March 2009 

NASA/TM-2009-215699 

Trade Study of Multiple Thruster Options for 

the Mars Airplane Concept 

Christopher A. Kuhl, Steven W. Gayle, Craig A. Hunter, Patrick S. Kenney, Salvatore Scola, 

David A. Paddock, Henry S. Wright, and Joseph F. Gasbarre 

Langley Research Center, Hampton, Virginia 
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March 2009 

NASA/TM-2009-215699 

Trade Study of Multiple Thruster Options for 

the Mars Airplane Concept 

Christopher A. Kuhl, Steven W. Gayle, Craig A. Hunter, Patrick S. Kenney, Salvatore Scola, 

David A. Paddock, Henry S. Wright, and Joseph F. Gasbarre 

Langley Research Center, Hampton, Virginia 
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 

Trade names and trademarks are used in this report for identification only. Their usage does not 

constitute an official endorsement, either expressed or implied, by the National Aeronautics and Space 

Administration. 
Trade Study of Multiple Thruster Optionsfor the Mars Airplane Concept 

NASA/TM-2009-215699

i

Table of Contents 

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

2.0 Review of ARES Baseline Propulsion System .......................................................................................... 2

2.1 Single 62N Thruster System Overview......................................................................................... 2

2.2 Single 62N Propulsion System Schematic.................................................................................... 3

2.3 Single 62N Baseline Thruster - Aerojet 14lbf (62N) AJ10-220 ................................................... 4

2.4 Airframe Thruster Mounting Concepts......................................................................................... 5

2.5 Single 62N Thruster Mounting and Support System .................................................................... 7

2.6 Single 62N Thruster Thermal Analysis....................................................................................... 10

2.7 Single 62N Thruster Plume Analysis.......................................................................................... 11

2.8 Single 62N Thruster Mission Performance................................................................................. 12

2.9 Single 62N Thruster System Summary....................................................................................... 13

3.0 Triple 22N Thruster Propulsion System Options..................................................................................... 14

3.1 Bi-Propellant Thruster Options................................................................................................... 14

3.2 Triple 22N Propulsion System Schematic .................................................................................. 15

3.3 Triple 22N Thruster Selection..................................................................................................... 16

3.4 Triple 22N Thruster Mounting and Support System................................................................... 17

3.4.1 Triple 22N In-Line Thruster Mounting Configuration .................................................. 18

3.4.2 Triple 22N Triangular Thruster Mounting Configuration.............................................. 22

3.5 Thermal Analysis of Triple 22N Thruster System...................................................................... 26

3.5.1 22N Thruster Thermal Characteristics........................................................................... 26

3.5.2 Modeling and Assumptions ........................................................................................... 28

3.5.3 Thermal Loads and Boundary Conditions ..................................................................... 29

3.5.4 Thermal Analysis Results .............................................................................................. 30

3.5.4.1 In-Line Thruster Configuration Results.......................................................... 30

3.5.4.2 Triangular Thruster Configuration Results..................................................... 34

3.5.4.3 Triangular Thruster Configuration Results with Estimated Plume................. 38

3.5.5 Thermal Analysis Conclusion........................................................................................ 45

3.6 Triple Thruster Plume Analysis and Characteristics................................................................... 46

3.6.1 Triple 22N In-line Thruster Plume Simulation Results ................................................. 47

3.6.2 Triple 22N Triangular Thruster Plume Simulation Results ........................................... 50

3.6.3 Triple 22N Thruster Plume Analysis Conclusions......................................................... 53

3.7 Triple Thruster Flight Simulation and Performance ................................................................... 54

3.8 Triple 22N Thruster System Summary ....................................................................................... 55

4.0 Performance Comparison of the Triple Thruster and Single Thruster Systems....................................... 57

4.1 Performance Assumptions .......................................................................................................... 57

4.2 Performance Summary................................................................................................................ 57

5.0 Decision Criteria for Thruster System Down-Select................................................................................ 59

5.1 Decision Criteria ......................................................................................................................... 59

5.2 Propulsion System Selection....................................................................................................... 61

6.0 Appendix A – Other Thrusters Considered.............................................................................................. 62

6.1.1 Thruster Candidate #2 - Aerojet 5-lbf Thruster ............................................................. 62

6.1.2 Thruster Candidate #3 - Astrium (DASA) 22N Thruster............................................... 62

7.0 Appendix B – Structural Analysis Assumptions...................................................................................... 63

7.1 Structure Load Cases .................................................................................................................. 63

7.2 Load Case Definitions................................................................................................................. 64

7.3 Mass Acceleration Curve ............................................................................................................ 65

8.0 Appendix C - Thermal Analysis Assumptions and Load Cases .............................................................. 66

9.0 References................................................................................................................................................ 68
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List of Figures 

Figure 1. Single 62N Thruster Configuration. ....................................................................................................... 2

Figure 2. Single 62N Thruster Propulsion System Schematic. .............................................................................. 3

Figure 3. Aerojet AJ10-220 62N Thruster. ............................................................................................................ 4

Figure 4. Airframe Support Structure for Thruster Mounting................................................................................ 5

Figure 5. Airplane Support Composite Structure................................................................................................... 6

Figure 6. Single 62N Thruster Configuration on Airplane .................................................................................... 7

Figure 7. Aerojet AJ10-220 62N Thruster Orientation .......................................................................................... 7

Figure 8. Aerojet 62N Thruster Mounting Concept............................................................................................... 8

Figure 9. Time Varying Temperature Data for AJ10-220 (Ir-Re) Thruster3........................................................ 10

Figure 10. Single 62N Thruster Rocket Plume CFD Analysis4 ........................................................................... 11

Figure 11. Single Thruster Duty Cycles - No Turbulence ................................................................................... 12

Figure 12. Single Thruster Duty Cycles - Severe Turbulence ............................................................................. 12

Figure 13. Triple 22N Propulsion System Schematic .......................................................................................... 15

Figure 14. AMPAC-Isp 22N Thruster Photos...................................................................................................... 16

Figure 15. AMPAC-Isp 22N Thruster Details ..................................................................................................... 17

Figure 16. Triple 22 N Thruster In-Line Configuration on Airplane................................................................... 18

Figure 17. Triple 22N In-Line Thruster Orientation ............................................................................................ 18

Figure 18. Triple 22N In-Line Thruster Mounting and Assembly....................................................................... 19

Figure 19. Triple 22N In-Line Thruster Assembly and Airframe Mounting Interface ........................................ 20

Figure 20. Triple 22N Thruster Triangular Configuration on Airplane ............................................................... 22

Figure 21. Triple 22N Triangular Thruster Orientation ....................................................................................... 22

Figure 22. Triple 22N Triangular Thruster Mounting and Assembly.................................................................. 23

Figure 23. Triple 22N Triangular Thruster Assembly and Airframe Mounting Interface ................................... 24

Figure 24. Thermocouple locations during test firing, and corresponding data10. ............................................... 27

Figure 25. Airframe Structural Characteristics .................................................................................................... 28

Figure 26. Heat Shield Options for Airplane Thermal Protection. ...................................................................... 29

Figure 27. In-line thruster configuration temperature profile .............................................................................. 30

Figure 28. Temperature profile, lower fuselage, in-line configuration baseline study. ....................................... 31

Figure 29. Temperature profile, support structure, in-line configuration baseline study..................................... 31

Figure 30. Temperature profile, lower fuselage, in-line configuration, low ε paint. ........................................... 32

Figure 31. Temperature profile, lower fuselage, in-line configuration, low ε paint. ........................................... 32

Figure 32. Temperature profile, lower fuselage, in-line configuration with heat shield...................................... 33

Figure 33. Temperature profile, lower fuselage, in-line configuration with heat shield..................................... 33

Figure 34. Triangular thruster configuration temperature profile ........................................................................ 34

Figure 35. Temperature profile, lower fuselage, triangular configuration baseline study. .................................. 35

Figure 36. Temperature profile, support structure, triangular configuration baseline study................................ 35

Figure 37. Temperature profile, lower fuselage, triangular configuration and low ε paint.................................. 36

Figure 38. Temperature profile, support structure, triangular configuration and low ε paint. ............................. 36
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Figure 39. Temperature profile, lower fuselage, triangular configuration with heat shield................................. 37

Figure 40. Temperature profile, lower fuselage, triangular configuration with heat shield................................. 37

Figure 41. Triangular thruster plume configuration............................................................................................. 38

Figure 42. Temperature profile, lower fuselage, triangular configuration, 700°C plume.................................... 39

Figure 43. Temperature profile, support structure, triangular configuration, 700°C plume. ............................... 39

Figure 44. Temperature profile, lower fuselage, triangular configuration, 400°C plume.................................... 40

Figure 45. Temperature profile, support structure, triangular configuration, 400°C plume. ............................... 40

Figure 46. Temperature profile, lower fuselage, triangular config., 700°C plume, low ε paint. ......................... 41

Figure 47. Temperature profile, support structure, triangular config., 700°C plume, low ε paint....................... 41

Figure 48. Temperature profile, lower fuselage, triangular config., 400°C plume, low ε paint. ......................... 42

Figure 49. Temperature profile, support structure, triangular config., 400°C plume, low ε paint....................... 42

Figure 50. Temperature profile, lower fuselage, triangular config., 700°C plume, heat shield........................... 43

Figure 51. Temperature profile, support structure, triangular config., 700°C plume, heat shield. ...................... 43

Figure 52. Temperature profile, lower fuselage, triangular config., 400°C plume, heat shield........................... 44

Figure 53. Temperature profile, support structure, triangular config., 400°C plume, heat shield. ...................... 44

Figure 54. Triple 22N In-line thruster semi-span CFD model. ............................................................................ 46

Figure 55. Triple 22N Triangular thruster semi-span CFD model....................................................................... 46

Figure 56. ARES OML superimposed over centerline slice through CFD simulation, in-line case.................... 47

Figure 57. ARES OML superimposed over slice through CFD simulation, in-line case..................................... 48

Figure 58. ARES OML superimposed over vertical slice through CFD simulation, in-line case........................ 48

Figure 59. ARES OML superimposed over vertical slices through CFD simulation. ......................................... 49

Figure 60. ARES OML superimposed over vertical slices through CFD in-line plume temperature profile. ..... 49

Figure 61. ARES OML superimposed over centerline slice through CFD simulation, triangular case............... 50

Figure 62. ARES OML superimposed over 5.88° slice through CFD simulation, triangular case...................... 51

Figure 63. ARES OML superimposed over 9.38° slice through CFD simulation, triangular case...................... 51

Figure 64. ARES OML superimposed over vertical slice through CFD simulation, triangular case. ................. 52

Figure 65. ARES OML superimposed over vertical slices through CFD simulation. ......................................... 52

Figure 66. ARES OML superimposed over vertical slices through CFD triangular plume temperature profile. 53

Figure 67. Duty Cycles of Triple Thruster System - No Turbulence Conditions ................................................ 54

Figure 66. Duty Cycles of Triple Thruster System - Severe Turbulence Conditions .......................................... 55

Figure 67. System Mass Breakdown Comparison ............................................................................................... 57

Figure 68. Comparison of Flight Range for Fixed Fuel Load.............................................................................. 58

Figure 69. Comparison of Flight Range for Fixed Wet Mass .............................................................................. 58

Figure 70. Decision Criteria Scores ..................................................................................................................... 60

Figure 71. Recap of Trade Study Thruster Options ............................................................................................. 61

Figure 72. Photos of Alternate Thruster - Aerojet 5-lbf Thruster ........................................................................ 62

Figure 73. Model of Alternate Thruster - Astrium 22N....................................................................................... 62

Figure 74. Flight vehicle folded in aeroshell........................................................................................................ 64
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List of Tables 

Table 1. Aerojet AJ10-220 62N Thruster Parameters2........................................................................................... 4

Table 2. Airplane Thruster Support Structure Mass Breakdown ........................................................................... 6

Table 3. Mass Breakdown of Mounting Concept for AJ10-220 62N Thruster...................................................... 9

Table 4. Hot Fire Testing Results for AJ10-220 62N Thruster3........................................................................... 10

Table 5. Single Thruster System Performance Summary .................................................................................... 13

Table 6. Single 62N Thruster System Mass Breakdown...................................................................................... 13

Table 7. Mars Airplane Commercial Thruster Options6 ...................................................................................... 14

Table 8. ARC 22N Thruster Characteristics7 ....................................................................................................... 16

Table 9. Triple 22N In-Line Mounting Component Mass Breakdown................................................................ 21

Table 10. Triple 22N Triangular Mounting Component Mass Breakdown ......................................................... 25

Table 11. Triple Thruster Propulsion System Performance Summary – 40 kg Fuel Load .................................. 56

Table 12. Triple 22N Thruster Option - Propulsion System Mass Breakdown ................................................... 56

Table 13. System Performance Comparison Summary........................................................................................ 58

Table 14. Decision Criteria and Weighting Factors............................................................................................. 59

Table 15. Supporting Data for Decision Criteria Scores...................................................................................... 59

Table 16. Decision Criteria Scores....................................................................................................................... 60

Table 17. Mission Structural Load Cases............................................................................................................. 63

Table 18. Source and value for all surface thicknesses used in the model........................................................... 66

Table 19. Sources for material properties used in this study................................................................................ 66

Table 20. Boundary conditions used in thermal analysis..................................................................................... 66

Table 21. Loads used in thermal analysis ............................................................................................................ 67
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1

1.0 Introduction 

The ARES baseline airplane, proposed in 2003 to the Office of Space Science for the Mars Scout 

Program1

, required a thruster that could produce roughly 40 N of thrust and have a 50% margin to 

accommodate climbs and accelerations. The Aerojet 14-lbf (62N) AJ10-220 thruster was selected for the 

propulsion system, which is the only thruster available in this range of thrust. With concerns about the 

availability of tooling for manufacturing this thruster, as well as the non-recurring costs that would be 

imposed on the program, other options needed to be considered. 

