March 2009 

NASA/TM-2009-215700 

Design of a Mars Airplane Propulsion System 

for the Aerial Regional-Scale Environmental 

Survey (ARES) Mission Concept 

Christopher A. Kuhl 

Langley Research Center, Hampton, Virginia 
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March 2009 

NASA/TM-2009-215700 

Design of a Mars Airplane Propulsion System 

for the Aerial Regional-Scale Environmental 

Survey (ARES) Mission Concept 

Christopher A. Kuhl 

Langley Research Center, Hampton, Virginia 
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 

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Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 1

TABLE OF CONTENTS 

1.0 Introduction .................................................................................................................................................3

2.0 Mission Description.....................................................................................................................................3

2.1 Concept of Operations ..............................................................................................................................3

2.2 Target Region............................................................................................................................................4

2.3 Atmospheric Flight System (Airplane) .....................................................................................................5

3.0 Propulsion Subsystem Trade Studies ........................................................................................................6

3.1 Propeller Development .............................................................................................................................6

3.2 Electrical Propulsion Systems...................................................................................................................6

3.3 Combustion Engine Systems.....................................................................................................................7

3.4 Rocket Systems.........................................................................................................................................7

3.5 Airplane Propulsion Subystem Selection..................................................................................................8

4.0 Bipropellant Rocket Subsystem Design.....................................................................................................9

4.1 System Schematic and Layout ..................................................................................................................9

4.2 High Pressure Helium Storage Tank.......................................................................................................12

4.3 Propellant Tanks .....................................................................................................................................12

4.3.1 Tank Design........................................................................................................................................12

4.3.2 Propellant Management Device (PMD) Design .................................................................................13

4.4 Bipropellant Thrusters.............................................................................................................................15

4.4.1 22N Satellite Reaction Control Thruster Options...............................................................................15

4.4.2 AMPAC-ISP 22-N Columbium Thruster ...........................................................................................16

4.4.3 Thruster Mounting and Orientation ....................................................................................................17

4.5 Thruster Valves.......................................................................................................................................18

4.6 Propellant Isolation and Flow Control Hardware ...................................................................................19

4.6.1 Pyrotechnic-actuated Valves (Pyrovalves) .........................................................................................19

4.6.2 Line Filters..........................................................................................................................................19

4.6.3 Check Valves......................................................................................................................................19

4.6.4 Helium Regulator................................................................................................................................20

4.6.5 Flow Venturis .....................................................................................................................................20

4.6.6 Service Valves ....................................................................................................................................20

4.6.7 Fluid Field Joints ................................................................................................................................20

4.7 Command and Instrumentation...............................................................................................................20

4.7.1 Temperature Sensors ..........................................................................................................................21

4.7.2 Pressure Transducer............................................................................................................................21

4.8 Hardware Integration ..............................................................................................................................22

5.0 Master Equipment List .............................................................................................................................23

6.0 Propulsion Subsystem Performance........................................................................................................24

6.1 Thruster Firing – Thermal Impact...........................................................................................................24

6.2 Thruster Plumes ......................................................................................................................................26

6.3 Thruster Firing Logic..............................................................................................................................27

6.4 Airplane Flight Simulation......................................................................................................................27

6.5 Airplane Flight Duration and Range.......................................................................................................27

7.0 Operational Procedures ............................................................................................................................29

7.1 Propellant Loading Procedures ...............................................................................................................29

7.2 System Activation Sequence...................................................................................................................29

7.3 Component Qualification Requirements.................................................................................................30

7.4 Ferric Nitrate Mitigation .........................................................................................................................31

8.0 Acknowledgements....................................................................................................................................32

9.0 Acronyms ...................................................................................................................................................32

10.0 References ..................................................................................................................................................33
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 2

TABLE OF FIGURES 

Figure 1. Mission Concept Overview........................................................................................................................3

Figure 2. Target Region.............................................................................................................................................4

Figure 3. ARES Airplane Aero-shell Configuration .................................................................................................5

Figure 4. ARES Airplane Geometry..........................................................................................................................5

Figure 5. ARES Airplane Propulsion Subsystem Schematic ..................................................................................10

Figure 6. Helium Composite Over-wrapped Pressure Vessel .................................................................................12

Figure 7. Propellant Tank and PMD Design ...........................................................................................................14

Figure 8. AMPAC-ISP 22N Thruster Details..........................................................................................................16

Figure 9. Triple Thruster Orientation and Alignment .............................................................................................17

Figure 10. Triple 22N In-Line Thruster Mounting and Assembly ..........................................................................18

Figure 11. Pyrovalve - Conax Model 1832-205 ......................................................................................................21

Figure 12. Liquid Filter – Vacco Model F1D10558-01 ..........................................................................................21

Figure 13. Gas-Side Filter – Vacco Model F1D10286-01 ......................................................................................21

Figure 14. Check Valves – Vacco Model V1D10856-02........................................................................................21

Figure 15. Helium Regulator - Vacco Model 64720-1............................................................................................21

Figure 16. Service Valves – Moog 50E740.............................................................................................................21

Figure 17. Pressure Transducer – GP:50 Model 7200.............................................................................................21

Figure 18. Propulsion System / Airframe Integration .............................................................................................22

Figure 19. Hot Fire Test Data for AMPAC-ISP 22N Thruster9. ............................................................................24

Figure 20. Temperature profile for lower fuselage without heat shield. .................................................................25

Figure 21. Temperature profile of lower fuselage with heat shield.........................................................................25

Figure 23. Rear view of Airplane showing airplane in coast (above) and with thrusters firing (below) ................26

Figure 22. Side view of Airplane showing airplane in coast (above) and with thrusters firing (below).................26

Figure 24. Thruster Duty Cycles for No Turbulence Conditions ............................................................................28

Figure 25. Thruster Duty Cycles for Severe Turbulence Conditions ......................................................................28

 

TABLE OF TABLES 

Table 1. ARES Airplane Parameters .........................................................................................................................5

Table 2. Comparison of Propulsion System Options ................................................................................................6

Table 3. Rocket System Propellant Candidates.........................................................................................................7

Table 4. Propulsion System Components..................................................................................................................9

Table 5. Propulsion Component Functional Descriptions.......................................................................................11

Table 6. Fuel Extraction Concepts ..........................................................................................................................13

Table 7. Mars Airplane Commercial 22N Thruster Options ...................................................................................15

Table 8. AMPAC-ISP 22N Thruster Characteristics...............................................................................................16

Table 9. Propulsion System - Master Equipment List.............................................................................................23

Table 10. Propulsion System Simulated Performance Summary............................................................................27

Table 11. Airplane Propellant Loading Procedure ..................................................................................................29

Table 12. Component Qualification Requirements .................................................................................................30

Table 13. ARES Ferric Nitrate Mitigation Recommendations................................................................................31
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 3

1.0 INTRODUCTION 

The Aerial Regional-Scale Environmental Survey (ARES) is a Mars exploration mission concept with the 

goal of taking scientific measurements of the atmosphere, surface, and subsurface of Mars by using an 

airplane as the payload platform. ARES team first conducted a Phase-A study for a 2007 launch 

opportunity, which was completed in May 2003. Following this study, significant efforts were 

undertaken to reduce the risk of the atmospheric flight system, under the NASA Langley Planetary 

Airplane Risk Reduction Project. The concept was then proposed to the Mars Scout program in 2006 for 

a 2011 launch opportunity. This paper summarizes the design and development of the ARES airplane 

propulsion subsystem beginning with the inception of the ARES project in 2002 through the submittal of 

the Mars Scout proposal in July 2006. 

2.0 MISSION DESCRIPTION 

The Aerial Regional-scale Environmental Survey (ARES) mission expands upon other Mars exploration 

missions, including Viking, Mars Global Surveyor (MGS), Odyssey, and the Mars Exploration Rovers 

(MER), to examine the structure and evolution of Mars’ atmosphere, surface, and interior. The ARES 

atmospheric flight system will fly above the Southern Highlands to study crustal magnetism, atmospheric 

boundary layer composition, chemistry and dynamics, and also to explore for near-surface water. 

2.1 CONCEPT OF OPERATIONS 

The ARES concept has a baseline launch date from Kennedy Space Center in October, 2011 on a Delta II 

2925, placing it on a 10.4 month trajectory to Mars. ARES will utilize a Type II interplanetary trajectory 

with a series of five planned trajectory correction maneuvers (TCMs) with a final direct entry into Mars. 

The entry system (aeroshell and atmospheric flight system) will separate from the carrier spacecraft 

before entering Mars atmosphere. The carrier spacecraft performs a divert maneuver to a Mars flyby 

trajectory, enabling it to relay to Earth the science and engineering data collected during the airplane 

flight. Mars Reconnaissance Orbiter (MRO) will serve as a redundant data return path for critical data. 

An overview of the mission concept is shown in Figure 1. 

Figure 1. Mission Concept Overview 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 4

2.2 TARGET REGION 

The Science Target Area (STA) is a region selected by the ARES science team that provides an optimal 

location for achieving the science objectives1

. The STA is defined by the intersection of two regions of 

interest; one region with highly variable crustal magnetism and the other region where water initiated 

atmospheric chemistry can be studied. The combination of these two zones yields a large STA spanning 

20° to 60° S latitude and 150° to 210° E longitude (See Figure 2). 

