Anthony J. Colozza

Analex Corporation, Brook Park, Ohio

Comparison of Mars Aircraft

Propulsion Systems

NASA/CR—2003-212350

May 2003

https://ntrs.nasa.gov/search.jsp?R=20030056588 2018-05-15T16:38:52+00:00Z
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Anthony J. Colozza

Analex Corporation, Brook Park, Ohio

Comparison of Mars Aircraft

Propulsion Systems

NASA/CR—2003-212350

May 2003

National Aeronautics and

Space Administration

Glenn Research Center

Prepared under Contract NAS3–00145
Available from

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

The Propulsion and Power Program at

NASA Glenn Research Center sponsored this work.

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

Trade names or manufacturers’ names are used in this report for

identification only. This usage does not constitute an official

endorsement, either expressed or implied, by the National

Aeronautics and Space Administration.
Table of Contents

LIST OF FIGURES ............................................................................................................................................ v

LIST OF TABLES ............................................................................................................................................ vii

INTRODUCTION................................................................................................................................................ 1

ELECTRIC PROPULSION SYSTEM ............................................................................................................. 2

ELECTRIC MOTOR AND CONTROLLER............................................................................................................... 2

GEARBOX........................................................................................................................................................... 3

PROPELLER......................................................................................................................................................... 3

DRIVETRAIN....................................................................................................................................................... 8

FUEL CELL PROPULSION SYSTEM ..................................................................................................................... 8

BATTERY PROPULSION SYSTEM...................................................................................................................... 12

COMBUSTION PROPULSION SYSTEM .................................................................................................... 16

FUEL AND PROPELLANT CANDIDATES ............................................................................................................ 16

PISTON EXPANDER........................................................................................................................................... 24

CONVENTIONAL INTERNAL COMBUSTION ENGINE......................................................................................... 30

COMBUSTION ENGINE SUMMARY ................................................................................................................... 35

ROCKET PROPULSION SYSTEM............................................................................................................... 39

PROPELLANT AND PERFORMANCE .................................................................................................................. 39

ROCKET SYSTEM LAYOUT: SIMPLE BLOWDOWN SYSTEM............................................................................. 43

ROCKET SYSTEM: REGULATED ....................................................................................................................... 44

NOZZLE SIZING................................................................................................................................................ 60

THRUSTER CYCLING........................................................................................................................................ 65

PROPULSION SYSTEM COMPARISON .................................................................................................... 69

APPENDIX A.—TRL LEVEL DEFINITION............................................................................................... 73

REFERENCES................................................................................................................................................... 74

NASA/CR—2003-212350 iii

Fuel/Oxidizer System ................................................................................................................................. 17

Oxidizer ..................................................................................................................................................... 19

Fuel ............................................................................................................................................................ 17

Off Design Rocket Engine Operation........................................................................................................ 61

Monopropellant ......................................................................................................................................... 21

List of Figures

Figure 1 Cl versus Cd for SD8000-PT airfoil at 60,000 Re [2] ........................................ 5

Figure 2 Cl versus Alpha for SD8000-PT airfoil at 60,000 Re [2]................................... 5

Figure 3 SD8000-PT Airfoil Cross Section [2].................................................................. 6

Figure 4 Propeller Blade Twist and Chord Length............................................................ 6

Figure 5 Propeller Performance Curves (2.8 m Diameter, 130 m/s Flight Speed,

6 Blades) [1] .......................................................................................................... 7

Figure 6 Fuel Cell Propulsion System ..............................................................................11

Figure 7 Battery Powered Propulsion System..................................................................15

Figure 8 Bipropellant Piston Expander Engine................................................................25

Figure 9 Mini-Sniffer Aircraft...........................................................................................26

Figure 10 Monopropellant Piston Expander Propulsion System ......................................27

Figure 11 Bipropellant Piston Expander Propulsion System ............................................28

Figure 12 Fuji BT-86 Two Cycle Engine [19]...................................................................30

Figure 13 O.S. Engines FF320 4 Cycle Engine [20] .........................................................31

Figure 14 Combustion Engine System Mass Breakdown for 1 Hour

Duration Mission.................................................................................................36

Figure 15 Combustion Engine System Mass Breakdown for 2 Hour

Duration Mission.................................................................................................37

Figure 16 Combustion Engine System Mass Breakdown for 4 Hour

Duration Mission.................................................................................................38

Figure 17 Density of MON25 and MMH as a Function of Temperature [3]...................43

Figure 18 Blowdown Bipropellant Rocket System............................................................46

Figure 19 Bipropellant System with No Active Control ...................................................47

Figure 20 Bipropellant System with On / Off Control Rocket..........................................48

Figure 21 Bipropellant System with Throttle Control .......................................................49

Figure 22 Monopropellant System with No Active Control .............................................50

Figure 23 Monopropellant System with On / Off Control ................................................51

Figure 24 Monopropellant System with Throttle Control .................................................52

Figure 25 Combustion and Exit Pressures vs Thrust for Off Design

Engine Operation ................................................................................................63

Figure 26 Mass Flow and Isp vs Thrust for Off Design Engine Operation......................64

Figure 27 Estimate of the Effect of Pulse Duration on Isp [16]........................................66

Figure 28 Estimate of Ignition Delay as a Function of Design Thrust [16] .....................67

Figure 29 Acceletometer trace of 10N Kaiser Marquardt Startup [17] ............................68

Figure 30 Mass Comparison Between the Various Types of Propulsion Systems..........70

NASA/CR—2003-212350 v

List of Tables

Table 1 Component efficiencies for Electric Powered Aircraft ........................................ 2

Table 2 Drivetrain Component Masses............................................................................... 8

Table 3 Hydrogen and Oxygen Masses and Volumes for Various Mission Durations... 9

Table 4 Fuel Cell Power Generation System Component Masses for Various Flight

Durations [4, 5, 6].................................................................................................10

Table 5 Battery Specifications for Various Battery Types [7,8,9,10,11,12] ..................13

Table 6 Battery Mass, Volume and Effective Specific Energy for Various Mission

Durations [7,8,9,10,11,12] ...................................................................................14

Table 7 Battery Propulsion System Mass Breakdown.....................................................15

Table 8 Fuels and their Phase Change Temperatures [13,14,15]....................................19

Table 9 Oxidizers and their Phase Change Temperatures [13,14,15].............................21

Table 10 Monopropellants and their Phase Change Temperatures [13,14,15] ................23

Table 11 Combustion Properties and Characteristics for Candidate Fuels and

Oxidizers [13, 16,17]............................................................................................24

Table 12 Fuel and Oxidizer Specification for Piston Expander Engine ...........................28

Table 13 Piston Expander Engine System Masses (kg) (note * indicates pressure

regulator instead of fuel pump)............................................................................29

Table 14 IC Engine Characteristics [19,20]........................................................................30

Table 15 2 and 4 Cycle Engine Fuel and Oxidizer Specifications....................................32

Table 16 2 Cycle Engine System Masses (kg) (note * indicates pressure regulator

instead of fuel pump)............................................................................................33

Table 17 4 Cycle Engine System Masses (kg) (note * indicates pressure regulator

instead of fuel pump)............................................................................................34

Table 18 Low Temperature Propellant Candidates [13,14,15] .........................................40

Table 19 Propellant Mass for 35N Thrust...........................................................................41

Table 20 Rocket System Mass Breakdown for 1 Hour Flight Duration

[3,4,24,25,26,27,28,29,30,31]..............................................................................54

Table 21 Rocket System Mass Breakdown for 2 Hour Flight Duration

[3,4,24,25,26,27,28,29,30,31]..............................................................................55

Table 22 Rocket System Mass Breakdown for 4 Hour Flight Duration

[3,4,24,25,26,27,28,29,30,31]..............................................................................56

Table 23 Rocket System Volume Breakdown for 1 Hour Flight Duration

[3,4,24,25,26,27,28,29,30,31]..............................................................................57

Table 24 Rocket System Volume Breakdown for 2 Hour Flight Duration

[3,4,24,25,26,27,28,29,30,31]..............................................................................58

Table 25 Rocket System Volume Breakdown for 4 Hour Flight Duration

[3,4,24,25,26,27,28,29,30,31]..............................................................................59

Table 26 Ideal Nozzle Geometry for Mars Atmosphere and 35 N Thrust........................61

Table 27 Propulsion System Development Level for a Mars Aircraft Application.........71

NASA/CR—2003-212350 vii

Introduction

For all aircraft, the characteristics and capabilities of the propulsion system are key

elements in establishing the aircraft’s feasibility and flight envelope. This is especially

true for an aircraft that is to fly on Mars. The Martian environment, as well as the launch

from earth and transit in deep space, produces additional obstacles to an aircraft’s

performance capabilities. Aircraft volume is critical since the aircraft must be stowed and

fit into an aeroshell capsule for transit and entry into the Mars atmosphere. Therefore

propulsion system volume is also critical and any components (such as a propeller) that

are large would also need to be stowed.

Because of the thin atmosphere on Mars, where the surface atmospheric density is

comparable to that at 30 km (110 kft) here on Earth, the ability to generate lift is fairly

difficult. To compensate for this thin atmosphere, the aircraft would need to fly fast and

limit its mass as much as possible. Therefore minimizing propulsion system mass is a

critical factor in enhancing the capabilities of the aircraft. Also due to the high speed at

which the aircraft will need to travel and the rocky surface of Mars, landing the aircraft is

not a realistic option. This means then that the aircraft will be able to fly only as long as

the propulsion system can operate. Since mission duration will depend on the efficiency

of the propulsion system, the more efficient and lightweight the system is the longer the

mission and the more capable the aircraft.

Because of this critical impact that the propulsion system has on the aircraft’s capabilities

as well as mission viability, a detailed examination of various propulsion systems was

performed to determine what type of system is most applicable for a desired mission

duration. Three main types of propulsion systems were evaluated, each with various sub

types of systems. The systems examined are listed below.

o Electrical Propulsion Systems

o Fuel Cell Power Propulsion System

o Battery Powered Propulsion System

o Combustion Engine Systems

o Piston Expander Combustion Engine Propulsion System

o 4 Cycle Internal Combustion Engine Propulsion System

o 2 Cycle Internal Combustion Engine Propulsion System

o Rocket Systems

o Bipropellant Rocket Propulsion System

o Monopropellant Rocket Propulsion System

Each of these systems was evaluated to determine its mass and performance

characteristics for three mission durations. The mission durations chosen were 1 hour, 2

hours and 4 hours. These durations provide a fairly large range in aircraft capabilities.

The thrust level that the propulsion systems were designed to was 35N. This thrust level

was based on an estimate for an aircraft with an approximate 5m wingspan and was

deemed representative of the thrust level requirements for proposed near-term Mars

aircraft.

NASA/CR—2003-212350 1
Electric Propulsion System

The first step in sizing an electric power system to provide propulsion power for the

aircraft is to determine the total amount of power needed to produce the desired amount

of thrust. Simplistically this power (P) required by the power system is equal to the thrust

(T) desired times the flight velocity (V) divided by the system efficiency (ηt).

P= TV/ηt (1)

For a basic electric system the efficiency is given by the following.

ηt

 =ηe ηm ηg ηp (2)

 Equation 2 represents the efficiencies of the main power system components, which

include the electronic controls, drive motor, gearbox and propeller. The operational

efficiency associated with each of these components is given in table 1. These

efficiencies are representative approximations, which are subject to change based on a

more detailed system and component design.

Component Efficiency

Control Electronics ηe 0.98

Motor ηm 0.90

Gearbox ηg 0.85

Propeller ηp 0.85

Total System Efficiency ηt

 0.637

Table 1 Component Efficiencies for Electric Powered Aircraft

Based on the efficiencies listed in table 1, in conjunction with a flight speed of 130 m/s

and a thrust level of 35 N, the total power required is 7142 W. To increase the system

performance and reduce the power consumption there is the possibility of eliminating the

gearbox from the system. However, this would require that the motor spin at the RPM of

the propeller. If the propeller is large this RPM would be low (under 1000) and it may be

difficult to find a motor that can achieve this while still maintaining a high efficiency.

Electric Motor and Controller

The power level which the electric motor must operate at is about twice that of high-end

power levels of present day model aircraft. Using the high power electric motors for

model aircraft as well as some larger (much higher power) motors for electric vehicles, a

scaling factor of 1291 W/kg was produced for the motor and 6233 W/kg for the motor

controller. Based on these scaling values the motor and motor controller masses would be

5.52 kg and 1.15 kg respectively.

NASA/CR—2003-212350 2
Aside from the overall mass of the motor there are some additional factors that must be

considered when evaluating which motor to use. The atmosphere of Mars is very thin

(similar to the atmospheric pressure at 30 km altitude on Earth) and composed mostly of

CO2. Because of this thin atmosphere, arcing from electrical devices can become a

serious problem. Regardless of the type of motor is desired, testing within environmental

conditions that simulate the Mars atmosphere would be necessary to insure that arcing

will not occur. The two main types of electric motors that can be used for this type of

application are a brushless DC motor and a brushed DC motor. A brushless motor

reduces the risk of arcing that may be seen with a brushed motor while operating at high

power levels within a rarefied atmosphere, while a brushed motor can produce better

efficiency at higher power levels (current levels).

Another issue is the cooling that the motor and motor controller may require. Cooling is

particularly a concern due to the very thin Martian atmosphere and the extended period of

operation at high power levels. Typical model aircraft motors are very lightweight but are

designed to operate for short periods of time (on the order of 10s of minutes). Because

they are so lightweight they do not have a lot of thermal mass. Therefore they will tend to

heat up fairly rapidly and require significant convective cooling to keep them from

overheating. Normally this would come from the air passing over the motor. However, in

this application it will be operating within the Martian atmosphere and the convective

cooling by the atmosphere may not be sufficient. A potential solution would be to place a

sleeve over the motor that has fins for enhancing the heat transfer to the air stream or to

use a liquid cooling system to keep the motor cool. Either of these approaches has a mass

and drag effect on the aircraft. Another issue is the structural integrity of a very small

motor connected to a fairly large propeller. There can be significant bending loads placed

on the drive shaft. This increases the risk of a mechanical failure of the motor.