A trade study was performed at NASA Langley Research Center under the Planetary Airplane 

Risk Reduction (PARR) project (2004-2005) to examine the option of using multiple, smaller thrusters in 

place of a single large thruster on the Mars airplane concept with the goal to reduce overall cost, schedule, 

and technical risk. 

The 5-lbf (22N) thruster is a common reaction control thruster on many satellites. Thousands of 

these types of thrusters have been built and flown on numerous programs, including MILSTAR and 

Intelsat VI. This study has examined the use of three 22N thrusters for the Mars airplane propulsion 

system and compared the results to those of the baseline single thruster system. 

This report covers: 

 Overview of the single thruster and multiple thruster propulsion system options. 

 Thruster to airframe mounting concepts. 

 Thermal analysis of the thrusters and heat soak-back into the airframe. 

 Plume calculations for the various thruster arrangements. 

 System mass estimates and overall performance of the thruster options. 

 Direct comparison of the 62N thruster option and the multiple 22N thruster options. 

 Thruster system down-select methodology and results. 
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2.0 Review of ARES Baseline Propulsion System 

2.1 Single 62N Thruster System Overview 

 The ARES baseline propulsion system was designed to use a bi-propellant, pulsed control with 

MMH fuel and MON-3 oxidizer. Propellant is stored in two identical titanium tanks each with a screenand-channel propellant management device (PMD) for fuel and oxidizer extraction. The fuel flow is 

controlled through a regulated helium blow-down system where helium is stored at high pressure in a 

composite over-wrapped pressure vessel. Propulsion system components integrated into the airframe are 

shown in Figure 1. 

Figure 1. Single 62N Thruster Configuration. 
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2.2 Single 62N Propulsion System Schematic 

 The ARES airplane baseline propulsion system is a regulated bi-propellant system with pulsed 

thruster control. Major components of this system consist of fuel and oxidizer tanks, helium pressure 

tank, and a thruster. Fuel and oxidizer tanks are isolated before flight using normally-closed pyrotechnic 

valves and isolated during operation using redundant parallel check valves. The schematic layout of the 

ARES propulsion subsystem is shown in Figure 2. 

Figure 2. Single 62N Thruster Propulsion System Schematic. 
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4

2.3 Single 62N Baseline Thruster - Aerojet 14lbf (62N) AJ10-220 

 The 14lbf Aerojet AJ10-220 bi-propellant 

thruster, shown in Figure 3, is derived from 

Lockheed's “Bus-1”, a previously classified US 

Air Force spacecraft. "Bus-1" is a modular bus 

developed by Lockheed that is qualified for STS 

and ELV's. The bus provides propulsion, GN&C, 

Communications, & Data Management. The Bus1 unit consisted of 6 reaction control thruster 

modules (RCTM) each of which contained two 14 

lbf thrusters, making a total of 12 thrusters per 

unit. These thrusters were qualified over a 10-22 

lbf (44.5 – 97.9 N) thrust range and have a 

demonstrated chamber life of greater than 40 

hours of operation. At the 14 lbf level, the 

specific impulse is approximately 285s. The 

thruster incorporates fuel film cooling and series 

redundant valves to provide a redundant leak 

protection. A total of 54 production RCTMs were 

delivered for the “Bus-1” program from 1987 to 

1992. Table 1 lists the performance parameters of 

the AJ10-220 thruster. 

A follow-on thruster design was conducted by Aerojet to improve the life of the thruster by using 

an Iridium/Rhenium chamber as opposed to the production Columbium chamber. The Ir/Re 14 lbf engine 

has demonstrated through hot-fire tests to deliver a specific impulse of 305 s and a predicted life over 40 

hours. The Ir/Re 14 lbf is similar to the production model with the exception of an Ir/Re chamber, a 

heater circuit, and the thrust plate and frame are modified to accommodate different spacecraft interfaces. 

Some, but not all, of the tooling exists in storage at Aerojet for the Columbium production 

engine. The current thruster design uses shunts and a frame to direct heat away from the valve during 

hot-fire, and valve overheating will occur unless the chamber heat is shunted to some other structure. 

Current thruster design uses a frame with vibration isolators to manage launch vibration loads and any 

change in thruster mounting will require a new dynamic analysis. Depending on the magnitude of 

changes from the qualified thruster, non-recurring costs start at approximately $1M for a Mars Airplane 

program. 

Figure 3. Aerojet AJ10-220 62N Thruster.2

Table 1. Aerojet AJ10-220 62N Thruster Parameters2

Fuel Monomethyl Hydrazine (MMH)

Oxidizer Nitrogen Tetroxide (MON3) 

Inlet Pressure 140 – 375 psia 

Propellant Inlet Temp 40°F – 110°F 

Thrust @ 220 psia 14 ± 0.7 lbf (62 ± 3 N) 

Thrust @ 140 – 375 psia 9.5 – 22.0 lbf (42 – 98N) 

SS Isp @ 220 psi 285 nominal, 265 min 

MR @ 220 psia 1.65 ± 0.05 

Expansion Ratio 75:1 

Weight 4.3 lb (1950 grams) 

Length 7.5 inches (19 cm) 

Diameter 2.3 inches (5.8 cm) 

Valve Voltage 28 ± 4 VDC 
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2.4 Airframe Thruster Mounting Concepts 

 To mount the thrusters to the airplane, additional support structure was designed for this study 

that was not part of the original ARES Step-2 baseline design. The airframe structure concept for 

mounting of thruster assemblies into existing Mars Airplane airframe design is shown in Figure 4. The 

thruster ribs and bulkheads are composite sandwich structure similar to existing Mars Airplane airframe 

design. Figure 5 shows a typical rib or bulkhead cross-section. The upper and lower flanges of the 

thruster ribs and bulkheads are bonded directly to the center-body skins. The aft thruster bulkhead is 

mechanically joined to the center-body main ribs. The thruster ribs are mechanically joined to the aft 

thruster bulkhead and the center-body aft spar. The thruster forward bulkhead is mechanically joined to 

the thruster ribs. Loads sustained by the thruster assembly are passed directly into the center-body 

airframe. 

Figure 4. Airframe Support Structure for Thruster Mounting 

(New structure shown in light-green) 
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 The composite structure was analyzed with Pro-Engineer Mechanica Finite Element Analysis 

(FEA). Preliminary lay-up schedules for laminates and honeycomb sizing have been determined to 

enable mass estimates for composite support structure. The analysis approach was to conservatively size 

the composite structure to mount either of the three thruster configurations considered in this study. 

Launch loads, cases 1a and 1b from Appendix B, were determined to be the critical load cases the 

composite structure would sustain, therefore they were applied to the FEA model. After multiple design 

iterations, a final preliminary design was developed based on the required design factor of safety of 2.0 

for composite materials, show in Appendix B. Preliminary mass estimates are shown in Table 2. 

Graphite/Epoxy Facesheet 

Honeycomb (1/4 in. thk.) 

Upper Flange 

Lower Flange 

Figure 5. Airplane Support Composite Structure 

Table 2. Airplane Thruster Support Structure Mass Breakdown 

Component Material 

Weight

(lbf) 

Mass 

(kg) 

Aft Thruster Bulkhead Composite Sandwich 0.289 0.131 

Forward Thruster Bulkhead Composite Sandwich 0.158 0.072 

Thruster Ribs Composite Sandwich 0.180 0.082 

Total 0.627 0.284 
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2.5 Single 62N Thruster Mounting and Support System 

The orientation of the 62N thruster in relation to the airplane is shown in Figure 6. The thruster 

tilts 5.88° from the horizontal reference so that the thrust vector is directed through the airplane center of 

gravity, shown in Figure 7. 

The AJ10-220 thruster mounting bracket is titanium alloy bolted directly to the thruster injector 

flange, shown in Figure 8. Mounting scheme is typical for this thruster. The thruster bracket is mounted to 

the airframe similar to the 22N thruster assemblies. The mounting bracket is bolted directly to composite 

bulkhead. A structural analysis was performed on the 62N thruster mounting bracket with Pro-Engineer 

Mechanica FEA. Load case 1a and 1b were applied to the FEA model. 

Figure 6. Single 62N Thruster Configuration on Airplane 

Side View 

5.880

W.L.0.00 

Horizontal Ref. Rear View

10.25” 

Figure 7. Aerojet AJ10-220 62N Thruster Orientation 

(Thruster is tilted 5.88° to direct thrust vector through the center of gravity) 
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Airframe Interface Mounting Bracket 

62N Thruster 

Injector Flange Airframe 

Mounting Plate 

Figure 8. Aerojet 62N Thruster Mounting Concept 
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 Similar to the composite structure, launch load cases (1a and 1b in Appendix B) were determined 

to be the critical load cases applied to the 62N thruster due to the large “g” loads seen during launch. The 

mass of a single 62N thruster is equal to 1.95 kg. Based on Figure 77, Appendix B, the MAC value for 

the thruster is 50. Therefore the “g” force applied to the thruster for load case 1a in the axial direction 

would be 52.8 and for load case 1b would be 50 in any direction. Based on F = ma the force on each 

thruster for the axial direction of load case 1a would be 227 lbf and for load case 1b 215 lbf. The 

preliminary mounting bracket design was sized according to design factors of safety. Preliminary mass 

estimates for mounting hardware are shown in Table 3. 

Table 3. Mass Breakdown of Mounting Concept for AJ10-220 62N Thruster 

Component Material 

Density

(lbf/in3) 

Weight

(lbf) 

Mass 

(kg) 

 Thruster Bracket Titanium Alloy 0.16 0.159 0.072 

Mounting Hardware CRES 0.29 0.100 0.045 

Total 0.259 0.117 
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2.6 Single 62N Thruster Thermal Analysis 

 As part of the acceptance testing of the 14 lbf production engine, thrusters were hot-fired three 

times for 100 seconds. The time requirement for these thrusters was 214 seconds at 14 lbf thrust level. 

Table 4 shows results of hot fire testing at a duty cycle similar to that of the Step-2 ARES mission. 

Maximum temperature of the valve face never exceeds 110°F. Figure 9 shows the temperature data for a 

100 sec hot-fire steady-state test of the Ir/Re engine3

. The flange temperature (TF2) reaches a maximum 

temperature of 160°F at 95 seconds and climbs to 255°F after a few seconds of coast due to thermal 

soakback. 