During a science flight traverse of more than 70 minutes, the ARES rocket-propelled airplane will fly 

over 608 km obtaining measurements of crustal magnetism and the atmosphere. The reference traverse 

begins at a nominal delivery location of (51.2° S, 180.0° E) and travels >1.7° south prior to performing an 

aeromagnetic survey defined by two additional parallel North-South tracks, each of length 100 km, shown 

in Figure 2. Data is relayed to the fly-by spacecraft and by a redundant path to the MRO orbiter. At the 

end of the traverse, the airplane impacts the surface, ending the atmospheric flight segment of the 

mission. No data is assumed to be transmitted from the airplane after surface impact. 

60°S

20°S

220°E

150°E

Gusev Crater

Deployment 

Ellipse

Figure 2. Target Region. 

-20° to –60° Latitude, 180° Long (+/- 30°)
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 5

2.3 ATMOSPHERIC FLIGHT SYSTEM (AIRPLANE) 

The ARES Atmospheric Flight System (AFS) includes the airplane, its subsystems and the science payload. 

Continued analyses and focused trade studies, augmented by experimental results from wind-tunnel and flight 

testing, have matured the airplane and subsystem designs well beyond the fundamental conceptual level. A robust 

airplane configuration has been achieved through several design cycles, with the airplane outer mold line frozen 

since 2004, resulting in significant vehicle performance and reliability improvements. The airplane geometry is 

depicted in Figure 4. Delivery of the ARES airplane into the atmosphere of Mars requires an entry aero-shell. 

The primary configuration constraint is packaging of the airplane into the entry aero-shell in a way that 

accommodates the environments of launch and entry (Figure 3). The airplane shape is derived from the aero-shell 

packaging constraints and the need for subsonic flight required by the science measurements. Table 1 provides 

the overall mass and geometric properties for the Atmospheric Flight System1

. 

Table 1. ARES Airplane Parameters 

Parameter Value 

Wing Span 6.25 m 

Overall Length 4.45 m 

Overall Height 0.7 m 

Reference Wing Area 7.0 m2

Mean Aerodynamic Chord 1.25 m 

Dry Mass (CBE) 112 kg 

Propellant Load 45 kg 

Cruise Speed 145 m/s 

Figure 3. ARES Airplane Aero-shell Configuration 

Figure 4. ARES Airplane Geometry 

Wing Hinge 

Airplane Extraction 

System Aeroshell 

Thruster 

Bank 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 6

3.0 PROPULSION SUBSYSTEM TRADE STUDIES 

Propulsion subsystem capabilities are critical to establishing an airplane’s performance and flight envelope. This 

is especially true for an airplane that must be stowed in an aero-shell (See Figure 3), which imposes strict mass 

and volume limits on the airplane and subsystems. In selecting a propulsion subsystem technology for the 

airplane, two competing criteria were used to evaluate each system; flight range and implementation risk. The 

first criterion for propulsion subsystem technology selection is the ability of the airplane to achieve ascience range 

requirement of at least 480 km of flight. The second criterion relates to implementation risk of the propulsion 

system. The cost constraints and risk posture desired for Mars Scout missions require maximizing the use of 

existing, flight proven hardware. Extraction and deployment of the airplane from the aero-shell increases mission 

risk and any additional deployments necessitated by a propulsion subsystem, such as propeller blades, are heavily 

considered in the final propulsion subsystem selection. Various propulsion subsystem options, listed in Table 2, 

were examined to determine the candidate that best balanced the competing criteria of maximizing flight range 

and minimizing technical risk2

. To stay within mass constraints of the 185 kg airplane, the propulsion subsystem 

total mass was constrained to a maximum of 63 kg; 18 kg for hardware and about 45 kg for propellants. (During 

system design iterations, the amount of fuel varies slightly depending on available mass margin on the airplane). 

Results of the propulsion system analyses are listed in Table 2 and described below3

. 

Table 2. Comparison of Propulsion System Options 

3.1 PROPELLER DEVELOPMENT 

Electrical and combustion engine systems require the use of a propeller for generating the needed thrust. 

Propeller technology for use in the Mars atmosphere is an immature technology (TRL~3). Primary limitations are 

the aerodynamic performance needed to generate thrust, the likely need for supersonic tips speeds, and the 

potential for blade erosion or performance losses. Additionally, propeller blade diameters will be significantly 

large relative to the airplane, resulting in challenges to packaging, unfolding, and deployment of the propellers 

prior to flight. 

3.2 ELECTRICAL PROPULSION SYSTEMS 

A simplified fuel cell system can be used to generate electrical power, which in turn can be used to power an 

electric motor and spin a propeller to generate thrust. This system can achieve flight of 1512 km, carrying 15 kg 

of fuel with a 48 kg dry mass. The relative newness of the technology (TRL~4) and lack of experience in utilizing 

a fuel cell as the main power system in an airplane makes it a very high risk item. 

A battery propulsion system is the simplest of all systems considered and requires fewer active controls and has 

no mechanical components, aside from the drive train. However, batteries with high specific power have low 

discharge rates. This low rate forces the battery capacity higher and adds significant mass to the system. The 

battery systems analyzed failed to satisfy the baseline science requirement, achieving a range of only 202 km. 

Propulsion Category System Description 

Dry 

Mass 

[kg] 

Fuel 

Load 

[kg] 

Flight 

Hours

Range 

[km] 

TRL Development 

Effort 

Electrical Propulsion Fuel Cell 48 15 3.0 1512 4-5 HIGH 

Systems Battery 63 0 0.4 202 4-6 MED 

Combustion Engine Piston Expander Engine 19 44 2.0 1008 4-5 HIGH 

Systems 4 Cycle Combustion Engine 21 42 4.1 2066 4-5 HIGH 

Bipropellant Rocket 15 48 1.2 605 7-9 LOW Rocket Systems Monopropellant Rocket 11 52 0.9 454 7-9 LOW 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 7

3.3 COMBUSTION ENGINE SYSTEMS 

The concept of using a piston expander engine as a primary propeller driver dates back to the first studies of Mars 

aircraft design in the late 1970’s. The design is based on the Mini-Sniffer, high altitude, unmanned vehicle 

program4. The piston expander engine analysis shows a capability of 1008 km of flight carrying 44 kg of 

hydrazine and 19 kg of propulsion hardware. Piston expander technology maintains a TRL of ~4-5 and 

significant development would be needed to be implemented to produce an engine that can operate within the 

Mars environment. 

The 4-cycle internal combustion engine option for Mars flight is similar to a conventional airplane engine but 

would require design changes to operate within the thin atmosphere of Mars using a bi-propellant fuel and 

oxidizer. This option resulted in the longest flight range of 2066 km, with a dry-mass of 21 kg and a hydrazine 

fuel load of 42 kg. The combustion engine technology maintains a low TRL for a planetary flight application and 

would require a significant amount of design, testing and qualification for use in a Mars airplane. 

3.4 ROCKET SYSTEMS 

Rocket propulsion systems have a long heritage in space flight applications and in rocket-powered airplanes. 

There are a number of propellants that can power a rocket propelled airplane within the Mars environment. 

Several propellant combinations were considered, based on specific impulse, a stable liquid state during 

interplanetary cruise, and spaceflight heritage3

. Potential propellant combinations were narrowed down to one 

monopropellant and one bi-propellant option, listed in Table 3. Hydrazine has been used in spacecraft for several 

decades, typically in low thrust applications like satellite station keeping. Hydrazine is decomposed through a 

catalytic reaction with a metal oxide. The bi-propellant combination monomethyl hydrazine (MMH) and nitrogen 

tetroxide (MON3) is commonly used in the satellite industry for reaction control thrusters. Both hydrazine and 

MON3 have freezing points above expected temperatures during interplanetary cruise, and would therefore 

require active thermal control. The combination of MMH and MON3 has a much higher specific impulse than 

hydrazine but bi-propellant systems are typically more complicated than monopropellant systems. 

Table 3. Rocket System Propellant Candidates 

To evaluate the rocket propulsion system options, a number of designs were considered. These ranged from 

passive blow-down systems to regulated systems that were controlled either by on / off pulsing or by controlling 

the propellant flow to the thruster. For each type of control scheme, both monopropellant and bipropellant 

systems were considered. The analysis of rocket systems resulted in a flight range of 605 km for the bi-propellant 

system and 454 km for the monopropellant system. Rocket propulsion systems, due to extensive use in military 

applications and the satellite industry, have a high technology readiness level between 7 and 9, which is the 

desirable technology maturation level for a Mars Scout mission. 

Two rocket systems were viable options for a Mars airplane application, a MMH/MON3 bi-propellant system and 

a hydrazine mono-propellant system. Brief consideration was given to use of a MON25 oxidizer because of its 

lower freezing temperature, but introducing this technology with relatively low flight heritage was inconsistent 

with the risk posture of the Scout program. The final selection was driven by a system analysis that examined the 

 

Type Propellant Type 

Freezing 

Point (°C)

Density @

20°C 

(gm/cm3) 

Comb. 

Temp. 

(°C) Isp (sec) 

Monopropellant Hydrazine (N2H4) Fuel 1.5°C 1.008 633 199-230 

BiPropellant 

Monomethyl Hydrazine (MMH) (N2H6C) 

Nitrogen Tetroxide (MON3) 

(97% N2O4, 3% NO by weight) 

Fuel 

Oxidizer 

-52.5°C 

-11.2°C 

0.874 

1.39 

3122 285-310 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 8

difference in available fuel load to support the flight range of the two systems. Since the monopropellant system 

is less complex, a potentially lower dry-mass would permit additional fuel load. However packaging constraints 

of the airplane require dual propellant tanks for the monopropellant system, resulting in essentially the same 

hardware configuration as the bipropellant system. Dry-mass difference between the two systems is negligible, 

and overall performance is based solely on the specific impulse of the propellants listed in Table 3. The 

monopropellant system cannot carry enough fuel to achieve the baseline science requirement of 480 km or greater 

flight range. 