Gearbox

The function of the gearbox is to step down the shaft RPM from the motor speed to the

desired propeller RPM. Based on this initial analysis it seems that a 10 to 1 reduction

ratio will be needed. This will take the motor RPM of about 10,000 down to the propeller

RPM of about 1,000. The 10 to 1 reduction is achievable in a single step. This is about as

high a reduction that can be reliably achieved with a single step down. Based on some

lightweight gearboxes that have been previously built for aircraft applications a scaling

number of 0.305 kg/kw for the mass as a function of power transmitted was devised. The

power transmitted by the gearbox is the power required to fly (thrust multiplied by

velocity: 4550 W) divided by the propeller efficiency (85%). This yields a gearbox mass

of 1.63 kg.

Propeller

The propeller design was done in order to get an estimate of the propeller size and mass

needed to produce the thrust required by the aircraft. The design was not optimized and

the optimal geometry can differ from what is presented here. However, this gives a

starting point from which to build and represents one configuration that should be

capable of meeting the mission needs.

NASA/CR—2003-212350 3
For this preliminary design a vortex theory code with low Reynolds number corrections

was used to generate the estimated performance [1]. The airfoil selected for the propeller

was the SD8000-PT low Reynolds number airfoil [2]. This airfoil was chosen because of

its very good lift-to-drag characteristics at low Reynolds numbers as well as its post stall

lift generation capability. A curve of airfoil lift coefficient (Cl) vs. drag coefficient (Cd)

and lift coefficient (Cl) vs. angle of attack (alpha), at a Reynolds number of 60,000 are

shown in figures 1 and 2. The geometry of the airfoil is shown in figure 3.

The propeller was assumed to be a fixed pitch propeller. This was done to reduce the

complexity of the system. However with a fixed pitch propeller some capability and

performance had to be sacrificed. It has to be determined what the best operating blade

angle is for the propeller. The blade angle is akin to various gearing ratios in a car. The

higher the blade angle (or gear) the lower the operational RPM (for a given thrust) but the

greater the torque required. The geometry of the propeller, which includes the twist and

the variation of chord with radius, is shown in figure 4. Since the blade is twisted the

pitch angle will vary along the blade length. Therefore, the pitch angle specified for the

blade represents the angle the blade is at the 3/4 radius point. For this design the pitch

angle was 38°. The propeller has six blades and the diameter was chosen to be 2.8 m.

This diameter and blade number was sufficient to achieve the thrust level of 35N at 1 km

altitude on Mars.

Determining the operating point meant selecting the blade angle, RPM, power required

and thrust generated. The selection of these parameters will also influence the capabilities

of the other components of the system, such as motor RPM and power available.

Propeller performance curves were generated for the geometry described. These

performance curves are shown in figure 5. This configuration requires 5353 watts of shaft

power, produces an efficiency of 85% and has an operational RPM 1172.

Based on the geometry specified and using a hollow carbon section for the propeller, the

mass of each blade could be calculated. The mass of each blade is 1.63 kg. Therefore the

total propeller mass is 9.8 kg.

NASA/CR—2003-212350 4
0.01

0.02

0.03

0.04

0.05

0.06

-0.2 0 0.2 0.4 0.6 0.8 1 1.2

Cd

Cl

Figure 1 Cl versus Cd for SD8000-PT Airfoil at 60,000 Re [2]

-0.4

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

-10 -5 0 5 10 15

Cl

Alpha

Post Stall

Figure 2 Cl versus Alpha for SD8000-PT Airfoil at 60,000 Re [2]

NASA/CR—2003-212350 5
Figure 3 SD8000-PT Airfoil Cross Section [2]

Figure 4 Propeller Blade Twist and Chord Length

40

45

50

55

60

65

70

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6

Blade Angle (degrees)

Chord Length (m)

Radial Station (3.5 m diameter)

Blade Twist

Chord Length

NASA/CR—2003-212350 6
0

605040302010

121086420

1100 1200 1300 1400 1500 1600 1700 1800

Thrust (N) 

Power (kw)

RPM

Thrust

Thrust

Thrust

Thrust

Thrust

Thrust

Power

Power

Power

Power

Power

Power

38°

40°

42°

44° Design Pont 35N

34°

36°

Figure 5 Propeller Performance Curves (2.8 m Diameter, 130 m/s Flight Speed, 6 Blades) [1]

NASA/CR—2003-2123507
Drivetrain

The drive-train component masses are listed in the following table. These are based on

the 35N thrust requirement and a 130 m/s flight velocity.

Component Mass (kg)

Motor Controller 1.15

Motor 5.52

Gearbox 1.63

Propeller 9.8

Total Drivetrain 18.10

Table 2 Drivetrain Component Masses

Fuel Cell Propulsion System

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

be used to power the electric motor to spin a propeller generating thrust. This system is

shown in figure 6. The system consists of a fuel cell stack, pressure regulators, check

valves, filters, vent valves, flow and pressure sensors, controller, hydrogen and oxygen

pressure tanks and associated piping between the components. The gas flows from the

tanks are regulated to the desired pressure and filtered as the gases are passed to the fuel

cell. The system is controlled by monitoring the pressure and flow rate in both the

hydrogen and oxygen lines and adjusting the flow regulators and exit stream vent control

valves to maintain the desired reactant conditions within the fuel cell. The fuel cell will

normally operate with the vent flow control valves closed. These valves will need to be

opened to purge the fuel cell of water buildup or regulate an imbalance in the flow of one

of the reactant gases.

Based on the power required, the fuel cell stack must produce 7142 W of output power

during nominal operation. The conversion efficiency of fuel to electricity is

approximately 50%. Therefore the flow rate of hydrogen (M) needed can be determined

by the following equation, where Ptot is the total power output of the fuel cell, η is the fuel

cell efficiency estimated to be 50% and F is based on Faradays constant, 96484 (A

s/equivalent), which for hydrogen is 192968 (A s/g mole).

M P

F

tot = 2

1 2 1000 . η (3)

This translates into a hydrogen consumption rate of 0.484 kg/hr for the 7.791 kW of

power required. Mission durations of 1, 2 and 4 hours were considered. The

corresponding hydrogen mass for each mission duration considered is given in table 3.

The amount of oxygen needed to react with the hydrogen can be determined from the

basic chemical reaction of combining hydrogen and oxygen to from water. This reaction

is given below.

NASA/CR—2003-212350 8
2H2(g) + O2 (g) → 2H2O (4)

Based on this reaction 8 grams of oxygen are consumed for each gram of hydrogen.

Therefore the reaction mass flow of oxygen is 8 times that of the hydrogen flow. The

oxygen requirement for the 1, 2 and 4 hour missions is given in table 3.

For a storage pressure of 20.7 Mpa (3000 psi) and a factor of safety of 1.8 for the tanks,

the specifications on the hydrogen and oxygen system is given in table 3. The gas

volumes for both hydrogen (Vh) and oxygen (Vo) were calculated using the equation of

state with a compressibility factor for the hydrogen gas. The volumes for both the

hydrogen and oxygen are given in table 3. The tank mass is based on a spherical carbon

composite tank (3500 Mpa, 507 kpsi yield and 1608 Kg/m3, 0.058 lb/in3 density) with a

thin metal liner that is resistant to the gas migration out of the tank [3].

Vh = 4157.2 Z MhTh/ Ph (5)

Z = 0.99704 + 6.4149E-9 Ph (6)

Vo = 259.83 Mo To / Po (7)

1 hour 2 hours 4 hours

Hydrogen Mass 0.484 kg 0.968 kg 1.936 kg

Oxygen Mass 3.872 kg 7.744 kg 15.488 kg

Hydrogen Volume 0.026 m3 0.053 m3 0.11 m3

Oxygen Volume 0.010 m3 0.021 m3 0.041 m3

Table 3 Hydrogen & Oxygen Masses & Volumes for Various Mission Durations

It should be noted that no allowance was made for either the hydrogen or oxygen gas

leaking from the tank during the mission. This is particularly a concern for the hydrogen

gas due to the high storage pressure and the relatively long transit time to Mars.

Cryogenic storage was also considered but the boil off during transit was too great to

make this a viable option.

A breakdown of the fuel cell system components and mass estimates for the three mission

durations examined are given in table 4. This estimate is based on the 35 N thrust level.

The fuel cell stack mass is based on a scaling factor of 1 kg/kW. This is an estimate based

on projections of near term fuel cell performance. Using the mass breakdown and power

output of the system, a specific energy (W-hr/kg) for each mission duration can be

calculated. This specific energy is given in table 4. As the mission duration increases, the

specific energy of the fuel cell system will increase. This is because only the tank and

reactant masses have to increase duration. As the desired mission duration increases a

fuel cell system will become more and more attractive from a mass standpoint.

NASA/CR—2003-212350 9
Component 1 Hour Duration 2 Hours Duration 4 Hours Duration

Fuel Cell 7.8 kg 7.8 kg 7.8 kg

Hydrogen Tank 1.35 kg 2.75 kg 5.71 kg

Oxygen Tank 0.52 kg 1.09 kg 2.13 kg

Check Valves (2) 0.88 kg 0.88 kg 0.88 kg

Pressure Sensors (2) 1.8 kg 1.8 kg 1.8 kg

Flow Sensors (2) 1.8 kg 1.8 kg 1.8 kg

Filters (2) 0.52 kg 0.52 kg 0.52 kg

Flow Regulators (2) 4.8 kg 4.8 kg 4.8 kg

Control Valves (2) 3.0 kg 3.0 kg 3.0 kg

Propellant Lines 1.5 kg 1.5 kg 1.5 kg

Controller Unit 1.5 kg 1.5 kg 1.5 kg

Power System

Total Dry Mass

25.47 kg 27.44 kg 31.44 kg

Power System

Total Wet Mass

29.83 kg 36.15 kg 48.87 kg

Power System

Specific Energy

(W-hr / kg)

239 395 584

Drivetrain 18.1 kg 18.1 kg 18.1 kg

Propulsion System

Total Mass

47.93 kg 54.25 kg 66.97 kg

Table 4 Fuel Cell Power Generation System Component Masses for Various

Flight Durations [4, 5, 6]

NASA/CR—2003-212350 10
P

Filter

Controller

Output Input

Fuel Cell

F

Check Valve

Flow Control Valve

Pressure Sensor

Flow Sensor

Vent

F P

Hydrogen Tank

Oxygen Tank

Pressure Regulator

Gearbox Electric Motor

Propeller

Motor Controller

Figure 6 Fuel Cell Propulsion System

NASA/CR—2003-212350 11
Battery Propulsion System

A battery propulsion system, shown in figure 7, is much simpler then the fuel cell system

discussed previously. The battery system requires less active controls and no mechanical

components in order to operate. This is a benefit for both risk reduction as well as cost

and implementation. One of the advantages of the battery system is that it could be easily

spread throughout the aircraft without significantly affecting its operation or complexity.

Present day battery systems cannot achieve the high energy density (W-Hr / kg) that is

achievable with the fuel cell system for the mission durations analyzed. Table 5 lists

present and near-term batteries and their operational characteristics. A number of battery

types were looked at, and the data presented are from “off the shelf” production batteries.

With some development these performance numbers can be expected to improve. Mass,

volume and effective specific energy values for each of the batteries considered for the 1,

2 and 4 hour mission durations are given in table 5.

By looking at the data in table 5 it is obvious that some types of batteries will not be

applicable to a Mars aircraft mission. This is mainly due to either the battery specific

energy being too low or the inability to extract power from the battery at a rate

commensurate with the mission duration. Most of the lithium-based batteries, although

they have very high specific energies, have very low power discharge rates. Because of

this, excess battery capacity must be carried in order to achieve the power level required

by the aircraft. This excess capacity effectively reduces the specific energy of the battery

for the particular mission duration being considered. Of the batteries listed, the lithium

sulfur chloride and lithium manganese dioxide show the most promise. In addition to

these lithium batteries, the silver zinc batteries may also be applicable. It should be

remembered that this table is based on present day technology and the batteries had not

been optimized for aircraft operation. With some development it is reasonable to assume

that the mass and volume of the battery can be reduced on the order of 10% to 15% for

this application. If a reduction beyond that is necessary for the battery to be applicable, it

would not be a good choice for this application.

The masses, volumes and specific power numbers listed in table 5 are based on 7142 W

of power required for the mission duration. The mass of the battery will scale with the

duration as long as the battery can be discharged within that timeframe. For example the

mass of the silver zinc batteries can be reduced in half by reducing the mission duration

from 4 hours to 2 hours. This is because this type of battery has a high rate of discharge

and will not be effected by the shortened mission duration. On the other hand shortening

the duration does not have an effect on the lithium batteries because they cannot provide

power at a faster rate. For example, using the lithium sulfur chloride battery to produce

7142 W of power requires a capacity of 8+ hours no matter how long the mission really

is.

Another issue that should be mentioned that has not been considered in the preliminary

design is the heat removal from the battery system. The batteries will generate

considerable heat during operation and most of this heat will need to be removed. The

removal of heat from the batteries could be more difficult then the heat removal from the

fuel cell system. This is because the heat generated from the battery system will be

distributed throughout the complete battery pack. If the batteries are distributed

NASA/CR—2003-212350 12
throughout the aircraft a significant amount of heat transfer material will be needed to

channel the heat from the batteries to a heat exchanger.

Total propulsion system masses are given in table 7. The battery mass used for each

mission duration is the one from table 6 that had the lowest mass for that duration.

