NASA/TM—2013–216620 

 

Mars Sample Return: Mars Ascent Vehicle 

Mission & Technology Requirements 

Jeffrey V. Bowles 

Systems Analysis and Integration Branch 

Ames Research Center, Moffett Field, California

Loc C. Huynh 

Science and Technology Corporation, Hampton, VA 

Veronica M. Hawke 

Science and Technology Corporation, Hampton, VA 

Xun J. Jiang 

Science and Technology Corporation, Hampton, VA 

 Click here: Press F1 key (Windows) or Help key (Mac) for help 

November 2013
NASA STI Program ... in Profile 

Since its founding, NASA has been dedicated 

to the advancement of aeronautics and space 

science. The NASA scientific and technical 

information (STI) program plays a key part in 

helping NASA maintain this important role. 

The NASA STI program operates under the 

auspices of the Agency Chief Information Officer. 

It collects, organizes, provides for archiving, and 

disseminates NASA’s STI. The NASA STI 

program provides access to the NASA 

Aeronautics and Space Database and its public 

interface, the NASA Technical Reports Server, 

thus providing one of the largest collections of 

aeronautical and space science STI in the world. 

Results are published in both non-NASA channels 

and by NASA in the NASA STI Report Series, 

which includes the following report types: 

TECHNICAL PUBLICATION. Reports of 

completed research or a major significant 

phase of research that present the results of 

NASA Programs and include extensive data 

or theoretical analysis. Includes compilations of significant scientific and technical 

data and information deemed to be of 

continuing reference value. NASA counterpart of peer-reviewed formal professional 

papers but has less stringent limitations on 

manuscript length and extent of graphic 

presentations. 

TECHNICAL MEMORANDUM. 

Scientific and technical findings that are 

preliminary or of specialized interest, 

e.g., quick release reports, working 

papers, and bibliographies that contain 

minimal annotation. Does not contain 

extensive analysis. 

CONTRACTOR REPORT. Scientific and 

technical findings by NASA-sponsored 

contractors and grantees. 

CONFERENCE PUBLICATION. 

Collected papers from scientific and 

technical conferences, symposia, seminars, 

or other meetings sponsored or 

co-sponsored by NASA. 

SPECIAL PUBLICATION. Scientific, 

technical, or historical information from 

NASA programs, projects, and missions, 

often concerned with subjects having 

substantial public interest. 

TECHNICAL TRANSLATION. 

English-language translations of foreign 

scientific and technical material pertinent to 

NASA’s mission.

Specialized services also include organizing 

and publishing research results, distributing 

specialized research announcements and 

feeds, providing information desk and personal 

search support, and enabling data exchange 

services. 

For more information about the NASA STI 

program, see the following: 

Access the NASA STI program home page 

at http://www.sti.nasa.gov 

E-mail your question to help@sti.nasa.gov 

Fax your question to the NASA STI 

Information Desk at 443-757-5803 

Phone the NASA STI Information Desk at 

443-757-5802 

Write to: 

STI Information Desk 

NASA Center for AeroSpace Information

7115 Standard Drive 

Hanover, MD 21076-1320 

This page is required and contains approved text that cannot be changed. 
NASA/TM—2013–216620 

 

Mars Sample Return: Mars Ascent Vehicle 

Mission & Technology Requirements 

Jeffrey V. Bowles 

Systems Analysis and Integration Branch 

Ames Research Center, Moffett Field, California 

Loc C. Huynh 

Science and Technology Corporation, Hampton, VA 

Veronica M. Hawke 

Science and Technology Corporation, Hampton, VA 

Xun J. Jiang 

Science and Technology Corporation, Hampton, VA 

 

 Click here: Press F1 key (Windows) or Help key (Mac) for help 

National Aeronautics and 

Space Administration

Ames Research Center 

Moffett Field, CA 94035-1000 

November 2013
Acknowledgments 

This work was performed by the AUS (Systems Analysis and Integration) Branch at NASA 

Ames in conjunction with AEMMS contract NNA10DF25C with Science and Technology 

Corporation. 

 Click here: Press F1 key (Windows) or Help key (Mac) for help 

Available from: 

 Click here: Press F1 key (Windows) or Help key (Mac) for help 

This report is also available in electronic form at 

http://

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 
1

 

Table of Contents 

1.0 Introduction ............................................................................................................................................. 4 

2.0 Objective and Technical Approach ......................................................................................................... 4 

3.0 Mission Requirements ............................................................................................................................ 4 

4.0 Assumptions and Ground Rules .............................................................................................................. 5 

5.0 Initial Baseline Configuration Selection ................................................................................................. 6 

5.1 Baseline Engine .................................................................................................................................. 6 

5.2 Candidate Body Geometries ............................................................................................................... 8 

5.3 Aerodynamics & Aerothermodynamics ............................................................................................ 10 

5.4 General Arrangement and Tank Configuration................................................................................. 13 

5.5 Mass Estimating Relationships ......................................................................................................... 15 

6.0 Trajectory .............................................................................................................................................. 16 

7.0 Baseline Configurations ........................................................................................................................ 21 

7.1 Vehicle Closure ................................................................................................................................. 21 

7.2 Orbital Architecture .......................................................................................................................... 21 

7.3 C3=0 Architecture ............................................................................................................................. 23 

8.0 Trade Studies ........................................................................................................................................ 25 

8.1 Number of Engines ........................................................................................................................... 26 

8.2 Pressure-Feed Propulsion System ..................................................................................................... 27 

8.3 Cryogenic Propellant Trade .............................................................................................................. 32 

9.0 Final Configurations ............................................................................................................................. 38 

9.1 Inert Capsule to low Mars orbit rendezvous ..................................................................................... 38 

9.2 Optimized ΔV Split Architecture ...................................................................................................... 38 

9.3 SEP Architecture ............................................................................................................................... 39 

10.0 Conclusions & Recommendations ...................................................................................................... 40 

Appendix A: Engine Trades ........................................................................................................................ 42 

Appendix B: Propellant Trades ................................................................................................................... 46 

Appendix C: CFD Solutions & Aeroheating .............................................................................................. 48 

References ................................................................................................................................................... 54

 
2

Table of Figures 

Figure 5.1.1 Burn-out mass fraction versus engine specific impulse ...................................................... 7

Figure 5.1.2 Propellant fraction and overall body length versus engine expansion ratio ........................... 7

Figure 5.1.3 Gross liftoff mass and dry mass versus engine expansion ratio ........................................... 8

Figure 5.2.1 Parametric body shapes ................................................................................................. 9

Figure 5.2.2 Overall vehicle length and total wetted area versus body fineness ratio ............................... 9

Figure 5.2.3 Overall vehicle length and total wetted area versus body fineness ratio ............................. 10

Figure 5.3.1 Drag coefficient versus angle-of-attack and Mach number ............................................... 11

Figure 5.3.2 CBAero predicted zero-lift drag coefficient versus Mach number and free stream dynamic 

pressure ....................................................................................................................................... 12

Figure 5.3.3 CBAero predicted drag coefficient versus angle-of-attack and Mach number for dynamic 

pressure = 0.005 bars .................................................................................................................... 12

Figure 5.3.4 CBAero predicted lift coefficient versus angle-of-attack and Mach number for dynamic 

pressure = 0.005 bars .................................................................................................................... 13

Figure 5.4.1 MAV general arrangement .......................................................................................... 14

Figure 5.4.2 MAV internal arrangement .......................................................................................... 14

Figure 5.5.1 Inter-tank and aft compartment structure mass estimate model (taken from Reference 5)..... 15

Figure 5.5.2 Rocket engine weight-to-thrust ratio versus engine thrust ................................................ 16

Figure 6.1 MAV ascent trajectory: Altitude versus time .................................................................... 17

Figure 6.2 MAV ascent trajectory: Altitude versus relative velocity .................................................... 18

Figure 6.3 MAV ascent trajectory: Dynamic pressure versus time ...................................................... 18

Figure 6.4 MAV ascent trajectory: Acceleration versus time .............................................................. 19

Figure 6.5 MAV ascent trajectory: Dynamic pressure versus relative velocity ...................................... 19

Figure 6.6 Orbital relative velocity loss breakdown .......................................................................... 20

Figure 6.7 Escape relative velocity loss breakdown .......................................................................... 20

Figure 7.2.1 Orbital architecture closure point versus GLOM and payload mass ................................... 22

Figure 7.2.2 Orbital Architecture, 200 kg payload MAV dimensions .................................................. 23

Figure 7.3.1 C3=0 architecture closure mass versus GLOM and payload mass...................................... 24

Figure 7.3.2 C3=0, 120 kg payload MAV dimensions ....................................................................... 25

Figure 8.2.1 Propellant fraction required versus chamber pressure ...................................................... 28

Figure 8.2.2 Overall vehicle length versus chamber pressure.............................................................. 28

Figure 8.2.3 Tank, pressurization and engine mass versus chamber pressure ........................................ 29

Figure 8.2.4 GLOW and dry mass versus chamber pressure ............................................................... 29

Figure 8.2.5 GLOW and dry mass versus chamber pressure ............................................................... 30

Figure 8.2.6 Vehicle overall versus chamber pressure ....................................................................... 31

Figure 8.3.1 Relative gross liftoff mass versus relative engine specific impulse .................................... 35

Figure 8.3.2 Relative dry mass versus relative engine specific impulse ................................................ 36

Figure 8.3.3 Relative vehicle length versus relative engine specific impulse ........................................ 36

Figure 8.3.4 Relative gross liftoff mass versus relative engine mass .................................................... 37

Figure 8.3.5 Relative dry mass versus relative engine mass ................................................................ 37

Figure 8.3.6 Relative overall vehicle length versus relative engine specific mass .................................. 38

Figure A1. Engine specific impulse versus chamber pressure and nozzle expansion ratio. ..................... 43

Figure A2. Engine thrust per unit nozzle exit area versus chamber pressure and nozzle expansion ratio. .. 43

Figure A3. Engine mass versus chamber pressure and nozzle expansion ratio. ..................................... 44

Figure A4. Engine diameter versus chamber pressure and nozzle expansion ratio. ................................ 44

Figure A5. Engine thrust-to-weight versus chamber pressure and nozzle expansion ratio. ...................... 45

Figure B1. Engine specific impulse versus nozzle expansion ratio for various hypergolic fuels with NTO.

 .................................................................................................................................................. 46

Figure B2. Engine specific impulse versus nozzle expansion ratio for various hydrocarbon fuels with liquid 

oxygen ........................................................................................................................................ 47

Figure C1. Cart3D predicted subsonic Mach = 0.6 surface pressure distribution for the Rev0 geometry . 49
3

Figure C2. Cart3D predicted subsonic Mach = 0.5 surface pressure distribution and Mach number 

contours for the Rev2 geometry...................................................................................................... 49

Figure C3. Cart3D predicted Mach = 0.95 surface pressure distribution and Mach number contours for the 

Rev2 geometry ............................................................................................................................. 50

Figure C4. Cart3D predicted Mach = 2.0 surface pressure distribution and Mach number contours for the 

Rev2 geometry ............................................................................................................................. 50

Figure C5. Cart3D predicted Mach = 2.0 and angle-of-attack = 10.0° surface pressure distribution and 

Mach number contours for the Rev2 geometry ................................................................................. 51

Figure C6. Zero-lift drag coefficient versus Mach number computed by CFD Euler code ...................... 51

Figure C7. Cart3D predicted pitching moment coefficient versus angle-of-attack and Mach number 

contours for the Rev2 geometry...................................................................................................... 52

Figure C8. Mission maximum convective heating distribution for Rev2 geometry ............................... 52

Figure C9. Mission integrated heat load for Rev2 geometry .............................................................. 53

Table of Tables 

Table 1 Ascent Velocities and mass fractions ................................................................................... 17

Table 2. Orbital architecture characteristics ..................................................................................... 23

Table 3. Escape architecture characteristics ..................................................................................... 25

Table 4 Number of Engine comparison ........................................................................................... 26

Table 5. 20kg payload pump feed versus pressure feed system comparison.......................................... 30

Table 6. 200 kg Payload mass pump feed versus pressure feed system comparison ............................... 31

Table 7. Propellant properties......................................................................................................... 33

Table 8. Orbital/200kg payload hypergolic and cryogenic performance comparison ............................. 34

Table 9. Orbital/200kg comparison between hypergolic and cryogenic propellants ............................... 34

Table 10. C3=0/120kg comparison between hypergolic and cryogenic propellants ............................... 35

Table 11. Final concepts mass statement and dimensions................................................................... 40

Revision History 

Original Release March 2013 NASA/TM-2013-216511 

RevA November 2013 
4

1.0 Introduction 

A Mars Sample Return (MSR) mission is the highest priority science mission for the next 

decade recommended by the recent Decadal Survey of Planetary Science, the key community 

input process that guides NASA’s science missions. A feasibility study was conducted of a 

potentially simple and low cost approach to Mars Sample Return mission enabled by the use of 

developing commercial capabilities. Previous studies of MSR have shown that landing an all up 

sample return mission with a high mass capacity lander is a cost effective approach. The 

approach proposed is the use of an emerging commercially available capsule to land the launch 

vehicle system that would return samples to Earth. This paper describes the mission and 

technology requirements impact on the launch vehicle system design, referred to as the Mars 

Ascent Vehicle (MAV). 