3.5 AIRPLANE PROPULSION SUBYSTEM SELECTION 

The final selection of a propulsion subsystem for the ARES airplane depended on options requiring minimal 

developmental effort and low technical risk while meeting science range requirements. For the electrical systems 

and combustion engines, a propeller is required to generate thrust, and there are significant challenges to 

operating a propeller in the Mars atmosphere. The propeller would need to operate near the sonic limit at its tip, 

which would place the blade operation in a low Reynolds number, high subsonic Mach number environment. 

This is outside the regime of conventional propeller design. Additionally, the size of the propeller would require 

it to be folded to fit within the aero-shell then deployed with the airplane during entry. Due to the aerodynamics 

and required deployment, the propeller would represent considerable risk for airplane operation. The technical 

risk caused by the need for a propeller eliminated the electrical and combustion engine systems and the 

monopropellant rocket was unable to meet the range requirement. Consequently, the bipropellant rocket system 

was selected as the only option that meets range and low technical risk requirements for the ARES mission. 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 9

4.0 BIPROPELLANT ROCKET SUBSYSTEM DESIGN 

The airplane propulsion subsystem design was derived from typical bipropellant systems used on such spacecraft 

as the Near Earth Asteroid Rendezvous and on the Mars Observer. System design has been iterated through 

several trade studies and iterations to arrive at the current baseline. The propulsion subsystem is a reliable singlestring architecture with redundancy existing only within the series-redundant check valves and within the thruster 

arrangement, where thrust margin and the ability to fly on two of three total thrusters exists. 

4.1 SYSTEM SCHEMATIC AND LAYOUT 

The airplane baseline propulsion subsystem is a regulated bi-propellant system with pulsed thruster control. 

Major components of this system consist of fuel and oxidizer tanks, a helium pressure tank, pyro valves, check 

valves, and thrusters. Due to cost and schedule constraints of the Scout missions, commercial hardware with 

significant spaceflight heritage is desirable to minimize developmental effort. ARES has meticulously researched 

and selected the commercially available “off-the-self” components and suppliers for the propulsion hardware that 

best meet the performance requirements and are compatible with ARES operational environment. Details of the 

off-the-shelf components are listed in Table 4. The schematic of this subsystem is shown in Figure 5. A 

description of the function of each component is provided in Table 5 

Table 4. Propulsion System Components. 

Component Vendor Heritage

Helium Pressurant Tank Lincoln Composites 220131-1 Aircraft/missile pressure vessel 

Propellant Tanks / PMDs Lockheed Martin Modified design of typical propellant tanks 

Bipropellant Thrusters AMPAC-ISP 5 lbf RCS Intelsat VI >1000 flown 

Thruster Valves Moog 51-178 BSS-601, BSS-701, Star-2, Muses-C, Astro-E/F 

NC Pyrotechnic Valves Conax 1832-205 MRO, Various 

Liquid Filter Vacco F1D10558 Various spacecraft/satellites 

GHe HP Filter Vacco F1D10286 Various spacecraft/satellites 

Check Valves Vacco V1D10856-2 Developed for Mars Ascent program 

Helium Regulator Vacco 64720-1 Various Satellite programs 

Flow Venturis Flow Systems, Inc. Various 

HP Service Valves Moog 50E740 Milstar derived 

Fluid Field Joints TBD Various 

Temperature Sensors Analog Devices AD90 Analog Devices 

Pressure Transducers GP:50 7200 EELV, Shuttle, TRW-ABL 

Thruster Valve Controller TBD (BRE or LMA) TBD 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 10

Figure 5. ARES Airplane Propulsion Subsystem Schematic 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 11

Table 5. Propulsion Component Functional Descriptions 

GHe Helium Supply Provides high pressure helium (3000psi), used to pressurize propellant tanks to 300 psi through regulator 

FVHE Helium fill valve Used for loading and off-loading helium. Capped before launch. 

PVHE Helium Pyrovalve Used to isolate high pressure helium tank and protect helium pressure regulator PRHE from long term

exposure to high pressure during cruise. 

SV1 Service Valve #1. Manual valve used for regulator functionality check at the system level. Capped before launch. 

FTHE Helium line filter Located downstream of pyrovalve to filter debris from pyro activation event. Select line size finer than that

of thruster valve inlets. 

PRHE Helium pressure 

regulator 

Regulates downstream pressure to propellant tanks to nominal 300 psig. Single regulator used since dual

would lead to increase propellant reserve required to cover mixture ratio uncertainty cause by possible

mismatched set points. 

SV2 Service Valve #2 Manual valve used for functionality and leak checking of the downstream series redundant check valves.

Capped before launch. 

FCK1 Fuel side check valve 

#1 Series redundant check valves used to alleviate large mass flow of MMH into upstream helium section. 

FCK2 Fuel side check valve 

#2 

Series redundant check valves used to alleviate large mass flow of MMH into upstream helium section.

A test port is located between the two check valves to confirm functionality. 

OCK1 Oxidizer side check 

valve #1 Series redundant check valves used to alleviate large mass flow of MON3 into upstream helium section. 

OCK2 Oxidizer side check 

valve #2 

Series redundant check valves used to alleviate large mass flow of MON3 into upstream helium section.

A test port is located between the two check valves to confirm functionality. 

SV3 Service Valve #3 Manual valve used for functionality and leak checking of the upstream series redundant check valves.

Capped before launch. 

SV4 Service Valve #4 Manual valve used for functionality and leak checking of the upstream series redundant check valves.

Capped before launch. 

PVO1 MON3 Upstream 

Pyrovalve 

Used to positively isolate the upstream helium piping to avoid mixing of hypergolic fluids. MON3 vapors

can migrate through the soft seals of the check valve over time. 

PVF1 MMH Upstream 

Pyrovalve 

Used to positively isolate the upstream helium piping to avoid mixing of hypergolic fluids. MMH vapors

can migrate through the soft seals of the check valve over time. 

FVO1 MON3 Fill/Drain Valve Manual valve upstream of tank used for loading and offloading of MON3 into tank. Capped before launch.

FVO2 MON3 Fill/Drain Valve Manual valve upstream of tank used for loading and offloading of MON3 into tank. Capped before launch.

FVF1 MMH Fill/Drain Valve Manual valve downstream of tank for loading and offloading of MMH into tank. Capped before launch. 

FVF2 MMH Fill/Drain Valve Manual valve downstream of tank for loading and offloading of MMH into tank. Capped before launch. 

MON3 MON3 Tank Titanium tank with internal PMD used to store MON3 oxidizer. 

MMH MMH Tank Titanium tank with internal PMD used to store MMH fuel. 

VTO MON3 Flow Venturi Used to tailor downstream pressure drop to achieve proper mixture ratio at thruster injector. 

VTF MMH Flow Venturi Used to tailor downstream pressure drop to achieve proper mixture ratio at thruster injector. 

PVO1 MON3 Downstream 

Pyrovalve Used to positively isolate the downstream piping to avoid mixing of hypergolic fluids. 

PVF1 MMH Downstream 

Pyrovalve Used to positively isolate the upstream helium piping to avoid mixing of hypergolic fluids. 

SV5 Service Valve #5 Manual valve used for functionality testing of engine valves, system leak check, and to allow for positive

pad pressure to be established after testing. Capped before launch. 

SV6 Service Valve #6 Manual valve used for functionality testing of engine valves, system leak check, and to allow for positive

pad pressure to be established after testing. Capped before launch. 

FTO MON3 Side Filter Located downstream of pyrovalve to filter debris from pyro activation event. Select line size finer than that

of thruster valve inlets 

FTF MMH Side Filter Located downstream of pyrovalve to filter debris from pyro activation event. Select line size finer than that

of thruster valve inlets 

FJ1-6 Field Joints for 

propellant lines 

Due to the OML of the airplane, thruster bank has to be installed separately from the rest of the propulsion

hardware. 

TV1-6 Thruster Valves Thruster valves control the flow of MON3 and MMH into the thruster injector chamber. 

THR1-3 22N Thrusters Bank of three thruster will be independently operated based on airplane control parameters. 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 12

4.2 HIGH PRESSURE HELIUM STORAGE TANK 

The helium storage tank is a conventional, composite-over wrapped pressure vessel (COPV) with a MEOP of 

3000 psi and minimum burst factor of safety of 1.5. To minimize non-recurring engineering costs, the tank design 

is based on an existing Aircraft / Missile COPV with a maximum operating pressure of 10,000 psig and a mass of 

3300 grams, manufactured by Lincoln Composites. This existing tank will be modified by shortening the barrel 

length from 39.4 cm to 28.1 cm and by reducing the amount of graphite over-wrap applied to the tank liner, 

thereby reducing the mass to approximately 1200 grams and reducing the maximum operating pressure to about 

3000 psig. Further mass reduction can be achieved by modifying the inner liner thickness but would require 

additional developmental efforts. External dimensions of the helium tank are shown in Figure 6. The tank has an 

internal volume of 4100 cm3 and will hold 137 grams of helium at operating pressure. 