Battery Type /

Manufacturer

Minimum Operating

Duration to 80%

DoD

Operating

Temperature

Range

Energy Density Full

Discharge

(W Hr / Kg)

Li Mn O2 /

Ultralife

4.8 hr –30 °C to 55 °C 240

Li Oxyhalide /

Electrochem

5 hr 0 °C to 150 °C 137

Li Sulfur Chloride

/ Electrochem

8.78 hr –32 °C to 93 °C 444

Li Bromine

Complex /

Electrochem

8.6 hr –55 °C to 72 °C 369

Li Thinyl

Chloride / Tadiran

20 hr –20 °C to 150 °C 361

Li Ion / Yardney 4 hr –20 °C to + 40 °C 145

Nickel Hydrogen /

Eagle Pitcher

4 hr or less –5 °C to + 35 °C 45

Ni-Cad

Eagle Pitcher

4 hr or less –5 °C to + 30 °C 36.5

Silver Zinc /

BST Systems

4 hr or less –20 °C to + 30 °C 150

Table 5 Battery Specifications for Various Battery Types [7,8,9,10,11,12]

NASA/CR—2003-212350 13
1 Hour

Duration

2 HoursDuration

4 HoursDuration

Battery Type Mass

(kg)

Volume

(m3)

Specific

Energy

(W hr/kg)

Mass

(kg)

Volume

(m3)

Specific

Energy

(W hr/kg)

Mass

(kg)

Volume

(m3)

Specific

Energy

(W hr/kg)

Li Mn O2 /

Ultralife

143 0.087 49.9 143 0.087 99.9 143 0.087 199.6

Li Oxyhalide /

Electrochem

260 NA 27.5 260 NA 54.9 260 NA 110

Li Sulfur Chloride

/ Electrochem

141 0.067 50.7 141 0.067 101.4 141 0.067 202.8

Li Bromine

Complex /

Electrochem

166 0.078 43.0 166 0.078 86.0 166 0.078 172

Li Thinyl

Chloride / Tadiran

396 0.205 18.0 396 0.205 36.0 396 0.205 72

Li Ion / Yardney 197 0.148 36.3 197 0.148 72.6 197 0.148 145.2

Nickel Hydrogen /

Eagle Pitcher

159 0.200 45 317 0.400 45 635 0.800 45

Ni-Cad

Eagle Pitcher

196 NA 36.5 391 NA 36.5 783 NA 36.5

Silver Zinc /

BST Systems

56 0.003 127 113 0.006 127 225 0.012 127

Table 6 Battery Mass, Volume and Effective Specific Energy for Various Mission Durations [7,8,9,10,11,12]

NASA/CR—2003-212350 14
Gearbox Electric Motor

Propeller

Motor Controller

Batteries

Figure 7 Battery Powered Propulsion System

Component 1 Hour 2 Hours 4 Hours

Battery 56 kg (Silver Zinc) 113 kg (Silver Zinc) 141 kg (Li-Sulfur

Chloride)

Wiring 1.5 kg 1.5 kg 1.5 kg

Controller 1.5 kg 1.5 kg 1.5 kg

Drivetrain 18.1 kg 18.1 kg 18.1 kg

Propulsion System

Total Mass

77.1 kg 134.1 kg 162.1 kg

Table 7 Battery Propulsion System Mass Breakdown

NASA/CR—2003-212350 15
Combustion Propulsion System

The combustion of fuel to generate thrust dates back to the beginning of aviation. This is

the standard method for producing thrust in almost all conventional aircraft. It is

accomplished though the use of an internal combustion engine, turbine or some other

mechanical device. One advantage almost all of the Earth based aircraft have is the

presence of oxygen within the atmosphere. On Mars there is no oxygen available and

therefore the combustion engines, which have such a rich aviation heritage here on Earth,

cannot be directly applied to Mars flight. Combustion engines, however, can still be

considered for Mars flight but the aircraft will need to either carry along an oxidizer or

use less energetic fuels that decompose through catalytic reaction.

Standard internal combustion engines can be considered for use with a fuel / oxidizer

system. To use a monopropellant through catalytic decomposition, an external

combustion engine or piston expander would be required.

The performance of a fuel-driven combustion engine is dependent on the type of fuel and

engine (thermodynamic cycle) used as well as the specific engine design. The need to

carry along an oxidizer poses a significant weight penalty on the aircraft. Therefore the

choice of the fuel / oxidizer combination is critical to optimizing the propulsion system.

Factors that need to be considered include: storage volume, environmental constraints

(such as freezing temperature and reactivity with other materials) and the amount of

energy produced during combustion.

Fuel and Propellant Candidates

The fuel selection will be based on the following criteria:

1. Ability of the fuel to meet the environmental conditions of the mission,

2. Ability of the fuel to be stored in the vehicle

3. Ability of the fuel to meet the performance goals of the aircraft.

The operational constraints on the fuel require it to be capable of being stored for

extended periods of time (up to 2 years) with little or no degradation and being capable of

withstanding the deep space environment during transit as well as the environment on the

surface of Mars. The main environmental issue during transit and on the Mars surface is

temperature. Assuming that there is no active thermal control or heating available, the

fuel must be capable of withstanding temperatures down to –40 °C for extended periods

of time. If the fuel can remain liquid at these temperatures then this greatly simplifies the

propellant delivery system as well as eliminates the need for power consuming heaters.

Also this reduces the overall risk of the mission by eliminating a failure mechanism

which can occur from improperly thawed fuel or a failed heater.

An overall list of potential fuels and oxidizers (and fuel oxidizer combinations) is given

in the following sections and included in tables 8, 9 and 10 [1,2].

NASA/CR—2003-212350 16
Fuel/Oxidizer System

Since there is no oxygen available from the environment of Mars the use of a

conventional combustion engine requires an oxidizer to be brought along as part of the

fuel system. Table 8 and 9 lists fuels and oxidizers that could potentially be used. Various

characteristics of the fuels and oxidizers have to be considered when selecting the

appropriate combination for used. Ideally they have a low freezing point (below –40°C)

to insure that they do not freeze during transit to Mars and would be in the liquid state

during storage. Being a liquid is critical due to volume constraints within the aircraft.

Other issues such as compatibility and ignition are a significant factor in the selection.

Certain fuel/oxidizer combinations will spontaneously react when combined (hypergolic).

An example of these combinations are monomethyl hydrazine (MMH) and

unsymmetrical dimethyl-hydrazine (UDMH) which are hydrazine derivatives that can

operate as the fuel in a bipropellant system. These fuels react instantaneously with certain

oxidizers (such as nitrogentetroxide (NTO)) and therefore are hypergolic and do not

require an ignition system. This characteristic reduces complexity by eliminating the need

for an ignition system and increases system reliability.

Fuel:

Hydrogen

Hydrogen is a stable, non-corrosive, non-toxic material. However in order to be useable

for this mission it must be kept in a liquid state. This requires cryogenic storage that

would significantly increase the complexity of the mission.

Ammonia

Ammonia is a stable compound that can be stored in Teflon, 18-8 stainless steel,

aluminum or polyethylene. It is mildly toxic but can be fatal in concentrated exposure.

The main issue with its use for this mission is that it is in the gaseous state under mission

conditions.

Hydrazine

Although its most common use is as a monopropellant, hydrazine can also be used as a

bipropellant. It has the same general properties as the monopropellant version however its

performance is significantly improved when utilized in combination with an oxidizer.

Monomethyl Hydrazine

Monomethyl hydrazine is fairly stable at lower temperatures, however it becomes

unstable above 260 °C (500 °F). It can be stored in 18-8 stainless steel, aluminum or

Teflon. It is toxic. Its liquid temperature range is well within the requirements for the

mission environment.

NASA/CR—2003-212350 17
Unsymmetrical Dimethyl-Hydrazine (UDMH)

Unsymmetrical dimethyl-hydrazine is stable at low temperatures but becomes violently

unstable at temperatures above 260 °C (500 °F). It can be stored in most materials

including mild steel, 18-8 stainless steel, aluminum, Teflon and polyethylene. It has a

lower level of toxicity then hydrazine but more then that of ammonia. Its liquid state

temperature range is well within that of the mission requirements.

RP-1

RP-1 is a stable fuel developed for space applications. It is stable up to 370 °C (700 °F)

and is compatible with all common metals as well as meoprene, asbestos, fluorocarbons

and epoxies. Its toxicity is comparable to that of kerosene. The liquid temperature range

for RP-1 is within the operating range for the mission. However, the freezing point is at

the estimated low temperature for the mission. To insure that RP-1 doesn't freeze during

the mission, some active thermal control would probably be required.

Methane

Methane is stable and compatible with all common metals as well as neoprene, asbestos,

fluorocarbons and epoxies. It is essentially non-toxic. The main issue with its use is its

low boiling point. This would require it to be used as a gas or stored cryogenically.

Propane

Propane essentially has the same properties as methane. It is stable and compatible with

all common metals as well as neoprene, asbestos, fluorocarbons and epoxies. The issues

with its use are the same as those of Methane.

Diborane

Diborane is a gas at room temperatures and will slowly decompose. At higher

temperatures it decomposes rapidly. It is compatible with most metals and some organic

materials. It has moderate toxicity. The issues with using diborane are significant. It

would need to be stored as a cryogenic liquid in order to provide for sufficient mission

duration as well as minimize the decomposition rate.

NASA/CR—2003-212350 18
Fuel Boiling

Point

(°C)

Freezing

Point

(°C)

Potential Oxidizers Density

(gm/cm3 at

20 °C)

Hydrogen (H2) –253 –259 Oxygen, Fluorine 0.071

(at –253 °C)

Ammonia (NH3) –33.4 –77.7

Oxygen, Fluorine, Nitrogen

Tetroxide, Chlorine

Trifluoride

0.611

Hydrazine (N2H4) 113.4 1.5

Oxygen, Fluorine, Nitrogen

Tetroxide, Chlorine

Trifluoride

1.008

Monomethyl

Hydrazine

(MMH) (N2H6C)

89.2 –52.5

Nitrogen Tetroxide, Chlorine

Trifluoride, Inhibited Red

Fuming Nitric Acid

0.874

Unsymmetrical

Dimethyl

Hydrazine

(UDMH)

(N2H8C2)

63.8 –57.2

Oxygen, Fluorine, Nitrogen

Tetroxide, Chlorine

Trifluoride

0.792

RP-1

(C11.74H21.83) 185 –40

Oxygen, Inhibited Red

Fuming Nitric Acid 0.801

Methane (CH4) –161 –183.9 Oxygen, Fluorine 0.415

(at –164 °C)

Propane (C3H8) –42.2 –187.1 Oxygen, Fluorine 0.585

(at –44 °C)

Diborane (B2H6) –92.6 –164.8 Oxygen Difluoride 0.435

(at –92.6 °C)

Table 8 Fuels and their Phase Change Temperatures [13,14,15]

Oxidizer:

Oxygen

Oxygen is highly reactive and non toxic. It is non-corrosive and is very stable in storage.

The main issue with its use is that it would be in the gaseous form under the mission

conditions. This will significantly limit the volume of oxygen which can be stored.

Liquid oxygen can be used however this brings up significant issues regarding the storage

and manufacture of a cryogenic liquid.

Fluorine

Fluorine is highly reactive with almost any material. It can be stored in 18-8 stainless

steel or copper but Monel is preferred. It is very important that all materials that come

into contact with fluorine are thoroughly cleaned so that there are no contaminating

NASA/CR—2003-212350 19
particles for the fluorine to react with. There are no nonmetallic materials that do not

react to some degree with fluorine. It is also highly toxic and corrosive to body tissue.

Like oxygen it is a gas at mission temperatures. Therefore it would need to be stored

cryogenically in order to be used in the mission.

Nitrogen Tetroxide

Nitrogen tetroxide is a stable compound. It is not highly reactive and can be stored in

mild steel, stainless steel, aluminum, Teflon and polyethylene. It's toxicity is comparable

to that of chlorine. Various formulations of nitrogen tetroxide are available. These

formulations vary the percent of nitric oxide in the formulation. This change in the nitric

oxide content can effect the freezing point for the propellant. The mixtures of nitrogen

tetroxide shown in the table above have varying amounts of nitric oxide. MON 25 (25%

mixed NO) has a significantly decreased freezing point over MON 3 (3% mixed NO).

This ability to lower the freezing point of nitrogen tetroxide makes it applicable to the

mission environment and would eliminate the need for thermal control of the propellant.

Chlorine Trifluoride

Chlorine trifluoride is a stable oxidizer that can be stored in 18-8 stainless steel, nickel

and Monel. However most common metals can be used if they are free of contaminants.

It is highly toxic with a toxicity comparable to fluorine. Its liquid state temperature range

is more then sufficient to meet the mission requirements.

Inhibited Red Fuming Nitric Acid (IRFNA)

IRFNA is subject to decomposition at elevated temperatures and its decomposition rate is

directly related to temperature. It can be stored in 18-8 stainless steel, polyethylene and

Teflon. It is toxic and corrosive to body tissue. The liquid temperature range of IRFNA is

sufficient to keep the oxidizer in a liquid state throughout the proposed mission duration.

Oxygen Difluoride

Oxygen Difluoride is stable at normal room temperature but becomes increasingly

unstable at elevated temperatures. It can be stored in 18-8 stainless steel, copper,

aluminum, Monel and nickel. Nonmetallic materials are generally not compatible. It is

highly toxic and corrosive to body tissue. The main issue is that it is a gas at mission

temperatures. In order to be useable it would need to be stored cryogenically in order to

be used in the mission. This would add significant risk and complexity to the overall

mission.

NASA/CR—2003-212350 20
Oxidizer Boiling Point (°C) Freezing Point (°C) Density

(gm/cm3)

Oxygen (O2) –183 –218.8 1.143

Fluorine (F2) –188.1 –219.6 1.505

Nitrogen Tetroxide (MON3)

(N2O4)

21.2 –11.2

1.45

Chlorine Trifluoride (CLF3) 11.8 –76.6 1.825

Inhibited Red Fuming Nitric

Acid (IRFNA)

(0.835HNO30.140NO20.020

H2O0.005HF)

~60 ~ –62.2 1.56

Oxygen Difluoride (OF2) –145 –223.9 1.521

Nitrogen Tetroxide

(MON25) (N2O4) –9 –54 1.45

Table 9 Oxidizers and their Phase Change Temperatures [13,14,15]

Monopropellant:

To reduce the complexity of the system, some monopropellant fuels can be used. These

fuels do not require an oxidizer but react by being passed over a catalyst. A common

example is hydrazine that decomposes through the following reaction.

3N2H4 → N2 + 4NH (8)

Although no oxidizer is needed, the heating value of these fuels is about one half that of a

conventional hydrocarbon fuel combusted with oxygen. It should be noted that because

of the need for a catalyst, hydrazine (as a mono-fuel) is not suitable for internal

combustion engines. It can however work in external combustion engines such as a

expander engine or Stirling cycle engine.