2.0 Objective and Technical Approach 

The objective of this study is to determine the mission and technology requirements impact 

on the MAV design by performing a conceptual design of several candidate configurations to 

establish baseline designs, and then executing sensitivity and technology trade studies on these 

baseline concepts. A set of mass estimating relationships (MERs) at the subsystem level were 

developed for this class of vehicle, and integrated into a vehicle synthesis code for computing 

mass and volume, and performing vehicle closure to meet mission requirements. These MERs 

included the expected elements such as structures, power system, propulsion system, nose 

fairing, thermal insulation, actuation devices, guidance and communication. 

 A parametric set of candidate outer mold-line (OML) configurations were defined and 

associated geometric characteristics determined. Aero and aero-thermal databases were 

developed for each OML configuration. Preliminary selection of the propulsion system was 

selected and initial trajectory optimization preformed to establish total ΔV requirements. 

Baseline deigns were established for two Mars Sample Return architectural approaches: 

launching an inert or powered Earth Return Vehicle (ERV) to low Mars orbit, and launching a 

powered ERV to escape velocity, with the ERV providing the remaining ΔV for trans-earth 

injection/capture. With these baseline designs determined, trade studies on mission 

requirements and alternate technology approaches were performed. 

3.0 Mission Requirements 

The mass and volume available for the launch stack (MAV and ERV) contained within a 

landing capsule depends on the possible mass/volume that can be landed. Preliminary analysis 

has indicated that a capsule could land up to 2+ mT at terrain elevations between -1 and +1.6 km 

elevation. Open literature diagrams of a representative capsule show that the central internal 

volume is approximately 1.2 m diameter and 4+ m length. The landed mass and size provide 

boundary conditions for the other elements of the study, and impose unique requirements and 

considerations on the MAV design 

The Mars Ascent vehicle (first stage) was designed for an assumed single stage to orbit or 

escape C3=0 or points in-between. The launch point was 0m MOLA altitude, 0o Latitude and 0o 

Longitude, heading due East. A 5% reserve for ascent propellants was assumed. 

Payload volume was based on an assumed payload density, which in turn was a function of 

the type of ERV propulsion concept. The MAV was sized to provide sufficient internal volume to 
5

accommodate propellants and payload volume, with an assumed packing efficiency to provide 

sufficient internal volume of other subsystems. 

4.0 Assumptions and Ground Rules 

The MAV was sized for each value of total ΔV and assumed payload mass using a set of 

assumptions and ground rules. These assumptions are based on best practices and experience 

for the conceptual design phase. 

Load bearing structures, including forward and aft compartments, and inter-tank were 

assumed to be high temperature thermo-plastic composites, skin-stringer stiffened semimoncoque construction. A design load of 5.0 axial Earth g’s (launch from Earth) was used to 

size the load bearing structure, with a safety factor of 1.4 on loads and a knock-down factor of 

0.80 on stiffness. Non-load bearing aeroshell structure was assumed to be min-gauge (5 plies) 

thermo-plastic. Secondary structure was assumed to be 10% of primary structure. 

Propellant tanks were also assumed to be high temperature thermo-plastic composites with a 

liner, and were sized using historical weight trends. A 20% reduction in tank mass was used to 

account for composite construction, as opposed to metal tanks used in the historical correlation 

for tank weights. A factor of safety of 2.0 was used for tank internal pressure. A 5% ullage 

volume was assumed for all propellant tanks. For hypergolic propellants, a low ullage pressure 

of .0345 MPa (5 psig) was used, while for cryogenic tanks a tank pressure of 0.172 MPa (25 

psig) was assumed. For pressure feed propellant systems, tank pressure was 25% higher than 

the thrust chamber pressure to account for total pressure losses the feed system. Because of 

low tank pressure, hypergolic tanks were assumed to have a dome eccentricity of 0.90 (relatively 

“flat” domes) to improve overall vehicle packing efficiency, while higher pressure tanks 

(cryogenic propellants and pressure-feed systems) used a tank dome eccentricity of 0.707. 

Because of the relatively benign aeroheating environments for Mars launch, minimum gauge 

P-45 cork was applied to the spherical nose cap region for thermal protection. Base closeout 

TPS consisted of 0.25 inches of P-45 cork. Cryogenic tank insulation was assumed to be lightweight closed-cell foam and was sized to provide 15 minute pre-launch hold time to prevent CO2

condensation on the tank exterior walls. 

Rocket engine performance was computed assuming chemical equilibrium in the combustor 

and up to the nozzle throat. Downstream of the throat, the nozzle flow was assumed to be 

chemically frozen (i.e. combustion product constituent mole fractions fixed at the throat values). 

Propellant initial temperature was assumed to be -10°C for hypergolic propellants and at normal 

boiling point for cryogenic propellants. Based on preliminary ascent trajectory optimizations, a 

liftoff thrust-to-weight ratio of 2.5 was used. Propellant oxidizer-to-fuel ratio was optimized for 

maximum engine specific impulse and nozzle expansion ratio for minimum launch mass. Startup 

propellant was computed assuming 2.0 seconds of engine ignition time. 

Lithium-ion batteries were used for prime power and assumed power density of 0.056 watts 

per pound with a 20% power margin applied. Five controller functions were used at 0.1kW per 

channel. A total ascent time of 300 seconds was used to size battery mass. Controller mass 

was assumed to be 0.45 kg (1.0 lb) each. Engine thrust vector control was achieved using a 

pressurized helium pneumatic system. For DHCC, a total mass allocation of 1.0 kg was 

assumed, while GN&C and Data Processing system mass was charged to the ERV. A weight 

growth allowance of 30% was used for all dry mass elements. 
6

5.0 Initial Baseline Configuration Selection 

As a starting point, an initial design was selected as the baseline vehicle, including engine 

type, propellant selection, structural arrangement and outer mold line configuration. In 

anticipation of overall vehicle length constraints imposed by the landing capsule, a single stageto-orbit architecture was selected. Integration of multi-stage configuration into a single stack 

could result in increased overall vehicle length even with the benefit of reduced total ascent 

propellant mass fraction. 

5.1 Baseline Engine 

The significance of a high performance rocket engine is demonstrated in Figure 5.1.1, where 

the burn-out mass fraction (essentially the vehicle dry mass plus payload) is plotted versus 

engine specific impulse for two ascent total required delta-V, orbital at 4150 m/sec and escape 

C3=0 at 5500 m/sec. As engine specific impulse is increased, the mass fraction for dry mass 

and payload increase. The higher the required total delta-V, the higher the propellant mass 

fraction, hence lower payload plus dry mass fraction. For the C3=0 case in particular, specific 

impulse on the order of roughly 340 seconds allows for achievable mass fractions on the order of 

20%, hence reasonable payload mass fractions. 

Based on the requirement for relatively high engine specific impulse and the consideration of 

single-stage-to-orbit dictated by overall vehicle length constraints, a hypergolic, pressure-feed, 

gas-generator cycle engine was selected as the baseline concept. As a reference engine, the 

XLR-132 engine was selected, with an engine specific impulse of 347 seconds, combustor 

pressure of 10.34 MPa (1500 psi) and an engine thrust-to-weight ratio of approximately 33. 

During the vehicle sizing process, the engine was scaled up or down to match the required lift-off 

vehicle thrust-to-weight ratio. 

An initial engine trade study was conducted to assess the impact of the overall nozzle 

expansion ratio on the vehicle closure mass. As the expansion ratio is increased, the engine 

performance as measured by specific impulse increases, however the nozzle becomes larger 

and the engine weight increases. In addition for this particular application, the higher expansion 

ratio results in a longer engine length, and hence potentially a longer over vehicle length, 

perhaps being constrained by the overall vehicle length limits imposed by the landing capsule. 

Upper limits on nozzle expansion ratio were imposed to prevent an over expanded nozzle on the 

surface of Mars. Appendix A presents impact of nozzle expansion ratio on engine performance 

characteristics, dimensions and weight. 

Figure 5.1.2 presents the propellant mass fraction and overall body length as a function of 

engine expansion ratio for a delta-V requirement of 4150 m/sec and a payload mass of 200 kg. 

As the expansion ratio is increased, the engine specific impulse increases due to more optimal 

nozzle performance resulting in lower required propellant fraction, while overall closed vehicle 

length also increases due to increased nozzle length. The resultant closed vehicle gross liftoff 

mass and vehicle dry mass is presented in Figure 5.1.3 as a function of engine nozzle expansion 

ratio. As the nozzle expansion ratio is increased, the vehicle dry mass also increases due to 

heavier engine mass associated with a larger nozzle. However there is a minimum in the gross 

liftoff mass with expansion ratio, with higher propellant mass and lower dry mass for lower 

expansion ratios versus reduced propellant mass and increased dry mass at higher expansion 

ratios. The overall trend in gross liftoff mass is fairly flat, with roughly a 1.5 % variation over the 

range of nozzle expansion ratios analyzed. If the overall vehicle length becomes a constraint, 

only a modest penalty in gross liftoff mass will occur. 
7

Figure 5.1.1 Burn-out mass fraction versus engine specific impulse 

Figure 5.1.2 Propellant fraction and overall body length versus engine expansion ratio 

Burn-Out Mass Fraction vs. Rocket ISP

0

5

10

15

20

25

30

35

40

280 300 320 340 360 380 400 420

Engine ISP, seconds

Mass Fraction, %

Orbital

C3 = 0

Propellant Fraction & Body Length vs Expansion Ratio

Delta_V=4150 m/sec -- Payload Mass = 200 kg

0.704

0.706

0.708

0.71

0.712

0.714

0.716

0.718

0.72

100 150 200 250 300 400

Engine Expansion Ratio

Propellant Fraction

3.4

3.5

3.6

3.7

3.8

3.9

4

4.1

4.2

4.3

4.4

Propellant Fraction

Overall Body Length, m

Body Length
8

Figure 5.1.3 Gross liftoff mass and dry mass versus engine expansion ratio 

5.2 Candidate Body Geometries 

In order to define the design space for the general configuration of the MAV, three fore body 

shapes were selected, with the body fineness ratio (defined as body diameter divided by body 

length) varied parametrically from a range of approximately 0.2 to 0.44, and are presented in 

Figure 5.2.1. For each shape, a surface grid was generated, consisting of a triangulated mesh 

ranging for roughly 5,000 to 15,000 triangles. Body diameter, wetted area and internal volume 

as a function of axial coordinate were computed and used in the vehicle sizing code closure 

process. 

An initial trade study was conducted to assess the impact of body fineness ratio on closed 

vehicle mass and dimensions. For a delta-V requirement of 4150 m/sec and a payload mass of 

200 kg, Figure 5.2.2 presents sized vehicle overall length and surface wetted area as a function 

of body fineness ratio. As the fineness ratio is increased, both closed vehicle length and wetted 

area decrease. The variation in closed gross liftoff mass and vehicle dry mass are shown in 

Figure 5.2.3, with a general trend in lower gross and dry mass with increasing fineness ratio. 

