figure 41. The profile drag does not include shaping for low boom.

3.3.3.2 Results

The total drag polar at Mach 1.6 and 50,000 ft altitude, a flight condition near mid-cruise, is shown in 

Figure 42(a). Figure 42(b) depicts a breakdown of profile drag by component (total profile drag at this 

condition is CD0 = 0.0843). Again, these results do not reflect shaping for low boom, as it is anticipated 

that the N+3 technologies described in Sections 9.0 and 10.0 will offset the aerodynamic penalties from 

doing so. Finally, the takeoff drag polars in and out of ground effect (IGE, OGE) are shown in Figure 43.

Figure 40.—Fuel packaging for final configuration.

Fore Fuselage

3,065 gal; 20,840 lb; 16%

Aft Fuselage

4,595 gal; 31,240 lb; 23%

Inboard Wing

8,292 gal; 56,380 lb; 42%

Outboard Wing

979 gal; 6,660 lb; 5%

Inverted V

2,765 gal; 18,810 lb; 14%

Total Fuel = 19,700 gal (133,900 lb)

• Long fuselage for low boom provides fuel volume forward of cabin, helps 

achieve greater range

• Fuel located for limited CG travel
NASA/CR—2010-216796 36

Figure 41.—N+3 cruise drag components.

Figure 42.—Aerodynamic performance at Mach 1.6, 50,000 ft.

Fuselage 

27%

Wing 34%

Fin 5%

Inverted V 

16%

Canard 3%

Nacelles 10%

Excrescences 5% 

-0.1

0.0

0.1

0.2

0.3

0.00 0.01 0.02 0.03

CD

CL

(a) Total Drag Polar (b) Profile Drag Breakdown

-0.05

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.000 0.005 0.01 0.015 0.020

CDi

CL

Mach 1.6

Mach 1.4

Mach 1.2

0.000

0.004

0.00

0.01

0.016

0.02

0.0 0.4 0. 1.2 1.6

Mach

CDo

Alt = 0

Alt = 20k'

Alt = 40k'

Alt = 60k'

0.000

0.001

0.002

0.003

0.004

0.005

0.006

0.007

0.8 1.0 1.2 1.4 1.6

Mach

CDwave

(a) Induced Drag (b) Profile Drag

(c) Wave Drag
NASA/CR—2010-216796 37

Figure 43.—Takeoff drag polars.

Figure 44.—Sized N+3 configuration weights statement.

3.3.4 Weights

3.3.4.1 Description

Weights of the sized N+3 configuration were based on LM Aero parametrics.

3.3.4.2 Results

A summary weights statement of the sized N+3 configuration appears in Figure 44.

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

0.00 0.04 0.08 0.12 0.16 0.20

CD

CL OGE

IGE
NASA/CR—2010-216796 38

3.4 System of System Analysis (WBS 3.6)

3.4.1 Purdue

Following N+3 configuration sizing, LM provided aircraft performance data to Purdue University, 

which performed a system-of-systems analysis of the vehicle’s effects on a future civil air transport 

system (ATS). The goals of this study were to quantify the SSTs’ contributions to and impacts on the 

ATS as a whole.

3.4.1.1 Description

Purdue’s system-of-systems study predicted the effects of adding SST’s to an all-subsonic fleet, 

beginning in 2030 and occurring at a constant annual rate through 2050. The baseline fleet, to which 

SST’s were added, consisted of six subsonic aircraft, each in service in 2005 and representing a unique 

capacity class, ranging from 20 seats (Embraer ERJ145) to over 300 (Boeing 777–200ER).The model 

ATS included those routes in the WWLMINET Network (Ref. 16) that either originated or terminated in 

the U.S., thereby including 180 airports. Passenger demand was modeled by increasing 2005 levels by a 

fixed annual percentage. Different demand scenarios were created by varying the yearly rates of STT 

addition to the fleet (either 25 or 50) and passenger demand growth (from 1 to 5 percent in 1 percent

increments).

For each future scenario, defined by specific levels of passenger demand and SST availability, Purdue 

investigated two allocations of aircraft: one maximized ATS productivity (passenger * knots), the other 

block-hours saved by SST passengers. The latter was calculated by comparing the SST’s block time on 

each route it flew to that of a subsonic aircraft. Fleet-wide carbon dioxide (CO2) and nitrous oxide (NOx) 

emissions were computed to gauge SSTs’ environmental impacts on the ATS. Finally, fleet direct 

operating costs (DOC) were calculated based on estimates of each aircraft’s DOC.

3.4.1.2 Results

Initial studies established the trade space between SST performance benefits and environmental 

impact, such as that illustrated in Figure 45. This graph shows the maximum productivity achievable as 

an allowable limit of CO2 is increased; they are shown in billions of passenger*knots and billions of lb, 

respectively. The values listed correspond to year 2050 in a scenario in which aircraft were allocated to 

maximize productivity based on 3 percent annual passenger demand growth SST’s added at 50 per year 

starting in 2025.

Subsequent studies investigated the increase in fleet DOC attributable to the SST’s as well as the 

routes to which they were assigned and the number of passengers on them. Figure 46 illustrates this DOC 

difference for a scenario based on 3 percent annual passenger demand growth. The left-hand graph 

(Figure 46(a)) is a plot of DOC versus year for fleets in which SST’s are added at 25 and 50 per year 

starting in 2025; DOC growth of an all-subsonic fleet is also included for comparison. The increase in 

DOC attributable to the SST’s was calculated as the difference between the DOC of a fleet including 

them to that of the baseline (all-subsonic) fleet. Figure 46(b) shows this DOC increase versus year from 

2025 to 2050. The nonsmooth nature of the DOC increase is due to their being allocated in different 

numbers and to different routes as they comprise more of the ATS fleet.

Figure 46 depicts DOC in millions of 2005 dollars and includes estimates of crew, landing and 

navigation fees, maintenance, depreciation, interest, and insurance costs based on a 1995 NASA report

(Ref. 17). Given the sensitivity of DOC to the underlying assumptions and the volatility of individual 

costs (esp. fuel), further research is necessary to refine these cost estimates and gauge them against the 

value of passengers’ time saved by using SST’s. 

Figure 47 contains information about the use of SST’s in the air traffic system. The left-hand plot, 

Figure 47(a), shows the number of routes to which SST’s were allocated and the number of passengers on 

them for the same scenario (3 percent annual passenger demand growth). The left-hand axis corresponds 
NASA/CR—2010-216796 39

to the former, while the right-hand axis corresponds to the latter and was based on an assumption of 90 

passengers per flight. 

Finally, Figure 47(b) shows the average range of SST routes within the ATS. While the routes and 

ranges varied as demand grew and more aircraft were added to the system. The average range was 

consistently between 3,000 and 5,000 nmi. This indicates that aircraft meeting the N+3 range goal of 

4,000 nmi will likely be well-suited for incorporation into the civil air transport system.

Productivity [×109 pax-kn]

Figure 45.—ATS productivity versus carbon dioxide.

Figure 46.—DOC growth and increase attributable to SST’s.

1.0

1.1

1.2

1.3

1.4

1.5

1.6

1.7

1.8

1.9

1.00 1.05 1.10 1.15 1.20 1.25 1.30 1.35 1.40

 

$0

$5

$10

$15

$20

$25

$30

$35

$40

$45

2025 2030 2035 2040 2045 2050

Year

DOC Increase due to SST's

(2005 $M)

25 SST/yr

50 SST/yr

$250

$300

$350

$400

$450

$500

$550

$600

2025 2030 2035 2040 2045 2050

Year

ATS Fleet DOC (2005 $M)

25 SST/yr

50 SST/yr

<No SST's>

(a) DOC versus Year (b) DOC Increase due to SST’s

CO2 [×109 lb]
NASA/CR—2010-216796 40

Figure 47.—SST trips, passengers, and range within ATC.

3.4.2 Transportation and Security Solutions

The impact of N+3 Supersonic Transports (SST’s) on airport terminal and en-route environments was 

investigated by LM Transportation and Security Solutions (LM TSS). This study assessed SSTs’ impact 

on airport and airspace capacity requirements, their interaction with subsonic traffic during arrival, 

departure, and subsonic climb and descent phases, and consequent increases in air traffic complexity and 

controller workload.

3.4.2.1 Description

LM TSS’s studies were based on the RAMS Plus air traffic modeling and simulation environment, 

which simulates traffic from a macro-to-micro level (gate-to-gate movements) based on the flights, 

sectors, airports, airspace and air traffic, controllers, procedures that define an ATC environment. This 

tool enabled an assessment of 4–D trajectory based operations in a future ATC environment in which 

controllers maintained required separation standards. The study was conducted as follows:

1. Potential routes for N+3 SST operation were identified.

2. Target SST Introduction date and anticipated supersonic operations were based on traffic 

forecasts

3. Modeled en-route air traffic, airport and terminal area operations.

4. Analyzed demand, conflict, workload and complexity assessment metrics using the RAMS Plus 

simulation tool.

5. Assessed new processes and aircraft types in the future National Air Transportation System

(NAS).

3.4.2.1.1 SST Climb Profiles

The N+3’s climb profile is defined by segments 2 to 6 of the sizing mission described in Section

3.2.2.2. They are expected to climb in a similar manner to standard aircraft (Boeing 777, Airbus 340 etc) 

during the initial climb out from the runway to around 10,000 ft. On reaching 10,000 ft the aircraft will 

level off and undergo a first acceleration leg lasting around 80 sec during which the aircraft will 

accelerate from around 290 to 450 kn.

Following the initial acceleration phase, the aircraft will resume its climb and continue to accelerate 

through the lower/middle airspace to reach 535 kn at around 24,000 ft. There it will level off a second 

time and enter the second acceleration phase which lasts approximately 125 sec, during which the aircraft 

will accelerate up to 719 kn (Mach 1.09). Thereafter the aircraft will climb rapidly towards its expected 

0

5

10

15

20

25

30

35

40

45

50

2025 2030 2035 2040 2045 2050

Year

# SST Roundtrips

0

5

10

15

20

25

30

35

40

45

SST Passengers (x1000)

25 SST/yr

50 SST/yr

(a) SST Trips and Passengers (b) SST Trip Range

0

1,000

2,000

3,000

4,000

5,000

6,000

2030 2035 2040 2045 2050

Year

Avg SST Trip Range (nmi)

25 SST/yr

50 SST/yr
NASA/CR—2010-216796 41

cruise altitude of 61,000 ft, accelerating during the entire climb portion to reach its top speed of 917 kn

(Mach 1.6).

A comparative analysis was done to ensure that the N+3 profile generated by RAMS Plus is 

consistent with the one developed by LM-Aero. Figure 48 and Figure 49 show that the two profiles are 

closely calibrated. Also for the purposes of comparison a representative climb profile for subsonic 

transport (A320) is depicted. As can be seen in Figure 51, from the performance standpoint the trajectory 

of N+3 intersects with conventional aircraft as N+3 climbs to supersonic speed and final cruise altitude. 

This could potentially create additional conflict as the flight overtakes the slower, conventional aircraft at 

higher flight levels.

Figure 51 indicates that RAMS Plus profile is closely calibrated with conceptual vehicle designed by 

LM Aero. It also shows comparison between N+3 and subsonic climb profiles.

3.4.2.1.2 Supersonic Descent Profiles

The N+3’s descent phase corresponds to segments 9 to 11 of the sizing mission (Section 3.2.2.2). LM 

SST’s analysis was based on a constant rate of 3500 ft/min from the end of cruise (this rate was 

subsequently adjusted to 4000 ft/min to maximize mission range). The aircraft decelerates rapidly during 

the initial descent between the Cruise level and the entry to the transitional (middle) airspace around 

36,000 ft, decreasing from supersonic cruise at 719 to 470 kn around 36,000 ft. Below that level, the AST 

will decelerate at a rate similar to that of conventional aircraft.

Comparative analysis ensured that the N+3 profile generated by RAMS Plus was consistent with the 

sizing mission; Figure 50 and Figure 51 show they are closely calibrated. A representative descent profile 

for subsonic transport (A320) is also depicted for comparison.

Figure 50 indicates that RAMS Plus profile is closely calibrated with N+3’s mission profile. It also 

shows comparison between N+3 and subsonic descent profiles.

Figure 48.—N+3 climb profile.

0

10000

20000

30000

40000

50000

60000

70000

0 200 400 600 800 1000 1200 1400 1600 1800

Altitude (Ft)

Mission Time (s)

"LM/Aero AST Profile"

"RAMS Profile"

"RAMS A320 Profile"

0

200

400

600

800

1000

0 200 400 600 800 1000 1200 1400 1600 1800

Speed (Kts)

Mission Time (s)

"LM/Aero AST Profile"

"RAMS Profile"

"RAMS A320 Profile"
NASA/CR—2010-216796 42

Figure 49.—N+3 mission profile (climb phase).

Figure 50.—N+3 descent profile.

0

50

100

150

200

250

300

0 200 400 600 800 1000 1200 1400 1600 1800

Distance (NM)

Mission Time (s)

"LM/Aero AST Profile"

"RAMS Profile"

"RAMS A320 Profile"

0

10000

20000

30000

40000

50000

60000

70000

7800 8000 8200 8400 8600 8800 9000 9200 9400 9600

Altitude (Ft)

Mission Time (s)

"LM Aero AST Profile"

"RAMS Profile"

"RAMS A320 Profile"

A320 'Adjusted'

Arrival Time

0

100

200

300

400

500

600

700

800

900

1000

1100

7800 8000 8200 8400 8600 8800 9000 9200 9400 9600

Speed (Kts)

Mission Time (s)

"LM Aero AST Profile"

"RAMS Profile"

"RAMS A320 Profile"

A320 'Adjusted'

Arrival Time
NASA/CR—2010-216796 43

Figure 51.—N+3 mission profile (descent phase).

3.4.2.2 Potential SST Routes

With expected technology improvements in the environmental and fuel efficiency of the N+3 SST, it 

is anticipated that the aircraft is not constrained to fly only oceanic routes and can fly supersonic over

land. This makes the aircraft ideally suited for economically viable domestic routes and major 

international routes. Additionally, the multi-mach N+3 aircraft will be capable of flying at speeds of up to 

2 Mach, will have a maximum range of 6000 nmi, and can carry up to 200 passengers.

Research from the 1980 U.S. OTA report on the Impact of Advanced Air Transport Technology, 

Chapter 3 (Variables Affecting a Supersonic Transport Market) (Ref. 1) suggests that “an aircraft’s 

product is passenger (/ cargo) miles”. On this basis N+3 can improve productivity using faster aircraft 

with the same number of ‘seats’ but with significantly shorter flying time thus achieving higher passenger 

miles per hour than a conventional operation. Additional factors also come into play, particularly relating 

to economic or social factors for key city-pairs or flight operations. For the purpose of this study it was 

decided to constrain the route allocation based on the aircraft productivity measure of passenger miles per 

hour, without distinguishing different passenger value.

In order to evaluate the value of operations, a passenger factor was produced comparing the passenger 

loads for two N+3 operations (a return trip) against a one-way conventional operation to produce a 

passenger productivity factor. In addition to the time for a return trip, a sufficient turn-around window of 

40 min was included in the calculation for the N+3. 

The passenger factor for any city pair could be calculated as follows:

• PaxFactor = [[2* SS_time] + 40-min] * SS_Pax_load / [Conventional_Time] * 

Conventional_Pax_load

• PaxFactor = Passenger factor

• SS_time = Time taken by supersonic transport between city pair

• SS_Pax_load = Supersonic passenger load

• Conventional_time = Time taken by conventional aircraft between city pair

• Conventional_Pax_load = Conventional passenger load

• 40 min = Supersonic turnaround time

1900

1950

2000

2050

2100

2150

2200

8000 8200 8400 8600 8800 9000 9200 9400 9600

Distance (NM)

Mission Time (s)

"LM Aero AST Profile"

"RAMS Profile"

A320 not included 

(still in En Route phase)
NASA/CR—2010-216796 44

Figure 52.—Potential N+3 route network.

As shown in Figure 52, the city-pairs (domestic and international) with the best passenger factor were 

selected based upon the results of the calculations. Once the city pairs were identified, arrival and 

departure times were adjusted to allocate suitable supersonic arrival and departure banks based on local 

departure and arrival times. Twenty major U.S. Airports and 11 international destinations were identified 

for potential N+3 operations. Not all of the possible origin-destination pairs are viable for N+3 due to

shorter route distance. The productivity of N+3 reduces with the shorter route distance to below 

60 percent at 1000 nmi due to time lost on surface, climb and descent (Ref. 1). For example Los Angeles 

to Newark and San Francisco to Newark are suitable but operations between Los Angeles and San 

Francisco are not. For each viable city-pair, between four and six daily operations are included depending 

on the arrival/departure banking calculations (adapted to local times) and N+3 journey time.

3.4.2.3 Modeling Air Traffic Operations

3.4.2.3.1 Scenario

To generate the scenario, Aircraft Situation Display to Industry (ASDI) data from August 24, 2007, 

containing arrivals and departures from top 35 U.S. airports was used as a starting point. ASDI is a subset 

of the Enhanced Traffic Management System (ETMS) and contains flight data from all scheduled carrier 

and business jets. To create a traffic scenario representative of 2030 traffic levels, additional operations 

were added based on the 2 percent annual growth rate predicted in MITRE’s Fleet Forecast. The N+3

operations were included based on the methodology described above. The resulting traffic sample 

included 906 N+3 supersonic operations among the 44,000 total flights (subsonic and supersonic, 

domestic and international) on a single day considered in the study.
NASA/CR—2010-216796 45

Figure 53.—Sector adaptation for U.S. airspace in 2008.

3.4.2.3.2 Airspace and Procedures

The airspace considered for these experiments is based on the 2008 NAS adaptation and contained 

960 En Route sectors as can be seen Figure 53. No assumptions about future airspace changes or 

modifications are considered in the scope of this study, for example big airspace (Ref. 18) that aim to 

consolidate terminal area with en route arrival sectors. In the terminal area we assumed precision RNAV

(P-RNAV) and RNP procedures will be in use. The N+3 will be capable of executing P-RNAV and will 

fly great circle distance between city pairs. Area Navigation (RNAV) is a method of Instrument Flight 

Rules (IFR) navigation that allows an aircraft to choose any course within a network of navigation 

beacons, rather than navigating directly to and from the beacons. This can conserve flight distance, reduce 

congestion, and allow flights into airports without beacons.

It was assumed in the study that NextGen airport management technology will be in place to permit 

traffic to achieve the airport demand being forecasted, through a combination of precision navigation, 

RNAV sequencing and merge capabilities, and Time/trajectory Based Flow Management tools. Given 

that supersonic traffic performs like conventional traffic in the arrival and departure phases, it was 

concluded that no special procedures would be required in the airport and terminal regions.

3.4.2.3.3 Air Traffic Controllers

Each of the En Route control sectors has a single Radar (R-Side) Controller allocated in the 

simulation. Planning Controllers or Assistants are not taken into consideration in the studies.
NASA/CR—2010-216796 46

3.4.2.3.4 Separations

Each ATC Controller in the model is allocated 5 nmi separation requirements and flight level 

separation using RVSM separation standards. The RAMS Plus simulation tool provides a fully adaptable 

set of ATC rules that can be used to provide separation assurance within any airspace volume, as well as 

to manage ATM-based flow/sequence management in airspace and arrival/departure/airport systems. 

Controller rules are allocated to each and every control element in the system, and in the airborne phases, 

uses projected 4–D airspace tubes to predict when conflicts occur. A forward-chaining production rule 

expert system provides the tool with a set of potential resolution actions which have been tuned to 

represent generic Air Traffic Controllers. In the event that separation problems are identified in the 

simulation involving supersonic and nonsupersonic traffic, the model has been tuned to allow supersonic 

aircraft to continue unimpeded to mimic performance based services concept.

3.4.3 Results

Analysis was performed on an air traffic simulation based on 2030 projected traffic levels including

N+3 aircraft. Additional supersonic transports are added through 2050 would be expected to yield 

additional arrival and departure operations. The results are described on the basis of phase of flight.

3.4.3.1 Airport and Terminal Area Operations

Despite a slight increase in the number of airport operations due to the introduction of N+3 traffic, we 

do not anticipate that they will have any significant impact on the airport capacity as their performance is 

very similar to subsonic transport. Research in the design of N+3 conducted by LM Aero suggests that 

they will fall into either medium or heavy wake category for runway separation standards. They will not 

have to be procedurally separated from other subsonic traffic in the terminal area as it is anticipated that 

N+3 will be equipped with precision—RNAV and capable to execute optimized arrival and departure 

procedures in the terminal area. In the arrival phase, N+3 will exploit advanced NextGen sequence and 

merge capabilities and perform in the same way as conventional traffic. Similarly, in the initial departure 

phase (up to 10000 ft) the N+3 traffic will operate with similar characteristics as conventional aircraft.

Since a subsonic operation might be replaced by 2 or 3 N+3 operations (the journey time is close to 

half in many cases) the arrival and departure rates increase slightly over the 24-hr period, particularly 

during peak arrival and departure periods.

A sliding view of departure demand is shown in Figure 54 and Figure 55 for selected U.S. airports 

where supersonic operation was defined. In both figures, blue bars represent conventional aircraft and 

orange represent N+3 operations. For example, the departure rate shown at 10:40 represents the number 

of departures for the period from 10:40 to 11:40 Zulu. As can be seen in most cases, the introduction of 

N+3 aircraft will require some additional departure capacity at major U.S. airports during the peak hours 

of conventional operations. The general trend suggests that most N+3 departures are scheduled for early 

morning and will utilize unused airport capacity as during that time the demand for conventional 

operations is relatively low. An exception was Honolulu airport (HNL) where introduction of N+3 will 

require a higher departure capacity during the peak hours of conventional traffic.