4.3 PROPELLANT TANKS 

Two propellant tanks are needed to hold the fuel (MMH) and oxidizer (MON3). These tanks will be one of the 

few components in the propulsion system that will be specifically designed and developed for the ARES mission. 

Additionally, the internal propellant management devices (PMDs) will be new designs specifically tailored to the 

mission. The tanks shells and PMDs are not based on any existing hardware, however, the design and 

development are based on conventional spacecraft propulsion system hardware. 

4.3.1 Tank Design 

The ARES fuel and oxidizer tanks are identical, cylindrical pressure vessels made of titanium, with a minimum 

wall thickness of 0.076 cm (0.03”), a maximum expected operating pressure (MEOP) of 300 psig, and minimum 

yield and burst factors-of-safety of 1.25 and 1.4, respectively. Tanks must be designed to endure vacuum loading, 

thereby requiring external stiffener rings. Internal volume constraints of the atmospheric flight system provide 

room for a tank diameter of 23.6 cm and a cylindrical barrel length of 32 cm. Each tank shell has a mass of about 

1.8 kg. While use of composite over-wrapped propellant tanks could offer mass savings, it was felt that such a 

development would be inconsistent with the risk posture desired for a Scout proposal. A volume of 

approximately 24 liters per tank would allow a propellant load of 17 kg of MMH and 28 kg of MON3 to be 

accommodated with approximately 10% ullage and 0.5% residual volume. 

 Aircraft / Missile COPV Vessel #220131-1 

 COTS Model: 3289 grams with 10,000 psig MEOP 

 Modified: 1200 grams with 3000 psig MEOP 

 Internal Diameter ~ 16 cm 

Figure 6. Helium Composite Over-wrapped Pressure Vessel 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 13

4.3.2 Propellant Management Device (PMD) Design 

Due to the unpredictable gravity and acceleration environment experienced by the airplane during flight, a 

Propellant Management Device (PMD) must be designed into the tanks to provide a continuous flow of oxidizer 

and fuel from the tanks to the thruster injection chambers. Continuous flow of liquid is required to maintain the 

appropriate fuel/oxidizer mixture ratio in the injection chamber and to avoid pockets of helium gas in the liquid 

lines that could be inadvertently heated causing a burst in the system. 

Several fuel extraction options were considered for the airplane, ranging from aerobatic airplane techniques to 

traditional spacecraft concepts5

, as described in Table 6. A traditional spacecraft screen and channel PMD was 

determined to be the most promising concept for the ARES airplane. 

The screen-and-channel type propellant management devices (PMDs) inside each propellant tank are used to 

ensure that propellant can be acquired under Mars gravity or any maneuvering or propulsive loads that may be 

encountered during airplane flight. A detailed design of the PMD concept resulted in a 6-channel totalcommunication PMD with 325x2300 stainless steel screens that will ensure gas-free propellant expulsion 

throughout the flight5

. Stainless steel screens in the oxidizer tank will pose additional challenges that must be 

considered in the system design. Since the oxidizer (MON3) will leach iron from stainless steel leading to the 

formation of ferric nitrate, there is the potential for this compound to precipitate out of the oxidizer and obstruct 

flow passages. Titanium does not have this incompatibility with MON3, however, the bubble point for titanium 

screens is less than the maximum expected hydrostatic pressure which will permit gas to enter the PMD channels, 

which is undesirable. Therefore, stainless steel screens must be used along with a ferric nitrate mitigation strategy 

(See Section 7.4). Propellant tank design with internal PMD is shown in Figure 7. 

Table 6. Fuel Extraction Concepts 

A dynamic propellant slosh analyses was conducted based on the expected acceleration environment of the 

airplane during extraction and flight5

. Along with the PMD channels, a center barrier acting as a baffle is 

included at the inside center of each tank to minimize sloshing during flight. A small screen window provides 

controlled movement of propellant between compartments. Residual propellant will exist in the tank when gas 

first enters the tank outlet. It is at this point when the propulsion system functionality is assumed to be over since 

gas in the liquid-side propellant lines could result in a burst in the system. The residual propellant in the tanks as 

well as propellant left in some of the PMD channels is estimated to be about 3-6% of the original fuel load, 

resulting in an expulsion efficiency ranging from 94-97%. 

Concept Description ARES Impact / Comments

Flop Tube Concept 

(Aerobatic Airplanes) 

A weighted tube forces the free end of the 

tube to move with varying acceleration so the 

tube remains wetted. 

Low expulsion efficiency and an impacting 

tube is undesirable. 

Header Tank Concept 

(Aerobatic Airplanes) 

Multiple tanks connected by check valves that 

reside above the engine and are gravity fed 

Concept is weight and space prohibitive for 

ARES 

Metallic Diaphragm Thin metal diaphragm pressurized on one side 

and “rolls” open as propellant drains 

Not suited for tanks with length to diameter 

> 2. Also not suitable for multiple loading 

and draining 

Bellows Provides reduction of internal volume of tank 

through expanding metal bellows 

Concept is usually heavy and new 

development will be required for titanium 

Elastomeric Diaphragm Internal membrane that is pressurized on one 

side to provide positive expulsion 

Not compatible with MON3 oxidizer. 

Screen and Channel 

PMD 

Concept has multiple enclosed flow paths that 

converge at tank exit. PMD relies on 

pressurized tank for expulsion 

Best option for ARES airplane, high 

expulsion efficiency 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept

NASA TM-2009-215700 

23.6 cm

32 cm (Barrel Length)

Propellant Out

Tank 

Mounting Hardware

PMD 

Channels

Helium Inlet

Solid Barr

Screen 

Window

Outer 

Figure 7. Propellant Tank and PMD Design 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 15

4.4 BIPROPELLANT THRUSTERS 

The ARES airplane requires 40N of thrust to maintain flight with a 50% margin to accommodate climbs, 

accelerations, and turbulent conditions. The first baseline airplane proposed in 2003 to the Mars Scout Program6

utilized a single 62N thruster (AJ10-220) manufactured by Aerojet, which is the only heritage thruster available in 

this range of thrust. With concerns about the availability of tooling for manufacturing this thruster, as well as the 

non-recurring costs that would be imposed on the program (since the last thruster was manufactured over 10 years 

ago), other options using more readily available thrusters were considered in a detailed trade study. This trade 

study was performed at NASA Langley Research Center under the Planetary Airplane Risk Reduction project in 

2005 to examine the option of using multiple, smaller thrusters in place of a single large thruster with the goal to 

reduce overall cost, schedule, and technical risk7

. The smaller thrusters that were considered included an array of 

various 5-lbf (22N) common reaction control thruster used on many satellites. 

4.4.1 22N Satellite Reaction Control Thruster Options 

A listing of readily available “off-the-shelf” 22N thrusters is shown in Table 77

. Thousands of these types of 

thrusters have been built and flown on numerous programs, including MILSTAR and Intelsat VI. The AMPACIn-Space Propulsion 22N Columbium thruster was chosen as the baseline thruster for the three-thruster system. 

This thruster was qualified under the Intelsat VI satellite program and over 1000 units have flown to date. An 

alternate thruster chosen was the Aerojet R-6C, which is an almost identical engine with similar production 

numbers and similar heritage. 

Table 7. Mars Airplane Commercial 22N Thruster Options 

Vendor Aerojet Aerojet 

AMPACISP (ARC) AMPACISP (ARC) Astrium - 

EADS 

Model R-6C R-6D 5 lbf 5 lbf S22-02 

Chamber Material C103 C103 C103 Pt/Rh Pt/Rh 

Status COTS Item Existing 

Design COTS Item Existing Design 

Existing 

Design 

Isp [sec] 293 294 291 300 285 

Area Ratio 150 150 150 150 150 

Mixture Ratio (nom) 1.65 1.65 1.65 1.6 1.65 

Feed Press [psia] 220 220 220 220 290 

Chamber Press [psia] 116 125 125 125 125 

Valve Model Mg 51-330 Mg 51-136 Mg 51-178 Mg 51-178 Mg 51-178 

Thruster Mass [kg] 0.54 0.4 0.45 0.47 0.49 

Valve Mass [kg] 0.14 0.27 0.19 0.19 0.19 

Total Mass [kg] 0.68 0.67 0.64 0.66 0.68 

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

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

Number Flown >600 0 >1000 0 0 

Heritage 

MILSTAR 

GOES ISRO 

Insat 2B 

Evolve from 

R-6C 

Thruster 

Telecom 

Intelsat VI 

Evolve 

from C103 

5lbf 

Thruster 

Astrium 10N 

Thruster 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 16

4.4.2 AMPAC-ISP 22-N Columbium Thruster 

AMPAC-In-Space Propulsion (formerly Bell Aerospace and Atlantic Research Corporation) developed a family 

of 5 lbf bipropellant engines using different valve configurations ranging from dual seat solenoid valves to single 

seat torque motor valves. These engines have demonstrated extremely long life of 100,000 seconds of steady 

state firing with high specific impulse ranging from 286-295s8

. These engines have also demonstrated wide 

flexibility of mission duty cycle operation by firing in any duty cycle ranging from 1-100% and accumulating 

over 325,000 firing pulses on multiple units with a minimum of 400,000 pulses on one unit. The engine uses a 

unique single element injector made of a coated columbium alloy orifice plate with titanium alloy feed tubes and 

a titanium support structure. The thrust chamber is made of columbium alloy and coated with R512E silicide 

internally and externally to protect from combustion products and to provide a surface with high emissivity for 

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

leakage. The valve is bolted to the injector using metal seals. The thrusters were designed, built, and qualified for 

the Intelsat VI series spacecraft. A diagram of the 22N baseline thruster used in this study is shown in Figure 8. 