Meeting the possible constraint on overall vehicle length can be addressed by going to a lower 

fineness ratio with only small impacts on gross liftoff and vehicle dry mass. 

GLOW & Dry Mass vs Expansion Ratio

Delta_V=4150 m/sec -- Payload Mass = 200 kg

1450

1460

1470

1480

1490

1500

1510

1520

1530

100 150 200 250 300 400

Engine Expansion Ratio

Gross Liftoff Mass, kg

135

140

145

150

155

160

165

170

175

Dry Mass, kg

GLOW

M_Dry
9

Figure 5.2.1 Parametric body shapes 

Figure 5.2.2 Overall vehicle length and total wetted area versus body fineness ratio 

Rev_0

D/L = 0.4375

Rev_2

D/L = 0.3125

Rev_1

D/L = 0.2065

Rev_0

D/L = 0.4375

Rev_2

D/L = 0.3125

Rev_1

D/L = 0.2065

Rev_0

D/L = 0.4375

Rev_2

D/L = 0.3125

Rev_1

D/L = 0.2065

Vehicle Length & Wetted Area vs. Body Fineness Ratio

Delta_V=4150m/s -- Payload Mass = 200kg

3.00

3.50

4.00

4.50

5.00

0.15 0.2 0.25 0.3 0.35 0.4 0.45

Body Fineness Ratio

Total Vehicle Length, m

5.00

6.00

7.00

8.00

9.00

10.00

Total Wetted Area, m2

Vehicle Length

Wetted Area
10

Figure 5.2.3 Overall vehicle length and total wetted area versus body fineness ratio 

5.3 Aerodynamics & Aerothermodynamics 

Using the surface grids for each of the three parametric shapes, aerodynamic and 

aerothermodynamic databases were generated for each shape. The aerodynamic database 

consisted of lift and drag coefficient as a function of Mach number, angle-of-attack and freestream dynamic pressure. Both engineering-based analysis (References 1 and 2) and Euler 

CFD codes (Reference 3) were used to compute the aerodynamic coefficients for each shape. 

The Euler code was used for subsonic, transonic and low supersonic Mach numbers, while the 

engineering code was run for super and hyper sonic flight conditions. For trade studies that 

involve forebody shape variations, the aerodynamic coefficients were linearly interpolated with 

body fineness ratio. 

Figure 5.3.1 presents drag coefficient as a function of angle-of-attack and Mach number. For 

Mach 2.0, both CBAero and Cart3D solutions are shown. The reference area is the crosssectional area of the body. There is good agreement between the two codes for Mach =2.0, with 

the drag coefficient higher for the engineering-based model due to viscous drag included. The 

aerodynamics predicted by CBAero are presented in Figures 5.3.2 through 5.3.4 across the 

supersonic/hypersonic Mach number range. The effect of the viscous drag at low dynamic 

pressure on the drag coefficient can be seen in Figure 5.3.2. The aerodynamic database was 

formatted and provided to the trajectory analysis model to compute lift and drag along the 

trajectory. As will be discussed below, the aerodynamic impacts on the ascent trajectory are 

small, due primarily to the low atmospheric density on Mars. Further details of the CFD solutions 

are presented in Appendix C. 

The aero-thermal database was generated using CBAero and consisted of surface pressure, 

surface shear, convective heat transfer coefficient and recovery enthalpy as a function of Mach 

number, angle-of-attack and free-stream dynamic pressure. Preliminary estimates of the aero 

GLOW & Dry Mass vs. Body Fineness Ratio

Delta_V=4150m/s -- Payload Mass = 200kg

1450

1460

1470

1480

1490

1500

0.15 0.2 0.25 0.3 0.35 0.4 0.45

Body Fineness Ratio

Gross Liftoff Mass, kg

150

152

154

156

158

160

Dry Mass, kg

Gross Liftoff Mass

Dry Mass
11

heating environment were conducted to assess the relative importance of the aero-thermal 

environment for ascent. The initial analysis indicated that the aero heating environment for the 

ascent trajectory was rather benign and there would be minimal thermal protection system 

requirements for the MAV, due again to the low atmospheric density. Sample aero heating 

results are presented in Appendix C. 

Figure 5.3.1 Drag coefficient versus angle-of-attack and Mach number 

0.00

0.25

0.50

0.75

1.00

1.25

1.50

0 5 10 15 20

CD

Angle of Attack (deg)

MSR MAV rev2

Cart3D Adjoint

Preliminary Results

Mach 0.5 Mach 0.95

Mach 1.05 Mach 1.5

Mach 2.0 CBAero Mach 2.0
12

Figure 5.3.2 CBAero predicted zero-lift drag coefficient versus Mach number and 

free stream dynamic pressure 

Figure 5.3.3 CBAero predicted drag coefficient versus angle-of-attack and Mach 

number for dynamic pressure = 0.005 bars 

Zero-Lift Aero Drag vs. Mach Number & Dynamic Pressure

Reference Area = 1.227 m2 

0.00

0.10

0.20

0.30

0.40

0.50

0.60

0.70

0.80

0.90

0.0 5.0 10.0 15.0 20.0

Mach Number

Drag Coefficient

qbar = 0.00005

qbar = 0.0005

qbar = 0.005

qbar = 0.01

qbar = 0.015

Drag Coeff vs. Mach & Alpha

Reference Area = 1.227 m2 / qbar = 0.005 bars

0.00

0.10

0.20

0.30

0.40

0.50

0.60

0.70

0.80

0.90

1.00

0.0 5.0 10.0 15.0 20.0 25.0

Angle-of-Attack, o

Drag Coefficient

Mach 1.2

Mach 1.5

Mach 2.0

Mach 3.0

mach 4.0

Mach 6.0

Mach 10.0

Mach 18.0
13

Figure 5.3.4 CBAero predicted lift coefficient versus angle-of-attack and Mach 

number for dynamic pressure = 0.005 bars 

5.4 General Arrangement and Tank Configuration 

With potential overall MAV body length constrained by the entry capsule internal volume, a 

nested tank configuration was selected for the internal arrangement. Similar to a common 

bulkhead design, the forward tank aft dome is accommodated within a recessed aft tank forward 

dome but with a purge gap between the two tanks. The inter-tank section is eliminated and the 

overall vehicle length is reduced. The mated tanks are structural members carrying pressure, 

axial and bending loads. Thrust loads from the rocket engine are transmitted through a thrust 

structure and introduced to a load bearing aft compartment, which in turn transmits the thrust 

loads into the aft tank. The general arrangement is presented in Figure 5.4.1. A forward skirt 

transitions from the domed tank section into the fore body, which is the fairing for the ERV 

payload, including a payload adapter. The nested tank internal configuration is shown in Figure 

5.4.2. The oxidizer tank is forward, resulting in forward axial center-of-gravity, reducing engine 

gimbal requirement for trim. 

High temperature polymer matrix composite thermo-plastic (PEEK) was chosen for the 

material for the aero-shell, tank (with a liner), aft compartment and thrust structure. The aft 

compartment and tanks were sized based on loads using mass estimating relationships (see 

below) and the forward aero-shell assumed to be minimum gauge skin-stringer stiffened 

structure, consisting of 5 skin plies + stiffeners, with an areal mass of 1.04 kg/m2

. 

Lift Coeff vs. Mach & Alpha

Reference Area = 1.227 m2 / qbar = 0.005 bars

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

0.0 5.0 10.0 15.0 20.0 25.0

Angle-of-Attack, o

Lift Coefficient

Mach 1.2

Mach 1.5

Mach 2.0

Mach 3.0

mach 4.0

Mach 6.0

Mach 10.0

Mach 18.0
14

Figure 5.4.1 MAV general arrangement 

Figure 5.4.2 MAV internal arrangement 

Nosecap

Forebody

Forward

Skirt Nested Tanks

Aft

Compartment

Thrust Structure

Engine

Nosecap

Forebody

Forward

Skirt Nested Tanks

Aft

Compartment

Thrust Structure

Engine
15

5.5 Mass Estimating Relationships 

The mass of the vehicle subsystems was computed using mass estimating relationships 

(MER’s), taken from various sources (References 4, 5, 6, 5 and 8). Generally, the MER’s are 

based on historical data correlated to pertinent design parameters for existing launch vehicles. 

The scale of the MAV concept was found to be typically at the low end of the correlated data, or 

completely outside the data range, resulting in some extrapolation for certain subsystem 

elements. An example (taken from Reference 5) is presented in Figure 5.5.1 for the unit areal 

mass of the inter-tank and aft compartment, based on unit axial load, body diameter and material 

property. The range of the axial loads for the MAV configuration is indicated in the lower left 

hand corner. Knowing the inertial load at the maximum 5.0 Earth’s axial acceleration and apply 

a factor-of-safety = 1.4, the areal weight of the aft compartment is computed from Figure 5.5.1 

and multiplied by the wetted area of the aft compartment to obtain its mass. 

In addition to structural and tank mass, MER’s were assembled from the above references 

for other subsystems, including induced environments, auxiliary systems (separation systems), 

main propulsion (main engines, feed/pressurization and controls), prime power systems 

(batteries and pneumatic systems for engine control), power conversion and distribution and data 

handling/communication/control (DHCC). Figure 5.5.2 presents the mass scaling model for the 

main propulsion engine, with engine thrust-to-weight ratio value plotted as a function of engine 

thrust. Generally the engine thrust-to-weight goes down in decreasing thrust level, reflecting the 

impact of minimum gauge on smaller engines. The engines shown in Figure 5.5.2 are generally 

pressure-feed OMS/RCS engines. The reference XLF-132 pump feed engine is also plotted. 

The mass of the scaled pump-feed engine was then computed using the required vacuum thrust 

and the shifted trend line of engine thrust-to-weight versus engine thrust (shifted to pass through 

the XLR-132 data point, but having the same slope). 

Figure 5.5.1 Inter-tank and aft compartment structure mass estimate model (taken from 

Reference 5) 

MAV Load Range
16

Figure 5.5.2 Rocket engine weight-to-thrust ratio versus engine thrust 

6.0 Ascent Trajectory 

The ascent trajectory for the MAV was simulated using POSTII (Reference 9). Aerodynamic 

tables of lift and drag coefficient were provided as a function of Mach number, dynamic pressure 

and angle-of-attack (Section 5.3). Rocket engine performance model consisted of engine 

vacuum specific impulse and nozzle exit area. The 3 DOF trajectory was optimized for maximum 

burn-out mass, with control parameters of angle-of-attack, bank angle and liftoff vehicle thrust-toweight. The Mars GRAM atmospheric model 2001 was used. The launch point was 0m MOLA 

altitude, 0o Latitude and 0o Longitude, heading due East. Constraints on maximum dynamic 

pressure, maximum q-alpha and final altitude were imposed. Three orbital ascent trajectories 

were computed: 100 X 100 km, 100 X 500 km, and 500 X 500 km and C3=0 at 500 km, the total 

ΔV calculated for each.

Figures 6.1 through 6.5 present the optimized ascent trajectory history for the 100 X 500 km 

orbital case. Figure 6.1 and 6.2 present altitude as a function of Mach number and time, 

respectively. Ascent time is approximately 265 seconds. The dynamic pressure time history is 

shown in Figure 6.3, with a maximum value of 250 Pa, which occurs at approximately Mach = 

1.6. The total acceleration as a function of time is presented in Figure 6.4, with a burn out 

maximum 5.3 Mars g’s. Figure 6.5 presents the optimal control for the MAV ascent trajectory, 

with angle-of-attack and bank angle shown as a function of time. The vehicle is banked to a 

“heads-down” attitude, initiated at roughly 30 seconds after liftoff.

Table 1 summarizes the total ΔV required for the three orbital and two escape cases. Also 

noted is the final-to-initial mass ratio required for each, assuming an engine specific impulse of 

343 seconds. Finally, the relative velocity loss breakdown for the 100 X 500 km orbital and the 

escape trajectories are presented in Figures 6.6 and 6.7. For both cases, the drag, thrust 

vectoring and atmospheric losses are all small. 