Similar to departure demand, a sliding view of arrival demand is shown in Figure 56 and Figure 57

for selected U.S. airports where supersonic operation was defined. Again, the blue and orange bars in both 

figures represent conventional and N+3 aircraft operations, respectively. For example, the arrival rate 

shown at 10:40 represents the number of arrivals for the period from 10:40 to 11:40 Zulu. As can be seen 

in most cases, the introduction of N+3 aircraft will require minimal additional arrival capacity at major 

U.S. airports with exception of HNL where higher arrival capacity might be required.
NASA/CR—2010-216796 47

Figure 54.—Departure rate for selected U.S. airports. (Blue bars represent conventional aircraft and orange represent 

AST operations.)
NASA/CR—2010-216796 48

Figure 55.—Departure rate for selected U.S. airports. (Blue bars represent conventional aircraft and orange represent 

AST operations.)
NASA/CR—2010-216796 49

Figure 56.—Arrival rate for selected U.S. airports. (Blue bars represent conventional aircraft and orange represent 

AST operations.)
NASA/CR—2010-216796 50

Figure 57.—Arrival rate for selected U.S. airports. (Blue bars represent conventional aircraft and orange represent 

AST operations.)

3.4.3.2 Transition Airspace Operations

Departures: The most significant impacts are expected to be observed in the transition airspace—in 

particular during four key phases of the supersonic flight profile in the departure phase: 

Phase 1: At around 10000 ft the N+3 will level off and undergo the first acceleration phase (from 280 

to 400 kn in around 75 sec).

Phase 2: Following the initial acceleration. N+3 will climb to a second transition level around 

24000 ft. 

Phase 3: At 24000 ft the aircraft undergo their second acceleration phase (from 500 to 750 kn in 

around 130 sec).

Phase 4: Steep supersonic climb to the ultra high airspace (typical cruise levels between 58000 and 

61000 ft).

Figure 58 shows the route network analysis across different departure phases (identified using 

different colors) for N+3 aircraft departing from the NAS-CONUS using 4-D trajectories. The potential 

route network was described previously in Section 3.3.2.2 and assumes that great circle routings for N+3

aircraft are enabled by NextGen. Figure 59 shows route network for east coast. This figure suggests that 

the route network is well separated across Phases 1, 2 and 3, and departures from one airport do not 

interfere with the others. The result is consistent for even busy east coast multiplex airspace (multiple hub 

airports in close proximity, for example JFK, EWR, PHL, BOS and IAD).
NASA/CR—2010-216796 51

Figure 58.—AST route network analysis across four AST departure phases.

Figure 59.—AST route network for the East Coast.
NASA/CR—2010-216796 52

3.4.3.2.1 Initial Acceleration Around 10,000 ft

Phase 1 acceleration occurs early in the N+3 profile at an altitude of 10,000 ft. As a result, during this 

phase the aircraft are in dedicated departure sectors which have managed throughput and present little or 

no risk of interaction with other traffic (flights follow one another out of the airport and are managed as a 

sequence into higher transition sectors). It is further assumed that in a NextGen environment, departure 

management capability such as Departure Flow Manager (DFM), currently in prototype stages, is 

available and assists controllers with the management of a variety of departure operations.

Phase 1 supersonic operations in the NAS are restricted to only 21 departure sectors. Figure 60 shows 

analysis of N+3 Phase 1 operations and conventional operations for selected departure sectors in New 

York, Los Angeles and Fort Worth. The Figure indicates that although some additional sector capacity 

will be required to handle N+3 operations, overall operations occur during periods with lower 

conventional traffic operations. This is due to slightly earlier departure times for N+3’s to allow for 

multiple return operations across a 24-hr period. This suggests that capacity exists in the departure sectors 

to allow for the management of Phase 1 operation of N+3’s. As can be seen the number of operations in 

these departure sectors are high as is typical of the low airspace and does not pose a significant problem 

as all flights are managed in a controlled sequence (unlike upper airspace where traffic can be mixed).

In order to evaluate the potential impact of air traffic complexity and controller workload due to N+3

operations, a RAMS Plus Controller workload model was used. The model is based on a set of generic 

ATC tasks generated during the simulation in response to discrete events that occur during the simulation 

(e.g. entry to an ATC sector will record a number of ‘real-world’ ATC tasks such as coordination, transfer 

of control, initial clearances etc.). Additionally task weights can be modified according to dynamic 

conditions. For example, a conflict between flights with a slow rate of convergence will require different 

workload than one with a high convergence rate. During the simulation, controller tasks of varying 

lengths are recorded at different times with metrics produced to evaluate the percentage loading in the 

next 60-min period every 15 min during the simulation.

Figure 61 shows mean and maximum air traffic complexity and controller workload for departure 

sectors impacted by N+3 operations. As can be seen, differences in air traffic complexity and controller 

workloads due to Phase 1 operations are minimal with average controller workload marginally higher 

across most sectors and small percentage increases in maximum percent load, which remains below 

80 percent.
NASA/CR—2010-216796 53

Phase 1 operations v hourly flights—ZLACS

Phase 1 operations v hourly flights—ZLALA

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Figure 60.—Phase 1 analysis for selected departure sectors.

 

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NASA/CR—2010-216796 54

Figure 61.—Phase 1 air traffic complexity and controller workload.

3.4.3.2.2 High-Speed Climb to Supersonic Acceleration Flight Level

Phase 2 of the N+3 operations consists of a high speed climb following the initial acceleration phase 

between 10000 and 23000 ft. The climb is contained within the same low level departure sectors as Phase 

1 operations. Analysis reveals that for all N+3 departures from U.S. CONUS airports, only a small 

percentage of departing traffic traverses more than 1 ATC sector during Phase 1 (acceleration) and Phase

2 (climb)—only 9 sectors more than the original 21 Phase 1 sectors have Phase 2 operations occurring in 

them. As shown in Figure 62, Phase 2 operations occur nearer periods of peak traffic load.

As is shown in Figure 63, workload impacts remain low during Phase 2 operations. Similar to Phase 

1, workload is generally 3 to 5 percent higher, with some larger increases in peak workload relating to 

management of several N+3 departures during the same departure bank. As these aircraft are procedurally 

separated and do not intersect with one-another, managing the N+3 traffic during Phase 2 may cause 

marginal increase in controller workload. In Phase 2 operations air traffic remains organized as departure 

management systems are expected to assist in the sequencing and management of flights as they are 

delivered to the departure transition airspace.

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NASA/CR—2010-216796 55

Phase 2 operations v hourly flights—ZLACS

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Figure 62.—Phase 2 analysis for selected departure sectors.

 

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NASA/CR—2010-216796 56

3.4.3.2.3 Phase 3: Acceleration toward Supersonic at 24000 ft

At 24,000 ft the N+3 aircraft levels off and accelerates from 500 to 750 kn in around 130 sec. This 

particular phase occurs in mid level en route sectors and the Phase 3 operations are dispersed across 

several NAS sectors. Figure 64 shows analysis of selected sectors in the Boston, Washington DC, New 

York and Houston En Route ATC centers. As can be seen in the figure, N+3 operations do not 

significantly increase the number of operations in these sectors over conventional operation numbers.

Figure 65 shows air traffic complexity and workload for sectors that are involved in Phase 3 

operations. The air traffic complexity and controller workload is considerably higher in these sectors with 

N+3 operations compared to the baseline. The increase in workload could be attributed to additional load 

involved with managing rapid acceleration of supersonic aircraft. 

Figure 63.—Phase 2 air traffic complexity and controller workload.

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NASA/CR—2010-216796 57

Phase 3 operations v hourly flights—ZBW46

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Figure 64.—Phase 3 analysis for midlevel en route sectors.

 

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NASA/CR—2010-216796 58

3.4.3.2.4 Phase 4: Supersonic Climb to Cruise Flight Level

Following Phase 3 acceleration, the N+3 climbs at supersonic speeds to its cruise altitude. In this 

phase, it climbs through other flight levels occupied by conventional traffic flying at significantly lower 

speeds. Figure 66 shows controller workload across all sectors in which N+3 supersonic climb occurs. As 

can be seen, the workload is considerably higher in the scenario with N+3 operations compared to the 

baseline. The results are consistent with research in the area of air traffic complexity (Ref. 19) and 

suggest that situations involving overtaking aircraft can cause a greater level of complexity. Sridhar, Seth 

and Grabbe (1998) (Ref. 20) suggest that speed differences of greater than 150 kn from average speed of 

all aircraft in the sector increases complexity.

Analysis of conflicts between N+3 operations and conventional traffic was conducted during Phase 3 

acceleration and Phase 4 supersonic climb phases. Figure 67 shows the number of conflicts identified 

during Phase 3 acceleration and Phase 4 supersonic climb using a standard 5 nmi separation. As expected, 

a majority of conflict situations involving the N+3 are found in the 34,000 to 39,000 ft. bands (“AST” in 

the figure title connotes “Advanced Supersonic Transport,” or the N+3 aircraft). These flight levels are 

most commonly used as cruising altitude for conventional aircraft. A comparison with the current number 

of subsonic conflicts is suggested as an area for further investigation.

For the 12 hr period being considered in the scope of this initial study, conflict counts are relatively 

low, peaking just over 50. The conflict situations are further classified by conflict characteristics, 

crossing, in track, or opposite. While this is encouraging, it is felt that further detailed studies of the 

interaction between N+3 and conventional traffic in these flight levels would be of great benefit. It is also 

important to further evaluate the type and geometry of conflict as these factors have different impact on 

complexity and controller workload.

Figure 65.—Phase 3 air traffic complexity and controller workload.

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NASA/CR—2010-216796 59

Figure 67.—Phase 4 conflict analysis.

3.4.3.3 En Route Operations—Phase 5

N+3’s will use flight levels significantly higher than conventional traffic (over 50,000 ft) as their final 

cruise altitude. At these altitudes the traffic density is expected to be relatively low therefore no 

significant impact on operations is anticipated. ATC activities during this phase are also expected to be 

low. For these reasons a detailed analysis of N+3 operations was not attempted during this study; one is 

recommended as future work.

3.4.3.4 Arrival Operations—Phase 6

In the descent phase, N+3 will descend rapidly from its supersonic cruise, decelerating until reaching 

the more conventional flight levels around 36,000 ft with speeds similar to conventional traffic at that 

level. Because the N+3 will have to merge with conventional traffic for arrival operations, arrival 

(Phase 6) interactions with those flights were considered.

Figure 68 shows the analysis of type of conflict by flight level bands for N+3 operations during arrival 

phase (again referring to the vehicle as an “AST”). As it descends into the conventional airspace (>FL 

39,000 ft) crossing or opposite conflicts are predominant. Sequence and merging conflicts (InTrack) become 

more pronounced as the N+3 descends toward the transition and arrival airspace below 29,000 ft. A 

comparison with the current number of subsonic conflicts is again suggested as an area for further 

investigation.

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NASA/CR—2010-216796 60

Figure 68.—Phase 6 conflict analysis.

3.4.4 Conclusions

An investigation of how N+3 operations impact the demand for airports and airspace, based on its 

performance characteristics, has been conducted. Results suggested that N+3 operations will require some 

additional airport capacity during peak hour hours at certain airports; however, most operations are likely 

to be scheduled during the periods of nonpeak conventional operations and will therefore utilize the 

unused airport capacity. Furthermore, N+3 route networks between various airports are fairly well 

separated (assuming they are able to fly great circle routings) even for busy air corridors. During 

departure, N+3 Phase 1 operations (climb to 10,000 ft) are contained within 21 National Air Traffic 

System (NAS) departure sectors and nine additional sectors for Phase 2 (high speed climb from 10,000 ft 

to 23,000 ft) operations. During Phase 1 and 2 of N+3 operations a marginal increase in air traffic 

complexity and controller workload is predicted. Phase 3 operations occur in mid-level, en-route sectors 

and are more dispersed across the NAS. Although the number of operations in these sectors increased 

marginally, air traffic complexity and controller workload was much higher. A significant increase in 

controller workload and complexity is anticipated during Phase 4 N+3 operations, which involves 

supersonic climb from 24,000 ft to cruising altitude. Analysis of conflict during supersonic climb, 

Phase 4, indicates that most conflicts occur between flight levels 34,000 to 39,000 ft; however, the overall 

numbers of conflicts were low for the duration of the study.

4.0 Tasks and Trade Studies—Propulsion Systems

Three propulsion system models were delivered to LM as part of this program. The first two were 

designed to meet all of the N+3 Supersonic goals, fuel efficiency, airport noise, and emissions. The last 

was “optimized” solely for cruise efficiency while ignoring the airport noise and emissions goals.

4.1 Technical Description

4.1.1 VCE Propulsion System

The advanced VCE uses advanced technologies aimed at an entry into service (EIS) date of 2035. The 

engine architecture includes variable cycle features and advanced thermal management. The high 

performance Axi-Plug exhaust includes advanced noise reduction technologies.

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NASA/CR—2010-216796 61

4.1.2 Conventional Mixed Flow Turbofan (MFTF) Propulsion Systems

The variable cycle features were eliminated from the turbofan that met the environmental goals along 

with the “Optimized” cruise engine that met or exceeded the performance goals without meeting any of 

the environmental goals. However these MFTF engines still had the other advanced technologies and the 

high performance Axi-Plug exhaust without the noise reduction technologies.

4.1.3 ATMS

The ATMS enables the engine to be designed at a higher OPR for improved thermal efficiency and 

enhanced engine life while operating at sustained high core temperature during cruise.

4.1.4 Advanced Low NOx Combustor

Also, as part of this program, advanced low NOx combustor concepts and technologies were 

investigated to determine the ability of the propulsion system to achieve the N+3 cruise emission goal.

The combustor concepts and technologies and the impact on cruise NOx emissions are discussed in 

Section 4.7. The impact of the combustor exit temperature (T4) on cruise NOx emissions, engine 

performance, and weight is also discussed in Sections 4.4.1.3 and 4.7.

4.2 Exhaust System

The exhaust system is very important for supersonic vehicles since the nozzle performance is critical 

to the efficiency of the propulsion system as well as enabling jet noise technologies.

4.2.1 VCE Exhaust

The exhaust system is critical for supersonic vehicles since the nozzle performance strongly impacts 

the efficiency of the propulsion system as well as enabling jet noise technologies. The VCE propulsion 

system also features an Axi-Plug exhaust system with advanced jet noise reduction features. The exhaust 

has variable area capability to ensure high performance throughout the flight envelope. The exhaust 

system also provides for a thrust reversing system. Ceramic Matrix Composites (CMC) and other 

advanced composite materials are used extensively in the exhaust to minimize cooling requirements and 

weight.

4.2.2 Conventional MFTF Exhausts

For the two other turbofan engines designed without VCE features, the cooling scheme and advanced 

materials for the exhaust were the same. However, the advanced jet noise reduction technologies are not 

used in the exhaust. For the engines without VCE features, the core and bypass streams are 

conventionally mixed and then exhaust through an Axi-Plug nozzle, so the convergent section tended to 

be longer to provide the additional mixing length required for this higher BPR engine. This extra length is 

beneficial for noise as the two streams have more mixing length and plenty of room for acoustic 

suppression. The Axi-Plug exhaust also helps with the noise due to the increased radius ratio.

4.2.3 Transonic Thrust Augmentation

The exhaust also has a feature that augments the mixed temperature of the exhaust by a moderate 

amount. This system has minimal impact on exhaust performance. Due to the low augmentation 

temperature and use of advanced materials, no additional cooling is required when the augmentation is 

on.

4.2.4 Exhaust Variable Geometry Features

The exhaust has variable throat and exit area capability to ensure high performance throughout the 

flight envelope. The translating cowl also provides for a thrust reversing system.
NASA/CR—2010-216796 62

4.3 Airframe Requirements

4.3.1 NASA N+3 Goals

For the NASA N+3 Supersonic transport (SST) study, LM used a Mach 1.6 cruising, 100 PAX 

vehicle of approximately 300 klb While the propulsion system does not have a strong impact on Sonic 

Boom other than minimizing the size of the engine and the overall aircraft, it does strongly impact airport 

noise, cruise emissions, and overall mission efficiency. 

To best reflect overall mission efficiency, NASA also developed a fuel efficiency Figure of Merit 

(FOM) around specific range and passenger load (PAX). The FOM is #PAX * (range / block fuel).

Range/block fuel is specific range or V0 / fuel flow. Specific range can also be presented in terms of 

V0 *(L/D)/(SFC)/Weight. The engine obviously has a direct impact on SFC, but the size of the engine 

can also impact aircraft L/D and weight.

4.3.2 Thrust Requirements

LM provided thrust requirements for their NASA N+3 SST design. For sizing the initial engine, 

takeoff and transonic thrust targets were set along with a Mach 1.6 TOC thrust target that would provide a 

300 ft/m ROC margin. Since the GE variable cycle propulsion system has additional noise suppression 

capability over a regular turbofan, the target Vjet at 10 percent Programmed Lapse Rate PLR was initially 

set higher than what would be set for a typical turbofan

4.3.3 Customer Installation Effects

LM requested that propulsion installation effects be included in the installed engine data. To model 

these losses LM provided inlet performance updates. LM provided data is used directly in the cycle model 

to model the inlet. Along with the inlet recovery table, LM provided inlet critical additive drag (Cdcrit).

The inlet drag is represented in the installed thrust.

The GE variable cycle propulsion system has the Axi-Plug exhaust with excellent nozzle performance 

and acoustic suppression capability. Since the nozzle has variable A9 capability, the area ratio can vary 

and therefore peak performance can be maintained at most all flight conditions. Tests and CFD analysis 

have shown that final exhaust performance is relatively unaffected by the noise reduction features in the 

exhaust. 

There should be very little throttle dependent drag from the Axi-Plug nozzle, so nozzle drag was left 

out of the installed thrust.

For customer offtakes, LM provided the table shown in Figure 69, 1.5 lb/s compressor interstage 

bleed, 1.25 lb/s Fan Stream bleed, and 80 hp taken from the HP shaft. The fuel lower heating value for the 

study was assumed to be 18400 Btu/lb.

Engine bleed and horsepower extractions—Nominal

(Per engine)

Horsepower 80

Engine bleed—Fan 1.25 lbm/sec

Engine bleed—Intermediate 1.5 lbm/sec

Figure 69.—LM NASA N+3 bleed and horsepower extraction 

requirements
NASA/CR—2010-216796 63

4.4 GE VCE Propulsion System

4.4.1 Propulsion System Design

4.4.1.1 Description

The core is initially sized to keep max cycle average T41 under the T41 limit throughout the flight 

envelope. The core pressure ratio (CPR) is set to keep the maximum cycle average T3 under the T3 limits.

Section 4.4.1.3 shows the propulsion system impact of dropping core temperatures and Section 4.7

discusses the impact of core design temperatures on the NOx emissions goal. As will be shown in Section 

4.4.1.3 the final design maximum cycle average T41 can be reduced by several hundred degrees to 

improve cruise NOx emissions, without having a significant impact on engine performance. 

The engine total airflow size is set to yield the Mach 1.6 TOC thrust requirement. The Transonic 

Thrust Augmentation device typically provides sufficient margin to meet transonic and takeoff thrust 

requirements. At max power, the fan is running at 100 percent corrected speed. The Transonic Thrust 

Augmentation device can either be on or off depending on demand. The Axi-Plug exhaust has partially 

variable A8 and A9, which can be adjusted to accommodate the modest temperature bump from 

augmentation and still maintain near peak efficiency across the operating range.

The initial design space exploration to determine the best FPR and CPR for the N+3 engine started 

with an engine from a previous internal study, which could meet noise requirements with no margins.

This engine was updated with an advanced high-pressure ratio core while still limiting max cycle average 

T3 to the N+3 timeframe limit.

To design the engine, thrust available was tracked at SLS takeoff, Mach 0.3 SL and 1,000 ft, 

transonic, and Mach 1.6 TOC. These were the critical flight conditions discussed in Section 3.2.2.2.

Maximum power with and without the Transonic Thrust Augmentation device is observed at all 

conditions other than Mach 1.6. Vjet at 90 percent takeoff power to represent a Programmed Lapse Rate 

(PLR) for sideline noise measurement is also tracked to check the engines ability to meet acoustic 

requirements.

4.4.1.2 Results

An initial design trade study was performed to examine the effects of FPR, and dry versus Transonic 

Thrust Augmentation device operation on engine size. Engines were designed over a range of FPR’s to 

meet thrust requirements with and without the Transonic Thrust Augmentation device. The CPR and core 

size were rescaled for each engine targeting the max cycle average T3 and T41. The study showed that 

without the Transonic Thrust Augmentation device, the engines were sized at the takeoff or transonic 

condition, providing excess thrust at TOC. Conversely, when using the Transonic Thrust Augmentation 

device, the engine was always sized at TOC and was smaller.

The results from these preliminary trade studies are shown in Figure 70. The previous GE internal 

studies had already shown that the highest FPR engine should be able to meet NASA’s N+3 acoustic 

goals. So, any lower FPR engine should also be able to meet the acoustic goals. All engines sized to meet 

Takeoff thrust at Mach 0.3 ISA +36 °F with or without the Transonic Thrust Augmentation device. The 

Transonic Thrust Augmentation device was always used to meet transonic thrust requirement, particularly 

at the lower FPR. Although sensitivities to assess overall mission impact were not available at this point 

in the studies, it was felt that the highest FPR engine would be an appropriate propulsion system for this 

program; it was the smallest and lightest engine. It did not appear that the SFC penalty for the higher FPR 

would have as near an impact as the size impact. Therefore, the higher FPR was chosen for the initial 

baseline VCE design. The cycle data is used to design the Flowpath for the engine and the exhaust.
NASA/CR—2010-216796 64

Figure 70.—Results from preliminary FPR Design Space Exploration Trade Study.