Exit diameter is 5.4 cm and the total length of the thruster and attached valve, including shown inlet tubing, is 

approximately 27.3 cm. The valve is shown with the Moog series redundant solenoid valve model Mg 51-178. A 

summary of performance characteristics of the 22N thruster is shown in Table 8. 

Table 8. AMPAC-ISP 22N Thruster Characteristics 

Fuel / Oxidizer MMH / MON3 

Thrust 22N 

Specific Impulse 293 s 

Area Ratio 150:1 

Inlet Pressure Range 94 – 400 psi 

Feed Temperature 70°F Nominal 

105°F Max SS Operation 

160°F Pulsed Mode SS 

Chamber Press (nom) 125 psia 

Max Impulse 2,668,800 N-s 

Min Impulse Bit 0.045 N-s 

Nozzle Exit Diameter 5.4 cm 

End to End Unit Length 27.3 cm (includes valve) 

Mass 0.8 kg 

Valves

Combustion Chamber

Nozzle

Valve Mounting Flange

Injector Flange

10.75 in

Figure 8. AMPAC-ISP 22N Thruster Details 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 17

4.4.3 Thruster Mounting and Orientation 

Airplane thrusters are mounted in an in-line configuration arrangement of three 22 N thrusters. Right hand and 

left hand thrusters were rotated about c.g. location 4.5° to allow for adequate clearance of the valves and 

propellant lines. Thrusters are tilted 5.88° from the horizontal reference to direct thrust vector through the 

airplane c.g., shown in Figure 9. 

Figure 9. Triple Thruster Orientation and Alignment 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 18

The aft and forward brackets are titanium alloy and are fabricated to place thrusters in correct orientation. The aft 

brackets are bolted to the valve flange using existing mounting holes (see Figure 10). The thruster valve flange 

was chosen for bracket interface, because it only reaches a maximum temperature of 200°F at steady-state 

operation. The thrusters are also supported at the forward bracket and captured by titanium retaining rings. Once 

the retaining rings are placed into existing groove in valves, they are fastened to the forward bracket. The forward 

and aft brackets are bolted directly to composite airframe support structure. The design approach was to enable 

the ability to readily remove and reinstall the thrusters from the airframe during integration and testing. 

4.5 THRUSTER VALVES 

Thruster valve chosen for the ARES airplane propulsion system is the Moog 51-178 series redundant solenoid 

driven valve, shown mounted to the thruster in Figure 8. Each thruster has two independent valves to control the 

flow of fuel and oxidizer to the combustion chamber. These valves provide on/off flow capability by energizing 

solenoids from voltage signals provided by the propulsion valve driver card. The valve consists of stainless steel 

and Teflon construction with a total mass of 0.19 kg and consumes 26 watts of power per valve when energized. 

The valve is normally closed with the Teflon® poppet and armature held against the valve seat by a suspension 

spring. When the valve coils are energized, the electromagnetic force overcomes the spring pre-load to lift the 

armature/poppet from the seat, providing a flow passage for the propellant. When the coil is de-energized, the 

spring forces the armature/poppet closed. 

Since this valve is driven by electromagnetic forces, and the ARES science payload includes a magnetometry 

experiment, magnetic cleanliness of this valve is a key consideration. A magnetic screening of this valve model 

was completed at NASA GSFC in June 2006 to measure the non-operating static field and the operating dynamic 

field. The test demonstrated a small static field (30 nT at 1 m; 0.2 nT at 2.6 m) and a small dynamic field (60 nT 

at 1m; 0.4 nT at 2.6 m) per coil when actuated. With a 2.6 m distance between the magnetometers and the 

thruster valves, and a 20 nT static and 5 nT dynamic magnetic cleanliness requirement, reductions of the thruster 

valve magnetic fields are desired but not required. Since these valves have series redundant solenoids, the valve 

can be successfully operated by reversing the current through one of the coils, which will result in cancellation of 

the magnetic field strength. This was successfully demonstrated in the screening resulting in a negligible reading 

of the dynamic field. Operating the valve in this manner may require a partial re-qualification to determine if any 

adverse effects arise. 

Figure 10. Triple 22N In-Line Thruster Mounting and Assembly 

Valve Flange 

AMPAC-ISP 22N Thruster

Airframe Interface 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 19

4.6 PROPELLANT ISOLATION AND FLOW CONTROL HARDWARE 

Since the MMH and MON3 propellants are hypergolic, it is critical to ensure their isolation to avoid inadvertent 

ignition in any part of the feed system. Normally-closed, pyrotechnically-actuated valves (“pyro valves”) (PVF1 

and PVO1 in Figure 5) are used to isolate the high pressure helium supply from both propellant tanks and to 

prevent propellant vapor migration prior to system operation. Redundant series check valves are used to ensure 

that there is no vapor or liquid transfer and mixing of the hypergolic propellants during the operation of the 

system. The possibility exists, during the long periods of inactivity, for fuel or oxidizer to seep through the check 

valves and condense in upstream sections of the feed system. To minimize trapping any condensed propellant in 

dead-end upstream sections, the series-redundant check valves for the fuel and oxidizer are located upstream of 

the pyro valves, which provide the long duration isolation function. Downstream of the propellant tanks, 

normally closed pyro valves PVF2 and PVO2 are also used to provide a redundant seal against external leakage of 

the propellant through the single-seat thruster valves. In addition to increasing reliability, such redundant seals 

are required for safety during ground handling. All manual service valves also incorporate redundant seals. 

4.6.1 Pyrotechnic-actuated Valves (Pyrovalves) 

ARES uses normally-closed pyrotechnic valves to positively isolate portions of the propulsion subsystem to avoid 

mixing of the hypergolic propellants in areas other than the combustion chamber. A total of four Conax Model 

1832-205 duel initiator pyrovalves (Figure 11) and one Conax Model 1832-192 high pressure, single initiator 

pyrovalve provide the propellant isolation function. The valves are constructed of stainless steel with a 500 psig 

and 9400 psig operating pressure respectively, and a 0.02 second response time once initiated. When a current is 

received by the firing circuit, the initiator ignites a charge within the valve body which forces a ram downward 

removing a metal closure and opening the flow path. For redundancy, the pyrovalve consists of two initiators 

which will be fired sequentially to assure proper functioning of the valve. 

The propellant tanks are isolated by two gas-side pyrovalves located just upstream (PVO1 and PVF1 from), and 

two liquid-side pyrovalves located directly downstream (PVO2 and PVF2). An additional high pressure 

pyrovalve (PVHE) is used to isolate the high pressure helium tank and to protect the helium pressure regulator 

(PRHE) from long term exposure to high pressure during the long cruise segment to Mars. 

4.6.2 Line Filters 

Small particulates in the flow stream can block narrow passages in areas such as the thruster injector chambers. 

Debris resulting from the firing of pyrovalves are a major source of such particulates, so line filters are placed just 

downstream the liquid side pyrovalves (PVO2, PVF2). ARES uses the Vacco 3/8” Low Pressure Filter Assembly 

F1D10558-01, which is a titanium etched disc, 25 micron filter with a 400 psig operating pressure and a mass of 

0.08 kg (Figure 12). Similarly, the helium regulator is susceptible to small particles blocking the passage ways, 

so a high-pressure gas filter is placed downstream of the helium-tank pyrovalve (PVHE) to filter debris after 

pyrovalve activation. The high pressure filter is Vacco Model F1D10286-01, which is also an etched disc filter 

but has an operating pressure rating of 4200 psig (Figure 13). 

4.6.3 Check Valves 

Check valves are used in the gas-side of the propulsion system to alleviate large mass flow of MMH or MON3 

into the helium section, once the pyrovalves have been activated. Since the propellant is hypergolic, any mixing 

of the fuel and oxidizer in the gas-side would result in combustion and most likely a pressure spike large enough 

to burst the tubing. Series redundant check valves were selected over parallel redundant or “quad” valves due to 

the single string fault policy of the airplane systems. The check valves in the propulsion system are Vacco 3/8” 

low pressure check valve Model V1D10856-02 (Figure 14), which is a titanium welded unit used on the Mars 

Ascent program. The valve has a mass of only 20 grams and an operating pressure of 550 psi. 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 20

4.6.4 Helium Regulator 

A regulator is used to reduce the pressure from 3000 psi at the helium tank to the propellant tank operating 

pressure of approximately 300 psi. Several options exist for a flight qualified helium regulator from both space 

applications and defense programs such as the THAAD5. ARES has baselined the Vacco Model 64720-1 single 

stage regulator (Figure 15). Over 400 of these regulators have been flown in satellite applications over the past 18 

years. The regulator has an inlet pressure range from 400 to 4200 psig and an outlet pressure range from 224 to 

4200 psig. The unit has a mass of 844 grams and can handle helium flow up to 8.5 scfm of GHe. 

4.6.5 Flow Venturis 

Flow venturis are used on the liquid side of both the MON3 and MMH system to limit the propellant flow and 

tailor the feed system pressure drop to achieve the proper mixture ratio of fuel and oxidizer in the engine 

combustion chamber. Venturis are custom built depending on the final propellant line pressure drop 

characteristics. Flow Systems Inc. will provide the venturis to the ARES project. 