Engine thrust-to-Weight vs Thrust Trend

Hypergolic Engines 

0.0

5.0

10.0

15.0

20.0

25.0

30.0

35.0

40.0

1 10 100 1000 10000

Engine Thrust, lbf

Engine Thrust-to-Weight

STS OMS

RS-25

RS-44

RS-21

RS-24

RS-28

RS-41

Lunar Asent

Trend Line

XLR-132
17

Table 1 Ascent Velocities and mass fractions 

Figure 6.1 MAV ascent trajectory: Altitude versus time 

Orbit (km) Velocity (m/s) V (m/s) m/mo

(100,100) 3503 4027.9 0.3019

(100, 500) (3597,3227) 4115.5 0.2942

(500, 500) 3318.1 4206.7 0.2862

C3-0 @100km 4951.2 5476.1 0.1962

C3=0 

@500km

4690 5578 0.1904

Orbit (km) Velocity (m/s) V (m/s) m/mo

(100,100) 3503 4027.9 0.3019

(100, 500) (3597,3227) 4115.5 0.2942

(500, 500) 3318.1 4206.7 0.2862

C3-0 @100km 4951.2 5476.1 0.1962

C3=0 

@500km

4690 5578 0.1904

MAV Ascent Trajectory

100x500 km Orbit

0

20000

40000

60000

80000

100000

0 50 100 150 200 250 300 350

Time, secs

Altitude, m
18

Figure 6.2 MAV ascent trajectory: Altitude versus relative velocity 

Figure 6.3 MAV ascent trajectory: Dynamic pressure versus time 

MAV Ascent Trajectory

100x500 km Orbit

0

20000

40000

60000

80000

100000

0 500 1000 1500 2000 2500 3000 3500 4000

Rel. Velocity, m/s

Altitude, m

MAV Ascent Trajectory

100x500 km Orbit

0

50

100

150

200

250

300

0 50 100 150 200 250 300

Time, secs

Dyn. Pressure, Pa
19

Figure 6.4 MAV ascent trajectory: Acceleration versus time 

Figure 6.5 MAV ascent trajectory: Dynamic pressure versus relative velocity 

MAV Ascent Trajectory

100x500 km Orbit

0.0

0.5

1.0

1.5

2.0

2.5

3.0

0 50 100 150 200 250 300

Time, secs

Acceleration, Earth Gs

MAV Ascent Trajectory

100x500 km Orbit

0

40

80

120

160

200

0 50 100 150 200 250 300

Time, secs

Bank, deg

0

4

8

12

16

20

Bank

Alpha, deg

AoA
20

Figure 6.6 Orbital relative velocity loss breakdown 

Figure 6.7 Escape relative velocity loss breakdown 

Relative Velocity Losses

Drag Loss

Gravity Loss

Thrust Loss

Aero Loss

Rel.V 3353 m/s

615 m/s

113 m/s

Total Loss = 762.7 m/s

V = 4115.5 m/s

Relative Velocity Losses

Drag Loss

Gravity Loss

Thrust Loss

Aero Loss

Rel.V 3353 m/s

615 m/s

113 m/s

Total Loss = 762.7 m/s

V = 4115.5 m/s

C3=0: Relative Velocity Losses

Drag Loss

Gravity Loss

Thrust Loss

Aero Loss

Rel.V

V=5578 m/s

4463 m/s

1023

113

Total Loss=1168 m/s

C3=0: Relative Velocity Losses

Drag Loss

Gravity Loss

Thrust Loss

Aero Loss

Rel.V

V=5578 m/s

4463 m/s

1023

113

Total Loss=1168 m/s

Atmos Loss 

Atmos Loss 
21

7.0 Baseline Configurations 

Two MAV architectures were initially examined, one to launch a prescribed payload to orbital 

flight conditions for Mars orbital rendezvous (inert payload) or phasing orbit for Earth return by an 

ERV, and a second to escape velocity conditions for Earth orbital rendezvous. The payload 

mass was initially treated parametrically to capture the impact of the payload mass and volume 

on the MAV design requirements. These baseline configurations were then used as the 

reference point for the trade studies. 

7.1 Vehicle Closure 

Initial vehicle sizing was conducted using the assumptions stated above in Section 4.0 for the 

baseline configuration and engine, as described in Section 5.1 and 5.2. Total ΔV’s for orbital and 

escape missions presented in Section 6.0 where used for determining propellant fraction 

required. A vehicle synthesis code was developed using the subsystem MER’s and a first cut 

geometry/packing model to compute propellant fraction available and required volume for all 

subsystems. For both the orbital and escape architectures, a range of payload masses were 

examined and the vehicle closed by matching propellant fraction required to propellant fraction 

available using the vehicle synthesis model. 

7.2 Orbital Architecture 

The Mar’s orbital architecture MAV places either an inert capsule (assumed mass of 20 kg for 

orbital rendezvous) or a chemical powered ERV (mass ranging from say 80 to 200 kg) into low 

Mars orbit, with the required total ΔV = 4150 m/sec. For either payload, an assumed payload 

density of 400 kg/m3 was used to compute the payload volume. The MAV was then sized to 

provide sufficient payload, propellant tankage and subsystem mass and volume to perform the 

mission. 

Figure 7.2.1 presents the vehicle closure plot, showing propellant fraction required and 

propellant fraction available as a function of gross liftoff mass for a fixed payload requirement. 

The propellant fraction required is a function only of the required total ΔV and the engine specific 

impulse. Due to the low density of the Martian atmosphere, hence low aerodynamic drag on 

ascent, the engine specific impulse and the vehicle specific impulse are almost the same and as 

a result the propellant fraction required becomes independent of the vehicle mass (size). The 

intersection of the propellant fraction required and propellant fraction available represents a 

closed vehicle design. As the payload requirement is increased the closure GLOM also 

increases, but so does the payload mass fraction, ranging from roughly 6% at 20kg payload 

mass to about 13% at 200 kg. 

An important characteristic of the vehicle design is the slope of the propellant fraction 

available at the intersection point. As mission requirements or engine performance 

characteristics change, the propellant fraction required may shift up or down. Subsystem 

technology or performance may also vary, resulting in a shift of the propellant fraction available 

curve. If the slope of the propellant fraction available curve at the intersection point is shallow, 

small changes in either the propellant fraction available or propellant fraction required will result 

in relatively large shifts in the intersection point, hence relatively large changes in the closure 

gross mass. The steeper the propellant fraction available curve at the intersection point, the less 

sensitive the vehicle closure mass to shifts in either propellant fraction curve. As seen in Figure 

7.2.1, the slope of the propellant fraction available curve tends to decrease with increasing 

GLOM. Hence, at higher payload mass requirements, both the required vehicle launch mass 

increases and the design becomes more sensitive to mission requirements or technology 

performance. 
22

Another measure of design robustness is often presented in terms of a “growth factor”: the 

amount the vehicle will grow for every additional kilogram of mass required to be carried to orbit. 

Table 2 summarizes the salient characteristics, including some subsystems masses, dry mass 

and gross liftoff mass, dimensions and growth factor of the orbital architecture for the various 

payload requirements. For the orbital architecture concepts, every kilogram added is roughly 6 

more kilograms of gross liftoff mass. For the baseline 200 kg payload design, Figure 7.2.2 

presents overall vehicle dimensions (not to scale). 

Figure 7.2.1 Orbital architecture closure point versus GLOM and payload mass 

Orbital Architecture: Delta_V = 4150 m/sec

Payload Density = 400 kg/m3

0.6

0.65

0.7

0.75

0.8

200 700 1200 1700 2200

Gross Liftoff Mass, kg

Propellant Fraction

Prop Fraction Req

PF_Avl @ 20 kg

PF_Avl @ 80kg

PF_Avl @ 120kg

PF_Avl@200 kg
23

Table 2. Orbital architecture characteristics 

Figure 7.2.2 Orbital Architecture, 200 kg payload MAV dimensions 

7.3 C3=0 Architecture 

The Mars’ escape architecture MAV accelerates an ERV (mass ranging from say 20 to 200 

kg) to C3 = 0, with the required total ΔV = 5500 m/sec. For this mission, it was assumed that the 

ERV would be solar-electric powered (SEP), rather than chemical powered used for the orbital 

architecture concept. An assumed payload density of 240 kg/m3 was used to compute the 

payload volume, reflecting the less compact SEP configuration. 

Dry Mass, kg 54.0 90.6 112.5 154.1

GLOW, kg 312.5 717.4 976.7 1484.

Length, m 2.19 3.02 3.38 3.95

Diameter, m 0.52 0.71 0.79 0.91

6.5

11.5

39.4

31.8

28.6

15.5

120 kg

6.8

17.3

24.1

19.4

9.4

5.7

20 kg

Dry Mass fraction, % 12.6 10.4

Growth Factor 6.7 6.3

Main Propulsion, kg 33.9 49.3

Main Engine, kg 27.4 39.9

Total Structure, kg 21.3 42.5

Tanks, kg 11.8 22.4

Payload Mass 80 kg 200 kg

Dry Mass, kg 54.0 90.6 112.5 154.1

GLOW, kg 312.5 717.4 976.7 1484.

Length, m 2.19 3.02 3.38 3.95

Diameter, m 0.52 0.71 0.79 0.91

6.5

11.5

39.4

31.8

28.6

15.5

120 kg

6.8

17.3

24.1

19.4

9.4

5.7

20 kg

Dry Mass fraction, % 12.6 10.4

Growth Factor 6.7 6.3

Main Propulsion, kg 33.9 49.3

Main Engine, kg 27.4 39.9

Total Structure, kg 21.3 42.5

Tanks, kg 11.8 22.4

Payload Mass 80 kg 200 kg

3.95 m

2.908 m

0.909 m 0.421 m

0.763 m
24

Figure 7.3.1 presents the vehicle closure plot, showing propellant fraction required and 

propellant fraction available as a function of gross liftoff mass for a fixed payload requirement. 

Again, the intersection of the propellant fraction required and propellant fraction available 

represents a closed vehicle design. As the payload requirement is increased, the closure GLOM 

also increases, but so does the payload mass fraction, ranging from roughly 2.6% at 20kg 

payload mass to about 6.2% at 200 kg. Table 3 summarizes the salient characteristics, including 

some subsystems masses, dry mass and gross liftoff mass, dimensions and growth factor of the 

orbital architecture for the various payload requirements. For the escape architecture concepts, 

the growth factor in on the order of 13 kg added for every additional 1 kg, roughly twice that of 

the orbital architecture concepts. For the baseline 120 kg payload design, Figure 7.3.2 presents 

overall vehicle dimensions. 

Figure 7.3.1 C3=0 architecture closure mass versus GLOM and payload mass 

C3=0 Architecture: Delta_V = 5500 m/sec

Payload Denisty = 240 kg/m3

0.700

0.725

0.750

0.775

0.800

0.825

0.850

200 700 1200 1700 2200 2700 3200 3700

Gross Liftoff Mass, kg

Propellant Fraction

Prop Fraction Req

PF_Avl @ 20 kg

PF_Avl @ 80 kg

PF_Avl @ 120kg

PF_Avl @ 200 kg
25

Table 3. Escape architecture characteristics 

Figure 7.3.2 C3=0, 120 kg payload MAV dimensions 

8.0 Trade Studies 

With the baseline MAV designs established, trade studies were conducted to look at alternate 

design approaches and technologies. The goal was to explore the design space and identify the 

most promising configurations and candidate technologies and determine design sensitivities. 

The trade studies were typically conducted for one payload class, rather than over the range of 

assumed payload masses. 