The engine matches the Mach 1.6 thrust requirement exactly. The Transonic Thrust Augmentation 

device is required to meet takeoff and transonic thrust requirement. The cycle data for this engine was 

used to estimate takeoff noise levels for this baseline configuration. The takeoff noise level prediction for 

this baseline configuration is discussed in Section 4.3.2.

An engine datapack with mission and takeoff data was generated covering the envelope LM

requested. Cycle performance data includes: 1.5 lb/s compressor interstage bleed, 1.25 lb/s fan stream 

bleed, and 80 hpx (60 kw) extracted from the HPC and GE’s Axi-Plug exhaust system performance with a 

variable A8 and A9.

4.4.1.3 Impact of Core Temperature on Engine Size and Performance

The engine core temperature has a significant impact on performance, weight, life, and NOx 

emissions. HPC discharge temperature (T3) and combustor exit temperature (T4) have a strong impact on 

combustor emissions, as shown in (Ref. 21). HPT rotor inlet temperature (T41) determines how much 

energy per pound of flow is available from the HPT and therefore core size. T3 and T41 set overall 

thermal efficiency and therefore strongly impact SFC. T3 has the strongest impact on SFC, typically the 

higher the better. There is also an optimum T41, above which T41 has little impact on SFC.

T3, T4, and T41 also set cooling flows for the core components. The ATMS and the advanced CMC 

materials system allow the engine to operate at high T3 with less cooling flow impact, but the T41 or T4 

still impact required cooling flows. T4 is the gas path temperature for the HPT nozzle inlet and T3 is the 

temperature of the cooling air. From these, the required cooling flow for the HPT nozzle is set. The 

amount of cooling flow and T3 then set turbine rotor inlet temperature (T41). T41 is a mass average of the 

cooling flow for the nozzle at T3 and gas flow through the nozzle at T4. At the extreme, if a material was 

available that could survive at T4 temperatures, then no cooling would be required and T4 would be equal 

to T41. T3 and T41 also set cooling flows for the HPT and T3 and T49 cooling flows for the LPT. Even 

though for an ideal Brayton cycle, higher core temperatures are better for thermal efficiency, the required 

cooling flows tend to diminish the benefit. Once the gas path temperature gets below the bulk metal 

temperature capability of the material, then the cooling flows will no longer change as T4 and T41 drop.

Figure 71 shows the impact of T4 on the VCE propulsion system concept, adjusting cooling flows 

appropriately as temperature drops. As can be seen, the T4 can be initially pulled back minimal impact on 

SFC. The T4 impact on propulsion weight tends to be somewhat linear. Although as T4 drops further, the 

propulsion penalties start to increase more significantly.

Sized to Dry Takeoff Thrust

-25%

-20%

-15%

-10%

-5%

0%

5%

11% 22% 32%

% FPR Increase

SFC

Diameter

Weight

Propulsion Wt

Diameter

Cruise SFC
NASA/CR—2010-216796 65

4.4.1.4 Impact of Plug Size in Axi-Plug Nozzle

LM had expressed some concern about the high nozzle closeout angles (boattail) on the GE Axi-Plug 

exhaust. It was noted that increasing the size of the plug could reduce the boattail angles on the exhaust. 

Since the Axi-Plug exhaust was leveraged form earlier internal studies, a CFD input file already existed 

that could be used to determine the impact of plug size on exhaust performance.

For this quick CFD study two flight conditions were analyzed, takeoff with the Transonic Thrust 

Augmentation device on and Mach 1.6 TOC. Two Plug sizes were investigated, the original and 

50 percent larger.

Overall the study showed that the larger plug might decrease CFG by a slight amount. Although, it 

was felt that reductions in this loss would be possible with more time to optimize the geometry. The study 

also showed that the CFG could be improved by a modest amount at takeoff cutback. 

4.4.2 Noise Assessment

4.4.2.1 Description

The noise goals for the supersonic N+3 application are 10 to 20 EPNdB cumulative margin below 

Stage 4 (or 20 to 30 EPNdB cumulative margin below Stage 3). LM provided the take-off gross weight 

(TOGW) target, number of engines and thrust requirements. GE made assumptions regarding the 

remaining information required to perform noise assessments. GE provided three different propulsion 

systems, two of which were intended to meet the noise goal and one which was designed solely for cruise 

efficiency. Noise assessments were only performed on the two propulsion systems intended to meet the 

noise goal. All of the noise assessments performed in this study are preliminary and no margins are 

applied. GE provides these as guidance only and implies no guarantees or commitments.

The first propulsion system developed was a VCE that enables the advanced jet noise reduction 

features in the exhaust. The exhaust system incorporates a number of proven noise reduction technologies 

for supersonic exhaust nozzles and combines them into one innovative exhaust system. 

Figure 71 —T4 impact on weight and SFC

0%

1%

2%

3%

4%

5%

6%

7%

8%

9%

10%

11%

12%

13%

14%

15%

16%

3000 3100 3200 3300 3400 3500 3600 3700 3800

Cycle Max Avg T4 F

% Change from Base

% Change SFC

% Change Wt

0%

1%

2%

3%

4%

5%

6%

7%

8%

9%

10%

11%

12%

13%

14%

15%

16%

3000 3100 3200 3300 3400 3500 3600 3700 3800

Cycle Max Avg T4 F

% Change from Base

% Change SFC

% Change Wt
NASA/CR—2010-216796 66

4.4.2.2 Results

Jet noise is typically the dominant propulsion system noise component for the take-off monitor 

locations and usually the lateral, sideline monitor location will be the most difficult noise level to achieve 

so the initial acoustic assessment was limited to this condition. Based on previous internal work with LM

an appropriate fan pressure ratio was chosen that should meet the noise goal. However, the previous 

studies did not include any thrust augmentation capabilities so the initial noise study looked at the 

propulsion system with and without augmentation during the sideline monitoring location. Figure 72

shows the results of this noise study, the points on this chart are all under the Sideline Stage 3 

requirement. The methodology followed to obtain the noise estimates are outlined below:

• Jet noise estimate of the equivalent ideally mixed exhaust 

– Assumes appropriate altitude, 1,476 ft sideline

– External velocity appropriate for take-off 

• Scaled jet noise levels from experimental database of three configurations for a range of mixed jet 

velocities:

– Conic nozzle equivalent ideally mixed exhaust 

– Jet Noise Reduction Technology 1

– Jet Noise Reduction Technology 2

• Determine benefit of the Jet Noise Reduction configurations by subtracting from equivalent 

ideally mixed exhaust conic nozzle

• Account for number engines 

• Account for engine-engine jet noise shielding based on installation

• Compare to the sideline Stage 3 requirement to obtain sideline margin 

 

Figure 72.—Total jet noise level for sideline monitor location for VCE engine 

for different power setting with and without augmentation. 

Final SL Jet Noise vs Power

40 50 60 70 80 90 100

Percent Engine Power (%)

Total Jet Noise EPNL (EPNdB)

Without Augmentation

With Augmentation

2 dB
NASA/CR—2010-216796 67

This figure shows the percentage of engine power versus the sideline jet noise estimate for the total 

propulsion on the aircraft. An important observation from this is that the augmentation provides a benefit 

over the higher range of engine power. This is because the jet noise reduction feature is highly dependent 

on the exhaust velocity range. As the velocity increases the benefit of this feature typically improves, such 

that although the primary jet noise associated with the higher exhaust velocity may be higher, the 

improved performance of the feature may result in a lower final jet noise level. Typically for the same 

thrust the augmentation results in a higher temperature and thus higher exhaust velocity. Figure 73

demonstrates this effect; showing scaled experimental data for a conic nozzle, and configurations with the 

jet noise reduction features. There is clearly a level of the exhaust velocity where the features show a 

benefit once above. This figure shows a typical result of scaling existing scale model data, for one engine 

and before any system corrections have been applied.

This study showed sufficient margin for the sideline or lateral monitor location jet noise with 

augmentation during take-off. A more detailed noise assessment was conducted for all three of the 

monitor locations, sideline, flyover or cutback, and approach. Propulsion noise estimates were made for 

the jet noise and the fan inlet radiated noise. The fan exhaust radiated noise was not included due to 

uncertainty in its prediction as well as the significant room for acoustic suppression treatment through the 

relatively complex exhaust. Airframe noise levels were crudely estimated using an existing preliminary 

design tool as no better estimates were provided. The methodology for the jet noise estimate has already 

been summarized. The fan inlet radiated noise component was estimated using a GE version of the 

Heidmann code that has been validated with a similar fan system under a previous program. This estimate 

includes all of the features of the VCE fan module, and a conservative estimate of acoustic suppression 

treatment in the inlet. Additional assumptions are listed below:

• Flyover/cutback

– Assumes thrust cutback to an appropriate level

– Assumes cutback altitude

– Appropriate external velocity

Figure 73.—Typical jet noise level for jet noise reduction features versus mixed 

velocity scaled experimental data.

Mixed Exhaust Velocity (ft/sec)

Jet EPNL (EPNdB)

Conic

Feature 1

Feature 2

Increasing

5 dB
NASA/CR—2010-216796 68

Based on these noise assessments the VCE propulsion system is expected to meet the program goal.

The take-off profile should be assessed to improve the cutback procedure and a better estimate of the 

airframe noise is needed. The airframe noise levels used in the current study contribute 4 to 5 EPNdB to 

the cumulative margin.

The fan inlet radiated noise also makes a significant contribution to the cumulative margin. Alternate 

propulsion/airframe integrations should be assessed to mitigate this. There are alternate exhaust 

configurations that could take advantage of airframe shielding and much more highly integrated 

installations.

4.5 Conventional Mixed Flow Turbofan (MFTF) Propulsion System

4.5.1 Propulsion System Design

4.5.1.1 Description

LM requested that GE design a moderate BPR Mixed Flow Turbofan (MFTF) without variable cycle 

features or the associated special acoustic features in the exhaust that would still meet the NASA N+3 

acoustic requirements for the program. The FPR for the MFTF was chosen based on previous internal 

studies. While the special acoustic features in the exhaust were no longer available without the variable 

cycle features, the Transonic Thrust Augmentation device was still deemed possible although a greater 

challenge due to the reduced fan stream temperatures.

4.5.1.2 Results

The initial configuration chosen was designed with the same core temperatures and technologies as 

the VCE baseline. The CPR was increased to make up for the loss of boost to get as much OPR as 

possible without adding a booster stage. Both the MFTF and VCE were designed to produce the required 

TOC thrust, but the MFTF makes more augmented thrust at takeoff and transonic, because of the greater 

thrust lapse.

The airflow capacity of the MFTF engine is 1/3 greater than the VCE engine, so the fan inlet is 

greater in diameter, the engine longer, and heavier. The nozzle exhaust duct is quite long to accommodate 

the large mixer that would be required for the higher BPR engine. In the end the MFTF engine showed 

only a slight improvement in SFC over the VCE concept. LM mission results indicated that the MFTF 

lost 240 nmi in range relative to the VCE.

4.5.2 Noise Assessment

4.5.2.1 Results

For the conventional MFTF noise assessment jet noise and fan inlet noise estimates were made both 

with and without augmentation, the airframe noise components are the same as the VCE engine since the 

vehicle weight didn’t change for the noise assessment. The assumptions for the monitor locations were 

also consistent. The jet noise assessment was fairly different from the previous assessment for the VCE 

propulsion system since the exhaust system was very different. The conventional MFTF uses 

conventional mixing of the core and bypass flows before reaching the Axi-Plug nozzle, shown in 

Figure 74. The methodology for estimating the jet noise was as follows:

• Jet noise prediction for fully mixed exhaust

• Appropriate penalty since lobed mixers generate high frequency noise and don’t result in fully 

mixed flow

• Appropriate noise reduction due to the plug

• Consistent engine and system corrections as VCE assessment
NASA/CR—2010-216796 69

Figure 74.—Conventional MFTF mixed flow and 

axisymmetric plug nozzle exhaust.

From this assessment both the cases, with and without augmentation, met the noise goal but the no 

augmentation case provided more cumulative margin. This is consistent since there are not technologies 

included in this exhaust system which perform significantly better at higher velocities. Therefore the 

conventional MFTF propulsion system would not have an acoustic benefit for using the augmentation 

during the take-off operations.

4.6 “Optimum” Cruise Conventional MFTF Propulsion System

4.6.1 Propulsion System Design

4.6.1.1 Description

GE was also requested to design an “Optimum” cruise engine ignoring all other program goals. The 

propulsion system development and refinement sought an N+3 Supersonic design with the highest cruise 

efficiency without regard to any noncruise design requirements, specifically, acoustics and emissions.

To design the engine the same advanced technology assumptions with a 2025 TRL-6 Technology 

Availability Date were used. The engine architecture has no variable cycle features but does includes 

other advanced technology features appropriate for the time period, including the Axi-Plug Exhaust with 

Transonic Thrust Augmentation device. The Optimized cruise engine is a conventional MFTF like the 

previous one, only with a higher FPR and lower BPR.

4.6.1.2 Results

For the Optimized cruise engine study, Lockheed sensitivities were provided in the form of range, 

drag, and weight sensitivity data in a spreadsheet. These were used to determine the impact on vehicle 

range due to variation in diameter, SFC and weight.

Figure 75 shows the results from the cruise engine optimization study. As part of the optimization 

process, it was discovered that the Fan inlet diameter could be reduced somewhat, which improved range 

by approximately 1 percent for the baseline.

Figure 75.—Impact to range based on LM sensitivities.

Largest gain in range

VCE

Baseline Moderate 

FPR

high 

FPR

higher

FPR

Inlet Diameter (inches) 0.0 -1.1 -4.6 -7.1

Range increase/decrease due to Diameter (nm) 0.0% 0.7% 3.2% 4.9%

SFC 0.0% -1.2% -0.5% 0.7%

Range increase/decrease due to SFC (nm) 0.0% 1.5% 0.5% -0.8%

Total Quote Weight (lb) 0.0% 2.7% -6.8% -13.1%

Range increase/decrease due to Weight (nm) 0.0% -0.4% 1.0% 2.0%

Total % Range increase/decrease 0.0% 1.8% 4.7% 6.1%

Impact On Range

MFTF
NASA/CR—2010-216796 70

The next step in the study was to define the cruise “optimized” MFTF with approximately the same 

mass averaged FPR as the baseline VCE, resulting in a moderate FPR MFTF. Since the previous 

conventional MFTF study showed a decrease in range performance with lower FPR, it was decided to 

investigate turbofans with higher FPR’s. The high FPR MFTF showed a significant gain in range over the 

moderate FPR MFTF. Next a higher FPR MFTF was designed and the sensitivities still showed an 

increase in range, although the increase was much less, indicating that the trade for size for SFC was 

leveling off. The higher FPR MFTF was chosen for the “Optimized” cruise engine to deliver to LM.

The max inlet airflow capability for the higher FPR MFTF is much less than the VCE baseline, 

because it makes the same thrust at a higher FPR. All engines were designed to make the same TOC 

thrust, but the ”optimized” MFTF makes less augmented thrust at takeoff and transonic, because of the 

smaller thrust lapse. With transonic augmentation, this engine is almost perfectly balanced at takeoff and 

TOC. The fan diameter is much smaller and the weight comes out lower than VCE baseline, but is not 

expected to meet the N+3 acoustic and emissions goals.

4.7 Emissions status

Cruise NOx emissions were estimated using an analytical method developed by GE Aviation. This 

method simulates the combustion process in the Twin Annular Premixed Swirler (TAPS) combustor 

(Figure 76) used in the propulsion system under study. It should be noted that this methodology was 

previously also used for estimating EINOx for the N+1 supersonic combustor program. It uses Chemkin 

to calculate a laminar lean premixed opposed flow flame to represent the well-mixed portion of the 

combustion, and a rich premixed opposed flow flame to represent the diffusion flame from the pilot. Jet-A 

was the fuel, and the chemical kinetic mechanism was obtained from GE Aviation. Figure 77 shows the 

flow configuration.

Figure 76.—GE Twin Annular Premixed Swirler (TAPS) combustor.
NASA/CR—2010-216796 71

Figure 77.—Opposed flow flame 

calculated using Chemkin with 

Jet-A.

Some assumptions have been made to estimate GE62 NOx levels for the N+3 time frame. Starting 

with assumptions being used in an ongoing Low Emissions Combustor Program for Supersonic Aircraft 

additional assumptions are made in terms of cooling flow, residence time, and pilot level for the N+3 

timeframe. Using these levels and adjusting the T4 level down based on the discussion in Section 4.4.1.3

the emissions goal of 5 EINOx is projected to be met.

4.8 Efficiency (FoM) status

Based on the propulsion system cycle data GE provided to Lockheed-Martin and the results of 

Lockheed-Martin’s sizing studies the final version of the GE propulsion system resulted in an FOM 

exceeding the N+3 goal.

4.9 Technology Trades and Studies

4.9.1 Low-Emissions Combustor

4.9.1.1 Description

Meeting the N+3 emission goal of cruise EINOx = 5 g/kg fuel using a TAPS-type combustor has 

been one of the most challenging aspects of the supersonic N+3 program. To achieve the NOx goal, even 

more aggressive technology development targeting an N+3 timeframe is necessary in regards to reducing 

combustor residence time, cooling flow, pilot, and fuel/air mixing time. Figure 78, Figure 79, and Figure 

80 show the variation of EINOx with different levels of mixer air, mixedness and amount of pilot, 

respectively, for different values of T4.

Figure 78.—Effect of amount of mixer air on EINOx with 

different values of T4.
NASA/CR—2010-216796 72

Figure 79.—Variation of EINOx with different levels of 

mixedness versus T4.

Figure 80.—Variation of EINOx with different levels of 

pilot air versus T4.

4.9.1.2 Results

A significant reduction in residence time is beneficial. To enable this, a significantly shortened 

combustor will be required. However, a shortened combustor will have high emissions of CO at low 

power (e.g. idle) conditions due to incomplete combustion. A shortened combustor may also not have 

sufficient volume for reliable operability. Additional technologies will be required to address this 

deficiency in operability. Therefore, a variable geometry combustor might be necessary, where the 

combustor is relatively long at low power conditions, and is then shortened during cruise.

A significant jump in cooling and material technologies is also required. The liner and dome cooling 

flows should be reduced significantly and more air can be directed to the mixer. Significant advances in 

material development would be necessary to achieve this.

4.9.2 Intercooler

The principal challenge for an intercooler (IC) is coming up with a lightweight low loss Hx system to 

remove heat from the compressor gooseneck (T25). The larger the temperature difference between the hot 

stream (T25) and the cooling source, the smaller the Hx required. The larger the Hx system, the greater 

the size, weight, and associated pressure losses.
NASA/CR—2010-216796 73

Figure 81.—Two notional types of heat exchangers.

4.9.2.1 Description

This study explored entitlement for the IC in the GE baseline VCE. Most of the complexities of the 

Hx system were ignored to see if the system would work with no losses and provide sufficient benefit to 

be worth a more detailed investigation. The analysis accounted for the amount of heat that needed to be 

transferred to obtain a certain drop in T25 and then the amount of Hx surface area required to achieve this 

heat transfer.

Typically, the purpose of an intercooler is to take heat out of the air entering the compressor (T25) to 

reduce core temperatures T3, T4, and T41. Once T3 is reduced the CPR can be increased driving T3 back 

up to further improve thermal efficiency. Also, with an intercooler, core size could be dropped increasing 

BPR and driving T4 and T41 back up, again improving thermal efficiency. For these supersonic N+3 

studies, one of the most difficult environmental goals to achieve is the cruise NOx emissions goal. T4 and 

somewhat T3 have a strong impact on NOx emissions. For this analysis, the IC was used only to reduce 

T25 dropping core inlet temperature to reduce NOx cruise emissions, while maintaining thermal 

efficiency.

To perform this study, the VCE was designed with IC at the inlet to the core. It was assumed all 

gooseneck air passes through the notional heat exchanger to maximize the amount of heat that can be 

transferred and the drop in T25.

In the study two types of Hx were investigated, a surface cooler and tube style Hx as shown in 

Figure 81. The surface cooler is the lightest heat exchanger with the minimum amount of losses, but 

overall it requires more room to get the same amount of Hx surface area as the tubed. The tube style Hx 

has more losses, because they actually obstruct the flow going through the array of tubes and the flow 

going though the tubes have losses due to the passage through the small tubes. The tubes can however get 

more Hx exchanger area into a smaller space.

4.9.2.2 Results

The cycle analysis showed that to achieve about a reasonable drop in T4 requires about a small F drop 

in T25. From an ideal analysis the GE baseline could achieve a small drop in T25, but the calculation for 

Hx surface area showed that the required size was prohibitive.

Another alternative to the air-to-air Hx is to use the fuel as a heat sink. This would require an Air/Fuel 

Heat Exchanger or Air/Oil/Fuel Exchanger. A simple calculation of the thermal load from the gooseneck 

air to get the small F T25 drop showed that simplistically the fuel temperature would have to increase a 

significant amount to take all the heat due to the small quantity of fuel relative to gooseneck air.
NASA/CR—2010-216796 74

Realistically this would not work as the fuel temperature increased the delta T would drop reducing the 

heat transfer and eventually the fuel temperature would reach T25.