0

200

400

600

800

1000

1200

1400

1600

0 20 40 60 80 100 120 140 160 180 200

Time [seconds]

Temperature [°F]

Nozzle Exit Temperature

Injector Flange Temperature

Figure 9. Time Varying Temperature Data for AJ10-220 (Ir-Re) Thruster3

Table 4. Hot Fire Testing Results for AJ10-220 62N Thruster3

Fuel Inlet 

Pressure 

[psia] 

Ox Inlet 

Pressure 

[psia] 

Burn 

Time 

[sec] Cycle:[sec] 

Coast 

[sec] 

Max Chamber 

Temp [°F] 

Max Flange 

Temp [°F] 

Max Valve 

Temp [°F] 

225 225 60 5 on /5 off 20 1300-1700 190-215 105 

145 170 60 5 on /5 off 20 1250-1650 140-190 100 

340 380 60 5 on /5 off 20 1450-1650 210-235 110 

375 375 60 5 on /5 off 20 1425-1625 200-230 110 
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2.7 Single 62N Thruster Plume Analysis 

 In order to evaluate the potential impact of the rocket plume on the aircraft’s tail structure, an 

isolated plume CFD simulation was conducted4

. Color contours of Mach number are shown in Figure 10, 

along with a line contour indicating the thermal edge of the plume at To=220K. The nozzle stagnation 

temperature is 3395K and the free-stream stagnation temperature is 199K. This analysis shows that there 

is sufficient clearance between the 220K edge of the plume and the tail surfaces and support booms to 

preclude any thermal impingement effects. In addition, the limited influence of the plume on the 

surrounding free-stream indicates that flow around the tail surfaces will not be changed dramatically by 

the presence of the plume. 

Figure 10. Single 62N Thruster Rocket Plume CFD Analysis4
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2.8 Single 62N Thruster Mission Performance 

 The flight simulation of the Mars airplane was evaluated using a 6 degree-of-freedom model from 

the Langley Standard Real-Time Simulation application framework5

. Control logic was incorporated to 

hold the aircraft’s speed while flying a mission profile. Since the airplane uses a pressure-blowdown 

propulsion system, the thruster is not throttled; instead duty cycle adjustments provide the needed thrust 

modulation. The thruster control law was also modified to accommodate the pulsing mode of the thruster. 

The law uses two key elements, a thrust period and a duty cycle. The two elements define how long the 

thruster is on (period * duty cycle) and off. Once a thrust period has ended the duty cycle is modified to 

reflect the error between the commanded Mach number and the average Mach number for the period. 

Two atmospheric conditions were simulated, one with no turbulence, and the other condition with severe 

atmospheric turbulence. Thruster duty cycles for the flight condition with no turbulence are shown in 

Figure 11. The histogram shows a predominant duty cycle between 45% and 60 % for the majority of the 

flight. During severe turbulence, the data shown in Figure 12 provide no predominant duty cycle which 

suggests that the flight controls are trying to keep up with the fast changing flight conditions. Duty Cycle (thrust on / thrust off)

0% 20% 40% 60% 80% 100%

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

0 500 1000 1500 2000 2500 3000 3500 4000

Flight Time [seconds] Histogram: Percent of Total Flight Time

Figure 11. Single Thruster Duty Cycles - No Turbulence 

Duty Cycle (thrust on / thrust off)

0% 20% 40% 60% 80% 100%

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

0 500 1000 1500 2000 2500 3000 3500 4000

Flight Time [seconds] Histogram: Percent of Total Flight Time

Figure 12. Single Thruster Duty Cycles - Severe Turbulence 
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2.9 Single 62N Thruster System Summary 

 The results of the simulation show that for the airplane with a single 62N thruster flying in noturbulence conditions, the flight range with 40 kg of propellant is about 527 km. In severe turbulence 

conditions for the same fuel load, the flight range is reduced to about 520 km (Table 5). The total mass 

of the single thruster propulsion system is broken down into components shown in Table 6. Based on 

commercially available components, estimates from vendors, and analysis, the current best estimate for 

the 62N thruster system is 13.18 kg. Incorporating the contingencies assigned to each component, the 

total mass increases to a growth mass of 15.95 kg. 

Table 5. Single Thruster System Performance Summary 

Parameter Value 

Total AFS Mass 145 kg 

Propulsion Mass 15.95 kg 

Propellant Load 40 kg 

Fuel 15 kg - MMH 

Oxidizer 25 kg - MON3 

Thrust 62 N 

Isp ~285 s 

No Turbulence

Flight Duration 65.75 min 

Range 527 km 

Avg Fuel Usage 13.2 km/kg 

Severe Turbulence

Flight Duration 64.47 min 

Range 520 km 

Avg Fuel Usage 13.0 km/kg 

Table 6. Single 62N Thruster System Mass Breakdown 

Common Components QTY 

CBE Each 

[kg] 

Total 

[kg] 

Cont 

[%] 

CBE + 

Cont [kg] Vendor Part Number 

Pressure Transducer 5 0.15 0.75 20% 0.90 GP50 7200 

Normally Closed Pyro Valve 5 0.15 0.75 15% 0.86 CONAX 1832-205 

GHe Filter 1 0.25 0.25 25% 0.31 Vacco F1D10636-01 

High Pressure Service Valve 11 0.113 1.243 10% 1.37 Vacco V1E10430-01 

Helium Regulator 1 0.363 0.363 10% 0.40 Vacco 66250 

Check Valve 4 0.02 0.08 25% 0.10 Vacco V1D10856-02 

Liquid Filter 2 0.18 0.36 25% 0.45 Vacco F1D10638-01 

GHe Tank 1 1.2 1.2 25% 1.50 Lincoln 220131-1 

Oxidizer Tank 1 2.87 2.87 25% 3.59 

Fuel Tank 1 2.87 2.87 25% 3.59 

Tank Support Brackets 4 0.1 0.4 25% 0.50 

Subtotal 10.74 13.07 

Single Thruster Option 

Thruster 1 1.09 1.09 10% 1.20 Aerojet AJ10-220 

Thruster Valves 1 0.86 0.86 10% 0.95 Moog 52-189B 

Flow Venturi 2 0.04 0.08 5% 0.08 

Misc Tubing, Brackets 1 0.25 0.25 25% 0.31 

PIA & PCA Plates 2 0.08 0.16 25% 0.20 

Engine Support Structure 1 0.117 0.117 25% 0.15 

Total Propulsion Mass 13.18 15.95 
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3.0 Triple 22N Thruster Propulsion System Options 

3.1 Bi-Propellant Thruster Options 

 An alternative to the single 14 lbf (62N) system is the concept of using three 22N bi-propellant 

thrusters. The options that exist are shown in Table 7. Based on the recommendation by Lockheed 

Martin following a contracted trade study6

, the AMPAC-In-Space Propulsion (ARC) 22N thruster was 

chosen as the baseline thruster for the three-thruster trade study. This thruster was qualified under the 

Intelsat VI satellite program and over 1000 units have flown to date. (It should be noted that the Aerojet 

R-6C is an almost identical engine with similar production numbers and similar heritage and can be 

considered as a backup for future development). 

Table 7. Mars Airplane Commercial Thruster Options6

Vendor Aerojet Aerojet Aerojet Aerojet 

AMPACISP (ARC) AMPACISP (ARC) Astrium 

- EADS 

Model AJ10-220 AJ10-220+ R-6C R-6D 5 lbf 5 lbf S22-02 

Chamber Material C103 Ir/Re C103 C103 C103 Pt/Rh Pt/Rh 

Status Existing 

Design 

Modified 

Design 

COTS 

Item 

Existing 

Design COTS Item Existing Design 

Existing 

Design 

Thrust 14 lbf 14 lbf 5 lbf 5 lbf 5 lbf 5 lbf 5 lbf 

Thrust 62 N 62 N 22 N 22 N 22 N 22 N 22 N 

Isp [sec] 282 282 293 294 291 300 285 

Area Ratio 75 75 150 150 150 150 150 

Mixture Ratio (nom) 1.65 1.65 1.65 1.65 1.65 1.6 1.65 

Feed Press [psia] 140 - 375 140-375 220 220 220 220 290 

Chamber Press [psia] 125 125 116 125 125 125 125 

Valve Model Mg 52-189 EKV ACS Mg 51-

330 

Mg 51-

136 Mg 51-178 Mg 51-178 Mg 51-178 

Thruster Mass [kg] 1.95 1.125 0.54 0.4 0.45 0.47 0.49 

Valve Mass [kg] 0.86 0.62 0.14 0.27 0.19 0.19 0.19 

Total Mass [kg] 1.09 1.063 0.68 0.67 0.64 0.66 0.68 

ROM Cost [M$] >$1M >$1M > $100k unkn <$100k <$200k unkn 

Qual Status Full In-Develop Full InDevelop Full In-Qual InDevelop

Number Flown >100 0 >600 0 >1000 0 0 

Heritage 

Lockheed 

BUS-1 

(DoD) 

Evolve from 

AJ10-220 

C103 

chamber 

MILSTAR 

GOES 

ISRO 

Insat 2B 

Evolve 

from R6C 

Thruster 

Telecom 

Intelsat VI 

Evolve 

from C103 

5lbf 

Thruster 

Astrium 

10N 

Thruster
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3.2 Triple 22N Propulsion System Schematic 

 The triple thruster propulsion system is identical to the ARES baseline 62N system shown in 

Figure 2 with the exception of the three thrusters, three sets of flow venturis, and a tubing manifold to 

route the fuel and oxidizer to the thruster valves of each thruster. A schematic of this system is shown in 

Figure 13. 

Figure 13. Triple 22N Propulsion System Schematic 
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3.3 Triple 22N Thruster Selection 

 AMPAC-In-Space Propulsion (formerly 

Bell Aerospace and Atlantic Research 

Corporation) developed a family of 5 lbf 

bipropellant engines using different valve 

configurations ranging from dual seat solenoid 

valves to single seat torque motor valves. These 

engines have demonstrated extremely long life 

of 100,000 seconds on multiple units with high 

vacuum specific impulse ranging from 286-

295s7

. These engines have demonstrated wide 

flexibility of mission duty cycle operation by 

firing in any duty cycle ranging from 1-100% 

and accumulating over 325,000 firing pulses on 

multiple units with a minimum of 400,000 

pulses on one unit. The engine uses a unique 

single element injector made of a coated 

columbium alloy orifice plate with titanium 

alloy feed tubes and support structure. The 

thrust chamber is made of columbium alloy and 

coated with R512E silicide coating for 

protection from the products of combustion internally and externally to allow for high emissivity for 

radiation cooling. The injector is electron beam welded to the thrust chamber to eliminate the potential of 

hot gas leakage. The valve is bolted to the injector using metal seals. The RCS thrusters were designed, 

built, and qualified for the Intelsat VI series spacecraft. A summary of performance characteristics of the 

22N thruster is shown in Table 8. 

 In 1996, AMPAC-ISP started an IR&D program to develop an improved 5 lbf (22N) bipropellant thruster for satellite station keeping. These thrusters have longer combustion chamber life by 

using a Pt/Rh alloy combustion chamber with high temperature capability and superior oxidation 

resistance instead of the Silicide-coated columbium thrust chambers that would not meet steady-state 

propellant throughput requirements. In 1997, AMPAC-ISP demonstrated the DST-4 Pt/Rh thruster and 

during the development tests a single Pt/Rh chamber accumulated 33 hrs of test time and over 900 kg of 

propellant throughput with no damage. The thruster has been tested with both torque motor and solenoid 

valves and has passed a qualification level vibration test with solenoid valves8

. 

Table 8. ARC 22N Thruster Characteristics7

Fuel MMH 

Oxidizer NTO 

Thrust 22N 

Specific Impulse 293 s 

Area Ratio 150:1 

Inlet Pressure Range 6.5 – 27.6 bar 

Feed Temperature 70°F Nominal 

105°F Max SS Operation 

160°F Pulsed Mode SS 

Chamber Press (nom) 125 psia 

Max Impulse 2,668,800 N-s 

Min Impulse Bit 0.045 N-s 

Mass 0.8 kg 

Engine Length ~270 mm 

Nozzle Exit Diameter 54 mm 

Total Thruster Mass 0.8 kg 

Figure 14. AMPAC-Isp 22N Thruster Photos 
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 A diagram of the 22N baseline thruster used in this study is shown in Figure 15. The total length 

of the thruster, including shown inlet tubing, is approximately 10.75 inches. The valve is shown with the 

Moog series redundant solenoid valve model Mg 51-178.