14.4

0.70

2.99

758

11.0

83.7

34.9

28.1

19.6

13.8

20 kg

Dry Mass, kg 144.1 181.6 251.4

GLOW, kg 1628 2187 3247

Length, m 4.03 4.52 5.26

Diameter, m 0.92 1.03 1.19

13.5

8.3

62.4

50.4

54.1

35.1

120 kg

Dry Mass fraction, % 8.8 7.7

Growth Factor 13.8 13.2

Main Propulsion, kg 52.2 80.9

Main Engine, kg 42.1 65.4

Total Structure, kg 40.7 79.1

Tanks, kg 27.1 49.7

Payload Mass 80 kg 200 kg

14.4

0.70

2.99

758

11.0

83.7

34.9

28.1

19.6

13.8

20 kg

Dry Mass, kg 144.1 181.6 251.4

GLOW, kg 1628 2187 3247

Length, m 4.03 4.52 5.26

Diameter, m 0.92 1.03 1.19

13.5

8.3

62.4

50.4

54.1

35.1

120 kg

Dry Mass fraction, % 8.8 7.7

Growth Factor 13.8 13.2

Main Propulsion, kg 52.2 80.9

Main Engine, kg 42.1 65.4

Total Structure, kg 40.7 79.1

Tanks, kg 27.1 49.7

Payload Mass 80 kg 200 kg

4.52 m

3.29 m

1.03 m 0.623 m

1.13 m
26

8.1 Number of Engines 

The baseline configuration consisted of a single gimbaled engine with pneumatic thrust 

vectoring control. An alternate configuration is to use multiple expansion nozzles with a common 

power head, and variable engine thrust to achieve pitch and yaw control. Roll control would 

continue to use gas-generator exhaust flow. A 15% penalty in power head mass was assumed 

for multiple engine feeds, and the weight of the thrust structure increases with the number of 

engines. The trade study was conducted for the orbital architecture. 

Table 4 presents a comparison between the baseline single engine and a 3 and 4 engine 

nozzle configuration for a fixed liftoff mass of 1500 kg. The total thrust for all three configurations 

is the same. The multiple engine design approach results in overall engine length being 

significantly shorter, but still allows engine integration into the base region of the vehicle without 

requiring any aft skirt fairing. Thrust structure and aft compartment mass increase with number 

of engines, while engine heat shield and base close-out TPS decrease due to reduced wetted 

areas. Engine total mass and propellant feed system mass show small variations. The biggest 

mass savings for the multiple engine configuration is in the elimination of the engine gimbal 

control and associated pneumatic prime power elements. For the fixed launch mass, the 3 and 4 

engine configurations save roughly 4% in dry mass, with the 3 engine configuration having a 

slight edge in propellant fraction available. From a closed mass perspective, the multiple engine 

configuration saves approximately 4% in dry mass and roughly 2% in liftoff mass. The last two 

rows of Table 4 present the closed vehicle dry mass and gross liftoff mass. 

Table 4 Number of Engine comparison 
27

8.2 Pressure-Feed Propulsion System 

The baseline engine concept is a gas-generator cycle with turbo-pump propellant feed. This 

design approach results in significant tank mass savings due to the lower required tank pressure. 

For the small payload mass (20kg inert capsule), the sized MAV vehicle is in 200 to 300 kg class 

range, resulting in engine thrust levels on the order of ~2000 N. This represents a scaling of the 

engine down to approximately 15% of baseline XLR-132 engine thrust level. Scaling a turbopump feed engine by a factor of 7 may incur significant performance penalties, which were not 

modeled in the current study. As a result, a pressure-feed propulsion system was studied and 

two vehicle classes were evaluated: 1) 20kg payload orbital vehicle, and 2) 200 kg payload 

orbital vehicle. For each design the effect of chamber pressure and nozzle expansion ratio were 

assessed. Both cases utilize hypergolic propellants. 

Appendix A presents the impact of chamber pressure and nozzle expansion ratio on engine 

specific impulse, thrust per unit exit area and engine mass. For the 20 kg payload mass, Figure 

8.2.1 presents the propellant mass fraction required as a function of the chamber pressure. 

There is a modest increase in the engine specific impulse with higher chamber pressure, hence 

a modest reduction of the propellant fraction required. For the closed vehicle, the overall vehicle 

length decreases with increasing chamber pressure ratio (see Figure 8.2.2), reflected directly to 

the increase in thrust per unit area at higher chamber pressure, hence a smaller, shorter engine. 

The trend flattens out at higher chamber pressures as a result of body length increases at the 

higher pressures. Figure 8.2.3 shows tank, pressurization system and engine mass as a 

function of thrust chamber pressure. As the pressure increases, the tank mass and 

pressurization system mass increase, while the engine mass decreases due to a physically 

smaller engine at higher chamber pressure, even though the unit area mass of the engine nozzle 

increases with chamber pressure. Finally, Figure 8.2.4 presents closed dry mass and gross 

liftoff mass as function of chamber pressure. The offsetting effects of tank and pressurization 

system mass versus engine mass results in a minimum of both dry and gross mass, occurring 

near a chamber pressure around 1.03MPa (150 psi). 

Table 5 summarizes the comparison of the pump-feed versus the pressure feed engine 

system. The tank mass and pressurization system are significantly higher for the pressure feed 

system compared to the pump feed design, with the main engine mass ~15% lighter. The 

resulting dry mass for the pressure feed system is 21% higher, with the propellant fraction only 

slightly higher. The closed gross liftoff mass is on the order of 26% heavier and 20% longer. For 

the 20 kg payload class, both concepts are comfortably below the maximum landed payload 

mass limit, and the pressure feed concept would be viable if the engine scaling issues 

associated with the pump feed system become limiting. 

The pressure feed trade study was repeated for a 200 kg payload mass requirement. At this 

scale, the pressure feed system is almost twice the gross liftoff mass as the equivalent pump 

feed system. As the propellant tanks become large, the tank mass increases rapidly and the dry 

mass and gross mass grow significantly. Figure 8.2.6 shows closed dry mass and gross liftoff 

mass as a function of chamber pressure. Dry mass is fairly insensitive to chamber (i.e. 

propellant tank pressure), again trading tank and pressurization system mass versus engine 

mass, and a minimum in gross mass occurs somewhat below 1.38 MPa (200 psi). Figure 8.2.7 

presents overall vehicle length as a function of engine chamber pressure, with significantly 

longer vehicles at the lower engine pressure. Table 6 summarizes the comparison of the two 

concepts for the 200 kg payload mass. The pressure feed systems becomes noncompetitive 

with the pump feed system and violates the maximum allowable landed mass and vehicle overall 

length by roughly 30% to 40%. 
28

Figure 8.2.1 Propellant fraction required versus chamber pressure 

Figure 8.2.2 Overall vehicle length versus chamber pressure 

Orbital Architeture: Payload = 20kg

Propellant Mass Fraction

0.72

0.725

0.73

0.735

0.74

0 100 200 300 400

Chamber Pressure, psi

Propellant Mass Fraction

Orbital Architeture: Payload = 20kg

Total MAV Length

2.00

2.50

3.00

3.50

4.00

0 100 200 300 400

Chamber Pressure, psi

Vehicle length, m
29

Figure 8.2.3 Tank, pressurization and engine mass versus chamber pressure 

Figure 8.2.4 GLOW and dry mass versus chamber pressure 

Orbital Architeture: Payload = 20kg

Tank, Press & Engine Mass

0

5

10

15

20

25

30

35

0 100 200 300 400

Chamber Pressure, psi

Tank, Press & Engine Mass, 

kg

Propellant Tanks

Pressurization

Main Engine

Orbital Architeture: Payload = 20kg

Dry & Gross Liftoff Mass

0

100

200

300

400

500

0 100 200 300 400

Chamber Pressure, psi

Dry Mass & GLOW, kg

Dry mass

GLOW
30

Table 5. 20kg payload pump feed versus pressure feed system comparison 

Figure 8.2.5 GLOW and dry mass versus chamber pressure 

Pump Feed Pressure Feed

Tank Pressure 0.034 MPa (5.0 psi) 1.30 MPa (188 psi)

Tanks 5.70 kg (13.7% Dry) 16.04 kg (31.8% Dry)

Structure 9.23 kg (22.2% Dry) 21.26 kg (42.2% Dry)

Main Engine 19.2 kg (46.8% Dry) 16.27 kg (32.3% Dry)

Pressurization 0.92 kg (2.2% Dry) 3.91 kg (7.8% Dry)

Main Propulsion 24.1 kg (46.8% Dry) 20.38 (40.4% Dry)

Dry Mass 54.0 kg (17.3% GLOW) 65.5 kg (16.6% GLOW)

Propellant Fraction 71.3% 73.1%

GLOW 312.5 kg 394.1 kg

Length 2.19 m 2.60 m

Diameter 0.520 m 0.566 m

Delta_VT = 4150 m/s: Payload = 200kg

Dry & Gross Liftoff Mass

0

500

1000

1500

2000

2500

3000

3500

0 100 200 300 400

Chamber Pressure, psi

Dry Mass & GLOW, kg

Dry mass

GLOW
31

Figure 8.2.6 Vehicle overall versus chamber pressure 

Table 6. 200 kg Payload mass pump feed versus pressure feed system comparison 

Delta_VT = 4150 m/s: Payload = 200kg

Total MAV Length

5.00

5.50

6.00

6.50

7.00

7.50

0 100 200 300 400

Chamber Pressure, psi

Vehicle length, m

Pump Feed Pressure Feed

Tank Pressure 0.034 MPa (5.0 psi) 1.72 MPa (250 psi)

Tanks 22.37 kg (18.8% Dry) 96.71 kg (33.2% Dry)

Structure 42.54 kg (35.9% Dry) 125.2 kg (43.0% Dry)

Main Engine 39.91 kg (33.7% Dry) 108.5 kg (37.3% Dry)

Pressurization 2.84 kg (2.4% Dry) 23.36 kg (8.0% Dry)

Main Propulsion 49.31 kg (41.6% Dry) 133.2 (45.8% Dry)

Dry Mass 154.1 kg (10.4% GLOW) 378.1 kg (14.4% GLOW)

Propellant Fraction 71.3% 73.1%

GLOW 1484 kg 2630 kg

Length 3.95 m 5.58 m

Diameter 0.909 m 1.044 m
32

8.3 Cryogenic Propellant Trade 

The baseline propellant system is space storable hypergolic propellant with a turbo-pump 

feed engine. A trade study was conducted to examine the impact of going to storable/cryogenic 

or all cryogenic propellants. Appendix B presents the engine performance comparison various 

storable and cryogenic propellants (Figure B2). Table 7 compares the physical properties for 

various hydrocarbon and hypergolic fuel/oxidizer mixtures. From an engine performance 

perspective, the methane/liquid oxygen combination provides the best engine performance and 

was selected for the trade study. RP-1/LOX has a lower engine specific impulse (about 4%), but 

has a bulk propellant density roughly 25% higher than the methane/liquid oxygen combination, 

and was also selected for study. 

Using the XLR-132 as the baseline engine and the 1-D rocket performance code, the 

hypergolic engine has an engine specific impulse of 343 seconds, compared to approximately 

370 seconds for the methane/liquid oxygen combination at the baseline engine chamber 

pressure and nozzle expansion ratio. The higher engine specific impulse will result in a lower 

fuel fraction required, however the lower aggregate propellant density will result in a larger 

vehicle to contain the propellant. The choice of the cryogenic fuels will also introduce design 

issues associated with the low temperature propellants. Transient heat transfer analysis 

indicated that 1.0 inch of a low-density rigid cryogenic insulation would prevent the external wall 

temperature from reaching the condensation temperature of the atmospheric carbon dioxide for 

approximately 15 minutes. 

The RP-1/LOX engine provides 354 seconds of engine specific impulse based on the 1-D 

rocket performance code, compared again to 343 seconds for the storable hypergolic 

propellants, and has a similar bulk propellant density to the hypergolic propellants. 

Table 8 presents a comparison between the hypergolic and selected cryogenic propellants 

for a 200 kg payload mass and a total ΔV of 4150 m/sec. Comparisons are also made for a

single engine and a 3 engine configuration. Engine nozzle expansion ratio was optimized for 

minimum gross liftoff mass. Due to lower propellant density for the methane/liquid oxygen 

design, the fore-body fineness ratio was increased to reduce overall vehicle length, but does not 

violate the maximum allowed body diameter limit of 1.2 m. For the RP-1/LOX design, the fore 

body diameter-to-length ratio was optimized for minimum gross liftoff mass. Table 9 presents a 

more detailed mass comparison for the various propellant combinations. 