The final conclusion was that using the IC to reduce T25 was not practical, the Hx would not fit in the 

engine. As was shown in the T4 core sizing trade study, the loss in performance for a reasonable drop in 

T4 was only a slight increase in SFC and small increase in weight. If the losses in the IC Hx system were 

fully accounted for there would likely be a much greater impact on SFC and the weight of the required Hx 

would be prohibitive.

4.9.3 Interturbine Combustor

Another attempt at reducing T4 for NOx cruise emissions was to introduce some burning between the 

HPT and LPT. A simple analysis performed by adding a small burner between the two turbines showed 

that T4 could be reduced. However burning in this lower pressure zone is much less efficient. The SFC 

penalty was significant. Again as in the intercooler nozzle, even this larger reduction in T4 could be 

obtained by simply designing the core to the same lower T4 with a smaller system impact.

4.9.4 Constant Volume Combustor (CVC) 

CVC technology is being investigated to improve the efficiency of gas turbine engines. These 

configurations rely on pressure-rise constant volume combustion rather than constant-pressure 

deflagration currently used in today’s gas turbine engines. Notional thermodynamic CVC cycles show 

increased thermodynamic efficiency greater than Humphrey (constant volume) and Brayton (constant 

pressure) cycles as a result of the pressure-rise associated with constant volume combustion.

4.9.4.1 CVC Modeling Approach

The following section describes the modeling approach of the CVC module within the engine. In the 

architecture considered, the regular deflagration-based (Brayton cycle) combustor is replaced with a CVC 

in the hot-gas path, Figure 82. The CVC would accept compressor-discharge air, detonate, and then 

exhaust into the high-pressure turbine through a nozzle. Such an architecture would potentially allow destaging the compressor (since some of the pressure-rise is achieved by the combustor), but would likely 

require a very different turbine design. Furthermore, the design of this transition piece, and in particular 

the area ratio, is very important in optimizing the performance of the engine.

The CVC was modeled in NPSS by modifying the standard deflagration burner module in the engine 

model to include constant-volume combustion pressure-rise multiplier, Figure 82. The CVC pressure-rise 

map was generated separately using a quasi-1–D gas-dynamic code and inserted into NPSS as a lookup 

table in the functional form PR = function (T3, T41). Previous trade studies showed that the pressure-rise 

from a CVC (once the geometry was defined, including the area ratio) was primarily a function of these 

two parameters. The baseline VCE engine was used as a starting point for the present study.

The performance of a CVC is strongly a function of the physical geometry of the combustor. For this 

reason, it is important to understand the assumptions used in generating the pressure-rise map. The 

assumptions in generating are as follows:

1. CVC with 4:1 area ratio nozzle.

2. Gaseous C2H4 fuel with reduced mechanism two-step chemistry.

3. Includes steady-flow pressure losses.

4. Adiabatic walls.

5. “Rubber” inlet geometry so that each point is optimized for “on-design” operation. 

6. Same compressor/HPT maps as simple cycle.
NASA/CR—2010-216796 75

Figure 82.—Schematic diagram showing 

implementation of PDE in NPSS model.

4.9.4.2 CVC Trade Study Results

The main impact of the CVC is to enable a higher OPR engine using a simpler, lower CPR 

compressor. The study was performed with and without the Transonic Thrust Augmentation device. The 

engine was sized using the same temperature limits as the VCE and thrust-matched at the cruise thrust 

requirement. Thrust matching at cruise resulted in excess thrust at take-off. Furthermore, the thrust benefit 

of the CVC was an adder on top of the Transonic Thrust Augmentation device in the baseline engine. The 

specific fuel consumption benefit is approximately 3.2 percent for the CVC across the flight envelope, 

and is also an adder to the Transonic Thrust Augmentation device. One of the main observations with 

constant volume combustion is that the pressure-rise achieved is a strong inverse function of the initial 

temperature (or compressor discharge temperature). In particular, as the compressor pressure ratio 

increases (and T3 increases as well), the benefit of the pressure-rise combustion is reduced

An additional case was considered where the compressor was intercooled, so that the engine could 

benefit from the increased P3, but the CVC could maximize its pressure-rise due to a lower intercooled T3.

This has a slight benefit over the CVC alone. The addition of the intercooler provides an additional SFC 

benefit, with a total benefit of 4.7 percent over the baseline VCE engine. The benefit of intercooling (with 

CVC) is seen to be reduced at lower OPR.

At this early stage, the CVC technology seems promising in that it can offer up to 5 percent

improvement in SFC above and beyond the advanced N+3 baseline VCE engine. The TRL level of CVC 

technology, however, is relatively low and further development work is required to be considered for 

implementation.

4.9.4.3 Future Work

The present trade study represents a preliminary assessment of incorporating CVC technology into an 

advanced N+3 supersonic engine architecture. CVC technology is still in early stages of development, 

and many of the methods and design tools require further work. Future assessments should include the 

following improvements:

1. Refined CVC pressure-rise map with Jet-A fuel properties.

2. Further trade studies on the impact of CVC geometry on pressure-rise and an optimization of the 

geometry within the engine architecture. In particular, nozzle area ratio is known to have a big 

impact.
NASA/CR—2010-216796 76

3. Improved turbine maps specific to CVC flows (when available).

4. Further parametric variations of T3, T41 and turbine efficiency.

5. Perform preliminary weight estimates and trade studies to understand impact of weight to SFC.

6. Impact of CVC operation at part-power conditions.

5.0 Technology Roadmap Development—Airframe (WBS 3.3)

The overall strategy to develop technology roadmaps began with technology identification. Each 

technology was assessed for viability, applicability, and duplication. The down selected technologies 

were then analyzed and ranked in terms of benefit, probability of success, and total program development 

costs. This resulted in a prioritized list of enabling technologies. A description and risk assessment, 

including a TRL maturation plan, is included for each prioritized technology in Section 9.0 New 

Technology.

5.1 Technology Identification (WBS 3.3.1)

The process used by LM, to identify enabling technologies for the N+3 vehicle, combined resources

and research activities to define and develop a comprehensive technology list. The LM technology 

identification process is highlighted in Figure 83. 

The process initiated with research on LM heritage, on-going NASA, and on-going LM resources to

identify potential enabling technologies. All used sources for technology identification include:

• Heritage QSST/QSP information—Configuration shaping and inlet technologies

• Subcontractor input—GE GRC, Penn State, Reed/Saric, Purdue University, and J. Hansman from 

the program kick-off meeting held on December 12, 2008

• NASA information from the 2008 Supersonic Funding Portfolio

• Revolutionary Technology Programs (RTP) technology capabilities

• Tech Fair—Jet Exhaust Manipulation and Webcore Technologies

• Brainstorming sessions—Three sessions conducted during Phase 1

Each resource was thoroughly researched by the N+3 team to identify potential technologies capable 

of enabling an N+3 advanced vehicle concept.

LM conducted additional research activities to identify technologies from small businesses and 

subject matter experts. On December 8, 2008, LM conducted a Tech Fair to engage small businesses and 

the academic community to further identify enabling technologies. LM sent out over 30 solicitations and 

ultimately received three participants within the fair. Each participant was required to submit a half hour 

presentation that outlined their specific technologies and the research that had been done to date. LM 

selected two out of the three technologies presented as potential enabling technologies. Those two 

participants include Webcore Technologies, based out of Miamisburg, Ohio and Dr. Mo Samimy from 

Ohio State University. Webcore Technologies specializes in innovative materials and composite 

manufacturing which aligned with the same 2030 vision required for the N+3 vehicle. Dr. Samimy’s 

current research is on the application of plasma in jet engine exhaust to modify flow levels for noise 

reduction. Three brainstorming sessions were conducted during the first quarter of the program to 

encompass all technology possibilities. These sessions were held with LM subject matter experts and 

other LM personnel to provide insight for identifying a comprehensive technology list.

Once all resources and activities were completed, each technology was categorized and captured in a 

comprehensive spreadsheet. The four technology categories include:

• Individual bolt-on techs

• Airframe and propulsion configurations
NASA/CR—2010-216796 77

• Tools and methods

• Materials and generic technologies

Each technology was divided by source, category, benefits, drawbacks, N+3 goals addressed, and 

readiness. Figure 84 demonstrates a sample of the technologies compiled.

Figure 83.—Technology identification process used to construct the technology roadmaps.

N+3 Supersonic Technology List

Configuration

Technology Source Category Description Benefits Drawbacks N+3 goals 

addressed

Readiness

High notch ratio 

planform with 

flow control

N+3 proposal Aero Wing with high notch ratio 

(high TE sweep)

Low induced and wing wave 

drag

Heavy, low speed higher 

vortex drag and pitch-up 

without attached flow control

Efficiency Idea. Would need to 

develop and test 

specific design

Synthetic cockpit 

vision

LM proposal 

brainstorming 

session

Avionics Use TV camera system to 

give pilot better visibility 

where nose is in way

Eliminates need for folding 

nose or any other design 

compromises necessary for 

the pilot to see the runway

None Efficiency Needs development 

but research activities 

are on going

Swing canard LM configuration 

brainstorming 

session

Configuration Deploy canard only at low 

speeds to control trim and 

stability

Lower drag, better boom Weight and complexity 

Of retraction 

Efficiency, 

boom

Would have to be 

designed in 

Wing tip tanks LM configuration 

brainstorming 

session

Configuration Strategically shaped wing 

tip tanks

Structural stability, span load 

alleviation, area ruling

Weight Efficiency, 

boom

TRL-9, airplanes 

have employed them 

before

Configuration 

shaping

LM proposal 

brainstorming 

session

Configuration Shape configuration to 

close fore and aft shock to 

acceptable levels using 

advanced MDO techniques

Constrain sonic boom while 

maintaining efficiency and trim

Efficiency is compromised if it 

has to be balanced against 

boom requirements

Boom, 

efficiency

Has been done for 

QSST, more research 

to be done

Variable sweep 

outboard wing 

panels

N+3 kickoff 

meeting 

brainstorm

Configuration Vary sweep of wing using 

pivots or other mechanical 

system

Easier to balance low speed 

takeoff requirements with high 

speed drag and boom goals, 

less weight than full variable 

sweep wings since it's just 

outboard panels

Large weight penalties in 

pivots and mechanical 

systems

Efficiency, 

noise, boom

TRL-9, done before 

on F14, F111, and B1

Quiet spike N+3 kickoff 

meeting 

brainstorm

Configuration Extending and retracting 

spike, extends for cruise to 

increase effective length of 

airplane, retracts for better 

landing

Reduces boom Weight and complexity of 

retraction mechanism

Boom, 

efficiency

Tested on F-15 for 

mechanical reliability

Knight's lance 

nose

N+3 proposal Configuration Similar to quiet spike but 

would be nonretractable

Reduces boom Weight, landing slap-down 

considerations

Boom, 

efficiency

Idea

Figure 84.—Technology catalogue excerpt used to define space of possible technologies, provide input for down 

selection, and disseminate ideas.
NASA/CR—2010-216796 78

5.2 Technology Down Selection (WBS 3.3.2)

Once all proprietary and nonproprietary enabling technologies had been identified, they were then 

down selected to provide a focused, manageable list of technologies to analyze for vehicle integration.

The overall process is illustrated in Figure 85.

The LM technology down selection process started with the catalogued N+3 technology set and was 

scrubbed with five factors in mind to reduce the overall list. Those factors include:

• Research—subject matter experts, other LM projects, literature

• Weighted Pugh Matrix Analysis

• N+0, 1, 2 technology duplication

• Subsonic/supersonic duplication

• Techs already supported elsewhere

Each technology was initially gauged by N+3 team members and LM subject matter experts to 

determine if the technology passed a “sanity” check. Technologies that were not applicable to the N+3 

goals, determined to be ineffective through other and heritage LM projects, or duplicated through N+1 

efforts were eliminated. Duplicated N+2 technologies were included in the next stages of the down 

selection process; however, they were noted as being N+2 technologies and were directly applicable to 

the N+3 goals. The remaining technologies were then put into a weighted Pugh Matrix to determine the 

overall benefit for an advanced, supersonic vehicle. 

The Pugh Matrix analysis weighed different measures of merit such as L/D, empty weight, SFC, 

boom, noise, emissions, safety, pax comfort, and cost to determine a net benefit value for each 

technology. This net benefit value was ultimately used to determine the overall prioritization set. Within 

the Pugh Matrix, numbers of –9, –6, –3, 0, +3, +6, or +9 were assigned to a measure of merit (efficiency, 

environment, and practicality) to determine if a technology “helped” or “hurt” and by how much. A value 

was not assigned unless the technology had a direct and significant impact on the measure of merit.

Differentiating and assigning a value to a specific measure of merit allowed an explicit analysis of each 

technology in regards to vehicle impacts. In addition, it was assumed that the technologies put through the 

analysis were “bolted” onto a baseline vehicle. This baseline vehicle was assumed to have been designed 

to meet the noise, emissions, and efficiency goals. As a result, configuration technologies were considered 

to be endemic or already installed on the baseline vehicle, and technologies would be “bolted-on” or 

simply added to the baseline design. The reasoning is that the configuration technologies were judged to 

be high-payoff with low impact and necessary to achieve the N+3 goals. Figure 86 illustrates the baseline 

configuration with endemic technologies highlighted. 

Figure 85.—Technology down selection process that produced 

manageable list of technologies to further evaluate and create 

roadmaps.
NASA/CR—2010-216796 79

Figure 86.—Configuration technologies assumed to be designed into the advanced vehicle concept.

Other Pugh Matrix ground rules and assumptions include:

• Number defines whether technology “helps” or “hurts” a specific vehicle measure of merit 

• Secondary technology effects were ignored unless they were significant 

• L/D and SFC refer to long-term cruise values, not short-term low speed

• Emissions includes NOx, CO2, and water contrails

• Noise is community noise on the ground at takeoff and landing, it does not include boom or cabin 

noise

• Cabin noise falls under the category of “passenger comfort”

• Some technologies are already considered to be on the baseline (e.g., Inverted V).

• Cost/Complexity refers to how much more or less “complicated” the vehicle becomes as the 

result of adding a technology

• Gray areas and exceptions were unavoidable due to the nature of assigning values to each 

technology 

• The baseline vehicle has been designed to meet the noise, boom and emission constraints as best 

as possible

Once the rankings of each technology had been accomplished, a weighted product function 

determined the overall net benefit values. Based on these values, we were then able to sort the technology 

based on high and low values. Figure 87 highlights the best and worst results from the Pugh Matrix effort.
NASA/CR—2010-216796 80

Figure 87.—Observations and insights from rankings contributed to prioritization.

Technologies rated the highest included adaptive/inflatable wing elements, distributed roughness, and 

the jet noise technology. Adaptive geometry produced a net Pugh Matrix benefit equal to +10. The reason 

adaptive geometry scored so high was due to the technologies broad range of potential applications. 

Control augmentation, inlet flow control, and adaptive OML were the major applications considered for 

adaptive geometry. This was the highest score followed by distributive roughness with a net benefit value 

equal to +9. Distributed roughness scored high because it had a positive impact on L/D with no negative 

impacts on other measures of merit. Further research and analysis proved that distributed roughness could 

provide a 1.0 to 1.5 point increase in L/D with minimal increase in system complexity. The jet noise 

technology set scored a net benefit value of +7 because the technologies are instrumental in reducing the 

overall noise levels to reach the N+3 goals. 

Technologies that were rated the worst include the high aspect ratio nozzle and plasma virtual length. 

The high aspect ratio nozzle scores low because it had a small noise benefit but has a high negative 

impact on weight and efficiency. The technology’s overall score equaled –12. Plasma virtual length 

scored a –6 because of the complexity issues associated with an additional system and its large power 

requirements. Most of the vehicle impacts from negatively scored technologies were small and only 7 out 

of 51 technologies resulted in values from –3 to –1. 

As a result, the following table (Table 11) demonstrates the nonprioritized list of technologies that 

were down-selected for further analysis. Each technology was selected based on a judgment call, the Pugh 

Matrix, and considerations to ultimately apply these technologies to an N+3 demonstrator. 

A technology that did not make the list, but is recommended for further study and analysis is aircraft 

take-off assist (Table 12). Take-off Assist, under aircraft operations, uses a subsonic vehicle to assist a 

supersonic vehicle during take-off in a similar manner to a glider being towed. The goal is to reduce the 

overall noise levels at take-off. Both vehicles would then land by themselves. This type of operation could 

also potentially increase efficiencies within the configuration design. This technology scored reasonably 

* click a category name to sort 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0

Technology Name Empty Weight L/D SFC Boom Noise Emissions Safety Pax Comfort Net Effect

adaptive/inflatable wing 

elements 6 3 1 10

distributed roughness 9 9

bleedless external compression 

inlet 1 6 7

with SiC fibers in efficient 

weaving or braiding process 1 6 7

morphable control surfaces 

(seamless gaps) 3 3 1 7

structural health monitoring 1 6 7

pre-mixed combustion -1 1 6 6

Goals

* click a category name to sort 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0

Technology Name Empty Weight L/D SFC Boom Noise Emissions Safety Pax Comfort Net Effect

quiet spike -3 -1 3 -1

variable sweep outboard wing 

panels -3 1 1 -1

non-round nozzles (rectangular 

or elliptical) -3 -1 1 -3

windowless cabin 6 -9 -3

swing canard -3 1 -1 -3

plasma virtual length -1 1 -9 3 -6

high aspect ratio nozzle -9 -6 3 -12

Goals

Worst

Best
NASA/CR—2010-216796 81

high (Pugh Matrix value = +5), but the impacts of this system within the future ATS are unclear. The 

system requires further research and analysis to understand the full impact.

TABLE 11.—NONPRIORITIZED OF PUGH MATRIX TECHNOLOGIES DOWN SELECTED FOR FURTHER ANALYSIS

Category Technology Source Description Benefits Drawbacks Benefit value 

(Pugh matrix 

analysis)

Laminar 

flow

Distributed 

roughness Reed/Saric

Distributed bumps on a/c 

surfaces. Creates acoustic 

disturbances out of phase with 

cross-flow disturbances and 

maintains laminar flow.

Lower drag at 

supersonic speeds

Manufacturing 

complexity +9

Adaptive 

geometry 

Inlet flow control LM RTP Vary inlet geometry to choke 

flow in the inlet at low speeds

Reduced inlet fan 

noise 

Complexity of 

using variable 

geometry

+3

Lift distribution 

control 

N+3 Kick-off 

brainstorming 

session

Ability to conform portions of 

configuration/ control lift 

surfaces

Ability to control lift, 

decreased drag, lower 

noise

Potential difficulties 

to create drag at 

low speeds

+7

Plasma Jet exhaust 

manipulation 

Tech Fair—

Ohio State

Plasma actuators on nozzle lip 

to play with jet characteristics 

to reduce noise

Jet noise reduced about 

2 to 3 dB

Power required for 

plasma +3

Pilot 

situational 

awareness 

Low boom 

violation cueing LM

Heads up display illustrating 

maneuver and acceleration 

limits for low boom operation

Mitigate boom and 

operate a/c within 

FAA limitations

+9

Boom/noise 

preplanning 

awareness 

LM Route planning for noise 

abatement

Intelligent route 

finding to apply noise 

to specific areas

Performance and 

efficiency decrease 

to abate noise/boom

+9

Tools/

Methods 

Boom shaping LM Shapes lifting surfaces to low 

boom targets Higher fidelity solution Computational Requirements +9

CFD-based 

MDAO 

LM/Stanford/

NASA

Shapes lifting surfaces to low 

boom targets Higher fidelity solution Computational Requirements +9

Integrated 

structural 

analysis 

LM

Integrated analysis that 

accounts for flutter and other 

ASE concerns

Higher fidelity/

integrated solution

Computational 

Requirements +9

TABLE 12.—RECOMMENDED ENABLING TECHNOLOGIES FOR FURTHER RESEARCH AND ANALYSIS

Category Technology Source Description Benefits Drawbacks Benefit value 

(Pugh matrix 

analysis)

Aircraft 

operations 

Take-off 

assist LM 

A minimalist autonomous takeoff 

assistor would tow the supersonic 

aircraft by cable through take-off up 

to launch altitude. The supersonic 

aircraft would have its engines at 

partial throttle during takeoff to stay 

within noise limits and to be 

controllable. The supersonic aircraft 

would land on its’ own.

Reduces 

community 

noise on takeoff

Potential 

operation 

problems 

within future 

ATS

+5

The technology down selection process produced a manageable list of promising technologies for 

further evaluation and identified enabling technologies suitable for the technology roadmaps. 

Nonproprietary results from the Pugh Analysis are posted in Appendix B.

5.3 Technology Prioritization (WBS 3.3.4)

Technology prioritization was done by calculating value in terms of the overall benefit of the 

technology to the air vehicle, probability of success, and the cost to develop that technology to a TRL 6. 

The equation used is highlighted below:
NASA/CR—2010-216796 82

Cost to Develop

Benefit Probability of Success Rank ∗ =

The results from the Pugh Matrix analysis were used to determine a benefit value for each 

technology, and referenced subject matter experts within LM were used to determine the probability of 

success and the overall program costs to mature each technology. Each technology was then prioritized 

and put into the following roadmap format, Table 14, to illustrate the enabling airframe and propulsion 

technologies required to realize an environmentally friendly vision vehicle.