The catalyst is the main element in the decomposition of hydrazine. The catalyst bed

enables the spontaneous and complete decomposition of hydrazine. Increased reliability

and simplicity are the main advantages of a single fuel system. This is achieved by not

requiring an ignition system and eliminating the oxidizer tank and feed system. The

catalyst is not consumed in the reaction and tests have shown that certain catalysts can

initiate decomposition of hydrazine at temperatures as low as –54 °C. The catalyst

carrier (material that holds the activating metal) has to have as large a surface area as

possible. The best and most versatile material for the active metal in the catalyst is

iridium. Present day catalysts have on the order of 133 m2

/gm of surface area with 32%

of its mass made up of the activation metal (iridium). The iridium catalyst is capable of

several hundred spontaneous cold starts and steady-state operation of several hours. This

NASA/CR—2003-212350 21
type of catalyst can decompose 40 gm/ cm2

 s of hydrazine [14]. The properties of the

various monopropellants are summarized in table 10.

Hydrogen Peroxide

Hydrogen peroxide has a flight heritage dating back to the 1930's, although until recently,

the technology has been dormant. It is seeing a revival of sorts, primarily as an oxidizer

in a bipropellant combination, but also as a monopropellant. The main advantage of

hydrogen peroxide is that it is non-toxic. However, the freezing point is higher then what

is necessary to perform the mission without thermal control. There are a few options for

dealing with this freezing issue. A passive thermal system may be used to maintain the

temperature above freezing. This may be possible since its freezing point is within 30 °C

of the expected environmental conditions. Hydrogen peroxide decomposes in the

presence of a catalyst such as carbon, steel or copper. For storage it doesn't react with

certain materials such as aluminum, tin, glass, polyethylene or Teflon.

Ethylene Oxide

Ethylene oxide remains liquid over a temperature range which is more then adequate to

meet the mission requirements. It is generally stable but the polymerization rate is

increased in the presence of some materials. Storage materials it is compatible with

include 18-8 stainless steel, aluminum, mild steels, copper, nylon and Teflon. This

material is relatively toxic and must be handled with caution.

Nitromethane

Nitromethane's temperature range is nearly within the range required by the mission. It

would require some insulation or thermal control to assure it would not freeze. It doesn't

react with 18-8 stainless steel, aluminum or polyethylene. These materials can be used for

storage. The main issue with its use is that it may detonate under conditions of

confinement, heating and mechanical impact. Any of which can possibly be experienced

during this mission. Also, it is relatively toxic and must be handled with caution.

n-Propyl Nitrate

n-Propyl Nitrate is capable of remaining liquid well within the temperature range of the

mission. It is relatively stable and insensitive to mechanical or thermal shock. It can be

stored in containers made of either 18-8 stainless steel, aluminum, polyethylene, Teflon,

nylon, Orlon, Dacron or Mylar. This material has no serious toxicity problems which

allows for easy handling.

Hydrazine

Hydrazine monopropellant has been used in spacecraft for the last 30 years, mainly for

low-thrust applications like satellite station-keeping. Decomposition is achieved by a

catalyzed reaction with a metal oxide. Materials which are compatible with hydrazine and

will not react include Teflon, 18-8 stainless steels, polyethylene and aluminum. The main

issues with using hydrazine are that it is highly toxic and has a high freezing point

NASA/CR—2003-212350 22
(approximately 1.5°C). Because of this high freezing point, significant heating would be

required throughout the mission in order to maintain the propellant in its liquid state.

Hydrazine Propellant Blend (HPB)

HPB represents a family of monopropellant formulations composed of hydrazine,

hydrazinium nitrate (HN) and water. The addition of HN and water serve to depress the

freezing point and increase the performance of plain hydrazine. Several HPB's were

developed and tested in the 1960's and 1970's, primarily for military applications. HPB's

are receiving renewed attention as a low freezing point monopropellant. Presently NASA

is sponsored HPB development work at Primex Aerospace.

Hydroxylammonium Nitrate (HAN)

HAN is a family of monopropellants composed of an oxidizer rich salt, a fuel component

and water. These types of propellants have been under development by NASA over the

last decade. HAN based monopropellants offer a high density low freezing point,

nontoxic alternative to hydrazine.

Monopropellant Boiling

Point

(°C)

Freezing

Point

(°C)

Density

(gm/cm3)

Combustion

Temperature

(°C)

Specific

Impulse

(Isp)

Hydrogen

Peroxide(0.9H2O20.1H2

O)

141.1 –11.5 1.39 757 148

Ethylene Oxide (C2H4O) 10.6 –112.8 0.87 1004 199

Nitromethane (CH3NO2) 101.2 –29 1.14

2193 245

n-Propyl Nitrate

(C3H7NO3) 110.5 –101.1 1.057 1078 210

Hydrazine (N2H4) 113.4 1.5 1.008 633 199

Hydrazine Propellant

Blend (HPB) HPB-1808

(18% HN, 8% water)

100 –20 na 230

61%

Hydroxylammonium

Nitrate (NH3OH)NO3

(HAN) 14% Glycine

(H2NCH2 CO OH)

100 –35 1.41 190

Table 10 Monopropellants and their Phase Change Temperatures [13,14,15]

NASA/CR—2003-212350 23
Of the fuel and oxidizers listed above, the following combinations were determined to be

the best candidates for use. The selection was made based on combustion properties,

environmental considerations and space flight heritage.

Fuel Oxidizer Heat of

Combustion

(kJ / kg)

Oxidizer /

Fuel Ratio

(by mass)

Density

Fuel / Oxidizer

(kg/m3)

Hydrogen

(H2)

Oxygen

(O2)

144,000 8 18.35 / 375.46

(at 3000 psi)

MonomethylHydrazine

(MMH)

(N2H6C)

Nitrogen

Tetroxide: 25%

mixed Oxides

(N2O4)

28,347 1.75 874 / 1450

Hydrazine

(N2H4)

NA 19,429 NA 1008

Unsymmetrical

dimethylHydrazine

(UDMH)

(N2H8C2)

Nitrogen

Tetroxide

25% mixed

Oxides

(N2O4)

39,960 2.25 792 / 1450

Table 11 Combustion Properties and Characteristics for Candidate Fuels and

Oxidizers [13, 16,17]

Piston Expander

The piston expander is an engine that can burn almost any mono or bipropellant. This

type of engine is used extensively by the Navy for torpedo propulsion. Due to its light

weight and ability to burn a number of different fuel types, this engine may be well suited

for a Mars aircraft application. Engines designs that are used by the Navy would not be

directly usable in a Mars aircraft. There are a number of development issues which would

need to be addressed in order to apply this type of engine to the proposed Mars aircraft.

The two main obstacles to the engine's use are cooling and scaling to the correct power

level. The torpedo motor is cooled during operation by circulating sea water through

passageways within the engine. This would not be possible for a Mars aircraft and an

alternate cooling scheme would need to be devised. The smallest production torpedo

motor produces orders of magnitude more power then what is needed by the proposed

Mars aircraft. Therefore the engine would need to be significantly scaled down to meet

the power and corresponding weight requirements for this type of aircraft and mission.

This scaling would require significant development time and cost. Utilizing a version of

the torpedo motor in an aircraft was previously examined under the ERAST program in

the mid 1990’s. A prototype engine was under development and some preliminary testing

was performed. A diagram of this engine is shown in figure 8.

NASA/CR—2003-212350 24
The concept of using a piston expander propulsion engine as the source of power for the

propeller dates back to the first attempts at designing a Mars aircraft in the late 1970’s.

This design was based on the Mini-Sniffer high altitude unmanned vehicle program.

Utilizing the Mini-Sniffer development, JPL tried to adapt this aircraft to a Mars

atmospheric flight vehicle. The Mini-Sniffer was a small remotely piloted aircraft

designed to fly at an altitude of up to 30 km (100,000 ft) on Earth. This was a propeller

driven aircraft with a hydrazine burning piston expander engine. The engine was

designed by James Akkerman from NASA Johnson Space Center and dubbed the

Akkerman engine.

Figure 8 Bipropellant Piston Expander Engine

A prototype engine was constructed and some testing was performed. The Akkerman

engine burns hydrazine fuel in a catalyst bed chamber. The prototype was wind tunneltested and flight-tested to altitudes of 20 kft. The engine was designed to provide 22kW

(30 hp) of power to the propeller. A scaled down version of this engine was proposed for

use in a Mars aircraft by Development Sciences Inc. under contract by JPL in the late

1970’s. Their estimate for the scaled Akkerman engine was that it would produce 11 kW

(15 hp) and weigh 6.8 kg (15 lbs), approximately 1.62 kW/kg for the engine. No

prototype of this scaled-down engine was ever produced. The best measured specific fuel

consumption for the original prototype engine was 2.7 kg/ kW-hr (4.5 lb/HP hr).

However by the end of the 1970’s the program was put on hold and all development on

the hydrazine engine was stopped. A picture of the Mini-Sniffer aircraft is shown in

Figure 9 [18].

NASA/CR—2003-212350 25
Figure 9 Mini-Sniffer Aircraft

A diagram of a piston expander propulsion system using a monopropellant fuel is shown

in figure 10 and using a bipropellant system is shown in figure 11.

Utilizing a heat of combustion for hydrazine (hc) of –19,429 KJ/Kg and the specific fuel

consumption (SPF) of 0.37 kW-hr / kg from the measured fuel flow given above,

produces an engine thermal efficiency (ηt

) of about 7%. This efficiency is calculated

from the following equation

ηt

 = SPF (3600) / (-hc) (9)

This measured value is for a prototype engine so it would be reasonable to estimate that

after some development, the engine efficiency can reach up to 10%. This low thermal

efficiency is characteristic of this type of combustion engine.

NASA/CR—2003-212350 26
Fuel Pump

Check Valve

Pyrovalve

From Power System Bus

Propeller

Catalytic Combustion

Chamber

Piston Expander Engine

Electrical Alternator

To Power System Bus

Figure 10 Monopropellant Piston Expander Propulsion System

The fuel requirements for the piston expander engine are given in table 12. These are

based on the properties of the fuels listed in table 11 and the flight power requirements

for the aircraft (7142 W). It is estimated that the combustion pressure within the

combustion chamber will be on the order of 13.8 MPa (2000 psi). Based on this pressure

and an assumed 1 second response time, the combustion chamber can be sized (68 cm3

for hydrazine, 105 cm3 for MMH + NTO and 83 cm3 for UDMH + NTO). The liquid

fuels and oxidizers are stored in titanium tanks at pressures of 690 kPa (100 psi) with a

tank structural factor of safety of 1.8. The gas storage tanks are composed of wound

carbon fiber with an inner metal liner with a storage pressure of 20.68 MPa (3000 psi)

and a structural factor of safety of 1.8.

A breakdown of the system component masses shown in figures 10 and 11 are listed in

table 13 for both the 2 and 4 hour mission duration.

NASA/CR—2003-212350 27
Fuel Pump

Check Valve

Pyrovalve

From Power System Bus

Propeller

Combustion

Chamber

Piston Expander Engine

Electrical Alternator

To Power System Bus

Fuel

Pyrovalve Oxidizer

Fuel Pump

Figure 11 Bipropellant Piston Expander Propulsion System

Duration

(hours) 1 2 4

Fuel /

Oxidizer

Flow

Rate

(kg/hr)

Volume

Fuel /

Oxidizer

(m3

)

Mass

Fuel /

Oxidizer

(kg)

Volume

Fuel /

Oxidizer

(m3)

Mass

Fuel /

Oxidizer

(kg)

Volume

Fuel /

Oxidizer

(m3)

Mass

Fuel /

Oxidizer

(kg)

MMH /

NTO

9.86 /

17.26

0.0113 /

0.012

9.86 /

17.26

0.023 /

0.0238

19.72

/34.51

0.0451 /

0.0476

39.45 /

69.04

UMDH /

NTO

7.02 /

15.79

0.0089 /

0.011

7.02 /

15.79

0.0177 /

0.0218

14.03 /

31.56

0.0354 /

0.0436

28.08 /

63.17

Hydrazine 14.43 0.014 14.43 0.0286 28.86 0.0573 57.71

Hydrogen

/ Oxygen

1.95 /

15.58

0.106 /

0.0415

1.93 /

15.58

0.2123 /

0.083

3.896 /

31.164

0.4245 /

0.166

7.791 /

62.328

Table 12 Fuel and Oxidizer Specifications for Piston Expander Engine

NASA/CR—2003-212350 28
Fuel / Oxidizer Hydrazine MMH and NTO UMDH and NTO Hydrogen and Oxygen

Duration 1 hr 2 hr 4 hr 1 hr 2 hr 4 hr 1 hr 2 hr 4 hr 1 hr 2 hr 4 hrFuel Tank 0.29 0.58 1.16 0.23 0.46 0.91 0.18 0.36 0.71 5.50 11.00 22.02Oxidizer Tank NA NA NA 0.24 0.48 0.96 0.22 0.44 0.88 2.15 4.30 8.62Pyro Valve 0.3 0.3 0.3 0.6 0.6 0.6 0.6 0.6 0.6 0.6 0.6 0.6Fuel Pump 0.35 0.35 0.35 0.7 0.7 0.7 0.7 0.7 0.7 4.8* 4.8* 4.8*Check Valve 0.44 0.44 0.44 0.88 0.88 0.88 0.88 0.88 0.88 0.88 0.88 0.88CombustionChamber

0.1 0.1 0.1 0.13 0.13 0.13 0.11 0.11 0.11 0.23 0.23 0.23

Engine 4.8 4.8 4.8 4.8 4.8 4.8 4.8 4.8 4.8 4.8 4.8 4.8Gearbox 1.63 1.63 1.63 1.63 1.63 1.63 1.63 1.63 1.63 1.63 1.63 1.63Propeller 9.8 9.8 9.8 9.8 9.8 9.8 9.8 9.8 9.8 9.8 9.8 9.8Total (dry mass) 17.71 18.00 18.94 19.01 19.48 20.41 18.92 18.92 20.11 30.39 38.04 53.38Total (wet mass) 32.14 46.86 76.65 46.13 73.71 128.90 41.73 64.91 111.36 47.90 73.10 123.50

Table 13 Piston Expander Engine System Masses (kg) (note * indicates pressure regulator instead of fuel pump)

NASA/CR—2003-212350 29
Conventional Internal Combustion Engine

The internal combustion (IC) engine option is similar to a conventional aircraft IC engine

except that it would need to be designed to operate on a bipropellant fuel system and

within a low atmospheric density environment. Unlike the piston expander engine

described above, a monopropellant would not be usable since the combustion of the fuel

needs to take place within the cylinder volume in order for the engine to operate. Piston

internal combustion engines are very reliable, well understood pieces of machinery.