3.4 Triple 22N Thruster Mounting and Support System 

 Two options for mounting the three thrusters on the airplane body were considered. The first is 

an “in-line” configuration where the outer two thrusters would be the primary driver and the center 

thruster used in instances where acceleration and climbing are needed. The second configuration consists 

of a triangular arrangement where the top two thrusters are primary and the bottom thruster is secondary. 

 Each thruster configuration requires that the thrust vector for each individual thruster, regardless 

of arrangement, passes through the specified flight vehicle’s center of gravity (c.g.). The location of the 

vehicle’s c.g. is centered between the flight vehicle’s fuel tanks (Figure 1). The ARES center-body 

location for each thruster assembly was determined from the base-lined ARES airframe configuration1

. 

 Launch load cases 1a and 1b (Appendix B) were determined to be the critical load cases applied 

to the 22 N thruster due to the large “g” loads seen during launch. The mass of a single 22 N thruster is 

equal to 0.8 kg. Based on Figure 77, the MAC value for the thruster is 60. Therefore the “g” force 

applied to the thruster for load case 1a in the axial direction would be 62.8 and for load case 1b would be 

60 in any direction. Based on F = ma the force on each 22 N thruster for the axial direction of load case 

1a would be 111 lbf and for load case 1b 106 lbf. 

Valve 

Valve Retaining Groove 

Nozzle 

Valve Mounting Flange 

Injector Flange

Combustion Chamber 

Figure 15. AMPAC-Isp 22N Thruster Details 
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3.4.1 Triple 22N In-Line Thruster Mounting Configuration 

 The triple thruster in-line configuration is an arrangement of three 22 N thrusters in a single row 

shown in Figure 16. Right hand and left hand thrusters were rotated about c.g. location 4.5° to allow for 

adequate clearance. The thrusters are tilted 5.88° from the horizontal reference to direct thrust vector 

through the airplane c.g., shown in Figure 17. 

Figure 16. Triple 22 N Thruster In-Line Configuration on Airplane 

Top View 

B.L. 0.00

4.50

4.50

5.880

W.L.0.00 

Horizontal Ref. 

Side View 

Rear View 

Figure 17. Triple 22N In-Line Thruster Orientation 
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 The aft and forward brackets are titanium alloy. Titanium alloy was selected because of its good 

strength-to-weight ratio, low density, and low coefficient of thermal expansion, good corrosion resistance, 

good toughness, and its performance at extreme temperatures (-320°F to 750°F). The brackets are 

fabricated to place thrusters in correct orientation. The aft brackets are bolted to the valve flange using 

existing mounting holes (see Figure 18). The thruster valve flange was chosen for bracket interface, 

because it reaches a maximum temperature of only 200°F at steady-state operation. The thrusters are also 

supported at the forward bracket and captured by titanium retaining rings. Once the retaining rings are 

Figure 18. Triple 22N In-Line Thruster Mounting and Assembly 

Forward Bracket 

Aft Brackets 

Retaining Rings

AMPAC-ISP 22N Thruster 

Valve Flange 

Airframe Interface 
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placed into existing groove in valves, they are fastened to the forward bracket. The forward and aft 

brackets are bolted to the composite airframe structure shown in Figure 19. The design approach was to 

enable the ability to readily assemble and disassemble the thrusters from the airframe and allow for 

repeatable location of thrusters. 

Figure 19. Triple 22N In-Line Thruster Assembly and Airframe Mounting Interface 
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 The triple 22N thruster in-line configuration was analyzed with Pro-Engineer Mechanica Finite 

Element Analysis (FEA). Load case 1a and 1b from Appendix B, were determined to be the critical load 

cases the thruster assembly would sustain, therefore it was applied to FEA model. After multiple design 

iterations, a final preliminary design was developed based on required factors of safety of 1.5 for ultimate 

strength and 1.25 for yield strength, see Appendix B. Preliminary mass estimates are shown in Table 9. 

Table 9. Triple 22N In-Line Mounting Component Mass Breakdown 

Component Material 

Density

(lbf/in3

) 

Weight 

(lbf) 

Mass 

(kg) 

RH Center Thruster Aft Bracket Titanium Alloy 0.16 0.060 0.027 

LH Center Thruster Aft Bracket Titanium Alloy 0.16 0.060 0.027 

RH Outboard Thruster Aft Bracket Titanium Alloy 0.16 0.048 0.022 

LH Outboard Thruster Aft Bracket Titanium Alloy 0.16 0.048 0.022 

Valve Retaining Rings (6 Total) Titanium Alloy 0.16 0.046 0.021 

Valve Support Bracket Titanium Alloy 0.16 0.170 0.077 

Mounting Hardware CRES 0.29 0.150 0.068 

Total 0.582 0.264 
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3.4.2 Triple 22N Triangular Thruster Mounting Configuration 

 The second thruster configuration is an arrangement of three 22N thrusters in a triangular 

orientation (Figure 20). Right hand, left hand, and bottom thrusters were rotated about c.g. location until 

adequate clearance was achieved. Details of the rotation and orientation of the thrusters are shown in 

Figure 21. 

Figure 20. Triple 22N Thruster Triangular Configuration on Airplane 

Figure 21. Triple 22N Triangular Thruster Orientation 

W.L.0.00 

Horizontal Ref.

5.880 3.50

Side View

2.20

2.20

B.L. 0.00 

Top View 

Rear View
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 The aft and forward brackets are titanium alloy. The brackets are fabricated to place each thruster 

in correct orientation. The aft bracket is bolted to the valve flange using existing mounting holes (see 

Figure 22). The thruster valve flange was chosen for bracket interface, because it reaches a maximum 

temperature of only 160°F at steady-state operation (Figure 9). The thrusters are also supported at the 

forward bracket and captured by titanium retaining clamps. Once retaining clamps are placed into 

existing groove in valves, the retaining clamps are fastened to forward bracket. Note: the forward thruster 

bracket for the triangular configuration incorporates an alternative approach for mechanically fastening 

Figure 22. Triple 22N Triangular Thruster Mounting and Assembly 

Valve Flange 

Airframe Interface 

Forward Bracket 

Aft Bracket 

Retaining Clamps 

AMPAC-ISP 22N Thruster 
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the thruster valves, than the row configuration forward bracket. Either method could be developed for 

each thruster configuration. The forward and aft brackets are bolted to composite airframe structure, 

shown in Figure 23. 

Figure 23. Triple 22N Triangular Thruster Assembly and Airframe Mounting Interface 
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 A structural analysis was performed on the triple 22N triangular thruster assembly with ProEngineer Mechanica Finite Element Analysis (FEA). Load case 1a and 1b, Appendix B, were determined 

to be the critical load cases the thruster assembly would sustain, therefore it was applied to the FEA 

model. After multiple design iterations, a final preliminary design was developed based on required 

design factors of safety of 1.5 for ultimate strength and 1.25 for yield strength. Preliminary mass 

estimates are shown in Table 10. 

Table 10. Triple 22N Triangular Mounting Component Mass Breakdown 

Component Material 

Density

(lbf/in3) 

Weight

(lbf) 

Mass 

(kg) 

Aft Thruster Bracket Titanium Alloy 0.16 0.154 0.070 

Forward Support Bracket Titanium Alloy 0.16 0.178 0.081 

Valve Clamp (6 Total) Titanium Alloy 0.16 0.010 0.005 

Mounting Hardware CRES 0.29 0.130 0.059 

Total 0.472 0.214 
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3.5 Thermal Analysis of Triple 22N Thruster System 

3.5.1 22N Thruster Thermal Characteristics 

 During operation, the thrusters reach temperatures significantly greater than the Mars ambient 

temperature of approximately -60°C. For continuous firing at room temperature, the combustion 

chambers of the thrusters reach approximately 1000°C in 45 seconds. This creates a temperature 

difference between the thrusters and outer fuselage of almost 1100°C, which will cause a large radiative 

heat transfer between the two components. 

 The fuselage of the aircraft is a composite structure designed by Aurora Flight Sciences 

Corporation9

 consisting of an aluminum honeycomb core surrounded by face sheets of carbon fabric. The 

adhesive holding the carbon face sheets to the aluminum honeycomb will begin to break down at 

approximately 250°C. The goals of the thermal analysis of these thrusters in simulated operation are: 

 To determine if there is any thermal benefit of using one configuration over the other 

 To determine if the fuselage will remain under 250°C, and 

 Suggest ways to lower the fuselage temperature if it is found to reach temperatures greater than 

250°C. 

 

 
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 During development of this thruster, Aerojet performed a test firing of a single thruster in order to 

determine the temperatures at different locations on the thruster10. Seven thermocouples were mounted to 

the thruster as shown in Figure 24. A temperature vs. time profile for each thermocouple was provided 

for 120 seconds of continuous firing in a vacuum at room temperature. The thruster appeared to reach 

thermal equilibrium after approximately 45 seconds. 

Chamber Max - T1 [°F]

Throat - T0 [°F]

Nozzle - T6 [°F]

Injector Frontface - T2 [°F]

Chamber Inner Wall - T3 [°F]

Injector Backface - T4 [°F]

Valve Flange - T5 [°F]

0

100

200

300

400

500

600

700

800

900

1000

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

0 20 40 60 80 100 120

Time [sec]

Temperature [°F]

Figure 24. Thermocouple locations during test firing, and corresponding data10. 

0.0 mm

T5

Valve

Mounting

Flange

? mm

T4

Injector Back

Face

+X

PNT 1

PNT 43

X0

Z0

27.381 mm

T3 - Chamber Inner Wall

24.638 mm

T2

Injector

Front Face

47.381 mm

T1

Film Dryout

201.287 mm

T0

Throat

? mm

T6

Nozzle
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 During flight, two thrusters will be continuously firing. These two thruster will reach an 

equilibrium temperature profile similar to the one provided by the test firing, however it will be slightly 

cooler due to the much lower ambient temperature on Mars and forced convection from the thrusters to 

the atmosphere. The third thruster will be turned on and off as needed during flight. 

3.5.2 Modeling and Assumptions 

 Each of the two configurations described in the previous sections consist of three AMPAC-Isp 

22N thrusters mounted to the airframe through the additional support structure shown again in Figure 25. 

The thrusters are connected to each other and to the inner spar structure of the aircraft by two titanium 

mounting brackets. These brackets are bolted to two separate bulkheads inside the vehicle. There is a 

rectangular cut out in the lower fuselage that allows the thrusters to protrude out of the aircraft; this can 

partially be seen in the right side of Figure 25. The nozzles are directly under a small section of the lower 

fuselage, and the combustion chambers are partially under this section, and partially under the cut out. 

Consequently, this allows for radiation transfer to the inner components of the aircraft as well as the lower 

fuselage. The locations of internal electronics, instruments and fuel tanks have not yet been decided and 

are not yet included in the design model. 

 A PRO-Engineer CAD model containing the ARES fuselage, inner support structure, and each 

thruster configuration described in this report was used for the thermal analysis. Locations for the 

instruments and other internal components had not yet been defined. An MSC.Patran.Thermal model of 

the aircraft had been built for a previous ARES thermal study1

 for the Step-2 proposal activity and this 

model was used as a starting point for a new MSC.Patran.Thermal model for the thruster investigation. 

 The fuselage of the aircraft had not changed since the previous study; however a cutout needed to 

be placed in the lower fuselage to accommodate the thrusters. The cutout was positioned to match the 

location of the cutout in the CAD model. The rest of the components in the model had to be removed; the 

inner support structure had changed, and the locations of the instruments and other internal components 

need to be redefined. 

 The fuselage and old support structure had been modeled as three separate surface layers in order 

to simulate the full composite structure; an inner carbon face sheet, an aluminum honeycomb core, and an 

outer carbon face sheet. From the CAD geometry, a new MSC.Patran support structure was created in 

this same fashion, with one layer for each part of the composite. According to the Aurora assembly 

drawings9

, the outer and inner faces of the carbon sheets were connected through the caps of the cchannel. The c-channels were therefore not directly modeled, and a box was created for the carbon layers 

with a surface inside for the honeycomb. 

Figure 25. Airframe Structural Characteristics 
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 An MSC.Patran version of each thruster configuration was created next, and imported into the 

model. Components of the thrusters were created as solids where possible, and the rest were 

approximated with surfaces. All fuel/oxidizer pipe lines and other internal components were removed. 

The thrusters were placed in the model at approximately the same vertical distance from the bottom of the 

fuselage as in the CAD geometry. 