The down selected technologies have applications to N+2 and N+3 air vehicles. Different 

technologies, including low boom violation cueing and preplanning awareness, first have application to 

the N+2 vehicle but will also be used on the N+3 vehicle to achieve the N+3 performance and 

environmental goals. It was reasonably assumed that these N+2 technologies, highlighted in Table 13, 

would be developed during subsonic N+2 programs and would not be a focus for future N+3 technology 

development. As a result, these technologies were prioritized and technology roadmaps were developed 

for completeness. However, it is assumed that these technologies will be developed in other efforts first 

and are not recommended for future N+3 development efforts. 

We recommend that future airframe development should concentrate on N+3 specific 

tools/methodologies, laminar flow, and adaptive geometry technologies. The order of importance is 

outlined in the table and follows the priority ranking provided. 

TABLE 13.—PRIORITIZED LIST OF ENABLING TECHNOLOGIES NECESSARY TO 

ACHIEVE N+3 PERFORMANCE AND ENVIRONMENTAL GOALS

Priority 

no.

Category Prioritized, 

enabling N+3 

technology

Addressed 

N+3 goal

Vehicle benefit Current 

TRL 

level

Vehicle 

application

Availability

timeframe 

(TRL = 6)

1 Tools/

methods 

Boom shaping Boom 

mitigation

Optimized boom 

configuration capable of 

achieving 70 PLdB

1-2 N+2 and N+3 

supersonic 

application

~2015

2 Tools/ methods CFD-based 

MDAO 

Boom 

mitigation

Optimized boom 

configuration capable of 

achieving 70 PLdB

2-3 N+3 ~2016

3 Laminar flow Distributed 

roughness with 

plasma aug.

Fuel efficiency, 

range

L/D +1 to 1.5, Increased 

efficiency at higher Machs, 

less friction/drag, reduction 

in aircraft wt.

3 N+2 and N+3 

supersonic 

application

~2022

4 Pilot 

situational 

awareness 

Low boom 

violation cueing 

Boom 

mitigation

In-flight boom intelligence, 

reduced sonic boom, 

operation within FAA 

limits

1-2 N+2 ~2018

5 Pilot 

situational 

awareness

Boom/noise 

preplanning 

awareness 

Boom 

mitigation

Intelligent/Boom 

conscience route finding, 

operation within FAA 

limits

1-2 N+2 ~2020

6 Tools/ methods Integrated 

structural 

analysis 

Efficiency Optimized empty weight 

design

1-2 N+3 ~2017

7 Adaptive 

geometry

Lift distribution 

control 

Efficiency, 

noise

Adaptive airfoils to 

optimize range

2-3 N+3 ~2018

8 Plasma Jet exhaust 

manipulation 

Noise Reduced nozzle PLdB 

noise by 6-8 dB

2-3 N+3 ~2016

11 Adaptive 

geometry 

Inlet flow 

control 

Noise, 

efficiency

Reduced inlet fan noise 

with improved efficiency 

(+10% inc in take-off thrust 

and transonic thrust. Plus 

10% subsonic SFC 

improvement)

2-3 N+3 ~2018
NASA/CR—2010-216796 83

6.0 Technology Roadmap Development—Propulsion

6.1 Technology Identification—Potential Enabling Technologies

6.2 Technology Down Selection/Prioritization

Early in the program a list of propulsion technologies were compiled and prioritized, shown in

Table 14. This prioritization was in terms of engineering judgment as to the benefit to the program, the 

chance of the technology succeeding if it was fully developed, and the cost of development of the 

technology. This prioritized list was used as the program developed to guide the choice of technologies to 

be evaluated in the propulsion system. As the assessments relative to the goals became clearer and as the 

configurations LM was focusing on became evident, some of these technologies made more sense and 

some were no longer appropriate to include. As discussed in Section 4.1, the advanced VCE, as well as 

the conventional MFTF use many advanced technologies aimed at 2025 TRL 6 Technology Availability 

Date and EIS date of 2035. The main advanced technology features of the propulsion system broken out 

for more detailed evaluation and trade studies are the VCE architecture engine, the ATMS, the Axi-Plug 

Low Noise Exhaust, and the Transonic Thrust Augmentation device. There is also an Advanced Low 

NOx Combustor concept capable of meeting the N+3 cruise NOx emission goal. There were also some 

additional technologies evaluated that were not part of the baseline propulsion system. The additional 

technologies aimed at reducing T4 to meet the cruise NOx emissions goal were intercoolers, interturbine 

combustion, and to improve SFC, constant volume combustion. These trade studies will be summarized 

in the following sections.

TABLE 14.—PROPULSION SYSTEM TECHNOLOGY MATRIX

Priority Technology Benefit to 

program

Chance of 

success

Cost to 

develop

Ranking

1 Exhaust noise reduction technologies 9 9 1 81

1 ATMS 9 9 1 81

1 Transonic thrust augmentation 9 3 1 27

2 Premixed combustion 9 3 1 27

2 Active control combustion 9 3 1 27

1 VCE 9 9 9 9

1 Advanced low NOx combustor 9 3 3 9

1 High OPR gas turbine 9 9 9 9

5 Over-wing mounted engines 3 3 1 9

7 Constant volume combustion 9 3 3 9

13 Cross-shafted auxiliary fans 3 3 3 3

3 Ceramic matrix composite exhaust 3 3 3 3

4 Inter-cooler 3 3 3 3

6 Embedded engines 3 3 3 3

8 Counter-rotating fan 3 3 3 3

10 Indirect Brayton cycle 3 3 3 3

12 Auxiliary turbofans 9 1 9 1

9 = high, 3 = medium, and 1= low

7.0 GOTCHA Analysis (WBS 3.3.3)

Using a NASA tool, each enabling technology was evaluated as to how it addressed an N+3 goal, and 

this path was then mapped using GOTCHA charts. Each N+3 goal (sonic boom, airport noise, cruise 

emissions, cruise speed, range, payload, and fuel efficiency) was laid out with corresponding technical 

challenges and resulting approach to solve those objectives. The results, which prove traceability for each 

technology to an N+3 goal, are shown in Figure 88 and Figure 89. 
NASA/CR—2010-216796 84

Figure 88.—GOTCHA successfully traces prioritized technologies to N+3 environmental goals.
NASA/CR—2010-216796 85

Figure 89.—GOTCHA successfully traces prioritized technologies to N+3 performance goals.

Payload

100-200 pax

SOA = 92-120 

pax

Cruise Speed

Mach 1.3 – 2.0 low 

boom flight

Mach 1.3 – 2.0 

unrestricted

SOA = Stg 1 Concorde 

@ M=2.2

Determine 

Optimum 

Aircraft/Wing 

configuration and 

technologies 

required to 

achieve 4000-

5500 nm Range. 

Use adaptive 

geometry to 

opt. Airfoils 

for a range of 

Mach #’s and 

control 

deflections 

for better lift 

distribution to 

inc. range

Base lining a 

variable cycle 

engine

Supersonic 

design is 

hindered by 

boom shaping

Using bleedless external 

compression 

inlets to lighten 

inlet weight

Reduce overall 

wt. of a/c to 

account for pax. 

– Adv. Materials, 

compression 

inlets, increase 

in L/D, laminar 

flow tech, adv 

analysis 

methodologies

Sizing a 

vehicle for 

pax safety 

and comfort

Turbofans 

sufficient at 

noise 

suppression 

and 

performance 

efficiency

GOALS

TECHNICAL CHALLENGES

APPROACHES

OBJECTIVES

Configuration 

Shaping

Range

4000-5500 nm

SOA = Stg 1 Concorde 

@ 3,900 nm

Determine 

optimum Cruise 

Speed for N+3 

vehicle

Determine 

optimum 

Cruise 

Speed for 

N+3 

vehicle

Laminar flow 

technology to 

increase L/D 

by1.0 – 1.5 

Fuel Efficiency

3.5-4.5

(pax-nm/lb fuel)

SOA = Stg1 Concorde @ 

1.75 pax-mi/lb fuel

Design high 

efficiency, 

small, light 

weight 

engine

Optimize 

system to 

achieve 3.5-

4.5 PAXnm/lb fuel

Low noise 

acoustic 

goal – low 

Vjet

Low cruise 

NOx 

emissions 

– low T 

and T41

Small light 

weight 

engine 

with low 

SFC

Maximizin

g thrust at 

tran-sonic 

pinch point

VCE with 

ATMS

Advanced 

low NOx 

Combustor

Noise 

Reduction 

Technolog

ies in 

Exhaust

High 

Performance 

Axi-Plug 

exhaust

Determine 

best engine 

for cruise 

efficiency

Light weight, 

low loss 

thrust 

augmentation
NASA/CR—2010-216796 86

8.0 New Technology—Airframe (WBS 3.3.3, 3.3.4)

The following section describes each prioritized technology and establishes which N+3 goal the 

technology addresses. Waterfall charts were also generated for each enabling technology to demonstrate 

the maturation plans for Phase II and Phase III efforts. All technology plans illustrate the maturation cycle 

up to TRL 6 and are grouped according to category. Each technology was assessed by considering the 

technology’s current TRL and an understanding of the resources that would likely be needed to raise the 

TRL. The results show that all technologies can achieve TRL 6 within the 2030 to 2035 timeframe and 

major risk items are identified.

8.1 Tools /Methodologies—Low Boom Shaping Fidelity

8.1.1 Addressed N+3 Goals

• Boom mitigation

8.1.2 Description

Current low sonic boom design and analysis methods have less fidelity (to various degrees) than is 

needed to achieve low boom loudness goals with the best aerodynamic performance simultaneously 

possible. Where possible, fast low fidelity methods will be replaced or modified with enhanced accuracy 

methods including more faithful geometric modeling. Low fidelity methods will work incrementally from 

high-fidelity methods in mixed-fidelity and multi-fidelity frameworks. Improved investigation and 

optimization schemes will find better solutions and better Prato fronts for trading-off between objectives.

CAD incorporation is desired for parametric geometry modifications that maintain smoother and more 

producible shapes, are better for other disciplines, and still achieve near optimum aero/boom 

performance. Improvements in CFD analysis accuracy and speed will be implemented through coding 

automation, more efficient gridding, improved propulsion boundary conditions, wind tunnel and flight 

measurement validations and propagation beyond the near-field. Additional enhancements to propagation 

methodology include rounding, turbulence, maneuvering and additional flexibility in inverse 

minimization targets.

8.1.3 Roadmap

Figure 90 delineates a broad course of action necessary to mature and develop these boom shaping 

enhancements. The main maturation efforts include parallel work with NASA on multiple means and 

methods of fidelity improvements for modified-linear tools, multi-fidelity frameworks and CAD 

integration, CFD improvements and inverse design targets. Once complete, verification and validation 

needs to be performed to ensure usability by personnel within the aerospace industry.

Figure 90.—Roadmap for tools/methods—low boom shaping fidelity.
NASA/CR—2010-216796 87

Figure 91.—Roadmap for tools/methods—CFD-based MDAO.

8.2 Tools/Methodologies—CFD-Based Adjoint Shape Optimization

8.2.1 Addressed N+3 Goals

• Efficiency

• Sonic Boom

8.2.2 Description

CFD has been identified as the necessary tool for accurate low sonic boom prediction. The 

implementation of an effective adjoint of sonic boom in CFD will greatly improve the effectiveness of 

low boom CFD-based design. Achieving low sonic boom with the highest performance efficiency 

possible requires final refinement of configurations with CFD-based design. Strong integration with CAD 

geometry definition will also be part of the process to insure geometric fidelity and final shapes practical 

for all disciplines, like smooth shapes for efficient structure.

8.2.3 Roadmap

The waterfall chart in Figure 91 shows the general outline of CFD-Based sonic boom adjoint MDO 

maturation plans from TRL 2-3 to TRL 6. Development of the tool is ideally done in conjunction with 

Universities and NASA, and then it must go through the verification and validation process. Total code 

development verification matures the technology to TRL 6 in approximately 2016. 

8.3 Laminar Flow—Distributed Roughness with Plasma Augmentation

8.3.1 Addressed N+3 Goals

• Drag reduction

• Fuel efficiency

8.3.2 Description

The distributed roughness technology consists of distributed bumps on wing surfaces to create 

acoustic disturbances out of phase with cross-flow disturbances to maintain laminar flow on swept 

surfaces. This technology anticipates extending the laminar flow to the majority of the vehicle’s surface.

Laminar flow reduces friction drag by 80 percent for the portion of the surface where it can be attained, 

greatly increasing fuel efficiency and range. 

During the Phase II extension of the Quiet Supersonic Platform program conducted by DARPA, the 

Arizona State University research team headed by William Saric and Helen Reed conducted a series of 
NASA/CR—2010-216796 88

tests in the NASA Langley 4 by 4 Unitary plan tunnel (UPWT) culminating during the period of January 

to February 2004. The UPWT work was designated as Task 6 of the overall ASU effort. The purpose of 

this task was to achieve NLF on a swept wing at Mach 2.17 using spanwise-periodic discrete roughness 

elements (now known as DRE)—a technology pioneered by William Saric and used in both high-speed 

and low-speed flow. The results at higher unit Reynolds numbers are encouraging and justify a 

continuation of this program of using DRE for NLF in supersonic flow sometime in the future.

Further technology development of distributed roughness is currently being conducted with Helen 

Reed, Bill Saric, and the Air Force Research Lab (AFRL) using plasma generation and spacing control of 

DREs to maximize NLF on the Revolutionary Configuration for Energy Efficiency (RCEE) program. The 

plasma can be generated with any desired roughness spacing for optimal effect at all flight conditions. 

This technology could also be used to prevent separation with shock/boundary layer interactions to 

improve efficiency. In addition to increased control for maximizing NLF, the plasma generator has many 

good operational features. It requires little power, has no moving parts, is flush mounted, is compact in 

size, and can be varied at high frequency.

8.3.3 Roadmap

Figure 93 highlights the maturation plan for plasma assisted natural laminar flow. Currently, wind 

tunnel work on the RCEE program in conjunction with AFRL and Reed/Saric at Texas A&M augments 

the overall maturation plan. Three additional maturation tests are required to bring the TRL from 3 to a 6. 

The first step is a subsonic flight test at Texas A&M. The second is a transonic flight test using a wing 

glove on a business jet. The third test, occurring well into Phase 3, includes a large scale supersonic flight 

test to push to TRL 6. 

Figure 92.—Laminar flow technology benefits applied to advanced vehicle concept.

Figure 93.—Roadmap for laminar flow technology—plasma augmented distributed roughness.

Lower

Upper

Distributed Roughness (DR) & Flow Transition (FT): 

Distributed roughness is Applied to Leading Edges of the 

Wing, Tail & Canard surfaces for Increased Lifting Efficiency. 

MDAO Analysis Indicates that L/D will be Increased by 1 to 

1.5
NASA/CR—2010-216796 89

Figure 94.—Roadmap for pilot situation awareness—low boom violation cueing.

8.4 Pilot Situational Awareness—Low Boom Violation Cueing

8.4.1 Addressed N+3 Goals

• Boom Mitigation

8.4.2 Description

Low Boom Violation Cueing is a pilot situational awareness tool that is assumed to be developed first 

for the N+2 vehicle. It is included within this list because it has N+3 applicability’s. The tool provides 

warnings on a heads-up display about maneuvers, turning rate limits, and accelerations that could trigger 

a breakdown of the vehicle’s low boom characteristics. It also provides mach cut-off indication for flying 

without boom on the ground to further help minimize boom impact. This will enable lower speed 

supersonic flight even before acceptable low boom level regulations are implemented.

8.4.3 Roadmap

Figure 94 shows the general development plan for low boom violation cueing. The first step to mature 

the technology includes gathering atmospheric data, taking many ground measurements to quantify 

statistical scatter from predictions and adjusting PCBOOM code to calculate focused boom conditions. 

Results from that effort will be incorporated with airplane system interfaces. Lastly, verification and 

validation matures the code to TRL 6. 

8.5 Pilot Situational Awareness—Preplanning Awareness

8.5.1 Addressed N+3 Goals

• Boom Mitigation

• Noise

8.5.2 Description

Preplanning awareness is a route planning and pilot situation awareness tool that is assumed to be 

developed for the first low boom vehicle, N+2. It is included within this list because it has N+3 

applicability’s. Preplanning awareness provides information about noise levels and flight path data to 

reduce boom impact over land. The tool can optimize over land operation and help plan routes to 

minimize sonic boom noise annoyance. Later incarnations can also help optimize airport noise abatement 

procedures for specific community routing, runways and engine throttling.
NASA/CR—2010-216796 90

Figure 95.—Roadmap for pilot situation awareness—preplanning awareness.

8.5.3 Roadmap

Figure 95 highlights the maturation plan necessary for preplanning awareness. The first step is to 

develop prediction capabilities. Next, determine pilot interface and system interfacing requirements. 

Finally, pilot interfaces with ATC need to be developed, implemented and ultimately verified. 

8.6 Tools/Methods—Integrated Structural Analysis

8.6.1 Addressed N+3 Goals

• Efficiency

8.6.2 Description

Low sonic boom design and optimization is initially performed simultaneously with aerodynamic 

optimization. Any method that provides the lift needed for boom also provides the drag and moment 

needed for aerodynamic efficiency. While both boom and aerodynamics are greatly dependant on vehicle 

surface geometry, another discipline is also greatly dependant on that geometry—structures. Structural 

strength and stiffness are even higher power functions (third and fourth) of thickness than drag (second) 

and rolls up into overall efficiency through structural weight. Further, since length and weight are primary 

parameters affecting boom, structural optimization is not only of primary importance to surface geometry 

design but strongly interacts with boom and aerodynamics. Based on their strong geometry design 

dependence, sonic boom, aerodynamics and structures will benefit from and employ the most 

sophisticated optimization methodology possible. Based on their strong interactions, further benefit 

should be achievable by optimizing all three disciplines in an integrated framework. Additionally, the 

particularly slender configurations of low boom designs will often need to employ more sophisticated 

structural analyses for flutter suppression, aeroelastic control, aero-servo-elasticity (ASE and even APSE, 

which adds propulsion interactions in designs without a tail or braced nacelle mount). The best methods 

for these sophisticated analyses need to be integrated into the low boom design framework so that all 

critical structural limitations are accurately addressed. Past commercial aircraft program trends have 

shown that investments in better initial design phases tends to pay-off greatly in reduced rework, lower 

overall development cost, lower recurring cost and better aircraft performance.
NASA/CR—2010-216796 91

Figure 96.—Roadmap for tools/methods—integrated structural analysis.

8.6.3 Roadmap

Figure 96 illustrates a high-level maturation plan for integrated structural analysis. Much of the plan 

relies on the integration of structural tools developed outside of this supersonic program. Most structural 

tool development will be provided by third party software developers. They may need some guidance to 

insure that their tools can be integrated into the low boom design framework. Validation of software 

analysis needs to be performed on a representative low boom design. Following development needs to be 

requested from the third party software developers for any validation cases with unsatisfactory results. 

Additional support may be needed from the third party software vendor to integrate their tools effectively 

into the low boom framework. 

8.7 Adaptive Geometry—Lift Distribution Control

8.7.1 Addressed N+3 Goals

• Efficiency

• Noise

8.7.2 Description

Adaptive geometry—lift distribution control is performed using shape-memory alloys or piezoelectric 

actuators to bend flexible surfaces without gaps. This adaptive geometry can improve control efficiency 

and reduce the overall airframe noise. Airfoils could also be optimized for over-water operation for 

improved lift distributions to increase range. Future research is needed to develop the methods and 

materials.

8.7.3 Roadmap

Figure 97 demonstrates the developments needed to mature lift distribution control. The first step is to 

demonstrate autonomous adaptive wing materials including tension/compression structures and tunable 

modulus composites. Next, the development of integrated shear stress sensing materials along with a 

wind tunnel demo of the 2–D integrated system brings the technology to TRL 4. TRL 6 is achieved with a 

3–D system design, fabrication and a demonstration.
NASA/CR—2010-216796 92

Figure 97.—Roadmap for adaptive geometry—lift distribution control.

8.8 Plasma—Jet Exhaust Manipulation

8.8.1 Addressed N+3 Goals

• Noise

8.8.2 Description

Jet Exhaust Manipulation uses high bandwidth and amplitude plasma actuators on the nozzle lip to 

adjust flow in such a way to reduce jet noise (and perhaps sonic boom). The technology development is 

currently being conducted by Dr. Samimy from Ohio State University. Research is needed to develop the 

most effective system: reduce power requirements, determine best amplitude, flow stream depth and 

circumferential location, frequency and circumferential modulation, and optimize integration with 

acoustic liners. The plasma effectiveness on the exhaust flow and mixing is enhanced by matching the 

natural eddy formation frequency. Further enhancement is achieved with circumferential variations and 

variation in strength, nozzle position and depth of plasma location. These parameters are just starting to 

be investigated on a simple conical jet. Thereafter, development will need to advance to the 3-stream 

nozzle flow of proposed VCEs (with perhaps a greater effectiveness potential).

8.8.3 Roadmap

Figure 98 demonstrates jet exhaust manipulation maturation plans continuing the simple nozzle 

studies with improved hardware to investigate the full range of effective power, position and frequency. 

Parallel efforts will model the flow effects and help to identify the most optimum plasma configuration, 

particularly for transition to the proposed multi-stream nozzle flow and integration with acoustic liners. 

Testing with the multi-stream nozzle set-up follows with multiple entries and final testing at near flight 

exhaust conditions and temperature and Reynolds number.
NASA/CR—2010-216796 93

Figure 98.—Roadmap for plasma—jet exhaust manipulation.

Figure 99.—Roadmap for adaptive geometry—inlet flow control.