4.6.6 Service Valves 

The ARES propulsion system has a total of 11 services valves. One service valve (FVHE) is located on the 

helium tank and used for helium pressurizing. Four valves are used for propellant tank filling and venting, one 

on the gas side of each tank (PVO1 and PVF1) and one on the liquid side of each tank (PVO2 and FVF2). Two 

additional service valves located downstream of the liquid-side pyrovalves provide the ability to vent the volume 

between the pyrovalves and the thruster valves. Additional service valves are used for functionality checks of the 

helium regulator and check valves. Check valves are independently verified using both a service valve and a 

sealed test port. The high-pressure service valves are manufactured by Moog, part number 50E740 with flight 

heritage based on the Milstar program. These valves are manually actuated, titanium construction with a proof 

pressure of 3265 psig and a mass of 0.15 kg each (Figure 16). 

4.6.7 Fluid Field Joints 

It is desirable to have an all-welded propulsion system since field joints have a history of being unreliable, 

causing extra testing and care to prevent leakage or damage. The propulsion system is installed into the airplane 

through the upper hatch located on the fuselage. However, the three thrusters have to protrude from the bottom 

outer-mold-line of the airplane, making it impossible to have a completely welded system. Consequently, six 

field joints connecting the three thruster/valve assemblies to the propellant tubing are required so the thrusters can 

be mounted from underneath the airplane. 

4.7 COMMAND AND INSTRUMENTATION 

Propulsion system instrumentation consists of pressure and temperature sensors on the helium tank and on both 

the gas and liquid sides of the propellant tanks. Analog signals from the pressure and temperature sensors are fed 

into the State of Health Monitoring & Attitude Control (SMACI) module which is part of the ARES airplane 

command and data handling subsystem. A subset of these signals are split and routed externally through the 

aeroshell and into the spacecraft C&DH computer for state of health monitoring of the propulsion system during 

the cruise stage of the mission. 

Commanding of propulsion system occurs during the initial pressurization of the propellant tanks and during 

thruster firing. During initial activation, commands originating from the airplane computer power switching 

boards act to arm and then fire the five pyrovalves in various locations on the propulsion system using 28VDC 

discrete signals. Commanding of the thruster valves is accomplished by a 28VDC circuit controlled by the valve 

driver card. When the valve receives the 28VDC, the solenoids open a flow passage for the propellants. The 

valve driver card communicates with the flight computer via RS-422 interface. Each thruster has independent 

fuel and oxidizer control valves, so with the triple thruster system, there are 6 channels required to operate the 

thrusters. Further development is required for a new or modified valve-driver card designed specifically for 

compatibility with the ARES architecture. 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 21

4.7.1 Temperature Sensors 

The temperature sensors used to monitor various locations on the propulsion system are solid state sensors 

powered by 28VDC and provides an output linear to temperature of 1μA/K. Temperature measurement range is 

from -55 to 150°C with a repeatability of ±0.1°C and a long-term drift of ±0.1°C/month. The baseline component 

identified for the ARES mission is the AD590 Solid State Temperature Sensor manufactured by Analog Devices. 

4.7.2 Pressure Transducer 

Pressures in the helium and propellants tanks are monitored by the flight qualified GP:50 Pressure transducer 

Model 7200 (Figure 17). This device has significant flight heritage including Lockheed Martin Evolved 

Expendable Launch Vehicle (EELV) and Space Shuttle Upgrades. The transducer receives 28VDC excitation and 

provides an analog output signal to the SMACI board. Each pressure transducer has a mass of 142 grams. 

Figure 11. Pyrovalve - Conax Model 1832-205 Figure 12. Liquid Filter – Vacco Model F1D10558-01 

Figure 13. Gas-Side Filter – Vacco Model F1D10286-01 Figure 14. Check Valves – Vacco Model V1D10856-02 

Figure 15. Helium Regulator - Vacco Model 64720-1 Figure 16. Service Valves – Moog 50E740 

Figure 17. Pressure Transducer – GP:50 Model 7200 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 22

4.8 HARDWARE INTEGRATION 

Propulsion hardware must be packaged inside the limited volume of the airplane fuselage, along with all other 

hardware such as science instrumentation, avionics, communication components, power modules and support 

structure. Fuel and oxidizer tanks are each supported by two brackets with two mounting interfaces for each 

bracket that attach directly to the airplane structure. The helium tank is attached directly to the airframe through 

similar brackets. Propellant isolation and flow control hardware will be mounted to two separate honeycomb 

plates; the Propellant Isolation Assembly (PIA) Plate will consist of fuel and oxidizer pyrovalves, fill valves and 

service valves, and the Pressurant Control Assembly (PCA) Plate will consist of the helium tank pyrovalve, 

helium pressure transducer, check valves, and service valves. PIA and PCA plates will be mounted to the airframe 

structure using Torlon® thermal standoffs. Propellant and helium lines are supported by tube support assemblies 

which are bonded directly to the PIA and PCA plates and to the airframe structure. The propulsion hardware 

integrated within the airframe structure is shown in Figure 18. 

Figure 18. Propulsion System / Airframe Integration 

Pressurant Control Assembly 

(PCA) Plate 

Propellant Isolation Assembly 

(PIA) Plate 

Helium Tank 

MMH Tank 

Thrusters 

MON3 Tank 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 23

5.0 MASTER EQUIPMENT LIST 

The master equipment list of all propulsion system components, number of flight units, mass and power, is shown 

in Table 9. The propellant tanks (with PMDs) are the most massive components at 5275 grams total, making up 

about 1/3 of the total dry-mass of the system. The thruster valves use 26 watts each for a total of 156 watts of 

power required. Thruster valve power will not be continuous but rather cycled on and off when necessary to 

maintain flight. Additionally, not all thrusters will be firing simultaneously (see Section 6.4). Total CBE dry 

mass of the propulsion system is 16.1 kg with a peak power requirement of 167 watts. 

 

Table 9. Propulsion System - Master Equipment List 

Component # Flight Units Mass [grams] Power [W] 

Bipropellant Thrusters 3 1260 - 

Thruster Valves 6 1140 156.0 

Propellant Tank Shell & Stiffeners 2 3228 - 

Propellant Management Device (PMD) 2 2046 - 

Helium Pressurant Tank 1 1200 - 

Check Valves 4 80 - 

Helium Regulator 1 550 - 

Liquid Filter (MMH & MON3) 2 360 - 

HP Service Valves 11 1650 - 

GHe HP Filter 1 114 - 

NC Pyro. Valves 5 1020 5.0 

High Pressure Transducer 1 142 0.6 

Low Pressure Transducer 4 568 2.4 

Flow Venturis 2 80 - 

Propellant Isolation Assembly (PIA) Plate 1 100 - 

Pressurant Control Assembly (PCA) Plate 1 100 - 

Thruster Valve Controller 1 400 3.0 

Thruster Valve Controller Chassis 1 500 - 

Lines, Tubing, Fittings 1 739 - 

Tank Support Brackets 1 400 - 

Residual Propellant & Ullage 1 434 - 

Residual Helium Gas 1 21 - 

 

TOTALS 53 16132 167 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 24

6.0 PROPULSION SUBSYSTEM PERFORMANCE 

ARES has analyzed the performance of the propulsion subsystem through flight simulations and by characterizing 

its impact on other subsystems and airplane structures. The thermal environment created by the firing of the 

thrusters was examined as well as the potential impact of plumes on airplane structures and control surfaces. 

6.1 THRUSTER FIRING – THERMAL IMPACT 

During operation, the thrusters reach temperatures significantly greater than the expected -60°C Mars ambient 

temperature. For continuous firing at room temperature, the combustion chambers of the thrusters reach 

approximately 1050°C in 45 seconds (Figure 19)9

. This creates a significant temperature difference between the 

thrusters and outer fuselage, which will cause a large radiative heat transfer between the two components. The 

airplane fuselage is a composite structure consisting of an aluminum honeycomb core surrounded by face sheets 

of carbon fabric. Adhesive holding the carbon face sheets to the aluminum honeycomb will begin to break down 

at approximately 250°C. The valve flange, where the thruster bank will be mounted to the airframe, maintains a 

temperature near 100°C during steady state firing. During flight, two thrusters will be continuously firing. These 

two thruster will reach an equilibrium temperature profile similar to the one provided by the test firing, however it 

will be slightly cooler due to the much lower ambient temperature on Mars and forced convection from the 

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

Based on previous thermal analyses of the three thruster system, which included heat transfer from both the 

thruster nozzles and the plumes, the thruster support structure will experience a maximum temperature of 139°C7

. 

However, an analysis of the lower fuselage without a heat shield (Figure 20) shows temperatures reaching 284°C, 

which exceeds the critical temperature (125°C) of the face sheet to honeycomb film adhesive. Consequently, a 

thin polished aluminum heat shield will be required to reflect the heat away, which will result in significantly 

lower fuselage temperatures, shown in (Figure 21). Plume temperature for analyses purposes was assumed to be 

700°C. Depending on the actual plume temperature, two different heat shields can be used. A large heat shield 

will be required for high plume temperatures, however for lower plume temperatures, a combination of a smaller 

Chamber Max - [°C]

Nozzle - [°C]

Valve Flange - [°C]

0

100

200

300

400

500

600

700

800

900

1000

1100

0 20 40 60 80 100 120

Time [sec]

Temperature [°C]

Figure 19. Hot Fire Test Data for AMPAC-ISP 22N Thruster9. 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 25

heat shield with low emissivity paint will keep the fuselage temperatures closer to the ambient temperature. 