 Surfaces were created to model two 0.5 mm aluminum heat shields for three analysis cases. They 

are designed to cover the portion of the fuselage directly above the thrusters. A larger heat shield was 

required for analyses including the thruster exhaust plumes. Figure 26 shows both heat shield 

configurations. 

 Thicknesses of shell components in the model were taken directly from the CAD model. Those 

that could not be obtained from the model were taken from Aurora assembly schematics of the composite 

structure9

. Values from this source, such as carbon fabric thickness, are approximate; they were 

determined by multiplying the number of fabric sheets used in each face of the structure by 0.127 mm 

(thickness of each sheet). Table 18, Appendix C gives the thicknesses used in this investigation. Material 

properties for the carbon face sheets and honeycomb core were taken from the previous study. 

Component materials and material property sources are presented in Table 19, Appendix C. 

3.5.3 Thermal Loads and Boundary Conditions 

 All three thrusters are given the same steady state temperature profile derived from Figure 24. 

This is a conservative assumption since one thruster will be pulsing on and off which should reduce the 

maximum temperature that it reaches. Convection and radiation inside the aircraft was ignored for this 

study (except in the rear-most inner compartment, which was included in the radiation view factor 

analysis). Convection would help to reduce and spread out the temperature on the hotter components, 

therefore neglecting it is a conservative assumption. Convection on the thrusters is also ignored. A 

turbulent forced convection model would need to be derived in order to model it accurately. Convection 

would help to reduce the steady state temperature reached by the thrusters, so neglecting it is a 

conservative assumption. However, the thrusters may heat the air passing over them enough to cause a 

convection heat load into the lower fuselage. This has also been ignored, and may need to be considered 

in a future study. Radiation from the thruster exhaust plumes is ignored for the primary trials. An 

analysis with estimated plume properties is done for the triangular thruster arrangement. When the CFD 

plume analysis is completed, and the geometry and temperature profile has been defined, an updated 

thermal analysis should be done including this information. The exhaust plumes will be hot, and will 

Figure 26. Heat Shield Options for Airplane Thermal Protection. 

Small heat shield (upper left) and large heat shield (lower right).
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contribute to the radiation transfer to the underside of the fuselage, and also to the tail. Table 20 and 

Table 21, Appendix C show the loads and boundary conditions applied to the model. Most of these 

conditions are always used in the analysis; however some trials will not contain a low emissivity paint or 

heat shield. Which ones are turned off for various runs are described in the next section. 

3.5.4 Thermal Analysis Results 

 Three steady state analyses were done for each thruster configuration using the boundary 

conditions and loads given in Table 20 and Table 21, Appendix C respectively. The “View factor 

radiation calculations” in Table 21 are the only loads that are modified for each case, which are described 

below: 

• Baseline – no paints or heat shield on the lower fuselage, bare carbon face sheets 

• Low Emissivity Paint – A paint with an emissivity of 0.3 is applied to the fuselage directly 

above the thrusters and to the inner face sheet of the upper fuselage. A number of aluminum, and 

leafing aluminum paints can provide this emissivity, or lower. Chromeric Silver paint 586 can 

also provide this emissivity. 

• Heat Shield – A 0.5 mm thick polished aluminum sheet is placed on the lower fuselage. The 

inner face sheet of the upper fuselage is coated with low emissivity paint. 

 Results for these trials are documented in Sections 3.5.4.1 and 3.5.4.2, and do not include 

radiation from the exhaust plumes. Section 3.5.4.3 provides results for the triangular configuration with 

the estimated exhaust plume properties given in Table 21. Analyses were completed for each of the three 

cases described above, however the painted area and heat shield area needed to be enlarged to protect 

more of the fuselage surface. 

3.5.4.1 In-Line Thruster Configuration Results 

 Figure 27 shows the temperature profile across the thrusters. This profile is described in Table 

20, Appendix C, and is used for each in-line configuration case. 

Figure 27. In-line thruster configuration temperature profile 
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3.5.4.1.1 Baseline Study Results 

Figure 28. Temperature profile, lower fuselage, in-line configuration baseline study. 

(Max Temperature = 209.6°C). 

Figure 29. Temperature profile, support structure, in-line configuration baseline study. 

(Max Temperature = 73.8°C). 
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3.5.4.1.2 Low Emissivity Paint Study Results 

Figure 30. Temperature profile, lower fuselage, in-line configuration, low ε paint. 

(Max Temperature = 154.2°C). 

Figure 31. Temperature profile, lower fuselage, in-line configuration, low ε paint. 

(Max Temperature = 72.2°C). 
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3.5.4.1.3 Heat Shield Study Results 

Figure 32. Temperature profile, lower fuselage, in-line configuration with heat shield. 

(Max T = -6.17°C).

Figure 33. Temperature profile, lower fuselage, in-line configuration with heat shield. 

(Max T = 71.8°C). 
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34

3.5.4.2 Triangular Thruster Configuration Results 

 Figure 34 shows the temperature profile across the thrusters. This profile is described in Table 

20, Appendix C, and is used for each triangular configuration case. 

Figure 34. Triangular thruster configuration temperature profile 
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3.5.4.2.1 Baseline Study Results 

Figure 35. Temperature profile, lower fuselage, triangular configuration baseline study. 

(Max Temperature = 185°C). 

Figure 36. Temperature profile, support structure, triangular configuration baseline study. 

(Max Temperature = 41.4°C). 
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3.5.4.2.2 Low Emissivity Paint Study Results 

Figure 37. Temperature profile, lower fuselage, triangular configuration and low ε paint. 

(Max Temperature = 137.2°C). 

Figure 38. Temperature profile, support structure, triangular configuration and low ε paint. 

(Max Temperature = 37.02°C).
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3.5.4.2.3 Heat Shield Study Results 

Figure 39. Temperature profile, lower fuselage, triangular configuration with heat shield. 

(Max Temperature = -8.33°C).

Figure 40. Temperature profile, lower fuselage, triangular configuration with heat shield. 

(Max Temperature = 36.4°C).
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3.5.4.3 Triangular Thruster Configuration Results with Estimated Plume 

 The full plume structure used in the analysis was approximately one quarter of the full plume 

length. The reason for this was to save computational time running the view factor analysis. The plume 

closest to the fuselage structure will contribute the most heating, and so the closest quarter is sufficient to 

get a general idea of the steady state temperatures resultant on the aircraft. 

Figure 41. Triangular thruster plume configuration. 

(Plumes shown at 700 °C.) 
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3.5.4.3.1 Baseline Study Results, 700 °C and 400 °C 

Figure 42. Temperature profile, lower fuselage, triangular configuration, 700°C plume. 

(Max Temperature = 284°C). 

Figure 43. Temperature profile, support structure, triangular configuration, 700°C plume. 

(Max Temperature = 139°C). 
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Figure 44. Temperature profile, lower fuselage, triangular configuration, 400°C plume. 

(Max Temperature = 215°C). 

Figure 45. Temperature profile, support structure, triangular configuration, 400°C plume. 

(Max Temperature= 45°C). 
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3.5.4.3.2 Low Emissivity Paint Study Results, 700 °C and 400 °C 

Figure 46. Temperature profile, lower fuselage, triangular config., 700°C plume, low ε paint. 

( Max Temperature= 247°C). 

Figure 47. Temperature profile, support structure, triangular config., 700°C plume, low ε paint.

( Max Temperature= 66°C). 
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Figure 48. Temperature profile, lower fuselage, triangular config., 400°C plume, low ε paint. 

(Max Temperature= 167°C). 

Figure 49. Temperature profile, support structure, triangular config., 400°C plume, low ε paint. 

(Max Temperature= 41°C). 
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3.5.4.3.3 Heat Shield Study Results, 700 °C and 400 °C 

Figure 50. Temperature profile, lower fuselage, triangular config., 700°C plume, heat shield. 

(Max Temperature= -9.5°C). 

Figure 51. Temperature profile, support structure, triangular config., 700°C plume, heat shield. 

(Max Temperature= 41°C). 
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Figure 52. Temperature profile, lower fuselage, triangular config., 400°C plume, heat shield. 

(Max Temperature= -28°C). 

Figure 53. Temperature profile, support structure, triangular config., 400°C plume, heat shield. 

(Max Temperature= 39°C). 
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3.5.5 Thermal Analysis Conclusion 

 If the plume is ignored, it appears that the fuselage will not exceed 250°C. The highest 

temperature reached is 209°C for the in-line configuration baseline study. The triangular thruster 

configuration results in the smallest fuselage temperatures for all studies. This seems physically 

reasonable since all three thrusters in the linear configuration have direct views to the lower fuselage, 

while one of the triangular thrusters is partially blocked by the other two. This blocked thruster is also 

further away from the fuselage which also reduces radiation transfer. 

 Adding the plume to the analysis clearly shows a large increase in the steady state temperatures 

on the aircraft. Assuming 0.5 for plume emissivity, and a plume temperature is 700°C, a paint with an 

emissivity of 0.3 will not be enough to keep the temperature below critical. A heat shield, or paint with a 

lower emissivity should be used if this is the case (see Figure 43, Figure 48, Figure 50, Figure 52). Paint 

with an emissivity of 0.3 should be sufficient however if the plume temperature is only 400°C. This 

analysis should be considered preliminary, as values for the plume are estimates. 

 The heat shield used in the plume study needed to cover a larger area than the one used in the 

non-plume studies (Figure 26). Depending on the plume temperature, the smaller shield could be used in 

combination with low emissivity paint. This would keep the fuselage temperatures closer to the ambient 

temperature if desired. For a high plume temperature, the larger heat shield should be used. 

 The last iteration for this trade study resulted in the plume temperature profile to be significantly 

lower than what was estimated for the thermal analysis. The plume temperature does not exceed 330K 

(66°C), which compared to the assumed 400°C and 700°C is significantly less. This suggests that the 

results obtained in the analyses ignoring the thruster plumes will more closely represent the expected 

airplane skin temperatures. The cases reported in the previous sections including the higher temperature 

plumes can be taken as worst case scenarios. 

 The smaller heat shield weighs approximately 0.068 kg, and the larger shield weighs 

approximately 0.16 kg. Both of these masses would increase to include any bonding material and/or 

fasteners needed to attach the shield to the aircraft. If a paint or shield is to be used, it does not appear 

that there will be any thermal benefit of using one configuration over the other. The triangular thruster 

does keep the lower fuselage temperature cooler, but not by enough to disqualify the use of the in-line 

thruster configuration. 
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3.6 Triple Thruster Plume Analysis and Characteristics 

 Computational fluid dynamics (CFD) simulation was used to evaluate possible plume interactions 

in the triple-thruster systems. This work considers the 3D thruster systems in a representative freestream 

environment, but does not include the aircraft outer mold line (OML) or influence of aircraft external 

aerodynamic flows. 

 CFD grids were created using the Langley “AXB” software, custom FORTRAN 95 codes, and 

the commercial “GridGen” tool. Thruster arrangements were modeled assuming spanwise symmetry 

about the vehicle centerline symmetry plane (X-Z plane at Y=0). Figure 54 and Figure 55 show the grids 

for the in-line and triangular arrangements as modeled for CFD. 

Figure 54. Triple 22N In-line thruster semi-span CFD model. 

(Mesh shown at 1/2 resolution). 

Figure 55. Triple 22N Triangular thruster semi-span CFD model. 

(Mesh shown at 1/2 resolution). 
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 The in-line CFD grid consisted of 120 blocks and 13.7 million cells. The triangular grid 

consisted of 84 blocks and 16.5 million cells. In both cases, the thruster nozzles discharge into an 

ambient freestream region with a 65 nozzle exit diameter downstream run. Reynolds-Averaged NavierStokes (RANS) simulations were run with the Langley “PAB3D” flow solver, using two-equation 

temperature-corrected k-ε turbulence closure and a linear Reynolds stress model. Freestream flow was 

set to M=0.55 with an ambient temperature of 346R (representative of ARES cruise conditions on Mars). 

Nozzle inflow was specified with a stagnation temperature of 1450R and a nozzle pressure ratio (NPR) of 

2193 (corresponds to nozzle stagnation pressure of 125 psia and Mars ambient pressure 0.057 psia). A 

specific heat ratio of γ=1.23 (representative of thruster exhaust) was used everywhere. 