8.9 Adaptive Geometry—Inlet Flow Control (Choking)

8.9.1 Addressed N+3 Goals

• Noise 

• Efficiency

8.9.1.1 Description

Inlet Flow Choking uses variable geometry to choke the inlet at low speed. This technology reduces 

and prevents the fan noise from propagating out of the inlet. Development within LM has been conducted, 

but the technology has not been demonstrated. Research has also shown 10 to 20 percent increases in 

take-off and transonic thrust as well as 10 percent improvements in subsonic SFC. Additional research is 

required to mature and verify the technology.

8.9.2 Roadmap

Figure 99 illustrates a high-level maturation plan for the technology. First, a design requirements 

study guides the fabrication and testing of a demo. Next, concept refinement is done with additional 

design and testing. Ultimately, a full-scale demo brings the TRL to 6 in approximately 2018. 
NASA/CR—2010-216796 94

Figure 100.—Airframe technology roadmap for enabling supersonic technologies.

8.10 Technology Roadmap Overview

Figure 100 highlights the general timeline that each prioritized, enabling technology will reach a 

TRL 6. Assuming sufficient funding, all technologies are within an acceptable timeframe to apply to an

N+3 demonstrator.

9.0 New Technology—Propulsion

Based on the initial technologies prioritized in Table 15, the process of designing the variable cycle 

propulsion system, and the assessment of how the propulsion system does relative to meeting the 

environmental goals, several technologies need to be developed for the proposed propulsion system.

Technology development roadmaps are provided for the VCE, low noise/high performance nozzle, 

advanced low NOx combustor, ATMS, Transonic Thrust Augmentation device, and constant volume 

combustion. All of these technologies are included in GE’s baseline propulsion system except for the 

constant volume combustion system. CVC was included in the roadmaps because the trade study, 

although relatively simple, does show significant SFC benefits and the current study didn’t get into the 

system weight and installation portion, there is thought to be benefit to continuing to look at this 

technology.

9.1 VCE

9.1.1 Goals and Objectives

1. Develop advanced VCE for supersonic commercial application 

2. Develop ATMS 

3. The engine will meet the NASA emissions, efficiencies, and noise goals allowing the aircraft to 

meet sonic boom goals

4. The engine should be designed to achieve low SFC for supersonic while providing lightweight, 

high reliability and long life.
NASA/CR—2010-216796 95

Figure 101.—Roadmap for VCE propulsion system.

9.1.2 Technical Description

The required elements to mature this propulsion system are outlined in Figure 101; the primary 

elements include technology and component maturation, continued advances and developments in 

advanced aerodynamics and materials.

An N+3 engine demo program could potentially start around 2019 depending on funding availability.

It is anticipated that many of the required technologies may already be at TRL 6 before that, but some 

most likely will not. This N+3 engine demo program would provide the opportunity to mature the 

remaining technologies and verify the capability to proceed to the full system development program for 

the N+3 Supersonic program. The initial engine demo program is currently unfunded. The demo engine is 

the potential test bed for technology maturation to TRL 6. If this demo test bed does not occur then all the 

technologies will have to be matured to TRL 6 in the later N+3 engine demo.

9.2 Low Noise/High Performance Nozzle Exhaust

9.2.1 Goals and Objectives

1. Exhaust system that meets operability, performance, and noise goals, utilizing the cycle based on 

the VCE.

2. Exhaust system to meet NASA N+3 Airport Noise Goal—Stage 3—20 to 30 EPNdB

3. High nozzle performance throughout flight envelope.

4. Light weight, high reliability, variable geometry

9.2.2 Technical Description

Propulsion system studies have shown that an advanced VCE with an exhaust system comprised of an 

Axi-Plug nozzle with jet noise reduction features can meet the N+3 supersonic airport noise goals. The 

basic Axi-Plug nozzle provides high performance with variable geometry. The VCE enables the jet noise 

reduction features in the nozzle.

9.2.3 Roadmap

Figure 102 shows the roadmap for the low noise/high performance exhaust system.

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NASA/CR—2010-216796 96

Figure 102.—Low noise/high performance exhaust system roadmap.

The next step in the roadmap is to design, fabricate, and test an exhaust system consisting of the AxiPlug nozzle jet noise reduction technologies. This will verify how these technologies work together. After 

that an exhaust system could potentially be matured in the potential engine demo or an N+3 engine demo.

9.3 Advanced Low NOx Combustor

9.3.1 Goals and Objectives

1. Develop advanced combustor with enhanced premixing capable of achieving EINOx = 5 at 

supersonic cruise conditions.

2. Develop advanced materials and cooling technologies to reduce cooling flow requirements.

3. Evaluate and develop advanced piloting systems to approach entitlement emissions, and improve 

combustion dynamics/stability.

9.3.2 Technical Description

Various system level approaches need to be examined to achieve the aggressive N+3 emission goal of 

EINOx = 5 at cruise conditions. If a variant of the single-stage Twin Annular Premixed Swirler (TAPS) 

combustor will be used, substantial improvements in mixing effectiveness will be required to create a 

lean-burn premixed combustion system capable of achieving these low NOx levels. If a TAPS variant is 

not feasible, other advanced mixing concepts such as Lean Direct Injection (LDI) or Accelerating Swirler 

Passage (ASP) may be employed to achieve a near-perfectly premixed condition.

Advanced cooling technology must be developed, which will reduce the amount of cooling flow 

needed and direct more air to the mixer. New materials such as Ceramic Matrix Composites (CMC) can 

be used to construct the combustor to reduce cooling requirements as well. 

Advanced piloting systems will be evaluated and developed, which will eliminate the high NOx 

production from the traditional pilot diffusion flame, and improve combustion dynamics/stability.

9.3.3 Roadmap

The combustor roadmap is shown in Figure 103. At the start of the supersonic N+3 program, the 

combustor was identified as the key enabling technology to achieve the N+3 cruise NOx goal. A NASAfunded combustor program is already underway aimed at supersonic vehicles and will be able to provide 

important information towards achieving this goal. The supersonic N+1 program plans on delivering 

hardware by mid-2010.

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NASA/CR—2010-216796 97

Figure 103.—Low emission combustor roadmap.

Based on information obtained from these programs, the supersonic N+3 combustor program should 

develop advanced concepts in rapid mixing, reduced residence time, reduced cooling, reduced pilot 

contribution, and increased flame stability, by approximately 2022. To be in service by the 2035 timeframe, an 

engine demo (TRL–6) would be required by about 2025, followed by a flight test by about 2031. 

9.4 ATMS

9.4.1 Goals and Objectives

1. Develop ATMS that permits high OPR operation at supersonic cruise conditions

2. Minimal impacts on duct losses or overall engine performance.

3. Light weight, high reliability, long life

9.4.2 Technical Description

The propulsion system gets a significant SFC improvement by running very high OPRs, but in doing 

so the T3 gets very high. Analysis has shown the ATMS system to contribute a significant improvement 

in SFC.

9.4.3 Roadmap

Figure 104 shows the roadmap for the development and maturation for the ATMS system for the N+3 

Supersonics application. The first step is a core engine demonstration. After this a heat exchanger appropriate 

for the N+3 supersonic application could be developed in preparation for an engine demonstration.

9.5 Transonic Thrust Augmentation Device

9.5.1 Goals and Objectives

1. Develop advanced augmenter capable of moderate ΔT augmentation for thrust boost at critical 

thrust pinched flight points

2. Minimal impacts on exhaust or overall engine performance.

3. Requires no flame holders or spray bars

4. Light weight, high reliability, long life

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N+3 Engine Demo

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NASA/CR—2010-216796 98

Figure 104.—Cooled cooling air roadmap.

Figure 105.—Transonic thrust augmentation device roadmap.

9.5.2 Technical Description

Previous internal work has evolved a Transonic Thrust Augmentation device. This lightweight simple 

system is capable of producing a large localized temperature augmentation, which mixes out quickly with 

minimal hot streak footprints that increases the overall temperature of the exhaust gases by a moderate 

amount while producing almost negligible losses.

9.5.3 Roadmap

Figure 105 shows the roadmap for the Transonic Thrust Augmentation device. The next steps are hot 

component then hot sector tests leading to an engine test.

9.6 CVC Roadmap

9.6.1 Goals and Objectives

1. Develop CVC for VCE to improve SFC.

2. Minimal impact to system.

3. Light weight, high reliability and long life.

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NASA/CR—2010-216796 99

9.6.2 Technical Description

Due to the current low TRL and complexity the initial development of incorporating CVC in a gas 

turbine is expected to occur for a nonaviation application and then leveraged for aviation applications.

9.6.3 Roadmap

Figure 106 shows the roadmap for the development of CVC.

9.7 Technology Roadmap Overview

Figure 107 highlights the general timeline that each prioritized, enabling technology will reach a 

TRL 6. Assuming sufficient funding, all technologies are within an acceptable timeframe to apply to a

N+3 demonstrator.

Figure 106.—CVC roadmap.

Figure 107.—Technology roadmap overview for propulsion technologies.
NASA/CR—2010-216796 100

10.0 Conclusions

10.1 Tasks and Trade Studies—Airframe

10.1.1 Advanced Concept Vehicle Definition

An initial concept with an inverted V-tail and under the wing engine placements was drawn as a 

reference to compare alternative configurations. After exploring the configuration design trade space it 

was concluded that the inverted V-tail configuration with under the wing engine placements provided the 

best platform to achieve the N+3 goals. Figure 108 highlights the final configuration that was ultimately 

sized to have a gross takeoff weight was 285,000 lb, with a thrust to weight ratio of 0.34 and a wing 

loading of 92 psf.

The final wing, canard and tails were optimized to reduce the sonic boom with minimal aerodynamic 

efficiency reduction. This optimization resulted in a reduction of the 1.38 psf max shock to 0.29 psf, and a 

reduction in L/D from 10.1 to 8.0. Distributed roughness and plasma technologies are outlined in the 

technology roadmaps as solutions to increase the L/D and achieve the 4000 to 5500 nmi range and 3.5 to 

4.5 pax-nmi/lb-fuel fuel efficiency goals.

10.1.2 Rapid Conceptual Design Low-Boom Design Process or Design Space Trade Studies

LM’s Rapid Conceptual Design (RCD) framework was instrumental in demonstrating that a lowboom aircraft, with adequate technologies, is able to meet the N+3 noise and mission performance goals. 

RCD was used to generate the target signature corresponding to an average loudness of 70 PLdB, shape 

lifting surfaces to meet that signature, size the N+3 configuration, and investigate the benefits of key 

technologies.

Figure 108.—Final configuration three-view.
NASA/CR—2010-216796 101

Figure 109.—Ground signature before and after aero/boom shape optimization

RCD-based aerodynamic/boom shape optimization was used to shape the surfaces of the wing, 

inverted V-tail, and canard to ensure that the aircraft’s ground signature conformed to the established, 

70-PLdB target. This markedly improved the aircraft’s ground signature as shown in Figure 109, and was 

accomplished by adjusting the camber slope at 20 specific locations to redistribute the configuration’s 

equivalent area distribution. Although the boom-optimized shape exhibited a lift-to-drag ratio 

approximately 21 percent lower than that of a purely minimum-drag shape, aerodynamic technologies 

such as NLF can counteract this effect and enable the N+3 vehicle to maintain mission performance that 

complies with its payload, range, speed, and efficiency goals.

RCD was subsequently used to size the configuration for maximum range at 285,000 lb MTOW. This 

resulted in a thrust-to-weight ratio and wing loading of 0.34 and 92 psf, respectively. Sizing was based on 

the VCE with thrust augmentation technology, which enabled the aircraft to achieve a balanced field 

length of 9970 ft. This engine also enabled the N+3 aircraft to reach 40,000 ft in 22 min, permitting it to 

climb above existing subsonic traffic prior to accelerating to Mach 1.6. The sized configuration’s range 

was limited by available fuel volume; morphing structure technology could relieve this constraint and 

increase range.

10.2 Tasks and Trade Studies—Propulsion

Preliminary designs for a VCE and two nonvariable mixed flow turbofans (MFTF) were developed 

using the Numerical Propulsion System Simulation (NPSS) code as part of the NASA N+3 supersonic 

program. This program studied advanced airframe and propulsion concepts for a supersonic commercial 

passenger airplane in the 2035 timeframe. In particular, the program set challenging noise, emissions, and 

efficiency requirements for an environmentally friendly aircraft. The proposed engine concept 

incorporated advanced technologies such as ATMS permitting high OPR and thermal efficiency, an AxiPlug exhaust with Noise Reduction Technologies, and a Transonic Thrust Augmentation device. The two 

MFTF engines did not use the Noise Reduction Technologies in the exhaust. The studies showed that 

noise and performance goals could be met by the VCE and the first MFTF within the context of 

continuing development of advanced technologies. The emission goal is projected to be met with the VCE 

and first MFTF propulsion systems through a combination of a more advanced TAPS type combustor and 

a small decrease in the peak temperature in the engine. To take advantage of the full thermal capability of 

these propulsion systems would require even more advanced technologies not previously considered, and 

alternate combustor or combustion approaches. The final engine was an “Optimized” cruise engine, 

which ignored the airport noise and emissions goals and focused solely on fuel efficiency in the smallest 

lightest propulsion system.

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NASA/CR—2010-216796 102

GE developed preliminary designs for a VCE and a MFTF propulsion system to meet the thrust 

requirements and program metrics, provided by LM. The overall efficiency figure of merit for both is 

projected to be better than the goal. Airport noise and assessments were performed for these propulsion 

systems and both are projected to meet the goal. The third “Optimized” cruise MFTF engine designed at a 

higher FPR provides the best FOM but would not meet the acoustic and emissions goals. An NPSS 

customer deck was developed for the VCE Propulsion System.

Cruise emission assessments were performed and the maximum T4 of the engine must be reduced to 

meet the cruise NOx goal. This results in a relatively small impact to SFC and propulsion weight for the 

VCE propulsion system. Other combustor concepts may be developed to regain the maximum thermal 

capability of this VCE in the future. The cruise “optimized” engine was not designed to meet the 

emissions goal.

10.3 System of Systems Analysis 

The N+3 vehicle’s impact on the civil air transport system (ATS) was analyzed by Purdue University 

in terms of productivity (passenger * knots), environmental impacts (carbon dioxide and NOx emissions), 

and fleet direct operating costs (DOC). Optimal allocations of N+3 supersonic transports- defined as that 

which maximizes either productivity or passenger time saved- were determined for various future 

scenarios, which were established by projecting different rates of N+3 availability and passenger demand 

growth. Maximum fleet-wide productivity was calculated for different limits on fleet CO2, establishing 

the trade space between ATS productivity and emissions in the presence of N+3 SST’s. More 

importantly, SST’s were consistently allocated to routes of 3000 to 5000 nmi, which is consistent with the 

N+3 range goal of at least 4000 nmi.

The impact of N+3 operations on the demand for airports and airspace, assessed by LM

Transportation and Security Solutions, will require some additional airport capacity during peak hour 

operations. However, most operations are likely to be scheduled during the periods of nonpeak 

conventional operations and will therefore take advantage of currently unused airport capacity. 

Furthermore, N+3 route networks between various airports are fairly well separated (assuming they are 

able to fly great circle routings) even for busy air corridors. Departure climbs to 10,000 ft at low speed are 

contained within 21 current National Air Traffic System (NAS) departure sectors; higher subsonic climbs 

to 23,000 ft involve nine additional, current sectors. A marginal increase in air traffic complexity and 

controller workload is predicted during these phases. The most significant increases in air traffic 

complexity and controller workload are incurred by the N+3’s transonic acceleration and supersonic 

climb, which lead to potential conflicts with existing subsonic traffic. These conflicts are most likely to 

occur between 34,000 and 39,000 ft; therefore, they can likely be alleviated by using the VCE to move 

transonic acceleration to 40,000 ft.

10.4 Technology Roadmaps—Airframe

Technology prioritization was done by calculating value in terms of the overall benefit of the 

technology to the air vehicle, probability of success, and the cost to develop that technology to a TRL 6. 

Each technology was then prioritized and put into the following roadmap format, Table 15, to illustrate 

the enabling airframe technologies required to realize an environmentally friendly vision vehicle. Risk 

analysis and maturation plans demonstrated that each technology addressed an N+3 goal and was capable 

of maturation within the 2030 to 2035 timeframe. It was also determined that different technologies were 

applicable to N+1 and N+2 air vehicles. These technologies, including low boom violation cueing and 

preplanning awareness (highlighted in the table), were added to the final technology list, but it assumed 

that they will be developed under other N+2 program efforts.
NASA/CR—2010-216796 103

TABLE 15.—PRIORITIZED AIRFRAME TECHNOLOGY ROADMAP

Priority 

no.

Category Prioritized, 

enabling N+3 

technology

Addressed 

N+3 goal

Vehicle benefit Current 

TRL 

level

Vehicle 

application

Availability

timeframe 

(TRL = 6)

1 Tools/ 

methods 

Low boom 

shaping 

Boom 

mitigation

Optimized boom configuration 

capable of achieving 70 PLdB

1-2 N+2 and N+3 

supersonic 

application

~2015

2 Tools/ 

methods

CFD-based 

MDAO 

Boom 

mitigation

Optimized boom configuration

capable of achieving 70 PLdB

2-3 N+3 ~2016

3 Laminar flow Distributed 

roughness

Fuel 

efficiency, 

range

L/D +1 to 1.5, Increased efficiency at 

higher Machs, less friction/drag, 

reduction in aircraft wt.

3 N+2 and N+3 

supersonic 

application

~2022

4 Pilot 

situational 

awareness 

Low boom 

violation cueing 

Boom 

mitigation

In-flight boom intelligence, reduced 

sonic boom, operation within FAA 

limits

1-2 N+2 ~2018

5 Pilot 

situational 

awareness

Boom/noise 

preplanning 

awareness 

Boom 

mitigation

Intelligent/Boom conscience route 

finding, operation within FAA limits

1-2 N+2 ~2020

6 Tools/ 

methods

Integrated 

structural 

analysis 

Efficiency Optimized empty weight design 1-2 N+3 ~2017

7 Adaptive 

geometry

Lift distribution 

control 

Efficiency, 

noise

Adaptive airfoils to optimize range 2-3 N+3 ~2018

8 Plasma Jet Exhaust 

Manipulation 

Noise Reduced nozzle PLdB noise by 6 to 

8 dB

2-3 N+3 ~2016

11 Adaptive 

geometry 

Inlet flow 

control 

Noise, 

efficiency

Reduced inlet fan noise with 

improved efficiency (+10% inc in 

take-off thrust and transonic thrust.

Plus 10% subsonic SFC 

improvement)

2-3 N+3 ~2018

Recommended future technology development efforts for airframe technologies include:

• Concentration on supersonic specific technologies for the N+3 vehicle to ensure serviceability in

the 2030 to 2035 timeframe

– Tools/methods

– Laminar flow

– Adaptive geometry

• First, develop low cost, high impact tools and methodologies such as Multi-Fidelity Low Boom 

Shaping and CFD-based MDAO to achieve low boom levels with minimum aerodynamic 

efficiency impacts and the most overall, multi-disciplinary optimum design. 

• Second, advance distributed roughness with plasma augmentation to achieve the maximum 

laminar flow possible by the N+3 timeframe

• Third, promote adaptive geometry technologies and mature adaptive geometry applications to 

maximize N+3 fuel efficiency.

10.5 Technology Roadmaps—Propulsion

GE and Lockheed jointly assessed the most beneficial technology elements of the propulsion system.

These are highlighted in Table 16. Technology roadmaps were developed for these items, leading to 

appropriate technology readiness level progression for the N+3 timeframe.

A list of key items for future work is listed below:

• Currently funded DOD programs are continuing the development of many of the VCE

technologies.

• The Transonic Thrust Augmentation device is a critical technology to continue developing.
NASA/CR—2010-216796 104

• The currently funded NASA Supersonics Low Emissions combustor program will provide key 

validation data for high temperature NOx levels, and the maturation of some enabling 

technologies. Alternate combustor/combustion concepts need to be explored, as these propulsion 

systems are developed to take advantage of the full thermal capability of the system.

• Predictive and design tool development in many areas need to be continued to be developed.

Aero-acoustics, both fan and jet noise tools are not at a stage to provide direct quantitative design 

inputs. Combustion and Emissions is another area where continued tool development is critical

TABLE 16.—PRIORITIZED PROPULSION TECHNOLOGY ROADMAP

Priority 

no.

Prioritized, enabling 

N+3 technology

Addressed N+3 goal Vehicle benefit Current 

TRL 

level

Vehicle 

application

Availability

timeframe 

(TRL = 6)

1 Low noise/high 

performance nozzle

Airport noise, fuel 

efficiency Meet noise and efficiency goal 3 N+3 ~2015

2 VCE Airport noise, fuel 

efficiency, sonic 

boom

Smaller, lighter propulsion system that 

meets airport noise goals

3 N+3 ~2016

3 ATMS Fuel efficiency, sonic 

boom Higher temperature capability, higher 

OPR for smaller, lighter, more efficient 

propulsion system

3 N+3 ~2022

4 Transonic thrust 

augmentation device

Airport noise, fuel 

efficiency, sonic 

boom

Smaller, lighter engine, potential airport 

noise benefit

2 N+3 ~2018

5 Advanced low NOx

combustor

Cruise emissions Meet cruise emissions 3 N+3 ~2020

6 CVC Fuel efficiency Fuel efficiency 2 N+3 ~2017

10.6 Meeting NASA N+3 Goals

Our integrated airframe and propulsion system, along with identified technologies, is projected to 

meet or exceed all N+3 targets. Results of the environmental and performance characteristics of our 

advanced vehicle concept with technology inputs are summarized in Table 17. 