Further analyses and testing are required to obtain accurate plume temperature profiles for the expected airplane 

operating conditions. 

Figure 20. Temperature profile for lower fuselage without 

heat shield. 

(Max Temperature = 284°C). 

Figure 21. Temperature profile of lower fuselage with heat shield. 

(Max T = -6.17°C). 

Temperatures in degrees Celsius 

Temperatures in degrees Celsius 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 26

6.2 THRUSTER PLUMES 

Computational fluid dynamics simulations were used to evaluate possible plume interactions and aeropropulsive 

behavior of the airplane in the Mars atmosphere during coast and during thruster firing. The simulation modeled 

the 3D thruster system in a representative free-stream environment (using PAB3d and 6-DOF LaSRS). Figure 23 

shows a slice through the centerline plane of spanwise symmetry, with the center thruster plume visible in the 

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

flow. Figure 22 shows a slice through the CFD simulation from the rear of the airplane. The individual thruster 

plumes (shown in the lower picture) are well isolated and maintain angular separation, with no major interactions 

apparent. Sufficient clearance exists between the tail and plumes to avoid any adverse affects of flight controls or 

performance. 

Figure 22. Rear view of Airplane showing airplane in coast (above) and with thrusters firing (below) 

Figure 23. Side view of Airplane showing airplane in coast (above) and with thrusters firing (below). 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 27

6.3 THRUSTER FIRING LOGIC 

To control the thrust modulation in a pressure-blowdown propulsion system, on/off duty cycle adjustments are 

continuously generated by the control system. The thruster control law uses two key elements; a thrust period and 

a duty cycle. The two elements define how long the thruster is off and on (period * duty cycle). Once a thrust 

period has ended, the duty cycle is modified to reflect the error between the commanded Mach number and the 

average Mach number for that period. The firing logic for the thrusters is divided into two separate duty cycles. 

The primary duty cycle governs the firing of two thrusters simultaneously to provide the needed thrust to 

overcome drag while flying under normal conditions. The secondary duty cycle governs the firing of the third 

thruster and is invoked during instances where airplane acceleration or climbing is needed or under turbulent 

conditions when more than 44N of nominal thrust is required to maintain flight. 

6.4 AIRPLANE FLIGHT SIMULATION 

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

Standard Real-Time Simulation application framework10. Control logic was incorporated to hold the airplane’s 

speed while flying a specific mission profile. Two atmospheric conditions were simulated to bound the upper and 

lower expected atmospheric conditions; one with no turbulence, and the other with severe atmospheric turbulence. 

Thruster duty cycle for the conditions of no turbulence is shown in Figure 24. The primary duty cycle is shown 

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

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

shows that the third thruster is only slightly used during the initial stages of flight, when the airplane is the 

heaviest, and stays off during the entire remainder of flight. During severe turbulence, the data in Figure 25 

shows that the two primary thrusters are fully firing at 100% duty cycle for a significant portion of the flight, and 

that the third thruster, or secondary duty cycle, is sparsely invoked throughout the flight and remains off for most 

of the flight. 

6.5 AIRPLANE FLIGHT DURATION AND RANGE 

The results of the simulation show that the airplane flying in no-turbulence conditions will achieve a range of 

about 543 km and duration of 68 minutes on a propellant load 40 kg. In severe turbulence conditions for the same 

fuel load, the flight range is reduced to about 536 km and the duration reduced to 66 minutes (Table 10). This 

averages to roughly 13.5 km/kg of propellant for the complete flight mission. With fuel consumption of 13.5 

km/kg, the baseline mission using 45 kg of propellant will achieve a range of ~608 km and a duration of about 70 

minutes. 

Table 10. Propulsion System Simulated 

Performance Summary 

Parameter Value 

Total AFS Mass 145 kg 

Propulsion Mass 18.06 – 18.12 kg 

Propellant Load 40 kg 

Fuel 15 kg - MMH 

Oxidizer 25 kg - MON3 

Thrust Three - 22 N 

Isp ~291 s 

No Turbulence

Flight Duration 67.63 min 

Range 543 km 

Avg Fuel Usage 13.6 km/kg 

Severe Turbulence

Flight Duration 66.22 min 

Range 536 km 

Avg Fuel Usage 13.4 km/kg 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 28

0%

20%

40%

60%

80%

100%

0%

20%

40%

60%

80%

100%

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

0%

20%

40%

60%

80%

100%

0%

20%

40%

60%

80%

100%

0 500 1000 1500 2000 2500 3000 3500 4000

`

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

Figure 24. Thruster Duty Cycles for No Turbulence Conditions 

0%

20%

40%

60%

80%

100%

0%

20%

40%

60%

80%

100%

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

0%

20%

40%

60%

80%

100%

0%

20%

40%

60%

80%

100%

0 500 1000 1500 2000 2500 3000 3500 4000

`

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

Figure 25. Thruster Duty Cycles for Severe Turbulence Conditions 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 29

7.0 OPERATIONAL PROCEDURES 

7.1 PROPELLANT LOADING PROCEDURES 

In preparation for airplane closeout prior to launch, a series of steps must be taken for propulsion system checkout 

and propellant preparation and handling. Since MMH and MON3 are hypergolic, the loading of each into the 

airplane must be separated by sufficient time to ensure the system is leak-proof, to avoid inadvertent mixing of 

propellant vapors or liquid. Table 11 outlines a typical bipropellant loading schedule leading to final airplane 

closeout. 

Table 11. Airplane Propellant Loading Procedure 

Activity Schedule

Ground support equipment setup, staging and calibration 5 days 

Final thruster functional tests (leak test and trace) 

Transducer calibration and Blanket Pressure Load of thruster tubing 2 days 

Proof test Helium tank to 1.1 MEOP to ensure integrity 1 day 

Pressurize Helium tank to Flight Pressure 

Final High Pressure leak check of Service Valves 1⁄2 - 1 day 

MMH Loading: 

Sample MMH for testing and resample if necessary 

Measure flight quantity of MMH and load into fuel tank 

Closeout and Final Leak Check of Fuel Side 

2 days 

1⁄2 - 1 day 

1⁄2 day 

1⁄2 day 

MON3 Loading: 

Sample MON3 for testing and resample if necessary 

Measure flight quantity of MON3 and load into oxidizer tank 

Closeout and Final Leak Check of Oxidizer Side 

2 days 

1⁄2 - 1 day 

1⁄2 day 

1⁄2 day 

Install NASA Standard Initiators (NSI) onto pyrovalves 1 day 

Final Closeout of Airplane 1 day 

 

Total Propellant Loading Schedule 15 days 

7.2 SYSTEM ACTIVATION SEQUENCE 

The propulsion system is pressurized and activated at the end of the Entry, Descent, and Deployment sequence. 

Based on altitude at the completion of the pullout maneuver, the airplane determines when the propulsion system 

should be initiated. If it is determined that the altitude is higher than 2 km AGL, then the airplane continues to 

glide until reaching the 2 km threshold, at which time the propulsion activation sequence begins. The propulsion 

system is activated by first cycling the engine valves to evacuate the engine manifold and the lines downstream of 

the liquid-side pyrovalves (See Figure 5). Next, the liquid-side pyro-valves are fired followed by the gas-side 

pyrovalves. The propulsion system is pressurized by the firing of the high-pressure pyrovalve located on the 

helium tank, providing pressure to the propellant tanks, forcing liquid downstream to the thruster valves. Thrust 

is then controlled by actuating the individual valves on each thruster. This sequence prevents the propellant tanks 

from experiencing “slam-start” conditions when the gas-side pyrovalves are fired. The pressure regulator will 

experience a slam-start when the helium tank pyrovalve is fired, so the regulator will have to be tested and 

qualified for this operational condition. 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 30

7.3 COMPONENT QUALIFICATION REQUIREMENTS 

Due to the cost and schedule constraints imposed by Mars Scout missions, hardware with significant flight 

heritage is preferred to minimize efforts in preparing such hardware for flight readiness. The only components of 

the propulsion system that require original design, development, and qualification are the propellant tanks and 

PMDs. The helium tank is based on a modified existing design and will undergo a delta-qualification. Thrusters 

and the helium regulator will also undergo a delta-qualification to characterize performance based on the unique 

operating conditions of the ARES airplane. The thruster will also be hot-fire tested to determine specific 

performance profiles based on expected ARES mixture ratios, duty cycles, and operating pressures. All 

propulsion hardware will endure a series of pressure tests, flow and leak checks, as well as environmental tests 

including vibration and shock. The component qualification matrix is shown in Table 12. 

Table 12. Component Qualification Requirements 

Dev Qual ΔQual Proof Burst Flow Leak Vib Shock ATP Other 

Propellant tanks 

PMDs Bubble Point 

GHe Tank 

Thrusters Hot Fire 

Helium Regulator Set pt, slam start, lockup 

Pres. Transducers Calibration 

Service Valves 

Check Valves Lot acceptance test 

Pyro valves Lot acceptance test 

Dev – Full development Required 

Qual – Qualification Required 

ΔQual - Certain design changes require re-qualification 

Proof – Test components to 1.1 MEOP 

Burst – Test components to structural failure 

Flow – Test flowrates of propellants 

Leak – Leak check fittings and components 

Vib – Expose components to expected vibration levels and check performance 

Shock – Expose components to expected g-levels and check performance 

ATP – Acceptance testing
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 31

7.4 FERRIC NITRATE MITIGATION 

The oxidizer, MON3 (nitrogen tetroxide with mixed oxides of nitrogen), reacts with stainless steel causing low 

levels of dissolved iron into the oxidizer. This reacts to form an iron nitrate complex, which can become quickly 

saturated, and solid particles may precipitate from the liquid during abrupt temperature and pressure changes11. 