 Simulations were run in double precision on the 40-node ARES computer cluster, with 2.8GHz 

Intel Pentium4 CPUs and 1GB memory per node. Solutions were sequenced through coarse and medium 

mesh levels to evaluate grid dependence effects, and had run times of 350-1200 wall-clock hours per case. 

To accelerate convergence, simulations were initiated with local time-stepping, which drives towards a 

steady-state solution (steady RANS). The in-line case converged to a steady state solution after about 

79,000 iterations. The triangular case did not converge to a steady state with local time-stepping after 

72,000 iterations, and the simulation was switched to a constant-time-step “time-accurate” mode 

(unsteady RANS). After an additional 41,000 iterations (with 4 sub-iterations per time step) the 

triangular case converged to a stable steady-state solution. 

3.6.1 Triple 22N In-line Thruster Plume Simulation Results 

 Results from the in-line case are presented in this section. In the following figures, the ARES 

OML is superimposed over the plume CFD simulation, with color contours of Mach number shown. 

Figure 56 shows a slice through the centerline plane of spanwise symmetry, with the center thruster 

plume visible. Here, the plume is seen to slowly bend from its initial 5.88° incidence to align with the 

freestream flow. Figure 57 shows a slice through the CFD simulation in a plane 5.88° from the 

horizontal, aligned with the thruster incidence. The individual thruster plumes are well isolated and 

maintain angular separation, with no major interactions apparent. Figure 58 shows a vertical slice 

through the CFD simulation at the tail mid-chord location. There is plenty of clearance between the tail 

and plumes. Very minor plume interactions are evident in the weakly three-dimensional (ie, non-round) 

plume shapes. Multiple slices through the plumes are shown in Figure 59 to give an idea of plume 

development going downstream. As in previous images, the three plumes are seen to be isolated with no 

major interactions. Figure 60 shows the temperature profile of the plume for the in-line configuration, the 

temperature does not exceed 330K (66°C). 

Figure 56. ARES OML superimposed over centerline slice through CFD simulation, in-line case. 
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Figure 57. ARES OML superimposed over slice through CFD simulation, in-line case. 

Figure 58. ARES OML superimposed over vertical slice through CFD simulation, in-line case. 
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Figure 59. ARES OML superimposed over vertical slices through CFD simulation. 

(Locations at X=2.5, 3.0, 3.5, 4.0, 4.5, 5.0, 5.5, and 6.0m (nozzle exits at X~2.3m), in-line case. Contours of Mach 

number above M=0.62 are shown). 

Figure 60. ARES OML superimposed over vertical slices through CFD in-line plume temperature profile. 
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3.6.2 Triple 22N Triangular Thruster Plume Simulation Results 

 Results from the triangular case are presented in this section. . Plume interactions are evident in 

Figure 61, as the top row plumes bend to align with the freestream flow and come together. Figure 61 

shows a slice through the centerline plane of spanwise symmetry for the triangular case. Figure 62 and 

Figure 63 slice through the CFD simulation at incidence angles of 5.88° and 9.38°, aligned with the top 

and bottom rows of thrusters, respectivelyHere, the lower thruster plume is shown clearly, and the 

influence of the upper row plumes is visible at the centerline. Figure 64 shows a vertical slice through the 

CFD simulation at the tail mid-chord location for the triangular case. There is plenty of clearance 

between the tail and plumes. The plumes are still distinct, but are closer together and exhibit stronger 

interactions than the in-line case. 

 Multiple slices through the triangular plumes are shown in Figure 65 to give an idea of plume 

development going downstream. In this configuration, strong plume interactions start to develop around 

X=6m (approx. 1.5m downstream of the tail trailing edge) as the plumes begin to merge. Figure 66 shows 

the temperature profile of the plume for the in-line configuration, the temperature does not exceed 330K 

(66°C). 

Figure 61. ARES OML superimposed over centerline slice through CFD simulation, triangular case. 
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Figure 62. ARES OML superimposed over 5.88° slice through CFD simulation, triangular case. 

Figure 63. ARES OML superimposed over 9.38° slice through CFD simulation, triangular case. 
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Figure 64. ARES OML superimposed over vertical slice through CFD simulation, triangular case. 

Figure 65. ARES OML superimposed over vertical slices through CFD simulation. 

(Locations at X=2.5, 3.0, 3.5, 4.0, 4.5, 5.0, 5.5, and 6.0m (nozzle exits at X~2.3m), triangular case. Contours of 

Mach number above M=0.62 are shown). 
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3.6.3 Triple 22N Thruster Plume Analysis Conclusions 

 Both thruster configurations resulted in stable steady-state plumes, but the triangular case clearly 

showed stronger evidence of plume interactions. This is likely due to the fact that the top row thrusters 

were at a lower horizontal angle of incidence relative to the freestream (2.2°) compared to the in-line case 

(4.5°). This allowed the plumes to more easily bend to align with freestream flow, resulting in closer 

proximity going downstream. In contrast, the in-line plumes maintained an angular separation all the way 

downstream to the simulation boundary and had minimal interactions.

Figure 66. ARES OML superimposed over vertical slices through CFD triangular plume temperature profile. 
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3.7 Triple Thruster Flight Simulation and Performance 

 Flight simulation of the Mars airplane with a three-thruster system was evaluated using the six 

degree-of-freedom model from the Langley Standard Real-Time Simulation application framework5

. The 

firing logic for the thrusters was divided into two separate duty cycles. The primary duty cycle governed 

the firing of two thrusters simultaneously to provide the needed thrust to overcome drag while flying 

under normal conditions. The secondary duty cycle governed the firing of the third thruster and was 

invoked during instances where airplane acceleration or climbing was needed which required more than 

44N of thrust. Similar to the single 62N system, two atmospheric conditions were simulated, one with no 

turbulence, and the other condition with severe atmospheric turbulence. Thruster duty cycle for the 

conditions of no turbulence is shown in Figure 67. The primary duty cycle is shown on the top chart and 

the secondary on the bottom. For conditions of no turbulence, the primary duty cycle stays 

predominantly around 70%, varying between 60% and 85% for the majority of flight. The secondary 

duty cycle shows that the third thruster is only slightly used during the initial stages of flight and stays off 

during the entire remainder of flight. Primary Duty Cycle 

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100% Secondary Duty Cycle 

0% 20% 40% 60% 80% 100%

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

0 500 1000 1500 2000 2500 3000 3500 4000

Flight Time [seconds] Histogram: Percent of Total Flight Time

Figure 67. Duty Cycles of Triple Thruster System - No Turbulence Conditions 
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 During severe turbulence, the data in Figure 68 shows that the two primary thrusters are fully 

firing at 100% duty cycle for a significant portion of the flight, and that the third thruster, or secondary 

duty cycle, is sparsely invoked throughout the flight and remains off for most of the flight. 

3.8 Triple 22N Thruster System Summary 

 The results of the simulation show that for the airplane with three 22N thrusters flying in noturbulence conditions, the flight range with 40 kg of propellant is about 543 km. In severe turbulence 

conditions for the same fuel load, the flight range is reduced to about 536 km (Table 11). 

 The total mass of the triple thruster propulsion system is broken down into components shown in 

Table 12. Based on commercially available components, estimates from vendors, and analysis, the 

current best estimate for the triple thruster system is 16.31 kg for the in-line configuration and 16.26 kg 

for the triangular configuration. Incorporating the contingencies assigned to each component, the total 

mass increases to 19.40 kg for the in-line configuration and 19.34 kg for the triangular configuration. Primary Duty Cycle 

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100% Secondary Duty Cycle 

0% 20% 40% 60% 80% 100%

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

0 500 1000 1500 2000 2500 3000 3500 4000

Flight Time [seconds] Histogram: Percent of Total Flight Time

Figure 68. Duty Cycles of Triple Thruster System - Severe Turbulence Conditions 
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 Table 11. Triple Thruster Propulsion System 

Performance Summary – 40 kg Fuel Load 

Parameter Value 

Total AFS Mass 145 kg 

Propulsion Mass 18.06 – 18.12 kg 

Propellant Load 40 kg 

Fuel 15 kg - MMH 

Oxidizer 25 kg - MON3 

Thrust Three - 22 N 

Isp ~291 s 

No Turbulence

Flight Duration 67.63 min 

Range 543 km 

Avg Fuel Usage 13.6 km/kg 

Severe Turbulence

Flight Duration 66.22 min 

Range 536 km 

Avg Fuel Usage 13.4 km/kg 

Table 12. Triple 22N Thruster Option - Propulsion System Mass Breakdown 

Common Components QTY 

CBE Each 

[kg] 

Total 

[kg] 

Cont 

[%] 

CBE + 

Cont [kg] Vendor Part Number 

Pressure Transducer 5 0.15 0.75 20% 0.90 GP50 7200 

Normally Closed Pyro Valve 5 0.15 0.75 15% 0.86 CONAX 1832-205 

GHe Filter 1 0.25 0.25 25% 0.31 Vacco F1D10636-01 

High Pressure Service Valve 11 0.113 1.243 10% 1.37 Vacco V1E10430-01 

Helium Regulator 1 0.363 0.363 10% 0.40 Vacco 66250 

Check Valve 4 0.02 0.08 25% 0.10 Vacco V1D10856-02 

Liquid Filter 2 0.18 0.36 25% 0.45 Vacco F1D10638-01 

GHe Tank 1 1.2 1.2 25% 1.50 Lincoln 220131-1 

Oxidizer Tank 1 2.87 2.87 25% 3.59 

Fuel Tank 1 2.87 2.87 25% 3.59 

Tank Support Brackets 4 0.1 0.4 25% 0.50 

Subtotal 10.74 13.07 

Triple Thruster Components 

Thruster 3 0.8 2.4 10% 2.64 AMPAC AJ10-220 

Thruster Valves 6 0.19 1.14 10% 1.25 Moog 51-178 

Additional Battery String 12 0.089 1.07 20% 1.28 

Flow Venturi 6 0.04 0.24 5% 0.25 

Misc Tubing, Brackets 1 0.3 0.3 25% 0.38 

PIA & PCA Plates 2 0.08 0.16 25% 0.20 

Subtotal 5.308 6.00 

In-Line Row Configuration 

Engine Support Structure 1 0.264 0.264 25% 0.33 

Total Propulsion Mass 16.31 19.40 

Triangular Configuration 

Engine Support Structure 1 0.214 0.214 25% 0.27 

Total Propulsion Mass 16.26 19.34 
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4.0 Performance Comparison of the Triple Thruster and Single Thruster 

Systems 

4.1 Performance Assumptions 

 The primary driver for comparing the triple thruster system with the single thruster system relies 

on the ultimate performance, or total flight range, achievable by the aircraft. This attribute can be 

subdivided into two categories; total achievable flight range with a fixed fuel load, and total achievable 

flight range given a fixed wet mass (i.e. the fuel load is reduced on the heavier system). In the case of 

fixed fuel load, the two systems were simulated using 40 kg of total propellant under conditions of both 

no-turbulence and severe turbulence. Comparing the performances assuming a fixed wet mass requires 

taking into account the differences in system mass of the three propulsion options studied, and adjusting 

the fuel loads of each system to result in identical system wet masses. The baseline configuration of the 

62N thruster was assumed to carry a fuel load of 40 kg with a dry mass of 15.95 kg. Figure 69 recaps the 

mass breakdown of each of the systems being considered. Since the triple thruster “in-line” configuration 

has 3.4 kg more dry mass, the fuel load was reduced by this amount to 36.6 kg of fuel. Similarly, the 

triple thruster “triangular” configuration fuel load was reduced by the same amount. System mass and 

fuel loads for each case are summarized in Table 13. 

4.2 Performance Summary 

 The single thruster system was able to achieve a flight distances ranging from 520 km to 527 km. 

Both triple thruster configurations under fixed fuel load achieved the same flight distance ranging from 

536 km to 543 km. A graph of these results is show in Figure 70, and summarized in Table 13. The “inline” triple thruster configuration carries a fuel load of 36.56 kg and achieves a flight distance ranging 

from 490 km to 496 km. The “triangular” triple thruster configuration carries a fuel load of 36.62 kg and 

achieves a flight distance ranging from 491 km to 497 km. These results are shown in Figure 71. 