The propellant fraction is lower for the methane/liquid oxygen propellant combination, due to 

the higher engine specific impulse noted above. For the closed designs, the overall length of the 

cryogenic vehicle is similar to than that of the higher density hypergolic propellant, due mainly to 

closing at a lower gross liftoff mass and higher diameter-to-length ratio. The dry mass of the 

methane/liquid oxygen vehicle is higher compared to the storable propellant design, due to 

higher surface area and required cryogenic tank insulation. The closed gross liftoff mass is 

slightly lower for the methane/liquid oxygen propellant concept, by roughly only 4%. The RP1/LOX vehicle exhibits higher dry mass, but lower gross mass compared to the hypergolic 

propellant design. 

The propellant trade study was also repeated for the C3=0 architecture. Table 10 presents a 

comparison between the hypergolic and selected cryogenic propellants for a 120 kg payload 

mass and a total ΔV of 5500 m/sec. Similar trends are found as for the lower total ΔV mission, 

with the methane/liquid oxygen concept having a higher dry mass of nearly 23%, but a slightly 

lower gross liftoff mass, on the order of roughly 5%. The RP-1/LOX vehicle shows 10% higher 

dry mass and 3% lower gross mass. 
33

The gross and dry mass of the MAV vehicle is driven largely by the performance and weight 

characteristics of the rocket engine. By far, the performance parameter that has the biggest 

impact on the vehicle design is the engine specific impulse, regardless of propellant choice. 

Figures 8.3.1 through 8.3.3 present vehicle gross mass, dry mass and overall length sensitivity to 

engine specific impulse for the orbital architecture. For all studied propellant concepts, a1% 

change in specific impulse results in a 2% change in gross liftoff mass and 1% change in dry 

mass and vehicle length. The methane/oxygen vehicle shows a slightly higher sensitivity of the 

overall vehicle length to engine performance. Figures 8.3.4 through 8.3.6 present the gross 

mass, dry mass and overall vehicle length sensitivity as a function of the relative rocket engine 

mass. Gross liftoff mass and vehicle length have similar sensitivities for all propellant 

combinations. Vehicle dry mass shows the highest sensitivity to engine mass, but less than “onefor-one” in growth sensitivity. The storable hypergolic propellant design exhibits the largest dry 

mass sensitivity to engine mass variations. 

The cryogenic propellant designs offer the possibility of modest reduction in gross liftoff 

mass, with increase in dry mass on the order of 10% for the RP-1/LOX configuration and 20% for 

the CH4/LOX design for total ΔV mission requirements of 4150 and 5500 m/sec. The use of insitu resource utilization (ISRU) offers the potential of significant reduction on landed MAV mass, 

with one or both of the propellants being produced and stored on Mars. For the RP-1/LOX 

configuration, MAV landed mass can be reduced 50% for the orbital architecture, and to 60% for 

the C3=0 architecture. The CH4/LOX concept would provide even higher MAV landed mass 

(70% to 80%) savings, since both propellants could be produced on the surface of Mars. The 

total landed weight savings must of course account for the mass, power and volume required by 

the additional subsystems to produce and store cryogenic propellants on the Martian surface. 

Table 7. Propellant properties 

60

46

32

46

~175

114

72

58

44

16

Mol Wt

UDMH ‘ 790 2.613.06 1175 337 .137 3472

MMH ‘ 875 2.56/1.96 1185 364 .050 3387

N NTO 1021 1.44/1.25 1218 387 .010 3263 2H4

C “ 789 2.08/1.70 979 352 .0595 3352 2H5OH

RP-1 “ 807 3.4/2.25 1012 422 .020 3497

C “ 703 3.50/3.0 987 399 .0147 3613 8H18

C “ 621 3.55/3.0 943 309 .579 3601 5H12

C “ 601 3.57/3.0 932 272 1.0 3592 4H10

C “ 582 3.6/3.0 920 231 8.5 3577 3H8

CH LOX 423 4.0/3.0 800.7 111.4 58 3419 4

Comb

Temp

(oK)

V.P @ 

Room 

Temp 

(bars)

B.P

(oK)

Prop

Density

@ O/FOpt

O/FS & 

O/FOpt

Density

(kg/m3)

Fuel Oxidizer

60

46

32

46

~175

114

72

58

44

16

Mol Wt

UDMH ‘ 790 2.613.06 1175 337 .137 3472

MMH ‘ 875 2.56/1.96 1185 364 .050 3387

N NTO 1021 1.44/1.25 1218 387 .010 3263 2H4

C “ 789 2.08/1.70 979 352 .0595 3352 2H5OH

RP-1 “ 807 3.4/2.25 1012 422 .020 3497

C “ 703 3.50/3.0 987 399 .0147 3613 8H18

C “ 621 3.55/3.0 943 309 .579 3601 5H12

C “ 601 3.57/3.0 932 272 1.0 3592 4H10

C “ 582 3.6/3.0 920 231 8.5 3577 3H8

CH LOX 423 4.0/3.0 800.7 111.4 58 3419 4

Comb

Temp

(oK)

V.P @ 

Room 

Temp 

(bars)

B.P

(oK)

Prop

Density

@ O/FOpt

O/FS & 

O/FOpt

Density

(kg/m3)

Fuel Oxidizer
34

Table 8. Orbital/200kg payload hypergolic and cryogenic performance comparison 

Table 9. Orbital/200kg comparison between hypergolic and cryogenic propellants 

1417

11.25

159.4

1.004

3.52

0.699

0.3516

200

3

1448

11.44

165.6

0.996

3.86

0.700

0.3516

175

1

RP-1/LOX

Dry Mass 154.1 145.8 182.5 176.0

Opt D/LFB 0.3125 0.3125 0.3906 0.3906

Opt ε 200 200 200 225

GLOW 1484 1450 1425 1397

Dry Mass % 10.38 10.10 12.81 11.51

Diameter 0.909 0.914 1.093 1.103

Length 3.95 3.55 3.91 3.56

Propellant Fraction 0.715 0.714 0.685 0.684

No. Engines 1 3 1 3

MMH/NTO CH4/LOX

1417

11.25

159.4

1.004

3.52

0.699

0.3516

200

3

1448

11.44

165.6

0.996

3.86

0.700

0.3516

175

1

RP-1/LOX

Dry Mass 154.1 145.8 182.5 176.0

Opt D/LFB 0.3125 0.3125 0.3906 0.3906

Opt ε 200 200 200 225

GLOW 1484 1450 1425 1397

Dry Mass % 10.38 10.10 12.81 11.51

Diameter 0.909 0.914 1.093 1.103

Length 3.95 3.55 3.91 3.56

Propellant Fraction 0.715 0.714 0.685 0.684

No. Engines 1 3 1 3

MMH/NTO CH4/LOX

Propellants MMH/NTO RP-1/LOX CH4/LOX

Engine Thrust, N 13802 13468 13253

Engine Thrust-to-Weight 35.2 35.62 34.70

Engine Mass, kg 39.9 38.51 38.90

Fuel/Oxidizer Volume, m3 0.429/0.508 0.406/0.738 0.695/0.769

Fuel Tank Mass, m 10.73 10.22 16.11

Oxidizer Tank Mass, m 11.64 17.23 17.86

Tank Insulation Mass. kg 0.0 3.84 8.49

Structure, kg 42.54 48.97 55.81

Propellant Feed & Press Mass, kg 2.84 3.12 3.65

Main Propulsion Mass 49.31 48.04 48.98

Thrust Structure mass, kg 4.65 4.86 5.13

Prime Power Mass, kg 2.33 2.28 2.25

Power Conversion & Dist Mass, kg 3.52 3.52 3.56

Propellant Feed System mass, kg 1.33 1.37 1.52

Engine Gimbal/Control mass, kg 5.80 5.67 5.71

Dry Mass, kg 154.06 165.60 182.52

Propellant Mass, kg 1119.23 1072.12 1032.83

Overall Body Length / Diameter, m 3.95 / 0.909 3.86/0.996 3.91/ 1.093

GLOW, kg 1484 1448 1425

Propellant Fraction Available, % 71.3 70.0 68.5
35

Table 10. C3=0/120kg comparison between hypergolic and cryogenic propellants 

Figure 8.3.1 Relative gross liftoff mass versus relative engine specific impulse 

Propellants MMH/NTO RP-1/LOX LOX/CH4

Engine Thrust, N 20163 19531 19255

Engine Thrust-to-Weight 41.0 40.6 40.1

Engine Mass, kg 50.11 49.09 48.5

Fuel/Oxidizer Volume, m3 0.711/0.841 0.669/1.194 1.132/1.254

Fuel Tank Mass, m 16.67 15.81 24.56

Oxidizer Tank Mass, m 18.19 26.28 27.42

Tank Insulation Mass. kg 0.0 5.66 12.11

Structure, kg 52.74 61.85 72.00

Propellant Feed & Press Mass, kg 4.04 4.45 5.26

Main Propulsion Mass 62.36 61.26 61.45

Thrust Structure mass, kg 6.39 6.66 7.07

Prime Power Mass, kg 3.22 3.13 3.09

Power Conversion & Dist Mass, kg 3.66 3.67 3.71

Propellant Feed System mass, kg 1.79 1.85 2.05

Engine Gimbal/Control mass, kg 6.73 6.64 6.59

Dry Mass, kg 179.14 197.54 220.02

Propellant Mass, kg 1853.2 1766.9 1715.0

Overall Body Length / Diameter, m 4.38 / 0.993 4.34/1.090 4.39/ 1.207

GLOW, kg 2168 2100 2070

Propellant Fraction Available, % 80.9 79.6 78.4

Gross Mass

0.90

0.95

1.00

1.05

1.10

1.15

1.20

1.25

0.85 0.9 0.95 1 1.05 1.1

MAV Engine Relative ISP 

Relative Gross Mass

MMH/NTO

CH4/LOX

RP-1/LOX
36

Figure 8.3.2 Relative dry mass versus relative engine specific impulse 

Figure 8.3.3 Relative vehicle length versus relative engine specific impulse 

Dry Mass

0.90

0.95

1.00

1.05

1.10

1.15

1.20

1.25

0.85 0.9 0.95 1 1.05 1.1

MAV Engine Relative ISP 

Relative Dry Mass

MMH/NTO

CH4/LOX

RP-1/LOX

Total Length

0.90

0.95

1.00

1.05

1.10

1.15

1.20

1.25

0.85 0.9 0.95 1 1.05 1.1

MAV Engine Relative ISP 

Relative Body Length

MMH/NTO

CH4/LOX

RP-1/LOX
37

Figure 8.3.4 Relative gross liftoff mass versus relative engine mass 

Figure 8.3.5 Relative dry mass versus relative engine mass 

Gross Mass

0.90

0.95

1.00

1.05

1.10

0.7 0.8 0.9 1 1.1 1.2 1.3

MAV Engine Relative Mass 

Relative Gross Mass

MMH/NTO

CH4/LOX

RP-1/LOX

Dry Mass

0.90

0.95

1.00

1.05

1.10

0.7 0.8 0.9 1 1.1 1.2 1.3

MAV Engine Relative Mass 

Relative Dry Mass

MMH/NTO

CH4/LOX

RP-1/LOX
38

Figure 8.3.6 Relative overall vehicle length versus relative engine specific mass 

9.0 Final Configurations 

During the course of this MAV conceptual design and technology requirements study, a 

concurrent effort focused on the conceptual design of the ERV, including total ΔV requirements 

and two propulsion systems options: 1) all chemical propulsion and 2) solar-electric propulsion. 

Two ERV designs evolved and preliminary mass and scale estimates were provided to the MAV 

design team. For the chemical propulsion ERV, an optimal ΔV split of 50% MAV and 50% ERV 

was determined and the associated mass and volume of the ERV defined. The MAV design 

closure process was repeated for the two finalized ERV concepts and the results presented here, 

along with the pressure-feed 20kg inert payload mass MAV design. Table 11 presents a MEL 

weight statement and overall vehicle dimensions for the three MAV configurations. All concepts 

utilize storable hypergolic propellants. 