Sharp bends in propellant tubing or pressure drops caused by orifices or valve seats are areas prone to ferric 

nitrate precipitation. This precipitate can restrict flow by accumulating in components with small openings or 

orifices or may damage components containing moving parts with close tolerances, such as check valves or 

thruster valves. 

The amount of precipitate that can form in a given location is proportional to the amount of oxidizer throughput 

past that point. Since the ARES mission is of short duration, using only 28 kg of MON3, compared to typically 

100s of kilograms used during missions where Ferric Nitrate precipitation becomes a concern. Characterization 

and removal of ferric nitrate in the Space Shuttle primary reaction control system was addressed by Johnson 

Space center and White Sands Test Facility12. Recommendations from this publication and how they impact the 

ARES project are shown in Table 135

. 

Table 13. ARES Ferric Nitrate Mitigation Recommendations 

Shuttle PRCS Recommendation ARES Impact 

Maximize the concentration of NO in oxidizer. NO has a fairly dramatic 

effect on the solubility of metal nitrate in oxidizer, so the NO concentration 

could be increased to a level that would ensure the oxidizer retained 

excellent hypergolic properties yet would allow metal nitrate concentrations 

to be as high as 5 ppm (or even higher) without precipitation under usage 

conditions. 

MON-3 Is Proposed for ARES, 

Relatively High NO concentration 

Use “point-of-use” metal nitrate filters to minimize the metal nitrate in 

nitrogen tetroxide during loading at the launch facility. The molecular sieve, 

developed at WSTF and used by the oxidizer supplier, is typically 0.2 ppm 

or less in metal nitrate concentration at tank truck loading. Ensuring that a 

minimum metal nitrate concentration is initially loaded into the orbiters will 

increase the possibility for avoiding the saturation point. 

ARES will use highly filtered 

N2O4 (MON3) and sample to 

ensure a low initial concentration 

ARES will maintain tank temps 

above 0°C for Cruise. ~ 1 day 

pre-entry increase tank temps to 

above 27°C. Raises the 

precipitation point 

Ensure that RCS valve temperatures do not drop below the saturation 

temperature (around 27°C) during all operations including orbiter ferrying. 

Allowing the temperature of the oxidizer to drop below the saturation point 

should be avoided under all circumstances. WSTF testing has shown that 

the precipitation point of metal nitrate is proportional to temperature, 

dropping to below the 1.5 ppm metal nitrate level for specification grade 

oxidizer at around 60°F 

Heaters can be used to maintain 

valve temperatures before ARES 

propulsion operation begins. 

Better understanding of the basic (kinetic) chemical reactions occurring to 

produce metal nitrate. Water intrusion with subsequent nitric acid formation 

is thought to be the major cause for corrosion, yet no proof or rate of 

reaction data has been established. Minimizing water intrusion (by using the 

universal throat plug and better handling techniques) and leaving systems 

wetted with nitrogen tetroxide (not water!) should help reduce corrosion. 

An all-welded system leak tight to 

10-6 scc/hr GHe external leak rate 

will preclude water entry. 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 32

8.0 ACKNOWLEDGEMENTS 

The author would like to acknowledge and thank the following individuals for their contributions to the 

maturation of the Mars airplane propulsion system concept; Anthony J. Colozza (Analex Corporation, NASA 

GRC), Carl S. Guernsey (NASA JPL), and Henry S. Wright (NASA LaRC) for initial concept development and 

baseline propulsion system design. Mike McGee, Greg McAllister, Tim Fischer, and Carey Parish (Lockheed 

Martin) for in-depth propulsion system trade studies, propellant tank designs, PMD design, and overall system 

maturation. P. Sean Kenney (NASA LaRC) for flight simulation and propulsion system performance, Steven S. 

Gayle and David A. Paddock (NASA LaRC) for mechanical design and thruster mounting concept, Joseph 

Gasbarre and Salvatore Scola (NASA LaRC) for thruster thermal analyses, and Craig A. Hunter (NASA LaRC) 

for thruster plume CFD and aero-propulsion simulation. 

9.0 ACRONYMS 

AFS Atmospheric Flight System 

AGL Above Ground Level 

ARES Aerial Regional-scale Environmental Survey 

C&DH Command and Data Handling 

CFD Computational Fluid Dynamics 

COPV Composite Over-wrapped Pressure Vessel 

EELV Evolved Expendable Launch Vehicle 

GHe Gaseous Helium 

GRC Glenn Research Center 

GSFC Goddard Space Flight Center 

JPL Jet Propulsion Laboratory 

LaRC Langley Research Center 

LaSRS Langley Standard Real-Time Simulation 

MEOP Maximum Expected Operating Pressure 

MGS Mars Global Surveyor 

MMH Monomethyl Hydrazine 

MON3 Nitrogen Tetroxide with 3% Mixed Oxides of Nitrogen 

MRO Mars Reconnaissance Orbiter 

NO Mixed Oxide of Nitrogen 

PAB3d NASA Langley 3-dimensional CFD code 

PCA Propulsion Control Assembly 

PIA Propulsion Isolation Assembly 

PMD Propellant Management Device 

PPM Parts per million 

SMACI State of Health Monitoring & Attitude Control 

STA Science Target Area 

TCM Trajectory Correction Maneuver 

TDB To Be Determined 

THAAD Theater High-Altitude Area Defense\ 

TRL Technology Readiness Level 

WSTF White Sands Test Facility 
Design of a Mars Airplane Propulsion System for the Aerial Regional-scale Environmental Survey (ARES) Mission Concept 

NASA/TM-2009-215700 33

10.0 REFERENCES 

1. “ARES, an Aerial Regional-scale Environmental Survey of Mars,” proposed in response to the NASA Office 

of Space Science Announcement of Opportunity NNH06ZDA002O for Mars Scout 2006 Missions, July 21, 

2006. 

2. “Comparison of Mars Aircraft Propulsion Systems,” Colozza, A.J., NASA CR-2003-212350, May 2003. 

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

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

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

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

2004. 

4. Reed, R.D. (1978), “High-Flying Mini-Sniffer RPV: Mars Bound,” Astronautics and Aeronautics, Vol. 16, 

No. 6, pp 26-39. 

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

Astronautics, April 19, 2005. 

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

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

investigations, April 30, 2003. 

7. “Trade study of multiple thruster options for the Mars Airplane concept,” Kuhl, C.A., Gayle, S.W., Hunter, 

C.A., Kenney, P.S., Scola, S., Paddock, D.A., Wright, H.S., Gasbarre, J.F., NASA TM-2009-215699. 

March 2009. 

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

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

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

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

AIAA-2000-1071, Aerospace Sciences Meeting and Exhibit, 38th, Reno, NV, Jan. 10-13, 2000. 

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

“Unmanned Unlimited” Systems, Technologies, and Operations – Aerospace Conference, San Diego, 

California, September 15-18, 2003. 

11. “Investigation of iron nitrate solubility in nitrogen tetroxide oxidizers”, MELLOR, B.SMITH, C. L. (Royal 

Ordnance, PLC, Westcott, United Kingdom) KAELSCH, I. (ESTEC, Noordwijk, Netherlands) AIAA-1992-

3610 SAE, ASME, and ASEE, Joint Propulsion Conference and Exhibit, 28th, Nashville, TN, July 6-8, 1992. 

13 p. 

12. “Space Shuttle reaction control system thruster metal nitrate removal and characterization,” Saulsberry, R. 

L.; McCartney, P. A., National Aeronautics and Space Administration. White Sands Test Facility, NM. The 

1993 JANNAF Propulsion Meeting, Volume 2 p 103-113 (SEE N95-16324 04-20), November 1993. 
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Technical Memorandum

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Design of a Mars Airplane Propulsion System for the Aerial 

Regional-Scale Environmental Survey (ARES) Mission Concept

5a. CONTRACT NUMBER

 6. AUTHOR(S)

Kuhl, Christopher A.

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NASA Langley Research Center

Hampton, VA 23681-2199

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National Aeronautics and Space Administration

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NASA

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14. ABSTRACT

The Aerial Regional-Scale Environmental Survey (ARES) is a Mars exploration mission concept with the goal of taking 

scientific measurements of the atmosphere, surface, and subsurface of Mars by using an airplane as the payload platform. 

ARES team first conducted a Phase-A study for a 2007 launch opportunity, which was completed in May 2003. Following this 

study, significant efforts were undertaken to reduce the risk of the atmospheric flight system, under the NASA Langley 

Planetary Airplane Risk Reduction Project. The concept was then proposed to the Mars Scout program in 2006 for a 2011 

launch opportunity. This paper summarizes the design and development of the ARES airplane propulsion subsystem beginning 

with the inception of the ARES project in 2002 through the submittal of the Mars Scout proposal in July 2006. 

15. SUBJECT TERMS

Mars; Airplane; Bipropellant; Propulsion; Rocket

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