15.95

19.40 19.34

12.0

13.0

14.0

15.0

16.0

17.0

18.0

19.0

20.0

123

Total System Mass [kg]

PIA & PCA Plates

Misc Tubing, Brackets

Engine Support Structure

Flow Venturi

Thruster Valves

Thruster

Common Components

Single 62N 

Thruster

Triple 22N 

Thrusters

(In-Line Row)

Triple 22N 

Thrusters

(Triangular)

Figure 69. System Mass Breakdown Comparison 
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Table 13. System Performance Comparison Summary 

Fixed Fuel Load Fuel 

No Turb 

[km/kg] 

Range 

[km] 

Turb 

[km/kg] 

Range 

[km] 

Average 

Range 

Single Thruster 40 13.2 527.1 13.0 519.7 523.4 

Triple Thrusters (In-Line Row) 40 13.6 542.8 13.4 535.6 539.2 

Triple Thrusters (Triangular) 40 13.6 542.8 13.4 535.6 539.2 

 

Fixed System Wet Mass Fuel 

No Turb 

[km/kg] 

Range 

[km] 

Turb 

[km/kg] 

Range 

[km] 

Average 

Range 

Single Thruster 40 13.2 527.1 13.0 519.7 523.4 

Triple Thrusters (In-Line Row) 36.6 13.6 496.1 13.4 489.5 492.8 

Triple Thrusters (Triangular) 36.6 13.6 496.9 13.4 490.3 493.6 

450.0 470.0 490.0 510.0 530.0 550.0

1

2

3

Single 62N 

Thruster

Triple 22N 

Thrusters

(In-Line Row)

Triple 22N 

Thrusters

(Triangular)

Severe 

Turb

No 

Turb

Figure 70. Comparison of Flight Range for Fixed Fuel Load 

450.0 470.0 490.0 510.0 530.0 550.0

1

2

3

Single 62N 

Thruster

Triple 22N 

Thrusters

(In-Line Row)

Triple 22N 

Thrusters

(Triangular)

No 

Turb

Severe 

Turb

Figure 71. Comparison of Flight Range for Fixed Wet Mass 
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5.0 Decision Criteria for Thruster System Down-Select 

5.1 Decision Criteria 

 The criteria for evaluating the three propulsion system options were derived based on the impact 

of each system on the airplane and overall mission performance, and on Mars Program cost and schedule. 

Out of the six criteria chosen, four relate to technical performance and two relate to program elements. 

However, based on the chosen weighting factors, the sum total of all technical attributes is equal to the 

sum total of all programmatic attributes. The decision criteria and weighting factors are listed in Table 

15. There were other criteria considered in this evaluation process that could not be evaluated well 

quantitatively. A consideration was given to functional redundancy; the airplane is capable of flying 

under limited conditions with only two 22N thruster so the third thruster provides additional redundancy 

over the 62N single thruster system. Additionally, the impact on the center of gravity caused by 

additional thruster mass in the rear of the airplane was also considered, however it was judged to be 

insignificant. 

 The scores for each criterion were derived numerically with the exception of thruster heritage, 

which was derived by a combination of numerical and qualitative methods. Table 14 shows the 

supporting data used for determining the scores for each of the criterion. The scores are calculated based 

Table 14. Supporting Data for Decision Criteria Scores 

62N Single 22N Inline 22N Tripod 

Subsystem Dry Mass 15.95 19.40 19.34 

Flight Range (fixed wet mass) 523.42 492.76 493.60 

Thruster Non-Recurring Cost 1000 125 125 

Unit Cost 125 100 100 

Units Needed 3 7 7 

Total Thruster Costs 1375 825 825 

Thruster Valve Power (each) [watts] 23 26 26 

Total Number of Thruster Valves 2 6 6 

Peak Thruster Valve Power 46 156 156 

Total Thruster On Time (Primary) 2011.48 2744.73 2744.73 

Total Thruster On Time (Secondary) 0.00 250.23 250.23 

Energy Used by Thruster Valves [Whr] 25.70 82.91 82.91 

Thruster Developed 1 1 1 

Thruster Flight Qualified/Flown 1 1 1 

Current Thruster Design/Production 0 2 2 

Total Units Flown in Space ~100 ~1000 ~1000 

LOGARITHM of Total Units Flown 2 3 3 

Thruster Heritage 4 7 7 

Table 15. Decision Criteria and Weighting Factors 

Criteria 

Weight 

Factor 

Energy Used by Thruster Valves 1 

Peak Thruster Valve Power 2 

Subsystem Dry Mass 3 

Flight Range (fixed wet mass) 3 

Total Thruster Costs 4 

Thruster Heritage 5 
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on a ratio relative to the baseline single 62N thruster system. Regarding thruster heritage, the scores were 

derived by subdividing this criterion into four parts; Development: one point is given if the thruster has 

been previously designed and developed, Qualification: one point if the thruster has been flight qualified 

and flown on a mission, Current Production: two points if the design is current and the thruster is in 

production, one point if the design is relatively current but the thruster is not in production, and zero 

points if the design is not current and the unit is not in production, Total Flight Units: the score is 

determined by taking the logarithm of the total units that have flown, therefore one flight unit would equal 

a score of zero, 10 units a score of 1 and so on. 

 The final results are summarized in Table 16 and shown graphically in Figure 72. The total score 

for the 62N baseline system, which is simply the sum of the weight factors, is 18. The two tri-thruster 

systems both scored equally higher than the baseline at a total score of 21.6. An examination of the 

scores shows that there are four criteria having the largest divergence between the single thruster and 

triple-thruster systems; the baseline system scores better under Peak Thruster Valve Power and Energy 

Used by Thruster Valves and the triple-thruster systems scores better under Total Thruster Costs and 

Thruster Heritage. Figure 72 shows that the sum of five of the six scores for all three propulsion systems 

are relatively equal, with the Thruster Heritage being the decisive factor between the options considered. 

Table 16. Decision Criteria Scores 

Criteria 

Weight 

Factor

Baseline 

Single 

14-lbf 

Thruster 

Weighted 

Score 

TriThruster 

22N 

In-Line 

Weighted 

Score 

TriThruster 

22N 

Tripod 

Weighted 

Score 

Subsystem Dry Mass 3 1 3 0.8 2.5 0.8 2.5 

Flight Range (fixed wet mass) 3 1 3 0.9 2.8 0.9 2.8 

Peak Thruster Valve Power 2 1 2 0.3 0.6 0.3 0.6 

Energy Used by Thruster Valves 1 1 1 0.3 0.3 0.3 0.3 

Total Thruster Flight Unit Cost 4 1 4 1.7 6.7 1.7 6.7 

Thruster Heritage 5 1 5 1.8 8.8 1.8 8.8 

Total 18 21.6 21.6 

18

21.6 21.6

0.0

5.0

10.0

15.0

20.0

25.0

123

Total Weighted Criteria

Thruster Heritage

Total Thruster Costs

Energy Used by Thruster Valves

Peak Thruster Valve Power

Flight Range (fixed wet mass)

Subsystem Dry Mass

Single 62N 

Thruster

Triple 22N 

Thrusters

(In-Line Row)

Triple 22N 

Thrusters

(Triangular)

Figure 72. Decision Criteria Scores 
Trade Study of Multiple Thruster Optionsfor the Mars Airplane Concept 

NASA/TM-2009-215699 

61

5.2 Propulsion System Selection 

 The thruster options considered in this trade study are shown again in Figure 73. The triple 

thruster options score almost 21% higher on the decision criteria than the single thruster option. The 

triple in-line configuration option and the triple triangular configuration option scored essentially the 

same on the decision criteria. As stated in Section 3.6.3, both triple thruster configurations resulted in 

stable steady-state plumes, but the triangular case clearly showed stronger evidence of plume interactions 

and the in-line plumes maintained an angular separation all the way downstream to the simulation 

boundary and had minimal interactions. 

 

The results of this study conclude that the triple 22N in-line thruster configuration is 

recommended for the new baseline Mars Airplane propulsion system. 

Figure 73. Recap of Trade Study Thruster Options 
Trade Study of Multiple Thruster Optionsfor the Mars Airplane Concept 

NASA/TM-2009-215699 

62

6.0 Appendix A – Other Thrusters Considered 

6.1.1 Thruster Candidate #2 - Aerojet 5-lbf Thruster 

 The Aerojet Model R-6C 5.0 lbf earth storable bipropellant rocket engine (MON-3/MMH) is 

flight qualified and is currently operational on the Lockheed 

Martin MILSTAR satellite, NASA GOES, ISRO Insat 2B 

and others. The R-6C is distinguished from some other 

attitude control 5.0 lbf thrusters by its unique ability to 

operate at any duty cycle, including both pulse mode and 

steady state. This ability, enabled by a regeneratively cooled 

injector design, has been fully mapped over a broad spectrum 

of duty cycles and equilibrium long-term steady state firings.

 

 The Aerojet Model R-6DM 5-lbf (22N) bipropellant 

(nitrogen tetroxide oxidizer and monomethylhydrazine fuel) 

rocket engine (Figure 1) which incorporates a single seat 

Moog bipropellant torque motor valve was qualified for 

space flight in 1989. The engine uses a fuel boundary layer 

film cooled combustion chamber and an unlike impinging 

multi-doublet injector to provide a specific impulse of 

approximately 294 seconds (2884 N Sec/Kg). This design 

provides a 100% increase in combustion chamber life when 

compared to the previously qualified and flight proven R-6C 

configuration which uses a single unlike impinging injector and active cooling of the injector head with 

fuel. The R-6D engine design will also interface with the Aerojet produced flexure guided single seat (R6D), series redundant seat valves (R-6DSR) and the Moog series redundant torque motor valve. The R-6D 

rocket engine accumulated 61,818 seconds of firing time (341 full thermal cycles) during design 

verification tests at Aerojet and during the subsequent qualification program demonstrated 336,331 

pulses. 

6.1.2 Thruster Candidate #3 - Astrium (DASA) 22N Thruster 

Astrium is a company from Munich, Germany with a new 22 N thruster that is a non-coated bipropellant 

rocket engine for pressurized propellant feed system using 

the storable propellants N2O4, MON-1 or MON-3 as 

oxidizer, MMH as fuel and helium as pressurant. The 

thruster is designed for both long term steady state 

operation and pulse mode down to 8 ms on-times. It 

operates with regulated pressure as well as in blown down 

mode. In addition it can be pulse operated far outside its 

nominal operating range in certain pulse modes if e.g. 

system tank pressures are exceeding tolerable limits. This 

thruster is a consequent evolution of the well proven 

Astrium 10N Pt/Rh thruster of which Astrium has already 

delivered several hundreds for numerous satellite 

programs. 

Figure 74. Photos of Alternate Thruster - 

Aerojet 5-lbf Thruster 

Figure 75. Model of Alternate Thruster - 

Astrium 22N 
Trade Study of Multiple Thruster Optionsfor the Mars Airplane Concept 

NASA/TM-2009-215699 

63

7.0 Appendix B – Structural Analysis Assumptions 

7.1 Structure Load Cases 

 Table 17. Mission Structural Load Cases 

No. Description Value FS: 

 Metals

FS: 

Composite

0 Launch Natural Frequency Axial > 50 Hz 

Lateral > 20 Hz 

N/A N/A 

1a Launch (Use total thruster mass for MAC) Axial = MAC+2.8 g 

Lateral = 4.5 g 

Yield = 1.25 

Ult. = 1.5 

2.0 

1b Launch (Use total thruster mass for MAC) Any = MAC g Yield = 1.25 

Ult. = 1.5 

2.0 

2 3rd Stage Cutoff Axial = 15.3 g 

Lateral = 0.1 g 

Angular Vel = 80 RPM 

Yield = 1.25 

Ult. = 1.5 

2.0 

3a 3rd Stage Spin-Up Angular Vel = 80 RPM 

Angular Accel = 11 rad/s2

Yield = 1.25 

Ult. = 1.5 

2.0 

3b 3rd Stage Ignition Axial = 7.5 g 

Lateral = 0.1 g 

Angular Vel = 80 RPM 

Yield = 1.25 

Ult. = 1.5 

2.0 

4a Parachute deployment Axial = 15 g 

Angular Vel = 2 RPM 

Yield = 1.25 

Ult. = 1.5 

2.0 

4b Parachute deployment – off axis Axial = 10 g 

Lateral = 3 g 

Angular Vel = 2 RPM 

Yield = 1.25 

Ult. = 1.5 

2.0 

5 AES Impact Axial = 15 g Yield = 1.25 

Ult. = 1.5 

2.0 

6 Thruster Ignition 5 lbf (ARC SW0003) 

14 lbf (AJ10-220) 

Yield = 1.25 

Ult. = 1.5 

2.0 
Trade Study of Multiple Thruster Optionsfor the Mars Airplane Concept 

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64

7.2 Load Case Definitions 

• Orientation 

– Axial is defined as the longitudinal direction of the launch vehicle and the 

aeroshell (or perpendicular to the plane of the center-body) 

– Lateral is any direction orthogonal to the axial direction 

• g’s = 1 Earth g – 9.807m/sec2

• Ultimate Margin of Safety: 

– Margin shall be positive 

• Yield Margin of Safety: 

– Margin shall be positive 

Aeroshell outline 

Axial Direction 

Figure 76. Flight vehicle folded in aeroshell. 