Conventional aircraft engines are too large for this application. However small IC

engines are used regularly for model aircraft. Examples are shown in figures 12 and 13.

For this analysis, both 2 and 4 cycle engines were examined. Although they are both

classified as IC engines there are significant differences in their operation and

performance. In general, 4 cycle engines are more efficient but heavier than 2 cycle

engines. The specific power and efficiencies for each are given in table 14.

4 Cycle Engine 2 Cycle Engine

Specific Power 1.06 kW/kg 1.64 kW/kg

Thermal Efficiency 25% 15%

Table 14 IC Engine Characteristics [19,20]

Figure 12 Fuji BT-86 Two Cycle Engine [19]

NASA/CR—2003-212350 30
Figure 13 O.S. Engines FF320 4 Cycle Engine [20]

Fuel and oxidizer flow rates were calculated using the engine performance and fuel

specifications listed in table 11. From these flow rates the required mass and volume of

the fuel and oxidizer were calculated. These are shown below in table 15.

The layout of the internal combustion system is the same as that shown in figure 11 for

the bipropellant piston expander engine except the engine would be a 2 or 4 cycle internal

combustion engine. Based on that diagram, the IC engine performance specifications

given in table 14 and the fuel / oxidizer specifications shown in table 15; component

mass estimates were determined. These mass estimates are shown in table 16 and 17 for

the 2 and 4 cycle engines respectively. As with the piston expander system, the liquid

fuels and oxidizers are stored in titanium 690 kPa (100 psi) tanks with a factor of safety

of 1.8. The gas storage tanks are composed of wound carbon fiber with an inner metal

liner with a storage pressure of 20.68 MPa (3000 psi) and a factor of safety of 1.8.

NASA/CR—2003-212350 31
Duration (hours) 1 2 4

Fuel / Oxidizer Flow Rate

(kg/hr)

Volume:Fuel /Oxidizer(m3)

Mass: Fuel /Oxidizer(kg)

Volume:Fuel /Oxidizer(m3)

Mass: Fuel /Oxidizer (kg)

Volume:Fuel /Oxidizer(m3)

Mass: Fuel /Oxidizer (kg)

4 Cycle Engine

MMH / NTO 3.96 / 6.93 0.0046 /

0.0048

3.96 / 6.93 0.0091 /

0.0096

7.92 / 13.86 0.0181 /

0.019

15.84 / 27.72

UMDH / NTO 2.8 / 6.32 0.0036 /

0.0044

2.81 / 6.32 0.0071 /

0.0087

5.615 / 12.633 0.0142 /

0.0174

11.23 / 25.27

Hydrogen /Oxygen

0.78 / 6.24 0.043 /

0.017

0.78 / 6.24 0.085 /

0.033

1.56 / 12.47 0.17 / 0.066 3.12 / 24.93

2 Cycle Engine

MMH / NTO 6.60 / 11.54 0.0075 /

0.0080

6.60 / 11.51 0.015 /

0.0159

13.19 / 23.01 0.030 /

0.032

26.38 / 46.17

UMDH / NTO 4.68 / 10.53 0.006 /

0.0073

4.68 / 10.53 0.012 /

0.0145

9.36 / 21.06 0.024 /

0.029

18.72 / 42.12

Hydrogen /Oxygen

1.3 / 10.39 0.071 /

0.028

1.30 / 10.39 0.142 /

0.055

2.60 / 20.78 0.284 /

0.111

5.21 / 41.55

Table 15 2 and 4 Cycle Engine Fuel and Oxidizer Specifications

NASA/CR—2003-212350 32
Duration 1 hour 2 hours 4 hours

Fuel / Oxidizer MMH &

NTO

UMDH& NTO

Hydrogen&Oxygen

MMH&NTO

UMDH &NTO

Hydrogen&Oxygen

MMH &NTO

UMDH& NTO

Hydrogen&Oxygen

Fuel Tank 0.15 0.12 3.69 0.30 0.24 7.37 0.61 0.48 14.73Oxidizer Tank 0.17 0.15 1.43 0.32 0.29 2.85 0.65 0.58 5.75Pyro Valve 0.6 0.6 0.6 0.6 .06 .06 0.6 0.6 0.6Fuel Pump 0.7 0.7 4.8* 0.7 0.7 4.8* 0.7 0.7 4.8*Check Valve 0.88 0.88 0.88 0.88 0.88 0.88 0.88 0.88 0.88Engine 4.75 4.75 4.75 4.75 4.75 4.75 4.75 4.75 4.75Gearbox 1.63 1.63 1.63 1.63 1.63 1.63 1.63 1.63 1.63Propeller 9.8 9.8 9.8 9.8 9.8 9.8 9.8 9.8 9.8Total (dry mass) 18.68 18.63 27.58 18.98 18.89 36.78 19.62 19.42 42.94Total (wet mass) 36.79 33.84 39.27 55.90 49.31 56.06 92.17 80.26 89.70

Table 16 2 Cycle Engine System Masses (kg) (note * indicates pressure regulator instead of fuel pump)

NASA/CR—2003-212350 33
Duration 1 hour 2 hours 4 hours

Fuel / Oxidizer MMH &

NTO

UMDH& NTO

Hydrogen& Oxygen

MMH &NTO

UMDH &NTO

Hydrogen& Oxygen

MMH &NTO

UMDH& NTO

Hydrogen&Oxygen

Fuel Tank 0.10 0.08 2.21 0.19 0.15 4.41 0.37 0.29 8.82Oxidizer Tank 0.11 0.09 0.86 0.20 0.18 1.71 0.39 0.35 3.42Pyro Valve 0.6 0.6 0.6 0.6 0.6 0.6 0.6 0.6 0.6Fuel Pump 0.7 0.7 4.8* 0.7 0.7 4.8* 0.7 0.7 4.8*Check Valve 0.88 0.88 0.88 0.88 0.88 0.88 0.88 0.88 0.88Engine 7.3 7.30 7.30 7.30 7.30 7.30 7.30 7.30 7.30Gearbox 1.63 1.63 1.63 1.63 1.63 1.63 1.63 1.63 1.63Propeller 9.8 9.8 9.8 9.8 9.8 9.8 9.8 9.8 9.8Total (dry mass) 21.12 21.08 28.08 21.30 21.24 29.81 21.67 21.55 37.25Total (wet mass) 32.01 30.21 35.10 43.08 39.49 43.84 65.23 58.05 65.30

Table 17 4 Cycle Engine System Masses (kg) (note * indicates pressure regulator instead of fuel pump)

NASA/CR—2003-212350 34
Combustion Engine Summary

Figures 14 through 16 show the mass comparison of the 3 types of combustion engines

that were considered in this analysis for 1, 2 and 4 hour mission durations.

NASA/CR—2003-212350 35
Figure 14 Combustion Engine System Mass Breakdown for 1 Hour Duration Mission

50

5 04 54 03 53 02 52 01 51 0

Mass (kg)

Hydrazine MMH & NTO UMDH &

NTO

H2 & O2 MMH & NTO UMDH &

NTO

H2 & O2 MMH & NTO UMDH &

NTO

H2 & O2

1 Hour Mission Duration .

Oxidizer FuelPropellerGearboxEngineCombustion ChamberCheck ValveFuel PumpPyro ValveOxidizer TankFuel Tank

Piston Expander 2 Cycle IC Engine 4 Cycle IC Engine

NASA/CR—2003-212350 36
Figure 15 Combustion Engine System Mass Breakdown for 2 Hour Duration Mission

0

8 07 06 05 04 03 02 01 0

Mass (kg)

Hydrazine MMH &

NTO

UMDH &

NTO

H2 & O2 MMH &

NTO

UMDH &

NTO

H2 & O2 MMH &

NTO

UMDH &

NTO

H2 & O2

2 Hour Mission Duration . 

Oxidizer FuelPropellerGearboxEngineCombustion ChamberCheck ValveFuel PumpPyro ValveOxidizer TankFuel Tank

Piston Expander 2 Cycle IC Engine 4 Cycle IC Engine

NASA/CR—2003-212350 37
Figure 16 Combustion Engine System Mass Breakdown for 4 Hour Duration Mission

0

8 06 04 02 0

140120100

Mass (kg)

Hydrazine MMH &

NTO

UMDH &

NTO

H2 & O2 MMH &

NTO

UMDH &

NTO

H2 & O2 MMH &

NTO

UMDH &

NTO

H2 & O2

4 Hour Mission Duration . Oxidizer 

FuelPropellerGearboxEngineCombustion ChamberCheck ValveFuel PumpPyro ValveOxidizer TankFuel Tank

Piston Expander 2 Cycle IC Engine 4 Cycle IC Engine

NASA/CR—2003-212350 38
Rocket Propulsion System

Rocket propulsion systems are a very reliable and controllable means of generating

thrust. There is a long history of rocket powered aircraft such as the X-1, which was the

first aircraft to break the speed of sound, and the X-15, which was the most noted and

capable rocket powered aircraft. Due to their high fuel consumption, rocket-powered

aircraft have been mostly used as experimental aircraft or to accomplish a specific goal.

For Mars flight a rocket-powered aircraft provides some significant advantages over

more conventionally propelled aircraft. A rocket system should have no trouble working

within the rarefied atmosphere of Mars. The risk associated with a rocket system will be

the lowest of all the types of propulsion systems considered.

To evaluate the rocket propulsion system, a number of system 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.

Propellant and Performance

There are a number of propellants that can potentially be used to power a rocketpropelled aircraft within the Mars environment. It should be noted that the term

propellant means both the fuel and oxidizer combination for a bipropellant system.

Ideally a propellant with a low freezing point (below –40 °C) would be used in order to

eliminate the need for active heating of the propellant throughout the mission. Table 18

lists a number of bipropellants and monopropellants that may be capable of meeting the

temperature requirements as well as being stable throughout the mission. Because of the

long transit time to Mars, no cryogenic propellants were considered. Only propellants

normally in the liquid state during storage were considered.

To maximize performance, the highest specific impulse propellant should be chosen.

Using the specific impulse (Isp) of a given type of propellant, the mass flow required by

the rocket to generate a given amount of thrust can be calculated. The thrust generated

(T) is given by the mass flow (M) out of the rocket nozzle multiplied by the exhaust

velocity (ue).

T = M ue (10)

NASA/CR—2003-212350 39
Propellant Oxidizer /

Fuel Ratio

Specific

Impulse

Isp

Freezing

Point

(°C)

Combustion

Temperature

(°C)

Fuel: Monomethyl Hydrazine

Oxidizer: Nitrogen Tetroxide

(MON25)

1.625 295 –52.5: F

–54.0: O

3122

Fuel: Monomethyl Hydrazine

Oxidizer: Chlorine Trifluoride

3.00 283 –52.5: F

–76.6: O

3318

Fuel: Monomethyl Hydrazine

Oxidizer:IRFNA

2.50 274 –52.5: F

–62.2: O

2848

Fuel: UDMH

Oxidizer: Nitrogen Tetroxide

2.70 286 –57.2: F

–54.0: O

3162

Fuel: UDMH

Oxidizer: Chlorine Trifluoride

2.85 278 –57.2: F

–76.6: O

3306

Fuel: RP-1

Oxidizer: IRFNA

4.90 263 –40.0: F

–62.2: O

2881

Monopropellant: Hydrazine Na 199 1.5 633

Monopropellant: Hydrogen Peroxide na 148 –11.5 757

Monopropellant: Ethylene Oxide na 199 –112.8 1004

Monopropellant: Nitromethane na 245 –29.0 2192

Monopropellant: n-Propyl Nitrate na 209 –101.1 1077

Monopropellant: HPB-2517 na 220 –54.0 ----

Monopropellant:

hydroxylammonium nitrate (HAN),

glycine and water

na 190 –35.0 ----

Table 18 Low Temperature Propellant Candidates [13,14,15]

The exhaust velocity (ue) can be calculated with the Isp of the propellant and the

gravitational constant (g) of 9.81 m/s. It should be noted that the gravitational constant is

not dependent on the location of flight (i.e. Mars) but on the location where the Isp was

calculated. Therefore the Earth’s gravitational constant is used regardless of where the

rocket is actually operated.

ue = Isp g (11)

The above relationships assume that an ideal nozzle is used and that the flow is fully

expanded to atmospheric conditions upon exiting the rocket nozzle. Based on this brief

analysis, the mass flow and total mass of propellant required can be calculated for a thrust

of 35N and flight durations of 1, 2 and 4 hours. The results are shown in table 19.

NASA/CR—2003-212350 40
Propellant Specific

Impulse

Isp

ExhaustVelocity

(m/s)

Mass

Flow

(kg/s)

Total

Propellant

Mass (kg)

1 Hour

Flight

Total

Propellant

Mass (kg)

2 Hour

Flight

Total

Propellant

Mass (kg)

4 Hour

Flight

Fuel: Monomethyl Hydrazine

Oxidizer: Nitrogen Tetroxide (MON25)

295 2893 0.0121 43.5 87.0 174.0

Fuel: Monomethyl Hydrazine

Oxidizer: Chlorine Trifluoride

283 2776 0.0126 45.3 90.5 181.0

Fuel: Monomethyl Hydrazine

Oxidizer:IRFNA

274 2687 0.0130 46.8 93.5 187.0

Fuel: UDMH

Oxidizer: Nitrogen Tetroxide

286 2806 0.0125 44.8 89.5 179.0

Fuel: UDMH

Oxidizer: Chlorine Trifluoride

278 2727 0.0128 46.0 92.0 184.0

Fuel: RP-1

Oxidizer: IRFNA

263 2580 0.0135 48.8 97.5 195.0

Monopropellant: Hydrazine 199 1952 0.0179 64.5 129.0 258.0

Monopropellant: Hydrogen Peroxide 148 1451 0.0241 86.8 173.5 347.0

Monopropellant: Ethylene Oxide 199 1952 0.0179 64.5 129.0 258.0

Monopropellant: Nitromethane 245 2403 0.0146 52.3 104.5 209.0

Monopropellant: n-Propyl Nitrate 209 2050 0.0170 61.3 122.5 245.0

Monopropellant: HPB-2517 220 2158 0.0162 58.3 116.5 233.0

Monopropellant: hydroxylammonium nitrate

(HAN), glycine and water

190 1864 0.0188 67.7 135.4 270.7

Table 19 Propellant Mass for 35N Thrust

NASA/CR—2003-212350 41
MMH/MON25

Of the propellants listed in table 16, MON25 seems to be the best choice. Although not a

standard rocket fuel propellant, this propellant provides a high ISP (up to 297s in testing),

good thermal characteristics for the Mars environment and is usable in present day

thrusters [2].