9.1 Inert Capsule to low Mars orbit rendezvous 

For the 20kg inert capsule Mars orbit rendezvous architecture, a pressure-feed single engine 

concept was selected for the final design. The decision to go with pressure-feed engine was 

driven in part by the large down scaling range required for the baseline turbo-pump feed engine. 

The second column of Table 11 presents the MEL mass statement for the inert capsule MAV, 

including subsystem mass and dry mass or gross mass fraction. The high pressure propellant 

tanks dominate the total structural mass, with a dry mass fraction of 27%. Total dry mass is 64.7 

kg, which is 17% of the gross liftoff mass. Total propellant mass fraction is77.4%, which includes 

ascent propellant, reserves and engine start-up. The payload mass fraction is 5%, with a gross 

liftoff mass of 393 kg. The overall vehicle length is 2.74 meters and diameter of 0.56 meters. 

9.2 Optimized ΔV Split Architecture

For the chemical propulsion ERV, the MAV total ΔV is 4250 m/sec and the ERV mass is 230 

kg, with a volume of 0.575 m3

. The engine design is turbo pump-feed, 3 engine configuration 

Total Length

0.90

0.95

1.00

1.05

1.10

0.7 0.8 0.9 1 1.1 1.2 1.3

MAV Engine Relative Mass 

Relative MAV Body Length

MMH/NTO

CH4/LOX

RP-1/LOX
39

with a engine nozzle expansion ratio of 200. For the low pressure propellant tank concept, the 

structural dry mass fraction is 40% and the main propulsion system dry mass fraction is 39%. 

Total dry mass is 165 kg, which is ~10% of the gross liftoff mass. Total propellant mass fraction 

is76.3%, which includes ascent propellant, reserves and engine start-up. The payload mass 

fraction is 13.4%, with a gross liftoff mass of 1719 kg, below the estimated landed mass limit. 

The overall vehicle length is 3.75 meters and diameter 0.97 meters, also below the length and 

diameter constraints. 

9.3 SEP Architecture 

For the SEP ERV, the MAV total ΔV is 5500 m/sec and the ERV mass is 110 kg, with a 

volume of 0.275 m3. The engine design is turbo pump-feed, single engine configuration with a 

engine nozzle expansion ratio of 200. The structural dry mass fraction is 38% (tank dry mass 

fraction is 25%) and the main propulsion system dry mass fraction is 46%. Total dry mass is 170 

kg, which is 8.4% of the gross liftoff mass. For the C3=0 mission, the total propellant mass 

fraction is 85.5%, which again includes ascent propellant, reserves and engine start-up. The 

payload mass fraction is 5.4%, with a gross liftoff mass of 2025 kg, very close to the estimated 

landed mass limit. The overall vehicle length is 4.28 meters and diameter 0.97 meters. 

Trade studies on the number of engines (Section 8) suggest a reduction in gross liftoff mass 

on the order of 5% can be achieved by going to a multi-engine configuration. The GLOW of the 

SEP architecture MAV could possibly be reduced to roughly 1900 kg, leaving only 100 kg for 

other landed payload elements. Application of the cryogenic propellant could also provide further 

GLOW reduction. The vehicle length exceeds the landing capsule length constraint. Initial 

design sensitivity trade studies indicated low gross liftoff mass sensitivity to fore body diameterto-length ratio. The fore body fineness are could therefore be increased to reduce the overall 

vehicle length to meet the 4.0 meter length constraint, and meet the maximum diameter limit with 

modest gross liftoff mass increase. 
40

Table 11. Final concepts mass statement and dimensions 

10.0 Conclusions & Recommendations 

A preliminary study was conducted to assess the impact of mission requirements and 

technology performance on the mass and size of various classes of Mars Ascent Vehicles, 

limited by allowable mass and volume of the landing capsule. Trade studies were completed to 

define vehicle mass and scale sensitivity to engine/propellant type and configuration, fore body 

geometry, payload mass and total ΔV requirements.

The volume/length constraints imposed on the MAV design resulted in a single stage-to-orbit 

architecture, and in turn the requirement for high performance, high thrust-to-weight rocket 

engine. A high pressure turbo pump-feed, high expansion ratio engine utilizing storable 

hypergolic propellants was selected for the baseline main propulsion engine, with engine specific 

impulse of 340+ seconds and an engine thrust-to-weight ratio of 30+. A nested tank design was 

chosen to eliminate the fuel/oxidizer inter-tank section, saving weight and reducing overall 

vehicle length. Composite materials were selected for tanks, aft compartment and aeroshell 

construction based on potential weight savings. Aerodynamic and aerothermal considerations 

do not significantly impact MAV performance and closure mass. 

For the inert 20 kg class payloads to low Mars rendezvous orbit, a pressure-feed hypergolic 

engine easily meets the down mass and volume constraints imposed by the landing capsule. A 

pump-feed concept has lower gross liftoff mass, however scaling of the baseline engine to meet 

thrust-to-weight requirements resulted in an engine size only ~15% of the baseline engine. The 

Engine(s) and Thrust/Engine 3 / 1210 nts 3 / 5330 nts 1 / 18,925 nts

Main Engine(s) 18.76 (37.7%) 44.87 (35.4%) 48.17 (36.8%)

Feed & Pressurization 3.34 (6.7%) 3.07 (2.4%) 3.82 (2.9%)

Engine Control, Purge & Install 0.20 (0.4%) 0.87 (0.7%) 7.6 (5.8%)

Prime Power 0.40 (0.8%) 0.40 (0.3%) 3.05 (2.3%)

Power Conversion/Distribution 3.12 (6.3%) 3.49 (2.8%) 3.63 (2.8%)

DHCC 1.0 (2.0%) 1.0 (0.8%) 1.0 (0.8%)

Contingency 14.93 (30%) 38.05 (30%) 39.3 (30%)

Dry Mass, kg (%GLOW) 64.72 (16.5%) 164.88 (9.6%) 170.3 (8.4%)

Propellants, kg (%GLOW) 303.88 (77.4%) 1311 (76.3%) 1740 (85.5)

Residuals + Pressurants, kg (%GLOW) 3.85 (0.1%) 12.48 (0.7%) 15.4 (0.7%)

Payload, kg (%GLOW) 20 (5.1%) 230 (13.4%) 110 (5.4%)

Gross Liftoff Mass, kg 392.6 1719 2035

ΔVTOTAL/ERV Mass 4150 / 20 kg 4250 / 230 kg 5500 / 110 kg

Payload Adapter 1.05 (2.1%) 12.07 (9.5%) 5.77 (4.4%)

Induced Environments 2.10 (4.0%) 2.68 (2.1%) 4.64 (3.5%)

Auxiliary (Separation) Systems 2.61 (5.2%) 18.67 (14.7%) 9.49 (7.2%)

Main Propulsion 22.3 (44.8%) 48.82 (38.5%) 59.59 (45.5%)

3.75 / 0.97

3.44 (2.7%)

7.77 (6.1%)

25.71 (20.5%)

2.76 (2.2%)

51.76 (40.8%)

Chemical Prop

Aeroshell 1.18 (2.4%) 2.67 (2.0%)

Overall length/Diameter, m 2.74 / 0.56 4.28 / 0.97

Thrust Structure 1.96 (3.9%) 6.02 (4.6%)

Tanks 13.29 ( 26.7%) 32.96 (25.1%)

Aft Compartment + Secondary 0.85 (1.7%) 2.17 (1.6%)

Total Structure, kg (% Dry Mass) 18.33 (36.8%) 49.61 (37.9%%)

ERV Concept Inert SEP

Engine(s) and Thrust/Engine 3 / 1210 nts 3 / 5330 nts 1 / 18,925 nts

Main Engine(s) 18.76 (37.7%) 44.87 (35.4%) 48.17 (36.8%)

Feed & Pressurization 3.34 (6.7%) 3.07 (2.4%) 3.82 (2.9%)

Engine Control, Purge & Install 0.20 (0.4%) 0.87 (0.7%) 7.6 (5.8%)

Prime Power 0.40 (0.8%) 0.40 (0.3%) 3.05 (2.3%)

Power Conversion/Distribution 3.12 (6.3%) 3.49 (2.8%) 3.63 (2.8%)

DHCC 1.0 (2.0%) 1.0 (0.8%) 1.0 (0.8%)

Contingency 14.93 (30%) 38.05 (30%) 39.3 (30%)

Dry Mass, kg (%GLOW) 64.72 (16.5%) 164.88 (9.6%) 170.3 (8.4%)

Propellants, kg (%GLOW) 303.88 (77.4%) 1311 (76.3%) 1740 (85.5)

Residuals + Pressurants, kg (%GLOW) 3.85 (0.1%) 12.48 (0.7%) 15.4 (0.7%)

Payload, kg (%GLOW) 20 (5.1%) 230 (13.4%) 110 (5.4%)

Gross Liftoff Mass, kg 392.6 1719 2035

ΔVTOTAL/ERV Mass 4150 / 20 kg 4250 / 230 kg 5500 / 110 kg

Payload Adapter 1.05 (2.1%) 12.07 (9.5%) 5.77 (4.4%)

Induced Environments 2.10 (4.0%) 2.68 (2.1%) 4.64 (3.5%)

Auxiliary (Separation) Systems 2.61 (5.2%) 18.67 (14.7%) 9.49 (7.2%)

Main Propulsion 22.3 (44.8%) 48.82 (38.5%) 59.59 (45.5%)

3.75 / 0.97

3.44 (2.7%)

7.77 (6.1%)

25.71 (20.5%)

2.76 (2.2%)

51.76 (40.8%)

Chemical Prop

Aeroshell 1.18 (2.4%) 2.67 (2.0%)

Overall length/Diameter, m 2.74 / 0.56 4.28 / 0.97

Thrust Structure 1.96 (3.9%) 6.02 (4.6%)

Tanks 13.29 ( 26.7%) 32.96 (25.1%)

Aft Compartment + Secondary 0.85 (1.7%) 2.17 (1.6%)

Total Structure, kg (% Dry Mass) 18.33 (36.8%) 49.61 (37.9%%)

ERV Concept Inert SEP
41

engine performance and mass scaling models may not accurately reflect engine performance at 

that scale. 

For propulsive ERV concepts in the 100 to 200+ kg mass class, the pump-feed engine 

concept is enabling technology. Pressure-feed concepts are not competitive and significantly 

exceed mass and volume constraints. For the optimal ΔV split architecture, storable hypergolic 

and all cryogenic methane/oxygen and kerosene/liquid oxygen concepts are equally competitive, 

with a slight gross mass edge for the cryogenic systems and a lower dry mass for the hypergolic 

design. The C3=0 baseline architecture marginally meets the mass and vehicle length 

constraints, but would leave little mass availability of other mission systems. The overall impact 

of ISRU technology needs to be assessed, with the mass, power and volume requirements by 

the additional subsystems to produce and store cryogenic propellants on the Marian surface 

determined, balanced against potentially significant mass saving in the landed MAV mass. 

Engine performance and mass characteristics have the biggest impact on the MAV design. 

The availability of a high performance, low mass turbo pump-feed engine in the proper thrust 

size is critical for the ERV class missions. The application of accurate engine performance and 

mass estimation models are highly recommended to confirm and refine the present conceptual 

design performance predictions. The mass estimating relationships used in the vehicle closure 

process for critical subsystems were based on much larger scale launch vehicles and often 

resulted in extrapolation of the design data base. More physics-based analysis is also 

recommended for further preliminary design efforts for the MAV concepts. Finally, a refined 

estimate of the landed payload capability of the entry capsule is recommended to assess the 

viability of the SEP ERV architecture. 
42

Appendix A: Engine Trades

A quasi-one dimensional rocket performance code was developed to predict engine 

performance as a function of chamber pressure, oxidizer-to-fuel ratio and nozzle expansion ratio. 

The equilibrium thermo-chemical model is based on the NASA Glenn CEA code (Reference 10), 

with addition subroutines developed to predict the thermodynamic properties of the nozzle flow 

assuming fixed mole fractions, i.e. frozen flow. The mass, momentum and energy equations are 

solved over a finite area ratio control volume. Based on published engine performance 

characteristics, the best match to the engine performance was obtained assuming frozen flow 

starting at the nozzle throat. 