TABLE 17.—LM’S PREFERRED CONCEPT WITH CRITICAL 

TECHNOLOGY MEETS OR SURPASSES ALL N+3 GOALS

NASA N+3 Efficient 

MultiMach Aircraft 

(Beyond 2030)

N+3 Goal Status

Environmental Goals

Sonic Boom 65 to 70 PLdB 

low boom flight

75 to 80 PLdB 

unrestricted flight

70 to 76 PLdB

KEY GOAL

Airport Noise 20 to 30 EPNdB

(cumulative below stage 3)

18.4

(32.2 jet only)

KEY GOAL

Cruise Emissions 

(g/kg fuel)

<5 EINOx

Plus particular and water vapor mitigation

5 EINOx

Performance Goals

Cruise Speed Mach 1.3 to 2.0 

low boom flight

Mach 1.3 to 2.0 

unrestricted

Mach 1.6

Range 4000 to 5500 nmi 4850 nmi

Payload 100 to 200 pax 100 pax

Fuel Efficiency 3.5 to 4.5

(pax-nmi/lb-fuel)

3.64

(pax-nmi/lb-fuel)

KEY GOAL
NASA/CR—2010-216796 105

Appendix A.—Acronym List

AFRL Air Force Research Laboratory

ANSI American National Standards Institute

ASA Absorption of Sound by the Atmosphere

ASDI Aircraft Situation Display to Industry

ASE Aero-Servo Elasticity

ASP Accelerating Swirler Passage

AST Advanced Supersonic Transport

ATC Air Traffic Control

ATM Air Traffic Management

ATMS Advanced Thermal Management System

ATS Air Transportation System

BFL Balanced Field Length

BIMS Biologically Inspired Morphing Structures

BPR Bypass Ratio

CAD Computer Aided Design

CAEP Committee on Aviation Environmental Protection

CDP Compressor Discharge Pressure

CFD Computational Fluid Dynamics

CFG Gross Thrust Coefficient

CG Center of Gravity

CMC Ceramic Matrix Composite

CPR Core Pressure Ratio

CVC Constant Volume Combustion

DBD Dielectric Barrier Discharge

DDT Deflagration-to-Detonation Transition

DFM Departure Flow Manager

DOC Direct Operating Costs

DOD Department of Defense

DRE Distributed Roughness Elements

EBC Environmental Barrier Coatings

EINOx Emission Index for NOx

EIS Entry Into Service

EPNL Effective Perceived Noise Level

ETMS Enhanced Traffic Management System

FAA Federal Aviation Administration

FAR Fuel/Air Ratio

FHV Fuel Heating Value

FOM Figure of Merit

FPR Fan Pressure Ratio

GE General Electric
NASA/CR—2010-216796 106

GRC Global Research Center

GTOW Gross Takeoff Weight

HPT High Pressure Turbine

HTEX Heat Exchanger

HX Heat Exchanger

IML Inner Mold Line

IR&D Internal Research & Development

IRP Intermediate Rated Power

ISA International Standard Atmosphere

KCAS Knots Calibrated Airspeed

KEAS Knots Equivalent Airspeed

LCTR Large Civil Tiltrotor

LDI Lean Direct Injection

LLI Late Lean Injection

LM Lockheed Martin

LMCO Lockheed Martin Corporation

LPT Low Pressure Turbine

LTO Landing and Takeoff

MDAO Multi-Disciplinary Analysis and Optimization

MFR Mass Flow Ratio

MFTF Mixed Flow Turbofan

MN Mach Number

MTOW Maximum Takeoff Weight

NAS National Air Transportation System

NLF Natural Laminar Flow

NOx Nitrogen Oxides

NPR Nozzle Pressure Ratio

NPSS Numerical Propulsion System Simulation

NRA NASA Research Announcement

OD Outer Diameter

OEI One Engine Inoperative

OEW Operational Empty Weight

OML Outer Mold Line

OPR Overall pressure ratio

OTA Open Travel Alliance

PAI Propulsion/Airframe Integration

PAX Passenger

PC Power Code

PDC Pulse Detonation Combustor

PDE Pulse Detonation Engine

PLdB Perceived Loudness Decibels
NASA/CR—2010-216796 107

PLR Programmed Lapse Rate

PR Pressure Ratio

P-RNAV Precision Route Navigation

QSST Quiet Supersonic Transport

RCD Rapid Conceptual Design

RH Relative Humidity

RNAV Route Navigation

RNP Required Navigation Performance

ROC Rate of Climb

RQL Rich Quench Lean

RVSM Reduce Vertical Separation Minimum

SERN Single Expansion Ramp Nozzle

SF Scale Factor

SFC Specific Fuel Consumption

SLS Sea Level Static

SOA State of the Art

SST Supersonic Transport

TAPS Twin Annular Premixed Swirler

TOC Top of Climb

TOFL Takeoff Field Length

TOGW Takeoff Gross Weight

TRL Technology Readiness Level

TSS Transportation and Security Solutions

UAS Unmanned Aircraft System

UPWT Unitary Plan Wind Tunnel

VCE Variable Cycle Engine

VLJ Very Light Jets

WBS Work Breakdown Structure

Appendix B.—Enabling Technologies

Table B.1 through Table B.4 show enabling technologies. 

TABLE B.1.—N+3 SUPERSONIC TECHNOLOGY LIST—BOLT-ON TECHNOLOGIES

Technology Source Category Description Benefits Drawbacks N+3 goals 

addressed

Readiness 

Engine: GE Global Research Center: Propulsion, Noise, Emissions, Fuel

Adaptive cycle engine GE Propulsion Advanced turbofan 

engine with adaptive 

cycle features

Fuel burn, noise Complexity Efficiency, 

noise, 

emissions

TRL 3

Contrail elimination 

through microwaves

Literature Propulsion Emit microwaves into 

jet plume to heat water 

vapor and delay 

formation of ice crystals

Can reduce or 

eliminate contrails

Weight, 0.1% 

engine power 

(claimed)

Emissions Researched by 

Cranfield 

University, Rolls 

Royce filing patent 

for technology

Sounds 

25,000 

engine 

very ef

formati

High aspect ratio nozzle N+3 kickoff 

meeting 

brainstorm

Propulsion Increase frequency of jet 

noise to levels where it 

would be better 

attenuated in the 

atmosphere

Reduces noise Adds weight and 

volume for nozzle 

duct, thrust 

penalties

Noise High-TRL, can be 

done

Not the 

reducin

AR 10 

benefit 

Embedded engine 

accessories

LM 

configuration 

brainstorming 

session

Propulsion Engine accessories 

better packaged in 

engine architecture

Reduces nacelle size 

by eliminating 

accessory box

Efficiency, 

boom

Engine architecture 

design

May be 

the dist

Fluid shield GE Noise Asymmetric outer layer 

in jet plume, use lower 

surface bypass air to 

shield noise from core

Reduces noise 

without much penalty

Weight and 

efficiency penalties 

of ducting the flow

Noise Research phase, 

might have been lab 

tested

Some b

Actuated core exhaust 

chevrons

Peter Coen's 

tech list 

(NASA 

internal, Ron 

Noebe)

Noise Chevrons are actuated 

using shape memory 

alloys, retracted after 

takeoff to eliminate 

thrust penalties

Reduce noise with no 

cruise efficiency 

penalty

Complexity, maybe 

weight

Efficiency, 

noise

Being researched 

with NASA NRAs

Small b

Scarfed nozzles Penn State 

(McLaughlin 

& Morris)

Noise Nozzles are "beveled" 

or "scarfed" on the 

bottom to direct noise 

upward

Reduces jet noise 

from sources close to 

the nozzle exit

Noise Lab tested Low co

Would 

profile 

are clo

Tunable resonators Penn State 

(McLaughlin 

& Morris)

Noise Precisely positioned 

arrays of resonators that 

generate an anti-phase

secondary sound field at 

the tip of the rotor-stator 

interaction in axial flow 

high pressure fans. 

Reduces fan noise by 

cancelling tones at 

the blade passage 

frequency and higher 

harmonics

Weight and 

complexity

Noise Lab tested Resona

concep

cancell

Doesn'

howeve

NASA/CR—2010-216796 109
TABLE B.1.—N+3 SUPERSONIC TECHNOLOGY LIST—BOLT-ON TECHNOLOGIES

Technology Source Category Description Benefits Drawbacks N+3 goals 

addressed

Readiness 

Internally mixed nozzles Penn State 

(McLaughlin 

& Morris)

Noise Efficient mixing of the 

fan flow with the core 

flow can produce a jet 

that will produce less 

noise without sacrificing 

thrust.

The estimated 

benefits are of the 

order of 2 dB should 

the conditions be 

ideal.

Weight from mixer 

shroud, slight 

performance penalty

Noise Has been 

demonstrated in the 

laboratory and in 

small scale engines 

in some engine 

conditions.

Non-round nozzles 

(rectangular or elliptical)

Penn State 

(McLaughlin 

& Morris)

Noise Nozzles are rectangular 

or elliptical shaped and 

are of moderate aspect 

ratio (<15)

Reduces noise in the 

major axis plane

Slightly increases 

noise in the minor 

axis plane, weight 

and volume issues, 

small performance 

penalties

Noise Lab tested, F117, 

other military 

airplanes

Not mu

high as

increas

penalti

depend

the hig

nozzle

Low-emissions 

alternative fuels

LM proposal 

brainstorming 

session

Fuel Alternative fuels that 

produce low emissions

Reduce emissions, 

possibly cost in the 

long run

Possibly increase 

cost since 

infrastructure might 

be more expensive, 

may hurt efficiency 

too

Efficiency 

(maybe), 

emissions

Depends on fuel, 

many research 

efforts, we would 

have to pick a 

specific one 

(algae?)

Likely 

based f

and NA

Active combustion 

control

Peter Coen's 

tech list (Penn 

State 

(Santavicca))

Emissions Active Combustion 

Control for Supersonic 

Low Emission 

Combustion

Lowers emissions Complexity, 

possibly 

performance 

penalties

Emissions Being researched 

with NASA NRAs

Configuration

High notch ratio planform 

with flow control

N+3 proposal Aero Wing with high notch 

ratio (high TE sweep)

Low induced and 

wing wave drag

Heavy, low speed 

higher vortex drag 

and pitch-up w/o 

attached flow 

control

Efficiency Idea. Would need to 

develop and test 

specific design

Depend

Synthetic cockpit vision LM proposal 

brainstorming 

session

Avionics Use TV camera system 

to give pilot better 

visibility where nose is 

in the way

Eliminates need for 

folding nose or any 

other design 

compromises 

necessary for the 

pilot to see the 

runway

None Efficiency Needs development 

but research 

activities are 

ongoing

Synthe

long tim

visibili

will be 

superso

Swing canard LM 

configuration 

brainstorming 

session

Configuration Deploy canard only at 

low speeds to control 

trim and stability

Lower drag, better 

boom

Weight and 

complexity of 

retraction 

mechanism

Efficiency, 

boom

Would have to be 

designed in

Not sur

have to 

outwei

Might t

a very 

Wing tip tanks LM 

configuration 

brainstorming 

session

Configuration Strategically shaped 

wing tip tanks

Structural stability, 

span load alleviation, 

area ruling

Weight Efficiency, 

boom

TRL-9, airplanes 

have employed 

them before

NASA/CR—2010-216796 110
TABLE B.1.—N+3 SUPERSONIC TECHNOLOGY LIST—BOLT-ON TECHNOLOGIES

Technology Source Category Description Benefits Drawbacks N+3 goals 

addressed

Readiness 

Configuration

Configuration shaping LM proposal 

brainstorming 

session

Configuration Shape configuration to 

close fore and aft shock 

to acceptable levels 

using advanced MDO 

techniques

Constrain sonic boom 

while maintaining 

efficiency and trim

Efficiency is 

compromised if it 

has to be balanced 

against boom 

requirements

Boom, 

efficiency

Has been done for 

QSST, more 

research to be done

Best so

boom. 

plasma 

but the 

availab

Variable sweep outboard 

wing panels

N+3 kickoff 

meeting 

brainstorm

Configuration Vary sweep of wing 

using pivots or other 

mechanical system

Easier to balance low 

speed takeoff 

requirements with 

high speed drag and 

boom goals, less 

weight than full 

variable sweep wings 

since it's just 

outboard panels

Large weight 

penalties in pivots 

and mechanical 

systems

Efficiency, 

noise, boom

TRL-9, done before 

on F14, F111, and 

B1

Swing 

many y

new an

concep

technol

conside

panels.

Quiet spike N+3 kickoff 

meeting 

brainstorm

Configuration Extending and retracting 

spike, extends for cruise 

to increase effective 

length of airplane, 

retracts for better 

landing

Reduces boom Weight and 

complexity of 

retraction 

mechanism

Boom, 

efficiency

Tested on F-15 for 

mechanical 

reliability

Questio

anythin

Retract

much i

Knight's lance nose N+3 proposal Configuration Similar to quiet spike 

but would be nonretractable

Reduces boom Weight, landing 

slap-down 

considerations

Boom, 

efficiency

idea An extr

role to 

signatu

be as lo

is no te

discrete 

mitigat

Windowless cabin LM proposal 

brainstorming 

session

Passenger 

cabin

Eliminate windows and 

replace with TV flatscreens showing the 

outside

Huge weight savings 

from eliminating 

holes in structure, 

manufacturability 

benefits, cabin noise 

benefits

Passenger 

acceptance

Efficiency Could easily be 

implemented, 

passenger 

acceptance needs to 

be thoroughly 

assessed

A signi

would 

no win

suffice 

thing", 

claustro

be don

(howev

serious

Deployable prop LM proposal 

brainstorming 

session

Propulsion Prop deploys out of 

turbofan engine to use 

for takeoff, allows less 

compromise in engine 

cycle

Increase engine 

cruise efficiency and 

decrease takeoff 

noise

Increased weight 

and complexity

Efficiency, 

noise

Low TRL, just an 

idea but not too farfetched

Hard to 

outwei

determ

mechan

work

Thrust vectoring LM proposal 

brainstorming 

session

Propulsion Vector engine exhaust 

slightly downward 

during cruise to improve 

boom

Reduce aft shock 

without having to 

heavily constrain 

configuration

Thrust (only get 

cosine), trim, 

weight if thrust 

vector is variable

Boom, 

efficiency

Could easily be 

done

Slight b

issue

NASA/CR—2010-216796 111
TABLE B.1.—N+3 SUPERSONIC TECHNOLOGY LIST—BOLT-ON TECHNOLOGIES

Technology Source Category Description Benefits Drawbacks N+3 goals 

addressed

Readiness 

Configuration

Acoustic cancelling of 

engine noise

LM proposal 

brainstorming 

session

Noise Speakers around engine 

that broadcast out-ofphase tones to cancel 

engine noise sources

Decrease fan tones Weight, complexity, 

power, little overall 

noise benefit since 

jet dominates

Noise Systems in various 

stages of research

Would 

sources 

engine

broadb

since in

Faired landing gear N+3 proposal noise Fairing over landing 

gear

Reduces gear noise Weight penalties, 

possible integration 

issues

Noise Faired gear 

concepts in research 

and testing

Main w

noise

Adaptive geometry

Inlet choking N+3 proposal Noise Vary inlet geometry so 

you get choked flow in 

the inlet at low speeds

Reduces inlet fan 

noise since noise 

can't propagate 

upstream through the 

choked region

Weight and 

complexity of 

variable geometry 

mechanism, could 

possibly use 

advanced morphing 

materials

Noise Researched before 

but never 

demonstrated

Read a

engine

necessa

comple

Potenti

morphi

mainta

Independent throat and 

exit area nozzle control

Penn State 

(McLaughlin 

& Morris)

Propulsion, 

noise

Can independently 

control throat and exit 

areas

Reduces broadband 

shock noise in the jet 

plume and optimizes 

performance for a 

wide variety of 

operating conditions

Weight penalties for 

mechanical systems 

actuators, may 

investigate 

morphing materials

Noise, 

efficiency

Linked control 

deployed on 

military engines

Sounds 

comple

Investi

should 

more d

Morphable control 

surfaces (seamless gaps)

N+3 kickoff 

meeting 

brainstorm

Controls Use a technology like 

shape-memory alloys or 

piezoelectric actuators 

and flexible material to 

bend surfaces without 

gaps

Lowers drag slightly, 

potential lower 

weight depending on 

system used, reduces 

airframe noise

Would make it more 

difficult to generate 

drag at low speeds 

on approach

Efficiency, 

noise

Low to mid-TRL, 

some shape memory 

alloy and 

piezoelectrics 

employed for other 

technologies like 

helicopter blades

Noise b

for sub

by elim

Inflatable wing elements LM 

configuration 

brainstorming 

session

Configuration Could inflate outboard 

wing sections for 

takeoff and stow them 

for cruise, could also 

have inflatable leading 

edges to control 

separation at high alpha

Lower drag, better 

low speed control

Weight and 

complexity of 

retraction 

mechanism

Efficiency Would have to be 

designed in

Some p

its way 

radius 

Plasma technology

DBD plasma flow control LM RTP Aero Use DBD plasma 

actuators to clean up 

separation regions and 

reduce drag, maybe also 

achieve spanwise 

laminar flow

Drag reduction Weight, complexity, 

power 

Efficiency DBD actuators are 

mid-TRL, how to 

deploy them needs 

to be studied

Potenti

separat

lamina

NASA/CR—2010-216796 112
TABLE B.1.—N+3 SUPERSONIC TECHNOLOGY LIST—BOLT-ON TECHNOLOGIES

Technology Source Category Description Benefits Drawbacks N+3 goals 

addressed

Readiness 

Plasma technology

Plasma boom 

optimization

LM proposal 

brainstorming 

session

Systems Use plasma generators 

to change airflow at key 

locations on aircraft and 

dissipate shocks.

Reduce boom 

without having to 

heavily constrain the 

configuration

Power requirements Boom, 

efficiency

DBD actuators are 

mid-TRL, how to 

deploy them and 

whether or not they 

would be effective 

for boom needs to 

be studied

May w

require

Plasma virtual length LM proposal 

brainstorming 

session

Configuration Use plasma generators 

to manipulate the 

airflow in front of and 

behind the vehicle, 

boom is reduced since 

vehicle's apparent length 

changes

Reduce sonic boom Power consumption 

to generate plasma

Boom Low TRL, some 

analytical studies 

done

Previou

power 

increas

but not 

Jet exhaust manipulation 

with plasma

Tech Fair -

Ohio State 

(Samimy)

Noise Use plasma actuators on 

nozzle lip to play with 

jet characteristics to 

reduce noise

Jet noise reduced 

about 2-3 dB

Power required for 

plasma

Noise Research phase Could p

chevro

noise r

Piezoelectric synthetic jet 

actuators for flow control

N+3 proposal Aero Control separation in 

critical regions, 

especially low speed 

high lift

Enables higher alphas 

with less pitch-up 

concerns, smaller 

wing

Power, weight, 

complexity

Efficiency Synthetic jets are 

well studied and 

have been deployed, 

how to deploy them 

in these applications 

needs study

Similar 

Laminar flow

Distributed roughness Saric and 

Reed

Aero Distributed bumps on 

aircraft surfaces. Creates 

acoustic disturbances 

out of phase with crossflow disturbances and 

maintains laminar flow 

on swept wings

Lower drag at 

supersonic speeds

Manufacturing 

complexity

Efficiency Wind tunnel tested Great p

flow. M

wings.

Fuselage riblets N+3 kickoff 

meeting 

brainstorm

Aero Use micro-riblets in skin 

to achieve drag 

reduction

Lower drag Manufacturing 

complexity, 

possibly extra 

weight

Efficiency Some research done 

at lab-scale

Looks 

Active LE suction for 

laminar flow

N+3 kickoff 

meeting 

brainstorm

Aero Use engine bleed to 

achieve laminar flow by 

suction

Lower drag at 

supersonic speeds, 

active system 

effective

Requires significant 

engine bleed, hurts 

efficiency, trade 

needs to be done if 

drag benefits

outweigh propulsion 

efficiency penalties

Efficiency Easy to employ, 

some development 

needed to get it 

right

Cost/be

roughn

NASA/CR—2010-216796 113
TABLE B.1.—N+3 SUPERSONIC TECHNOLOGY LIST—BOLT-ON TECHNOLOGIES

Technology Source Category Description Benefits Drawbacks N+3 goals 

addressed

Readiness 

Laminar flow

Supersonic micro-array 

flow control

Peter Coen's 

tech list 

(NASA 

internal)

Propulsion Improves boundary 

layer health 

downstream of 

oblique shock

Complexity, engine 

bleed

Efficiency Being researched 

with NASA NRAs

Not sur

compre

probab

Materials

Supersonic turbine blades 

made with SiC fibers in 

efficient weaving or 

braiding process

Peter Coen's 

tech list 

(Teledyne)

Propulsion SiC turbine blades Lighter, possibly 

cheaper to make than 

Nickel alloy, better 

temperature 

threshold, more creep

resistant. 

Maybe more 

expensive to 

manufacture

Efficiency Being researched 

with NASA NRAs

Better 

weight 

turbine 

needs t

Aluminum-Lithium N+3 proposal Materials Efficiency

GLass-REinforced 

aluminum (GLARE)

N+3 proposal Materials Efficiency

Titanium-5553 alloy N+3 proposal Materials Efficiency

Fiber-reinforced polymer 

composites (FRPC)

N+3 proposal Materials Efficiency

Aircraft operations

Takeoff overspeed LM Ops Take advantage of lowspeed aerodynamics of 

supersonic 

configurations. Set 

takeoff speed higher 

than required to lower 

climb CL, improve 

climb L/D, increase 

climb rate, and be able 

to cut back deeper over 

the cutback certification 

point

Reduces community 

noise on takeoff by 

providing more 

flexibility on the 

climb flight path.