MON25 is the oxidizer which is made up of nitrogen tetroxide with a 25% NO content.

The highest specific impulse values (in combination with MMH) occur with mixture

ratios of approximately 1.55. The density (ρ) as a function of temperature (T) of MON25

and MMH is given by the following equations and shown in figure 17 [16,17,21].

MON25:

ρ = 1.4308 – 0.0021402T – 4.7391E-6 T2 (12)

MMH:

ρ = 0.895 – 0.001T + 1.2358E-11 T2 (13)

The specific heat (Cp) of hydrazine (MMH) can be represented by the following relation,

where T is the temperature in degrees Kelvin.

Cp = 2.7313-7.2316E-5 T+ 1.6377E-6 T2 [J / (g K)] (14)

The products of the reaction between MMH and NTO can be of interest due to potential

interaction with science instrumentation or other equipment on board the aircraft.

Generally if the aircraft is following a straight path there will be little or no

contamination by the exhaust gasses. However, if the aircraft circles back it could

potentially fly through its exhaust stream (albeit dispersed). This may be of concern for

the science instrumentation and data gathering.

The reaction of MMH and NTO yields methylamine, dimethylamine, Nnitrosomethylamine, N-methylformamide, methanol, methylhydrazinium nitrate, water,

nitric oxide, nitrous oxide, nitrogen and carbon dioxide [14].

These combustion products are for MON3 not MON25. Because MON25 is an

experimental oxidizer and not presently used in production spacecraft, its combustion

products are not well documented. However, since the differences between MON25 and

MON3 is the concentration of NO there may be differences in the concentration of

nitrogen containing compounds but the products should be similar.

NASA/CR—2003-212350 42
Figure 17 Density of MON25 and MMH as a Function of Temperature [3]

Rocket System Layout: Simple Blowdown System

The simple blowdown system is the most basic and simple type of rocket propulsion

system. It consists of propellant tanks (one or two depending on whether it is a

monopropellant or bipropellant system) pressurized with helium gas, pyrovalves, and the

thrusters. A layout of this system is shown in figure 18 for a bipropellant fueled system.

The system is launched with propellant up to the pyrovalves. Pyrovalves are used to

prevent vapor migration through the thruster and into the aeroshell (potentially a

catastrophic situation). When the system is activated, the normally-closed pyrovalves are

fired, bringing propellant to the thruster valves. The system would be controlled by

cycling the thruster valves on and off. Because the system is fully pressurized at the start

of the mission, the thrust will steadily decrease throughout the aircraft flight. The

propellant tanks would need to be large enough to accommodate the fuel as well as the

helium pressurant. They could be made smaller if higher pressures were used, but that

NASA/CR—2003-212350 43
means a higher initial thrust. To further reduce the system mass, the line filters,

thermocouples and pressure transducers can be eliminated. This, however, will increase

operational risk.

Because of the non-uniform thrust profile, which will cause the aircraft to be either over

or under powered throughout most of the flight, and increased risk, this type of system

was deemed not viable for the Mars aircraft application.

Rocket System: Regulated

The regulated system is more complicated and has greater mass then the blowdown

system but enables an even thrust level throughout the flight. This type of system, with

no active control, is shown in figures 19 and 22 for a bipropellant and monopropellant

system respectively. This system has pressurized helium that is used to pressurize the fuel

and oxidizer tanks. The helium flow into these tanks is controlled by a pressure regulator

which is located on the lines to the fuel and oxidizer tanks. A normally closed pyrovalve

is used to seal the helium pressurant from the propellant tanks and to seal the propellant

tanks from the thruster during transit to Mars. Check valves and filters are placed in the

fuel and oxidizer lines. The system has a transducer for the chamber pressure (thruster

performance) and thermocouples for the propellant tanks. The fill lines for the helium

and propellant tanks are sealed with normally open pyrovalves which are activated once

the system is serviced prior to launch, thereby sealing the access lines. There are

checkout lines that are located after each of the pyrovalves. These lines are used to

pressure check the system and are sealed manually after the testing is finished.

With the addition of on / off control valves on the thruster, the thruster can be pulsed.

This provides some control over the integrated thrust over a period of time. By cycling

the thruster on and off the aircraft would glide for a period of time and then resume

powered flight. This system is shown in figure 20 for a bipropellant system and figure 23

for a monopropellant system.

For greater control, a variable control valve can be used to vary the thrust of the rocket by

regulating the helium flow to the fuel / oxidizer tanks. This enables a system that can be

throttled providing variable thrust. This effectively controls the pressure within these

tanks and therefore the flow to the engine. However, the amount of thrust regulation

achievable will depend on the capabilities of the thruster. To achieve complete thrust

throttling capability a variable throat nozzle is required. This adds significant complexity

and cost to the design. The variable thrust system is shown in figures 21 and 25 for

bipropellant and monopropellant systems respectively

The propellant candidates chosen for analysis with the bipropellant and monopropellant

systems is MON-25 / MMH and Hydrazine respectively. The MON-25 / MMH fuel

oxidizer combination was chosen because it has a high ISP value and is capable of

remaining liquid and operating at temperatures down of less the –40 °C. The low

temperature operation eliminates the need for an extensive thermal control system to

maintain propellant temperature throughout the mission. The monopropellant hydrazine

was chosen because it is the standard monopropellant used for space applications. The

other monopropellants, although they have better performance, would require an engine

development effort that would add significant cost and time to any aircraft program. The

NASA/CR—2003-212350 44
main advantage of hydrazine is that it has good restart characteristics. However, due to its

high freezing point, a thermal control system will need to be utilized throughout the

mission to insure it does not freeze. A layout of the regulated system components is

shown in figures 19 through 23 for both bipropellant (MON-25 / MMH) and

monopropellant (hydrazine) systems with and without no / off control capability [22,23].

NASA/CR—2003-212350 45
Figure 18 Blowdown Bipropellant Rocket System

PP

Pressure Transducer P Line Filter

TT

Thermocouple T

Nitrogen Tetroxide (MON-25)Pressurized with He

Monomethyl Hydrazine Tank Pressurized with He

On / Off Latch Valve Service Valve

Normally ClosedPyrovalve

Normally OpenPyrovalve

NCNO

NCNC

NO

NO

NASA/CR—2003-212350 46
Monomethyl Hydrazine Tank 

Nitrogen Tetroxide (MON-25)

Helium Tank P

TT

PP

Line Filter

Thermocouple Pressure Regulator

Check Valve

T

Pressure Relief Valve

Service Valve

Bulkhead Pressure Burst Disk

Normally ClosedPyrovalve

Normally OpenPyrovalve

Pressure Transducer P

NCNO

NCNC

NC

NO

NO

NO

Manual Valve

NO

NO

Figure 19 Regulated Bipropellant System with No Active Control

NASA/CR—2003-212350 47
Line Filter

Pressure Transducer Thermocouple

Pressure Regulator

Check Valve

P T

Pressure Relief Valve Service Valve Bulkhead

Thruster On/Off ValvePressure Burst Disk

Monomethyl Hydrazine Tank 

Nitrogen Tetroxide (MON-25)

Helium Tank P

TT

PP

NC

NC

NC

NO

NO

NO

NO

NO

Normally ClosedPyrovalve

Normally OpenPyrovalve

NCNO

Manual Valve

Figure 20 Regulated Bipropellant System with On / Off Control Rocket

NASA/CR—2003-212350 48
Line Filter Pressure Transducer

Thermocouple Pressure Regulator

Check Valve P

T Variable Control Latch Valve

Pressure Relief Valve Bulkhead Service Valve Pressure Burst Disk

Monomethyl Hydrazine Tank 

Nitrogen Tetroxide (MON-25)

Helium Tank P

TT

PP

NCNC

NC

NO

NO

NO

NO

NO

Manual Valve

Normally ClosedPyrovalve

Normally OpenPyrovalve

NCNO

Figure 21 Regulated Bipropellant System with Throttle Control

NASA/CR—2003-212350 49
Line Filter

Pressure Transducer Thermocouple Pressure Regulator

Check Valve

P T

Pressure Relief Valve Service Valve Bulkhead

Normally ClosedPyrovalve

NCNO

Normally OpenPyrovalve

Hydrazine Tank Helium Tank P

T

P

NC

NC

NO

NO

Manual Valve

NO

Figure 22 Regulated Monopropellant System with No Active Control

NASA/CR—2003-212350 50
Line Filter

Thermocouple Pressure Regulator

Check Valve

T

Pressure Relief Valve

On / Off Valve

Service Valve Bulkhead

Hydrazine Tank Helium Tank P

T

P

NC

NC

NO

NO

Manual Valve

Normally ClosedPyrovalve

NCNO

Normally OpenPyrovalve

Pressure Transducer P

NO

Figure 23 Regulated Monopropellant System with On / Off Control

NASA/CR—2003-212350 51
Line Filter

Pressure Transducer Thermocouple

Check Valve

P T

Pressure Relief Valve Service Valve Bulkhead

Variable Control Valve

Hydrazine Tank Helium Tank P

T

P

NC

NC

NO

NO

Manual Valve Normally Closed

Pyrovalve

NCNO

Normally OpenPyrovalve

NO

Figure 24 Regulated Monopropellant System with Throttle Control

NASA/CR—2003-212350 52
A breakdown of the masses and volumes for the regulated bi-propellant and

monopropellant systems is given in tables 20 through 22 for flight durations of 1, 2 and

4 hours with a continuous thrust level of 35 N. The pressure of the helium tank was

assumed to be 20.684 MPa (3000 psi) and the propellant tank pressure was 2.068 MPa

(300 psi). The helium tank was assumed to be spherical in shape and made of Arimid

fiber composite with a thin metal liner. An additional 10% volume of helium was carried

to account for the line volume within the system. The propellant tanks were assumed to

be made of titanium and also spherical in shape. Carbon tanks can be used, however a

development and qualification program would need to be performed since carbon is not

the standard material used for space-qualified propellant tanks. The spherical shape was

chosen for simplicity in the analysis. The actual tanks may need to conform to a specific

shape within the aircraft fuselage. A factor of safety of 1.8 was used for all tank sizing.

System volume as well as mass is a concern for the feasibility of the propulsion systems.

This is especially true for the rocket system since the propellant volume can be

considerable. Tables 23 through 25 list the gross volume of the major components for the

rocket system for 1, 2 and 4 hour durations. This volume is based on the maximum

outside envelope of the component. The line volume is for 3 m of connection lines for the

bipropellant system and 1.5 m for the monopropellant system.

NASA/CR—2003-212350 53
Component Bipropellant

No Control

BipropellantOn / OffControl

BipropellantThrottleControl

MonopropellantNo Control

MonopropellantOn / OffControl

MonopropellantThrottle Control

Helium Tank 0.528 kg 0.528 kg 0.528 kg 0.889 kg 0.889 kg 0.889 kgMMH or Hydrazine Tank 1.06 kg 1.06 kg 1.06 kg 3.56 kg 3.56 kg 3.56 kgNTO Tank 1.06 kg 1.06 kg 1.06 kg NA NA NAPressure Transducer (3) 0.9 kg (3) 0.9 kg (3) 0.9 kg (2) 0.60 kg (2) 0.60 kg (2) 0.60 kgPyrovalve normallyclosed

(4) 1.2 kg (4) 1.2 kg (4) 1.2 kg (2) 0.60 kg (2) 0.60 kg (2) 0.60 kg

Pyro Valve normallyopen

(5) 2.15 kg (5) 2.15 kg (5) 2.15 kg (3) 1.29 kg (3) 1.29 kg (3) 1.29 kg

Pressure Regulator (fixedor variable)

(1) 0.34 kg (1) 0.34 kg (2) 1.5 kg (1) 0.34 kg (1) 0.34 kg (1) 0.75 kg

Check Valve (4) 0.16 kg (4) 0.16 kg (4) 0.16 kg (2) 0.08 kg (2) 0.08 kg (2) 0.08 kgThermocouple (2) 0.04 kg (2) 0.04 kg (2) 0.04 kg (1) 0.02 kg (1) 0.02 kg (1) 0.02 kgLine Filter (2) 0.16 kg (2) 0.16 kg (2) 0.16 kg (1) 0.08 kg (1) 0.08 kg (1) 0.08 kgThruster (1) 1.50 kg (1) 2.50 kg (1) 2.50 kg (1) 1.50 kg (1) 2.50 kg (1) 2.50 kgOn / Off Valve (3) 1.50 kg (3) 1.50 kg (3) 1.50 kg (2) 1.00 kg (2) 1.00 kg (1) 1.00 kgLines and Fittings ~1.0 kg ~1.0 kg ~1.0 kg ~0.50 kg ~0.50 kg ~0.50 kgPressure Relief Valve (2) 1.05 kg (2) 1.05 kg (2) 1.05 kg (1) 0.53 kg (1) 0.53 kg (1) 0.53 kgBurst Disc (2) 0.1 kg (2) 0.1 kg (2) 0.1 kg NA NA NA

Total Dry Mass 12.75 kg 13.75 kg 14.91 kg 10.99 kg 11.99 kg 12.40 kg

MMH or Hydrazine 16.57 kg 16.57 kg 16.57 kg 64.50 kg 64.50 kg 64.50 kgNitrogen Tetroxide 26.93 kg 26.93 kg 26.93 kg NA NA NA