For a mixture of hypergolic propellants, specifically MMH and NTO, the predicted engine 

performance, namely engine specific impulse and thrust per unit nozzle exit area is presented in 

Figures A1 and A2, respectively. The oxidizer-to-fuel ratio has been optimized to produce 

maximum engine specific impulse, and has a fuel-rich value of 2.0. As presented in Figure A1, 

engine specific impulse increases with both increasing nozzle expansion ratio and increasing 

thrust chamber pressure. Thrust per unit nozzle exit area is a nearly linear function of chamber 

pressure, and decreases with increasing nozzle expansion ratio. For a required amount of rocket 

engine thrust, the thrust per unit nozzle exit area is an indicator of the physical size of the engine, 

with engine size varying roughly inversely with nozzle expansion ratio. 

The reference engine selected was the Rocketdyne orbital-transfer XLR-132, a hypergolic, 

gas-generator, pump-feed engine. The chamber pressure is 10.34 MPa (1500 psi) and it has a 

nozzle expansion ratio of 400. The predicted performance of the engine is within 0.5% of both 

specific impulse and thrust per unit nozzle exit area. Weight scaling of the engine with 

expansion ratio was based on the XLR-132, with a nozzle unit mass of approximately 12.5 kg/m2

. 

Using the XLR-132 as the baseline engine, trades on engine mass, thrust-to-weight 

characteristics and engine nozzle exit diameter as a function of expansion ratio and chamber 

pressure were conducted. For a fixed engine thrust of approximately 12,000 N, the engine 

mass, engine thrust-to-weight and nozzle exit diameter are presented in Figures A3 through A5. 

As both chamber pressure and nozzle expansion ratio are increased, engine mass and nozzle 

diameter decrease, while engine thrust-to-weight increase. For the expansion ratio = 100, a 

minimum in the engine weight occurs around 4.1 – 5.5 MPa (600 to 800 psi). 
43

Figure A1. Engine specific impulse versus chamber pressure and nozzle expansion ratio. 

Figure A2. Engine thrust per unit nozzle exit area versus chamber pressure and 

nozzle expansion ratio. 

MMH/NTO: Engine Isp

320.0

325.0

330.0

335.0

340.0

345.0

350.0

355.0

0 100 200 300 400 500

Expansion Ratio

Engine Specific Impulse, secs

Chamber Pressure=150psi

Chamber Pressure=300 psi

Chamber Pressure=650 psi

Chamber Pressure=1000 psi

Chamber Pressure=1200 psi

Chamber Pressure=1500 psi

MMH/NTO: Thrust/AEXIT

0.0

1000.0

2000.0

3000.0

4000.0

5000.0

6000.0

7000.0

8000.0

9000.0

0 100 200 300 400 500

Expansion Ratio

Thrust/AEXIT, psf

Chamber Pressure=150psi

Chamber Pressure=300 psi

Chamber Pressure=650 psi

Chamber Pressure=1000 psi

Chamber Pressure=1200 psi

Chamber Pressure=1500 psi
44

Figure A3. Engine mass versus chamber pressure and nozzle expansion ratio. 

Figure A4. Engine diameter versus chamber pressure and nozzle expansion ratio. 

Engine Mass vs. Chamber Pressure

MMH/NTO : Engine Thrust = 2700 lbf

0

20

40

60

80

100

120

140

160

180

200

0 200 400 600 800 1000 1200

Chamber Pressure, psi

Engine Weight, lbs

Expansion Ratio = 100

Expansion Ratio = 150

Expansion Ratio = 250

Engine Diameter vs. Chamber Pressure

MMH/NTO : Engine Thrust = 2700 lbf

0.00

0.50

1.00

1.50

2.00

2.50

3.00

3.50

4.00

4.50

0 200 400 600 800 1000 1200

Chamber Pressure, psi

Engine Diameter, ft

Expansion Ratio = 100

Expansion Ratio = 150

Expansion Ratio = 250
45

Figure A5. Engine thrust-to-weight versus chamber pressure and nozzle expansion 

ratio. 

Engine Thrust/Weight vs. Chamber Pressure

MMH/NTO : Engine Thrust = 2700 lbf

0.00

5.00

10.00

15.00

20.00

25.00

30.00

35.00

0 200 400 600 800 1000 1200

Chamber Pressure, psi

Engine T/W

Expansion Ratio = 100

Expansion Ratio = 150

Expansion Ratio = 250
46

Appendix B: Propellant Trades 

The baseline propellant selected was a storable hypergolic mixture of MMH and NTO. 

Alternate hypergolic fuels were also considered, including hydrazine and UDMH. Figure B1 

shows the engine specific impulse for the three fuels with NTO at the optimum oxidizer-to-fuel 

mixture ratio as a function of nozzle expansion ratio. Although hydrazine has the highest specific 

impulse and the mean propellant density is similar for all three fuel/oxidizer combinations, MMH 

was selected as the baseline propellant because the reference engine (XLR-132) utilizes MMH 

and the predicted engine performance could be anchored to the actual engine performance. 

Alternate hydrocarbon-based fuels with liquid oxygen were also considered. Figure B2 

presents the engine specific impulse as a function of nozzle expansion ratio, at a chamber 

pressure of 4.48 MPa (650 psi) and an oxidizer-to-fuel mixture ratio of 3.0. Methane provides a 

significantly higher specific impulse compared to the higher molecular weight hydrocarbons and 

was selected as the fuel of choice for the hydrocarbon-based propellants. RP-1/LOX was also 

selected, based on mean propellant bulk density considerations. Engine performance of the RP1/LOX design was based on an oxidizer-to-fuel ratio of 2.25. 

Figure B1. Engine specific impulse versus nozzle expansion ratio for various 

hypergolic fuels with NTO. 

Hypergolic Fuels

Optimum O/F & Chamber Pressure = 1500 psi

330

335

340

345

350

355

360

50 100 150 200 250 300 350 400 450

Nozzle Area ratio Ae/At

Engine Specific Impulse, 

secs

Hydrazine

MMH

UDMH
47

Figure B2. Engine specific impulse versus nozzle expansion ratio for various 

hydrocarbon fuels with liquid oxygen 

HydroCarbon (Paraffin Series) + LOX

Pc = 650 psi & O/F = 3.0

325

330

335

340

345

350

355

360

365

370

375

50 100 150 200 250 300 350 400 450

Expansion ratio Ae/At

Engine Specific Impulse, 

secs

Methane

Propane

Butane

Pentane

Octane

RP-1 (Kerosene)
48

Appendix C: CFD Solutions & Aeroheating 

Cart3D was used to provide a preliminary aerodynamic analysis on the MAV vehicles. 

Cart3D solutions were computed for three candidate forebody shapes at subsonic, transonic and 

low supersonic flight conditions. Cart3D is a high-fidelity inviscid analysis package for 

conceptual and preliminary aerodynamic design. The adjoint mesh adaptive version of Cart3D 

was utilized allowing for a more automated set up of the inputs to the code and a more rapid 

arrival at reasonable solutions for a large range of flow conditions. Geometry is input via an 

unstructured triangular mesh elements with the option of tagging to denote different regions (ex. 

aerodynamic, propulsion, base etc...). Cart3D version v1.4.7_LINUX64—12.05.07 was used 

for the analysis in this report. The MAV Rev2 geometry was analyzed at Mach 0.5, 0.95, 1.05, 

1.5 and 2.0 at four angles of attack (0°, 2°, 5° and 10°) with a typical solution running to over 1.5 

million cells. Force and moment histories for each of the grid cycles was monitored for 

convergence quality for each case. 

Figures C1 through C5 present Cart3D solutions at various Mach numbers and angles-ofattack, and show surface pressure coefficient distributions and Mach number contours. 

Figure C6 summarizes the Euler solutions, plotting zero lift drag coefficient versus Mach 

number over the subsonic, transonic and low supersonic flight regimes. Figure C7 presents the 

pitching moment coefficient versus Mach number and angle-of-attack. For the selected moment 

reference (roughly 50% of body length), the vehicle is statically unstable in pitch. For the internal 

arrangement selected, with the oxidizer tank forward, the axial center-of-gravity shifts forward 

and the vehicle would be statically stable. 

A preliminary trajectory was computed and the resulting Mach number, angle-of-attack and 

free-stream dynamic pressure history was used to interpolate the aerothermal database to 

estimate the ascent heating environment. Figures C8 and C9 summarized the heating 

environment, showing mission maximum heating rate distribution and mission integrated heat 

load, respectively. Peak heating occurs at roughly Mach = 5.0 and dynamic pressure of 60 Pa. 

Stagnation point peak heating is on the order of only 0.5 W/cm2

, and the integrated heat load is 

merely 0.5 MJ/m2

. Minimum gauge TPS on the nose cap region only (e.g. P-45 Cork) will be 

adequate. Peak acreage radiation equilibrium wall temperatures are on the order of 390oK, well 

below allowable temperatures for the selected structural material. 
49

Figure C1. Cart3D predicted subsonic Mach = 0.6 surface pressure distribution for 

the Rev0 geometry 

Figure C2. Cart3D predicted subsonic Mach = 0.5 surface pressure distribution and 

Mach number contours for the Rev2 geometry 
50

Figure C3. Cart3D predicted Mach = 0.95 surface pressure distribution and Mach 

number contours for the Rev2 geometry 

Figure C4. Cart3D predicted Mach = 2.0 surface pressure distribution and Mach 

number contours for the Rev2 geometry 
51

Figure C5. Cart3D predicted Mach = 2.0 and angle-of-attack = 10.0° surface 

pressure distribution and Mach number contours for the Rev2 geometry 

Figure C6. Zero-lift drag coefficient versus Mach number computed by CFD Euler code 

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.0 0.5 1.0 1.5 2.0 2.5

CDmin

Mach
52

Figure C7. Cart3D predicted pitching moment coefficient versus angle-of-attack and 

Mach number contours for the Rev2 geometry 

Figure C8. Mission maximum convective heating distribution for Rev2 geometry 

-0.025

0.000

0.025

0.050

0.075

0.100

0.125

0.150

0123456789 10

CM

Angle of Attack (deg)

MSR MAV rev2

Cart3D Adjoint

Preliminary Results

Mach 0.5 Mach 0.95

Mach 1.05 Mach 1.5

Mach 2.0

Moment reference X 1.9425 m Y 0m Z 0.0 m
53

Figure C9. Mission integrated heat load for Rev2 geometry 
54

References 

1) Kinney, D.J., “Aero-Thermodynamics for Conceptual Design”, AIAA Paper 13382, January 

2004, Reno Nevada 

2) Kinney, D. “Aerothermal Anchoring of CBAERO Using High Fidelity CFD”, AIAA Paper 2007-

608, January 2007 

3) Cart3D documentation website http://people.nas.nasa.gov/!aftosmis/cart3d/ 

4) Lynn, Emory, “Integrated Rocket Sizing Program”, Version 4.2, User’s Manual, Oct. 2007.

5) Forbos, J. C. and Woodhead, G. E., “Conceptual Design ad Analysis of Hypervelocity 

Aerospace Vehicles”, Mass Properties, AFWAL Report WL-TR-91-6003, Vol I, Part I. 

6) Bowles, J. V. “Conceptual Studies Activities”, Proceedings of the Second National Aerospace 

Plane Symposium, Applied Physics Laboratory, Laurel MD, Nov. 1986. 

7) Whitehead, J. “Propulsion Engineering Study for small-Scale Mars Missions, Lawrence 

Livermore National Laboratory, UCRL-CR-122442, Sept. 1995. 

8) Whitehead, Pittenger, and Colella, “design and Flight Testing of a Reciprocating Pump Feed 

Rocket, AIAA 94-3032, July 2994 

9) G.L. Brauer, et al, “Capabilities and Applications of the program to Optimize Simulated 

Trajectories (POST),” NASA CR-2770, February 1977. 

10) McBride, B. J. and Gordon, S., "Computer Program for Calculation of Complex Chemical 

Equilibrium Compositions and Applications II. User's Manual and Program Description", 

NASA RP-1311-P2, June 1996