1 _ _ − ∗ = FS Stress

Allow Ult Stress MS

ult

Ult

1 _ _ − ∗ = FS Stress

Allow Yield Stress MS

Yield

Yield
Trade Study of Multiple Thruster Optionsfor the Mars Airplane Concept 

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65

7.3 Mass Acceleration Curve 

1

10

100

1 10 100 1000

Mass (kg)

Acceleration (g)

Figure 77. Mass Acceleration Curve (MAC) 
Trade Study of Multiple Thruster Optionsfor the Mars Airplane Concept 

NASA/TM-2009-215699 

66

8.0 Appendix C - Thermal Analysis Assumptions and Load Cases 

Table 18. Source and value for all surface thicknesses used in the model 

Component Thickness (mm) Source 

Carbon Face Sheets, fuselage skin outer 

& inner 

0.381 (used estimate of 3 fabric 

sheets, actually thicker or thinner 

in some locations) 

Aurora Assembly SchematicsError! 

Bookmark not defined.

Carbon Face Sheets, inner c-channel of 

spars, ribs & bulkheads 

1 (estimate, may be thicker in 

some spots) 

Aurora Assembly SchematicsError! 

Bookmark not defined.

Carbon Face Sheets, outer c-channel of 

spars, ribs & bulkheads 

4.2 Aurora Assembly SchematicsError! 

Bookmark not defined.

Carbon Face Sheets, c-channel caps 8.636 (this value is for the carbon 

sheets plus the carbon tape) 

Aurora Assembly SchematicsError! 

Bookmark not defined.

Honeycomb Core, all 6.35 Aurora Assembly SchematicsError! 

Bookmark not defined.

Thruster nozzle 1.04 CAD Geometry Measurement 

Thruster injector coupling 1.157 CAD Geometry Measurement 

Thruster valves 1.71 CAD Geometry Measurement 

Heat Shield 0.5 Picked value 

Table 19. Sources for material properties used in this study 

COMPONENT MATERIAL PROPERTY SOURCE 

Carbon Face Sheets Carbon Fabric, Newport Composites Previous ARES analysis1, 

Graphite/Epoxy example11

Honeycomb Core Aluminum Honeycomb, Hexcel Previous ARES analysis1, 

Aluminum Honeycomb, HexWeb, 

HexWeb Manual - 5056 Al, 3/8" cell 

Thruster nozzle, combustion 

chamber, & injector bracket 

Columbium Reference 11 

Thruster valve mounting flange and 

mounting brackets 

Titanium 5Al-2.5Sn Reference 11 

Thruster valves Stainless Steel 304 Reference 11 

Heat Shield Aluminum Patran Thermal Database 

Table 20. Boundary conditions used in thermal analysis 

Boundary Conditions Value Description/Source 

Martian Ambient Air Temperature -70 °C Taken from previous model1. MarsGRAM 2001, 

Dick Davis (NASA GSFC) 

Martian Ground Temperature -68 °C Taken from previous model1. MarsGRAM 2001, 

Dick Davis (NASA GSFC) 

Space Temperature -269 °C Standard 4 K space temperature 

Thruster temperature profile Varying, 

°C 

Using the data provided in Figure 24, two linear 

distributions were created for the thrusters, one from 

the T0 through T6 to the nozzle end, and one from 

T0 through T1 to the end of the combustion chamber. 

This may actually make the end of the combustion 

chamber hotter that it actually gets, but is a 

conservative estimate. The injector bracket faces 

were set at T2 and T4. The valve mounting flange 

was set to T5. All other components were not given 

fixed temperatures. 
Trade Study of Multiple Thruster Optionsfor the Mars Airplane Concept 

NASA/TM-2009-215699 

67

Table 21. Loads used in thermal analysis 

Loads Value Description/Source 

Contact conductance between Al 

honeycomb and carbon fiber face 

sheets, all components 

0.006299 W/mm2K Taken from previous model1. Based on a gap 

with k = 4E-4 W/mmK and L = 0.0635 mm, 

Contact conductance between support 

structure bonds (bulkheads to ribs, 

spars to ribs, mounting brackets to 

bulkheads, upper hatches to rest of 

fuselage) 

0.006299 W/mm2K Used same as skin core convection. Heat 

transfer will be through the titanium corner 

brackets and possibly through any glue or 

bonding material in the joint. Contact 

resistance for a titanium-carbon fiber joint is 

not readily available. A brass alloy – brass 

Alloy bond with the highest roughness has a 

contact conductance of 0.0025 W/mm2

K at 

low pressure12. The value being used is most 

likely high if there is no adhesive or low if 

there is adhesive. 

Contact conductance between all metal 

– metal bonds (valve mounting flange 

contact, mounting brackets to thruster 

connections) 

0.0033 W/mm2K Used SS 416 at 1 MPa, (502.5 pg 7 curve i)12. 

This value is for standard atmospheric 

pressure. On Mars the pressure will be much 

less so for the same contact pressure, this 

should be a conservative estimate. 

Mars albedo heating to bottom of 

aircraft 

4.513 x 10-5 W/mm2 Taken from previous model. 

Solar and diffuse heating to top of 

aircraft 

0.000285 W/mm2 Taken from previous model1. Heating value 

was calculated assuming a coating of white 

paint 293 (α = 0.19, ε = 0.85)13. 

Radiation to ground from aircraft 

bottom 

ε = 0.85 Common emissivity for carbon fabric. 

Radiation is to ground temperature BC which 

is set at an emissivity of 1. The Martian 

surface probably does not have an emissivity 

of 1, and future studies should take this into 

account. 

Radiation to space from aircraft top ε = 0.85 Common emissivity for carbon fabric. 

Radiation is to space temperature BC which 

has an emissivity of 1 

CO2 convection on fuselage varying convection 

coefficient based on 

boundary layer growth over 

the fuselage 

Taken from previous model1. Value for h was 

derived by Joseph Gasbarre (NASA Langley, 

STSB). 

View factor radiation calculations εcolumbium = 0.19 

εtitanium = 0.31 

εstainless = 0.31 

εfuselage = 0.85 

εlow ε paint = 0.3 (when used) 

εheat shield = 0.05 (when used) 

Applied to thruster components and portions 

of the fuselage. All metal values are from 

Reference 11. Polished aluminum heat shield 

value from Reference 13 The low ε coating is 

an estimate, can probably find one lower. 

Thruster exhaust plume properties, 

estimates 

T = 400 °C, 700 °C 

ε = 0.5 

The temperatures represent an estimated range 

for the plume. The emissivity is estimated as 

the highest value of H2O vapor at 700 °C 

contained in Figure 9-16 in Reference 14. 

Since the combustion products do not contain 

H2O vapor, this it is difficult to gage the 

accuracy of this assumption. 
Trade Study of Multiple Thruster Optionsfor the Mars Airplane Concept 

NASA/TM-2009-215699 

68

9.0 References 

1

 “ARES, an Aerial Regional-scale Environmental Survey of Mars – Concept Study Report,” proposed in response 

to the NASA Office of Space Science Announcement of Opportunity AO 02-OSS-02 for Mars Scout investigations, 

April 30, 2003. 

2

 “Aerojet 14 lbf Satellite Propulsion Program”, Aerojet Internal Document AP98AL313-F, Provided to NASA 

LaRC in support of Phase-2 Mars Scout Proposal, April 2003. 

3

 “Advanced Small Rocket Chambers, Option 1 – 14 lbf IR/Re Rocket Final Report,” NASA CR-191014, D. M. 

Jassowski, M.L. Gage, GenCorp Aerojet, August 1992. 

4

 “Liquid Rocket Propulsion for Atmospheric Flight in the Proposed ARES Mars Scout Mission,” C. Kuhl, H. 

Wright, C. Hunter, NASA Langley Research Center, Hampton, VA; C. Guernsey, Jet Propulsion Laboratory, 

Pasadena, CA; A. Colozza, NASA Glenn Research Center, Cleveland, OH; AIAA-2004-3696. 40th 

AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Fort Lauderdale, Florida, July 11-14, 2004 

5

 “Simulating the ARES Aircraft in the Mars Environment,” Kenney, P.S., AIAA-2003-6579, 2nd AIAA “Unmanned 

Unlimited” Systems, Technologies, and Operations – Aerospace Conference, San Diego, California, September 15-

18, 2003. 

6

 “Mars Airplane Propulsion Study – Final Report,” NASA Contract NAS1-00141, Lockheed Martin Astronautics, 

April 19, 2005. 

7

 “Long Life 5 LBF Bipropellant Engines,” AIAA-85-1378, Chazen, M.L., Bell Aerospace Textron, 

AIAA/SAE/ASME/ASEE 21st Joint Propulsion Conference, Monterey, California, July 8-10, 1985. 

8

 “Development Tests on a 5-LBF Bipropellant Thruster Using a Platinum/Rhodium Thrust Chamber,” Driscoll, 

R.J., Yager, J., Roy, M.J., Kammerer, H.G., Atlantic Research Corporation, AIAA-98-3357, 

AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 34th, Cleveland, OH, July 13-15, 1998. 

9

 HADD2 Assembly Drawings, H2-SF101, H2-SF102, H2-SF104, H2-SF105, June 24, 2003, Aurora Flight 

Sciences Corporation, Manassas, Va. 

10 “Transient thermal model of a film-cooled bipropellant thruster ,” G. P. Purohit, P. A. Donatelli, J. R. Ellison 

(Hughes Space and Communications Co., El Segundo, CA), and V. K. Dhir (California, Univ., Los Angeles), 

AIAA-2000-1072, Aerospace Sciences Meeting and Exhibit, 38th, Reno, NV, Jan. 10-13, 2000. 

11 TPSX Material Properties Database, Web Edition. http://tpsx.arc.nasa.gov/, NASA 

12 Heat Transfer and Fluid Flow Data Books. Genium Publishing Co., Schenectady, NY, 1985. 

13 Spacecraft Thermal Control Handbook., Gilmore, David G., The Aerospace Corporation Press, 1994. 

14 Radiative Heat Transfer. Modest, Michael F., McGraw-Hill Inc. 1993. 
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

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Technical Memorandum

 4. TITLE AND SUBTITLE

Trade Study of Multiple Thruster Options for the Mars Airplane Concept

5a. CONTRACT NUMBER

 6. AUTHOR(S)

Kuhl, Christopher A.; Gayle, Steven W.; Hunter, Craig A.; Kenney, 

Patrick S.; Scola, Salvatore; Paddock, David A.; Wright, Henry S.; 

Gasbarre, Joseph F.

 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Langley Research Center

Hampton, VA 23681-2199

 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

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Subject Category 20

Availability: NASA CASI (443) 757-5802

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email: help@sti.nasa.gov)

14. ABSTRACT

A trade study was performed at NASA Langley Research Center under the Planetary Airplane Risk Reduction (PARR) project 

(2004-2005) to examine the option of using multiple, smaller thrusters in place of a single large thruster on the Mars airplane

concept with the goal to reduce overall cost, schedule, and technical risk. The 5-lbf (22N) thruster is a common reaction 

control thruster on many satellites. Thousands of these types of thrusters have been built and flown on numerous programs, 

including MILSTAR and Intelsat VI. This study has examined the use of three 22N thrusters for the Mars airplane propulsion 

system and compared the results to those of the baseline single thruster system. 

15. SUBJECT TERMS

Airplane; Bipropellant; Propulsion; Rocket; Thruster

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