Total Wet Mass 56.25 kg 57.25 kg 58.41 kg 75.49 kg 76.49 kg 76.90 kg

Table 20 Rocket System Mass Breakdown for 1 Hour Flight Duration [3,4,24,25,26,27,28,29,30,31]

NASA/CR—2003-212350 54
Component Bipropellant

No Control

BipropellantOn / OffControl

BipropellantThrottleControl

MonopropellantNo Control

MonopropellantOn / OffControl

MonopropellantThrottleControl

Helium Tank 1.03 kg 1.03 kg 1.03 kg 1.78 kg 1.78 kg 1.78 kgMMH or Hydrazine Tank 2.06 kg 2.06 kg 2.06 kg 7.11 kg 7.11 kg 7.11 kgNTO Tank 2.06 kg 2.06 kg 2.06 kg NA NA NAPressure Transducer (3) 0.9 kg (3) 0.9 kg (3) 0.9 kg (2) 0.60 kg (2) 0.60 kg (2) 0.60 kgPyrovalve normallyclosed

(4) 1.2 kg (4) 1.2 kg (4) 1.2 kg (2) 0.60 kg (2) 0.60 kg (2) 0.60 kg

Pyro Valve normally open (5) 2.15 kg (5) 2.15 kg (5) 2.15 kg (3) 1.29 kg (3) 1.29 kg (3) 1.29 kgPressure Regulator (fixedor variable)

(1) 0.34 kg (1) 0.34 kg (2) 1.5 kg (1) 0.34 kg (1) 0.34 kg (1) 0.75 kg

Check Valve (4) 0.16 kg (4) 0.16 kg (4) 0.16 kg (2) 0.08 kg (2) 0.08 kg (2) 0.08 kgThermocouple (2) 0.04 kg (2) 0.04 kg (2) 0.04 kg (1) 0.02 kg (1) 0.02 kg (1) 0.02 kgLine Filter (2) 0.16 kg (2) 0.16 kg (2) 0.16 kg (1) 0.08 kg (1) 0.08 kg (1) 0.08 kgThruster (1) 1.50 kg (1) 2.50 kg (1) 2.50 kg (1) 1.50 kg (1) 2.50 kg (1) 2.50 kgOn / Off Valve (3) 1.50 kg (3) 1.50 kg (3) 1.50 kg (2) 1.00 kg (2) 1.00 kg (1) 1.00 kgLines and Fittings ~1.0 kg ~1.0 kg ~1.0 kg ~0.50 kg ~0.50 kg ~0.50 kgPressure Relief Valve (2) 1.05 kg (2) 1.05 kg (2) 1.05 kg (1) 0.53 kg (1) 0.53 kg (1) 0.53 kgBurst Disc (2) 0.1 kg (2) 0.1 kg (2) 0.1 kg NA NA NA

Total Dry Mass 15.25 kg 16.25 kg 17.41 kg 15.43 kg 16.43 kg 16.84 kg

MMH or Hydrazine 33.14 kg 33.14 kg 33.14 kg 129.00 kg 129.00 kg 129.00 kgNitrogen Tetroxide 53.86 kg 53.86 kg 53.86 kg NA NA NA

Total Wet Mass 102.25 kg 103.25 kg 104.41 kg 159.86 kg 145.43 kg 145.84 kg

Table 21 Rocket System Mass Breakdown for 2 Hour Flight Duration [3,4,24,25,26,27,28,29,30,31]

NASA/CR—2003-212350 55
Component Bipropellant

No Control

BipropellantOn / OffControl

BipropellantThrottleControl

MonopropellantNo Control

MonopropellantOn / OffControl

MonopropellantThrottleControl

Helium Tank 2.06 kg 2.06 kg 2.06 kg 3.55 kg 3.55 kg 3.55 kgMMH or Hydrazine Tank 4.11 kg 4.11 kg 4.11 kg 14.23 kg 14.23 kg 14.23 kgNTO Tank 4.11 kg 4.11 kg 4.11 kg NA NA NAPressure Transducer (3) 0.9 kg (3) 0.9 kg (3) 0.9 kg (2) 0.60 kg (2) 0.60 kg (2) 0.60 kgPyrovalve normallyclosed

(4) 1.2 kg (4) 1.2 kg (4) 1.2 kg (2) 0.60 kg (2) 0.60 kg (2) 0.60 kg

Pyro Valve normally open (5) 2.15 kg (5) 2.15 kg (5) 2.15 kg (3) 1.29 kg (3) 1.29 kg (3) 1.29 kgPressure Regulator (fixedor variable)

(1) 0.34 kg (1) 0.34 kg (2) 1.5 kg (1) 0.34 kg (1) 0.34 kg (1) 0.75 kg

Check Valve (4) 0.16 kg (4) 0.16 kg (4) 0.16 kg (2) 0.08 kg (2) 0.08 kg (2) 0.08 kgThermocouple (2) 0.04 kg (2) 0.04 kg (2) 0.04 kg (1) 0.02 kg (1) 0.02 kg (1) 0.02 kgLine Filter (2) 0.16 kg (2) 0.16 kg (2) 0.16 kg (1) 0.08 kg (1) 0.08 kg (1) 0.08 kgThruster (1) 1.50 kg (1) 2.50 kg (1) 2.50 kg (1) 1.50 kg (1) 2.50 kg (1) 2.50 kgOn / Off Valve (3) 1.50 kg (3) 1.50 kg (3) 1.50 kg (2) 1.00 kg (2) 1.00 kg (1) 1.00 kgLines and Fittings ~1.0 kg ~1.0 kg ~1.0 kg ~0.50 kg ~0.50 kg ~0.50 kgPressure Relief Valve (2) 1.05 kg (2) 1.05 kg (2) 1.05 kg (1) 0.53 kg (1) 0.53 kg (1) 0.53 kgBurst Disc (2) 0.1 kg (2) 0.1 kg (2) 0.1 kg NA NA NA

Total Dry Mass 16.27 kg 17.27 kg 18.43 kg 24.32 kg 25.32 kg 25.73 kg

MMH or Hydrazine 66.28 kg 66.28 kg 66.28 kg 258.00 kg 258.00 kg 258.00 kgNitrogen Tetroxide 107.72 kg 107.72 kg 107.72 kg NA NA NA

Total Wet Mass 190.27 kg 191.27 kg 192.43 kg 282.32 kg 283.32 kg 283.73 kg

Table 22 Rocket System Mass Breakdown for 4 Hour Flight Duration [3,4,24,25,26,27,28,29,30,31]

NASA/CR—2003-212350 56
Component Bipropellant

No Control

BipropellantOn / OffControl

BipropellantThrottle Control

MonopropellantNo Control

MonopropellantOn / Off Control

MonopropellantThrottle Control

Helium Tank 4,180 cm3 4,180 cm3 4,180 cm3 7,040 cm3 7,040 cm3 7,040 cm3MMH or HydrazineTank

19,000 cm3 19,000 cm3 19,000 cm3 64,000 cm3 64,000 cm3 64,000 cm3

NTO Tank 19,000 cm3 19,000 cm3 19,000 cm3 NA NA NAPressure Transducer 105 cm3 105 cm3 105 cm3 70 cm3 70 cm3 70 cm3Pyrovalve normallyclosed

257 cm3 257 cm3 257 cm3 129 cm3 129 cm3 129 cm3

Pyro Valve normallyopen

2,500 cm3 2,500 cm3 2,500 cm3 1,500 cm3 1,500 cm3 1,500 cm3

Pressure Regulator(fixed or variable)

938 cm3 938 cm3 1,876 cm3 938 cm3 938 cm3 938 cm3

Check Valve 128 cm3 128 cm3 128 cm3 64 cm3 64 cm3 64 cm3Thermocouple 30 cm3 30 cm3 30 cm3 15 cm3 15 cm3 15 cm3Line Filter 136 cm3 136 cm3 136 cm3 68 cm3 68 cm3 68 cm3Thruster 368 cm3 860 cm3 860 cm3 368 cm3 860 cm3 860 cm3On / Off Valve 369 cm3 369 cm3 369 cm3 246 cm3 246 cm3 246 cm3Lines and Fittings 216 cm3 216 cm3 216 cm3 108 cm3 108 cm3 108 cm3Pressure Relief Valve 32 cm3 32 cm3 32 cm3 16 cm3 16 cm3 16 cm3Burst Disc 64 cm3 64 cm3 64 cm3 NA NA NA

Total Volume 47,323 cm3 47,815 cm3 48,753 cm3 74,562 cm3 75,054 cm3 75,054 cm3

Table 23 Rocket System Volume Breakdown for 1 Hour Flight Duration [3,4,24,25,26,27,28,29,30,31]

NASA/CR—2003-212350 57
Component Bipropellant

No Control

BipropellantOn / OffControl

BipropellantThrottle Control

MonopropellantNo Control

MonopropellantOn / Off Control

MonopropellantThrottle Control

Helium Tank 8,140 cm3 8,140 cm3 8,140 cm3 14,080 cm3 14,080 cm3 14,080 cm3MMH or HydrazineTank

37,000 cm3 37,000 cm3 37,000 cm3 128,000 cm3 128,000 cm3 128,000 cm3

NTO Tank 37,000 cm3 37,000 cm3 37,000 cm3 NA NA NAPressure Transducer 105 cm3 105 cm3 105 cm3 70 cm3 70 cm3 70 cm3Pyrovalve normallyclosed

257 cm3 257 cm3 257 cm3 129 cm3 129 cm3 129 cm3

Pyro Valve normallyopen

2,500 cm3 2,500 cm3 2,500 cm3 1,500 cm3 1,500 cm3 1,500 cm3

Pressure Regulator(fixed or variable)

938 cm3 938 cm3 1,876 cm3 938 cm3 938 cm3 938 cm3

Check Valve 128 cm3 128 cm3 128 cm3 64 cm3 64 cm3 64 cm3Thermocouple 30 cm3 30 cm3 30 cm3 15 cm3 15 cm3 15 cm3Line Filter 136 cm3 136 cm3 136 cm3 68 cm3 68 cm3 68 cm3Thruster 368 cm3 860 cm3 860 cm3 368 cm3 860 cm3 860 cm3On / Off Valve 369 cm3 369 cm3 369 cm3 246 cm3 246 cm3 246 cm3Lines and Fittings 216 cm3 216 cm3 216 cm3 108 cm3 108 cm3 108 cm3Pressure Relief Valve 32 cm3 32 cm3 32 cm3 16 cm3 16 cm3 16 cm3Burst Disc 64 cm3 64 cm3 64 cm3 NA NA NA

Total Volume 87,283 cm3 87,775 cm3 88,713 cm3 145,602 cm3 146,094 cm3 146,094 cm3

Table 24 Rocket System Volume Breakdown for 2 Hour Flight Duration [3,4,24,25,26,27,28,29,30,31]

NASA/CR—2003-212350 58
Component Bipropellant

No Control

BipropellantOn / OffControl

BipropellantThrottle Control

MonopropellantNo Control

MonopropellantOn / Off Control

MonopropellantThrottle Control

Helium Tank 16,280 cm3 16,280 cm3 16,280 cm3 28,160 cm3 28,160 cm3 28,160 cm3MMH or HydrazineTank

74,000 cm3 74,000 cm3 74,000 cm3 256,000 cm3 256,000 cm3 256,000 cm3

NTO Tank 74,000 cm3 74,000 cm3 74,000 cm3 NA NA NAPressure Transducer 105 cm3 105 cm3 105 cm3 70 cm3 70 cm3 70 cm3Pyrovalve normallyclosed

257 cm3 257 cm3 257 cm3 129 cm3 129 cm3 129 cm3

Pyro Valve normallyopen

2,500 cm3 2,500 cm3 2,500 cm3 1,500 cm3 1,500 cm3 1,500 cm3

Pressure Regulator(fixed or variable)

938 cm3 938 cm3 1,876 cm3 938 cm3 938 cm3 938 cm3

Check Valve 128 cm3 128 cm3 128 cm3 64 cm3 64 cm3 64 cm3Thermocouple 30 cm3 30 cm3 30 cm3 15 cm3 15 cm3 15 cm3Line Filter 136 cm3 136 cm3 136 cm3 68 cm3 68 cm3 68 cm3Thruster 368 cm3 860 cm3 860 cm3 368 cm3 860 cm3 860 cm3On / Off Valve 369 cm3 369 cm3 369 cm3 246 cm3 246 cm3 246 cm3Lines and Fittings 216 cm3 216 cm3 216 cm3 108 cm3 108 cm3 108 cm3Pressure Relief Valve 32 cm3 32 cm3 32 cm3 16 cm3 16 cm3 16 cm3Burst Disc 64 cm3 64 cm3 64 cm3 NA NA NA

Total Volume 169,423 cm3 169,915 cm3 170,853 cm3 287,682 cm3 288,174 cm3 288,174 cm3

Table 25 Rocket System Volume Breakdown for 4 Hour Flight Duration [3,4,24,25,26,27,28,29,30,31]

NASA/CR—2003-212350 59
Nozzle Sizing

In order to maximize the rocket engine performance as well as minimize the impact of

the exhaust plume on the science experiments, the nozzle should be designed to fully

expand the flow. This will minimize the movement of the exhaust gases from the exhaust

stream once they leave the nozzle. To fully expand the exhaust gases the exit pressure

(pe) of the exhaust stream would need to equal the ambient atmospheric pressure (pa). The

size of the nozzle necessary to fully expand the flow within the Mars environment can be

estimated by the following basic analysis. This analysis assumes a constant specific heat,

ideal gas and isentropic flow through the nozzle. The desired exit Mach number (Me) can

be found from equation 15, where po is the stagnation pressure in the combustion

chamber, pa is the atmospheric pressure and γ is the ratio of specific heats for the exhaust

gas. Since it is assumed that the exit pressure (Pe) will equal the ambient pressure (Pa) by

fully expanding the exhaust flow through the nozzle then the Mach number of the exhaust

gases at atmospheric conditions (Ma) is equal to the exit Mach number (Me) [32].

Therefore, Ma in equation 15 is equivalent to Me.