Safety Noise Can be done Likely 

Lower cruise altitude LM proposal 

brainstorming 

session

Ops Cruise at lower altitude 

(<30 kft) to avoid 

contrails and lower NOx 

emissions

Lower emissions, no 

radiation concerns

Less propulsion 

efficiency 

(supersonic 

transports generally 

like to cruise at 

50 kft and above), 

weather, ride quality

Emissions Just design 

configuration to 

cruise below 30 kft

Helps w

also av

cruisin

engine 

be subs

effort t

be com

above t

better a

Helical takeoff profile N+3 kickoff 

meeting 

brainstorm

Ops Use corkscrew takeoff 

profile 

Reduces community 

noise by containing 

footprint in the 

airport boundary

Possible passenger 

discomfort, ATC 

issues, not a 

certifiable noise 

profile

Noise Can be done Would 

certific

issues. 

infrastr

accomm

NASA/CR—2010-216796 114
TABLE B.1.—N+3 SUPERSONIC TECHNOLOGY LIST—BOLT-ON TECHNOLOGIES

Technology Source Category Description Benefits Drawbacks N+3 goals 

addressed

Readiness 

Aircraft operations

Structural health 

monitoring

LM Systems Sensor devices that 

monitor integrity of key 

structural members

Allows structures to 

be designed with less 

margin and for 

maintenance checks 

to be reduced

Complicates design Efficiency Research and lab 

stages

This is 

eventua

aircraft

benefit 

their st

Automatic low-speed 

pitch-up control

LM 

configuration 

brainstorming 

session

Systems Use automatic control 

system to mitigate pitchup tendencies, this 

enables greater trailing 

edge sweep which is 

more efficient

Greater trailing edge 

sweep since margin 

doesn't have to be 

designed in to 

account for pilot 

response time, lower 

drag

Redundancy, weight 

of system

Efficiency, 

boom

Low-TRL Good i

designe

respon

TABLE B.2.—N+3 SUPERSONIC ALTERNATIVE CONFIGURATION OPTIONS

Configuration Source Category Description Benefits—baseline Drawbacks—baseline N+3 goals 

addressed

Joined and 

braced 

structure

LM Airframe Joined structure (wing-tail), blended structure and 

bracing structure to reduce weight, alleviate dynamic 

structural criticality and/or enable thinner and higher 

fineness designs (reducing wave drag).

Reduce weight, reduce 

flutter, reduce wave drag

Efficiency, 

boom

Mother/

Daughter 

configuration

LM Airframe Subsonic “mother” configuration attached to 

supersonic “daughter” configuration. Daughter 

configuration would be optimized for supersonic 

cruise and would incur none of the airframe and 

propulsion penalties of low speed. Mother 

configuration would use its engines to takeoff and 

carry Daughter configuration through takeoff to 

higher Q climb, releasing for transonic climb to 

supersonic cruise. Mother may or may not be required 

for landing. Mother configuration could also be more 

of a low sweep/high BPR wing pack that could attach 

and detach and fly to and from airports independently.

There could be a system of system advantage as one 

Mother could potentially serve several Daughters.

They would be piloted when together but Mother 

might return autonomously.

Easier to balance noise 

with efficiency and boom 

by not having to 

compromise these design 

considerations in one 

design. Benefit would 

come from a fleet 

depending on how many 

daughters could be served 

by one mother

Safety issues, extra pilots 

adding to cost (unless 

mother was autonomous 

or flown remotely by 

daughter pilots when 

joined), total system 

heavier than single 

airplane

Efficiency, 

boom, 

emissions, 

noise

Aggre

in a fl

multip

stack 

single 

thoug

be me

Low-speed 

autonomous 

pack

LM Airframe Similar to Mother/Daughter, only subsonic airplane is 

smaller and autonomous. Supersonic A/C would be 

designed for cruise and subsonic A/C would be a 

minimalist configuration with unswept wings, highbypass turbofans, landing gear, and minimal structure.

It would takeoff and land mated to the supersonic 

passenger A/C.

Same considerations as 

Mother/Daughter. 

Same considerations as 

Mother/Daughter. The 

subsonic "pack" is 

smaller though so less 

expensive.

Efficiency, 

boom, 

emissions, 

noise

Same 

NASA/CR—2010-216796 115
TABLE B.2.—N+3 SUPERSONIC ALTERNATIVE CONFIGURATION OPTIONS

Configuration Source Category Description Benefits—baseline Drawbacks—baseline N+3 goals 

addressed

Takeoff 

assistor

LM Airframe Similar to subsonic "pack" only now it is an 

autonomous takeoff "assistor" which would be a 

minimalist configuration with quiet turbofans which 

would tow the supersonic passenger aircraft on a 

cable through takeoff and to launch altitude. The 

supersonic A/C could have its engines at partial 

throttle during takeoff to stay within noise limits and 

still be controllable. Supersonic A/C would land on its 

own

Same considerations as 

Mother/Daughter only 

now supersonic A/C has 

to land on its own.

Substantially less 

drawbacks than 

Mother/Daughter. Safety 

issue is much less since 

no in-air mating is 

required. The subsonic or 

"extra" configuration is 

also as small as possible 

to minimize the extra 

expense.

Efficiency, 

boom, 

emissions, 

noise

Proba

soluti

is rem

A/C h

noise 

with t

this w

the je

solved 

efficie

could 

and th

the ov

Oblique Flying 

Wing

LM Airframe Similar to the elliptical design produced by NASA 

Ames. Takes off horizontally, and then turns to an 

oblique angle for transonic and supersonic speeds.

Better adapts to low 

speed and high speed 

optimum sweep. 

Packaging, windows, 

weight, complexity from 

rotation mechanisms for 

engines, stability and 

balance

Efficiency, 

boom

Benef

super

requir

highe

Oblique wing 

with fuselage

LM Airframe Similar to the AD-1. Oblique angle of wing could be 

adjusted in flight about a single pivot.

Better adapts to low 

speed and high speed 

optimum sweep. Less of a 

weight penalty than a 

standard swing wing

Weight from pivot, 

stability and balance, 

mechanism complexity, 

more difficult to design 

for low-boom since lift 

starts early and has to be 

symmetric on both sides 

of the wing

Efficiency, 

boom

Woul

penal

by va

Swing wing LM Airframe Similar to F-14, B-1, F-111, and Boeing SST.

Potentially newer materials and technologies could 

enable the benefits of a swing wing for less weight 

penalty. Swing could just be employed on the outer 

portions of the wing as well or be employed on the 

canard.

Better adapts to low 

speed and high speed 

optimum sweep. Boom 

shaping benefits

Weight from pivots, 

mechanism complexity

Efficiency, 

boom

It's di

its wa

mech

would 

weigh

Twin fuselage LM Airframe Two fuselages sharing a common wing. Has some 

drag benefits from favorable interference effects 

between the fuselages.

Busemann biplane effects 

between the fuselages, 

shocks from adjoining 

fuselages also can create 

high pressure air on aft 

facing surfaces fineness 

ratio of fuselages can be 

decreased by halving 

diameter for the same 

length, engines can be 

mounted in center wing 

eliminating need for 

inverted v tail

Length is good for boom 

and this configuration 

compromises that, 

aeroelastic effects from 

torqueing of the 

fuselages, doesn't work 

well for low passenger 

volumes

Efficiency Boom 

to be 

if it's 

aero a

given 

config

be red

seatin

space 

famili

NASA/CR—2010-216796 116
TABLE B.2.—N+3 SUPERSONIC ALTERNATIVE CONFIGURATION OPTIONS

Configuration Source Category Description Benefits—baseline Drawbacks—baseline N+3 goals 

addressed

Aft fuselage 

mounted 

engines

GE Propulsion GE engine architecture for variable cycle Allows efficient variable 

bypass in one engine

Complexity, possibly 

weight

Efficiency, 

noise, boom

Proba

archit

noise 

Distributed 

engines

LM Propulsion Thrust is provided by many small engines distributed 

about the airframe

Potential benefits include, 

greater overall reliability, 

lower cost, aero benefits, 

less integration issues, 

lower noise

Potential penalties 

include the same things

Efficiency, 

noise, 

emissions

Looke

is com

Many 

depen

Embedded 

engines

LM, GE Propulsion Engines are embedded in the airframe instead of 

hanging by pylons

Drag benefits, boom 

benefits

Not a lot of volume to 

store engines, weight 

increases and prop losses 

from ducting, possible 

cabin noise penalties and 

rotor burst considerations, 

harder to integrate new 

engines

Efficiency Podde

to be 

boom 

Embe

easier 

Auxiliary 

turbofans

LM, GE Propulsion Turbofans deployed for takeoff and stowed for cruise. Use turbofans for low 

speeds to meet noise 

constraints, allows cruise 

engines to be optimized 

for cruise and not have to 

deal with noise issues.

Also eliminates the need 

for aux inlets

Weight and complexity of 

retraction, turbofans dead 

weight during cruise 

unless they are utilized in 

an efficient way (i.e.,

embedded with a duct)

Efficiency, 

noise

Turbo

fusela

receiv

Stowi

diffic

thrust

benef

engin

more 

Over-wing 

mounted 

engines

LM, GE Propulsion Engines mounted over the wing instead of under Shield inlet radiated fan 

noise on approach and 

eliminate nacelle shock 

contribution to boom 

signature

Aero penalty, wing would 

have to be shaped or 

some sort of flow control 

device employed to give 

the engine good airflow, 

possible propulsion 

penalties from expanded 

air coming from wing 

upper surface

Noise, boom Trade 

config

boom 

moun

design

reflex

Cross-shafted 

auxiliary fans

NASA Propulsion Bleedless mixed compression inlet Zero bleed inlet with 

supersonic mixed 

compression

Improves propulsion 

efficiency, zero-bleed 

allows for near-perfect 

performance

Mutu

comp

shoul

archit

conce

techn

NASA/CR—2010-216796 117
TABLE B.3.—N+3 SUPERSONIC TOOLS AND METHODS LIST

Tool and 

method

Source Category Description Benefits N+3 goals 

addressed

Readiness 

CFD mesh adaption for 

mid-field pressure signature

NASA Aero, boom Uses mesh adaption schemes to 

resolve the grid around shocks and 

provide an accurate mid-field 

pressure signal for boom propagation

Enables coupling of CFD nearfield analysis and far-field 

boom propagation

Efficiency, 

boom

In development, 

NASA internal, NRA 

funded

CFD mesh adaption and 

error quantification

NASA Aero, boom Uses adjoint methods to quantify 

error in CFD computations and 

adjust the mesh accordingly

Enables user to specify a 

spatial discretization error 

tolerance. Useful for resolving 

mid-field pressure signals for 

boom propagation and aero 

analysis

Efficiency, 

boom

In development, 

NASA internal, NRA 

funded

CFD-based adjoint shape 

optimization

LM Airframe 

shaping

Use adjoint methods with CFDbased aero and boom analysis to 

optimize OML for low-boom and 

efficient cruise

Balances aero and boom 

demands

Efficiency, 

boom

Employed on QSST 

for aero optimization.

Boom implementation 

still needs work.

Somewhat still in the 

research phases

System of System Fleet 

Allocation Tools

Purdue Airspace, 

ops

Performs SoS optimization of an 

aircraft fleet operating over a route 

structure

Assesses impacts on cost, 

environment, and mobility of 

injecting a supersonic aircraft 

into a standard fleet mix

Research and 

application

Integrated engine, exhaust 

plume, airframe prediction 

capability

NASA Boom In development, 

NASA internal, NRA 

funded

Lift and Nozzle Change 

Effects on Tail Shocks 

(LaNCETS)

NASA Boom In development, 

NASA internal, NRA 

funded, flight test 

validation

Low-Boom Tools for 

plume/aft shaping

Wyle (Plotkin) Boom Develop interactive design tool 

based on a Generalization of the 

George-Seebass theory. New 

extended theory design tool will be 

coupled to PCBoom. Develops an 

engine exhaust plume model for 

inclusion in the generalized GeorgeSeebass method

In development, Wyle 

(Plotkin), NRA funded

Atmospheric Sonic Boom 

Prediction Model for 

PCBOOM

Wyle (Page) Boom Boom In development, Wyle, 

NRA funded

Prediction of boom 

signatures around buildings

Penn State 

(Sparrow)

Boom Tool that predicts interaction of 

boom signatures and building 

structures

Enables refined outdoor boom 

predictions in urban 

environments, helps with 

acceptability studies and policy

Boom In development, Penn 

State, NRA funded

NASA/CR—2010-216796 118
TABLE B.3.—N+3 SUPERSONIC TOOLS AND METHODS LIST

Tool and 

method

Source Category Description Benefits N+3 goals 

addressed

Readiness 

Modeling of boom in 

buildings

VPI 

(Burdisso), 

Wyle (Sizov)

Boom Models that predict sonic boom 

transmission through building 

structures

Enables predictions of boom 

effects on people inside 

buildings, helps with 

acceptability studies and policy

Boom In development, VPI 

and Wyle, NRA 

funded

Modeling of rattle from 

sonic booms

Wyle (Sizov) Boom Models induced noise (rattle) from 

sonic booms on objects

Enables predictions of 

secondary noise effects from 

sonic booms, helps with 

acceptability studies and policy

Boom In development, Wyle, 

NRA funded

Improved Nozzle Force 

Measurements

NASA Boom 

testing

The one-component force balance in 

the Nozzle Acoustic Test Rig 

(NATR) was refurbished and 

upgraded to improve force 

measurement accuracy and 

repeatability

In development, 

NASA internal, NRA 

funded

Multifidelity design 

methods for supersonic 

aircraft

Stanford 

(Kroo)

Design In development, 

Stanford (Kroo), NRA 

funded

Detection and Control of 

Instabilities

Georgia Tech 

(Seitzman)

Emissions Detection and control of instabilities 

and blowoff for low emissions

Emissions In development, 

Georgia Tech, NRA 

funded

Emissions by Modeling 

Combustion

Stanford 

(Pitsch)

Emissions Emissions prediction and modeling 

of supersonic vehicle combustion 

systems

Emissions In development, 

Stanford (Pitsch), 

NRA funded

Compu. Tools for 

Superheated Fuels

UTRC, UMass Emissions Validated computational tools for 

low emissions injector design using 

superheated/supercritical fuels

Emissions In development, 

UTRC and UMass, 

NRA funded

Control of Boundary 

Representation Topology in 

MDAO

MIT (Haimes) Geometry Develop geometry management tools 

and techniques to address the 

remaining obstacles to hands-off 

operation of interdisciplinary 

coupling in MDAO with CAPRI

Hi-Fi MDO: Software 

Infrastructure

M4 

Engineering 

(Baker)

MDAO Develop physics-based 

multidisciplinary analysis and 

optimization capabilities in the form 

of reusable modules in an objectoriented environment

In development, 

Stanford (Kroo), NRA 

funded

Comprehensive Model for 

the Prediction of Supersonic 

Jet Noise

Penn State 

(Morris), 

Boeing

Noise Enables better understanding of 

supersonic jet noise so that 

designs can be accurately 

evaluated

Noise In development, Penn 

State (Morris) and

Boeing, NRA funded

NASA/CR—2010-216796 119
TABLE B.3.—N+3 SUPERSONIC TOOLS AND METHODS LIST

Tool and 

method

Source Category Description Benefits N+3 goals 

addressed

Readiness 

Prediction and modeling of 

supersonic jet noise using 

large-eddy simulation

Stanford, 

Illinois (Lele)

Noise Enables better understanding of 

supersonic jet noise so that 

designs can be accurately 

evaluated

Noise In development, 

Stanford and Illinois, 

NRA funded

Exhaust plume reshaping 

techniques

UC Irvine 

(Papamoschou)

Noise Reduces supersonic jet noise by 

using external and internal vanes to 

redirect and concentrate cold fan 

flow beneath the hot core flow

Reduces jet noise Noise In development, UC 

Irvine (Papamoschou), 

NRA funded

High Speed Temperature 

Analysis Software and 

Support for TemporallyResolved Laser Raman 

Temperature Measurements 

in a Supersonic Jet

Aerodyne 

Research 

(Anan)

Noise Noise In development, 

Aerodyne, NRA 

funded

Computational bleed model 

for supersonic inlets

Boeing (Mace) Propulsion Computes bleed losses for 

supersonic inlets

In development, 

Boeing (Mace), NRA 

funded

High Fidelity Supersonic 

Exhaust System Design and 

Validation 

Rolls Royce 

Liberty Works 

(Jimerson)

Propulsion, 

noise

Efficiency, 

noise

In development, 

RRLW, NRA funded

Aero-Propulso-ServoElastic designLM, NASA Structures, 

flutter 

control

Design methods and approaches that 

use flexibility of the vehicle for 

control

Lighter structure, simplified 

controls 

Efficiency 

Aeroelastic tailoring Structures, 

flutter 

control

Tailoring composite skin design to 

provide directional stiffness for 

particular loads

Lighter structure, more 

unconventional wing designs

Efficiency Used on the X-29 and 

other programs

Reduced-Order Modeling 

for APSE design

LM, NASA Structures, 

flutter 

control

Mathematical techniques for efficient 

structural dynamics calculations

Enables efficient design and 

analysis of vehicle structural 

dynamics with sufficient 

fidelity

Efficiency Research stage 

TABLE B.4.—N+3 SUPERSONIC LIST OF MATERIALS AND GENERIC TECHNOLOGIES

Technology Source Description Potential applications Areas to 

research

Readiness 

Ceramic matrix 

composite nozzles

N+3 proposal Nozzles are made of CMC 

materials rather than metallic 

ones and can stand the high 

temperatures

Hot nozzle materials Manufacturing techniques, 

cost reduction

GE working on it.

Proposal claims that LM 

"demonstrated" them

Great id

what rea

any draw

Aluminum-Lithium N+3 proposal

GLass-REinforced 

aluminum (GLARE)

N+3 proposal

Titanium-5553 alloy N+3 proposal

NASA/CR—2010-216796 120
TABLE B.4.—N+3 SUPERSONIC LIST OF MATERIALS AND GENERIC TECHNOLOGIES

Technology Source Description Potential applications Areas to 

research

Readiness 

Fiber-reinforced polymer 

composites (FRPC)

N+3 proposal

Integrated structure and 

subsystems

LM proposal 

brainstorming 

session

Subsystems made from structural 

components and "manufactured/

integrated" in rather than "bolted 

on"

A number of subsystems 

could be manufactured this 

way, applications will 

emerge in detail design 

and subsystem studies

Applications, manufacturing 

processes, cost savings over 

conventional subsystems 

integration

Initial idea Integral 

growing 

of "integ

generall

airplane 

when it 

very det

Carbon nanotube 

composites

LM proposal 

brainstorming 

session

Composites constructed primarily 

from carbon nanotube fibers

Primary and secondary 

structure is possible. A 

number of detail design 

applications would 

probably come first

Research investment in this 

area is considerable, 

mechanical, electrical, and 

thermal properties.

Manufacturing, especially on 

a large scale, is a big area of 

research.

Largely in research stage 

(TRL <2), could be ready 

by 2030

A lot of 

theory b

done on 

These m

when st

they cou

liberties 

length, e

Highly Tailored 

Composite Structures 

(hits)

N+3 proposal Use directional composite ply 

layups to tailor structures, 

particularly for stiffness and 

buckling

A number of structural 

members where buckling 

is a concern, specific areas 

of the wing skins

Layup procedures, 

manufacturing techniques, 

cost reduction

Plasma boom 

optimization

LM proposal 

brainstorming 

session

Use plasma generators to change 

airflow at key locations on 

aircraft and dissipate shocks.

Reduce boom without 

having to heavily constrain 

the configuration

Configuration dependent.

How could plasma help 

dissipate shocks? What 

power requirements are 

needed? How would the 

system work?

DBD actuators are midTRL, how to deploy them

and whether or not they 

would be effective for 

boom needs to be studied

Structural health 

monitoring

LM Sensor devices that monitor 

integrity of key structural 

members, Allows structures to be 

designed with less margin and for 

maintenance checks to be 

reduced

Main application would be 

key members of primary 

and secondary structure 

that are subject to frequent 

maintenance checks.

Effectiveness of monitoring, 

application techniques

Research and lab stages This is a 

eventua

Superso

from it m

structur

NASA/CR—2010-216796 121
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NASA/CR—2010-216796 123
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13. SUPPLEMENTARY NOTES

14. ABSTRACT

The N+3 Final Report documents the work and progress made by Lockheed Martin Aeronautics in response to the NASA sponsored 

program "N+3 NRA Advanced Concept Studies for Supersonic Commercial Transports Entering Service in the 2030 to 2035 Period." The 

key technical objective of this effort was to generate promising supersonic concepts for the 2030 to 2035 timeframe and to develop plans for 

maturing the technologies required to make those concepts a reality. The N+3 program is aligned with NASA's Supersonic Project and is 

focused on providing alternative system-level solutions capable of overcoming the efficiency, environmental, and performance barriers to 

practical supersonic flight. 

15. SUBJECT TERMS

Aircraft configurations; Aircraft descriptions; Aircraft design; Aircraft development; Aircraft evaluation; Aircraft flight tests; Aircraft 

performance; Aircraft simulation; Aircraft structures; Aircraft testing 

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