P

P

o M

a

= + a

− − ( ) 1 1

2

γ 2 1

γ

γ (15)

Once the exit Mach number is determined from equation 15, then the area ratio for the

nozzle can be calculated. This area ratio is the ratio between the nozzle throat (A*, point

where the flow within the nozzle reaches Mach 1) and the exit (Ae) is given by equation

16.

A

A M

e M

a * e ( ) = ( ( ))

+

+ − +

1 2 −

1

1 1

2

2

1

2 1

γ

γ

γ

γ (16)

The stagnation pressure just before the nozzle converging section will be about 95% of

that at the inlet to the combustion chamber. This is for a combustion chamber that is

uniform in area along its length and about 2 times the diameter of the nozzle throat.

Common inlet pressure into the combustion chamber is usually on the order of 2.07 MPa

(300 psi), however the exact pressure will depend on the rocket system design. Therefore

the stagnation pressure within the chamber (po) for use in determining the exit mach

number is 1.97 MPa (286 psi). For this initial sizing it will be assumed that the ratio of

the specific heats of the exhaust gases (γ) is 1.23. This is representative of the specific

heat for hydrazine combustion with nitrogen tetroxide.

Once the area ratio is calculated, the actual throat area will depend on the mass flow

required to produce the necessary thrust (T). For a fully expanded nozzle this throat area

is given by equation 17 [32].

NASA/CR—2003-212350 60
A T

p p

p o e

o

*

( ) (() )

=

− + −

+

−

−

2

1

2

1

1

2 1

1

1

γ

γ γ

γ

γ

γ

γ

(17)

The required length of the nozzle can be determined by assuming a standard 15° cone

angle for the nozzle in conjunction with the throat and exit radius values. Utilizing the

values and equations given above, the resultant geometry for the nozzle with a thrust

level of 35N is summarized in table 26.

Parameter Value

Exit Mach Number 5.49

Nozzle Area Ratio (Ae/A*) 153.67

Throat Area (A*) and Diameter 0.0942 cm2 / 0.346 cm

Exit Area (Ae) and Diameter 14.5 cm2 / 4.292 cm

Nozzle Length 7.37 cm

Table 26 Ideal Nozzle Geometry for Mars Atmosphere and 35 N Thrust

Off Design Rocket Engine Operation

As propellant is consumed during the flight, less thrust is needed. Therefore it would be

advantageous to reduce the thrust level throughout the flight. The thrust level of the

engine can be reduced by lowering the propellant tank pressure. The rocket systems

shown in figures 21 and 24, for bipropellant and monopropellant systems respectively,

would be capable of reducing the propellant tank pressure and thereby reducing the

engine thrust accordingly. A throttling down of approximately 30% should be achievable

by the rocket motor. This estimate of thrust reduction is based on the capabilities of a

similarly designed 8.9 N engine. The limiting factor on reducing the thrust of the rocket

engine (for a bipropellant system) is the fuel injectors. These injectors are designed to

operate at a specific propellant pressure and their performance will degrade as the

pressure is lowered. Increasing the thrust of the rocket engine is not advisable. Increasing

the rocket thrust beyond the design point will stress the engine both pressure-wise and

temperature-wise and can cause it to fail. The following analysis estimates the engine

performance for a thrust reduction from 35 N to 25 N (approximately 29%). This

reduction in thrust corresponds to an Isp decrease of from 295 to 255, respectively.

Since the geometry of the nozzle will remain the same as the combustion pressure is

reduced, the flow will become overexpanded within the nozzle. This means that the exit

pressure will be less then atmospheric pressure. Compression of the gas back to

atmospheric pressure will take place outside the nozzle through a series of shock waves.

The combustion chamber pressure for a given thrust level is given by equation 18 [32].

The gas constant (R) and combustion temperature (to) for the MMH and NTO propellant

was estimated to be 311 J/K Kg and 3395 K, respectively. The Isp was assumed to

NASA/CR—2003-212350 61
decrease linearly with thrust and is given in equation 19. The Isp value is based on the

Earth’s gravitational constant (ge) of 9.81 m/s2

.

P T PA

A

Rt I g A

m

o

a e

o

sp e

e

e

= +

− +

+ −

+

−

−

*

( )

( )

γ

γ γ

γ

γ

γ

γ

2

1

1 1

2

1

1

2 1

(18)

I T sp = + 4 155 (19)

Utilizing the combustion chamber pressure from equation 18, the mass flow of propellant

and exit pressure can be calculated. These are given by equations 20 and 21, respectively

[32].

m A p

Rt

o

o

. * = ( )

+

+

− γ

γ

γ

2 γ

1

1

1

 (20)

P

P

o M

e

= + e

− − ( ) 1 1

2

γ 2 1

γ

γ (21)

Because the flow through the nozzle is choked, which means there will be a normal shock

at the nozzle throat for the range of combustion chamber pressures encountered in this

analysis, the exit Mach number is based on the geometry of the nozzle and is given in

equation 21. This Mach number does not change with the variation in combustion

chamber pressure. The results of the above analysis are shown in figures 25 and 26.

Figure 25 shows the combustion chamber and exit pressure change over the thrust range

and figure 26 shows the mass flow and Isp change over the thrust range.

NASA/CR—2003-212350 62
0

500000

2500000200000015000001000000

35 34 33 32 31 30 29 28 27 26 25

Thrust (N)

Combustion Chamber Pressure (Pa)

900850800750700650600550500450400

Exit Pressure (Pa)

PoPe

Figure 25 Combustion and Exit Pressures vs Thrust for Off Design Engine Operation

NASA/CR—2003-212350 63
0.01

0.0105

0.011

0.0115

0.012

0.0125

35 34 33 32 31 30 29 28 27 26 25

Thrust (N)

Mass Flow (kg/s)

300290280270260250240

Isp (1/sec)

Mass FlowIsp

Figure 26 Mass Flow and Isp vs Thrust for Off Design Engine Operation

NASA/CR—2003-212350 64
Thruster Cycling

Controlling the amount of thrust can also be achieved by cycling the rocket motor on and

off during operation. This type of operation is represented by the rocket system shown in

figures 20 and 23 for a bipropellant and monopropellant system respectively. This

cycling produces an integrated thrust value that is less then that of continuous operation.

The average thrust value over a period of time will depend on the cycling frequency.

There are, however, some performance and mission impacts to cycling the engine to

regulate thrust.

First, the performance of the rocket engine will be less than that of steady state operation.

This is because on each cycle the rocket engine goes through a startup phase in which

transients in injector and chamber pressure reduce performance.

The ability of the rocket motor to withstand the cycling is well established. This mode of

operation is common in satellite applications, where the thruster will go through many

on / off cycles. Most of the satellite reaction control thrusters have gone through

extensive cycling qualification testing. For example in the qualification testing of an

Atlantic Research Corporation (ARC) 10 N thruster, it went through over 1 million cycles

and over 1,000 cold starts. This testing is very representative of what could be expected

for the aircraft rocket system since satellite reaction control thrusters are around the same

thrust level as that required by the aircraft. The capabilities of this type of thruster are

well matched to this aircraft application. During the propellant evaluation that was done

for the Mars micro-mission aircraft program, pulse testing was performed on an ARC 10

N thruster using MMH / MON25 propellant. Tests were performed for on / off pulse rates

of 0.1s / 0.1s , 0.2s / 0.2s and 0.5s / 0.5s.

The 0.5s / 0.5s pulse test was terminated because ignition of the propellant was not

occurring. The other two pulse tests ran successfully. This lack of ignition would need to

be experimentally investigated if MMH and MON-25 were selected as the propellant.

The Isp for the 0.1s pulse test was 208.4 and for the 0.2s pulse test it was 224.1. This

compares to an Isp of 288.4 for steady state operation (60 s run time) under the same

operating pressure and propellant mixture ratio. The thrust levels for the pulse tests were

also significantly reduced. For the steady state case the thrust was 10.28 N compared to

0.81 N for the 0.1 s pulse test and 1.62 N for the 0.2 s pulse. This reduction in Isp is due

to the startup effects of cycling the thruster. As the cycle times get larger the effect will

diminish. Figure 27 shows an approximation of the effect on Isp from pulsing the engine.

This figure is based on the limited data given above.

In another series of tests performed by Kaiser Marquardt, also for the Mars micromission aircraft program, ignition delay was plotted versus thrust level. At the design

thrust level (10 N) of the engine, the ignition delay was 7.5 ms. As the operational thrust

was reduced this delay increased to 11.5 ms at 4.5 N thrust.

NASA/CR—2003-212350 65
Pulse Cycling ARC Engine

300290280270260250240230220210200

0.1 1 10 100

Pulse Time (s)

Isp

Figure 27 Estimate of the Effect of Pulse Duration on Isp [16]

NASA/CR—2003-212350 66
7

7.5

8

8.5

9

9.5

1 0

10.5

1 1

11.5

1 2

4 0 5 0 6 0 7 0 8 0 9 0 100

Percent of Design Thrust

Ignition Delay (ms)

Figure 28 Estimate of Ignition Delay as a Function of Design Thrust [16]

NASA/CR—2003-212350 67
This behavior should be consistent with higher thrust level engines closer to the 35N that

was assumed for this analysis. Figure 28 was generated from the data supplied by

Marquardt and is an estimate of how pulsing is affected by off-design operation. From

figure 28 it can be seen that as the engine is operated further from its design thrust level,

ignition delay increases. Therefore the reduction in performance due to pulsing the

engine will increase as the engine is operated further from its design thrust level.

An additional characteristic of pulsing the engine is the loading or vibration it places on

the vehicle. This does not affect the engine’s performance but it may have an impact on

the science instruments or other hardware on the aircraft. A plot of a start-up acceleration

profile is shown in figure 29. This figure was generated by Kaiser Marquardt during the

testing of their 10 N thruster. The higher thrust level of the 35N thruster may increase the

g-levels experienced, but the profile should be similar. From this figure it can be seen that

vibration g-loads of over 30 g’s were experienced within the first 2 ms of startup.

Time, msec

Acceleration, g

0 2 4 6 8 10 12 14 16 18 20 -50

-40

-30

-20

-10

0

10

20

30

40

50

Sampling Rate = 200,000 Hz

Figure 29 Accelerometer trace of 10N Kaiser Marquardt Startup [17]

NASA/CR—2003-212350 68
Propulsion System Comparison

The main method for evaluating the propulsion systems is by their mass. The aircraft

mass is a critical factor in the design feasibility and any reduction in mass can greatly

enhance the aircraft’s performance and capabilities. Since the systems considered are

very different from each other in their design, their technology readiness level (TRL),

their operational risk and in the way they affect the other aircraft systems; comparing

them is somewhat like comparing apples and oranges. The systems were broken down

into three categories: electrical systems, combustion engine systems and rocket systems.

The lightest mass system of each of these types is compared for the 1, 2 and 4 hour

mission durations. This comparison is shown in figure 30. To try and maintain some

consistency between the systems, MMH and NTO were chosen as the propellant for both

the bipropellant rocket system and the combustion engines.

Based on the mass comparison shown in figure 30 it can be seen that the 4 cycle internal

combustion engine produces the lowest mass system for all mission durations. The fuel

cell system is more competitive as the duration of flight increases and is almost equal to

the 4 cycle engine for the 4 hour case. Compared to the other systems, the rocket system

mass grows considerably as the mission duration increases. Based on this, the rocket

system should not be considered for mission durations greater then 2 hours.

In selecting a propulsion system, the desired duration is a critical factor. However other

aspects such as development time and cost must also be considered. The rocket system

has the lowest performance but has the highest TRL due to its extensive use in the

satellite industry. The combustion engine, although widely used on Earth, would require

a significant amount of development for use in a Mars aircraft. A new engine would

need to be designed, tested and qualified for this application. This is because of the

differences in the type of fuel that would be used and the environments of Mars and deep

space.

If development cost and time is not an issue, the 4 cycle internal combustion engine

propulsion system is the best candidate, for flight durations up to approximately 4 hours.

The fuel cell system should also be considered for longer durations (near and greater then

4 hours). Figure 30 indicates that for durations greater then 4 hours the fuel cell system

would be the lightest.

NASA/CR—2003-212350 69
Figure 30 Mass Comparison Between the Various Types of Propulsion Systems

0

8 06 04 02 0

200180160140120100

System Mass (kg)

1 Hour Duration 2 Hour Duration 4 Hour Duration

Bipropellant Rocket4 Cycle Combustion EngineFuel Cell

NASA/CR—2003-212350 70
Near-Term System Selection

For a near-term flight of moderate to short duration (approximately 90 minutes or less)

where extensive propulsion system development is not possible, a bipropellant fueled

rocket system would be the best choice due to its present high level of development

(technology readiness level (TRL) of 7-9) and relatively low implementation risk.

The system is based mainly on components from satellite reaction control systems. This

provides a lot of design heritage and eliminates the need for space qualification of most

of the components within the system. The oxidizer used in this system is nitrogen

tetroxide with 25% nitric oxide content (MON-25). This differs from the standard

nitrogen tetroxide oxidizer which utilizes 3% nitric oxide content (MON-3). The reason

for utilizing the MON-25 oxidizer is that its freezing temperature is –54 °C compared to

–11 °C for MON-3. This reduced freezing point is a significant benefit for the system

design and reliability. The estimated minimum storage / operating temperature is –40 °C

for the propulsion system. Therefore MON-25 eliminates the need for an active heating

system during the mission. This reduces the parasitic mass of the propulsion system and

its complexity. Also the overall aircraft reliability is increased since there is little to no

chance of the oxidizer freezing and thereby causing the propulsion system to fail.

An estimate of the present level of development for the propulsion systems that were

considered in the context of a Mars aircraft mission are shown in table 27. A definition of

each TRL level is given in appendix A.

System TRL

Bipropellant Rocket 7 to 9

Monopropellant Rocket 7 to 9

Battery 4 to 6

Fuel Cell 4

Piston Expander Engine 4 to 5

4 Cycle Combustion Engine 4 to 5

2 Cycle Combustion Engine 4 to 5

Table 27 Propulsion System Development Level for a Mars Aircraft Application

The monopropellant rocket system was not selected because its performance was less

than that of the bipropellant rocket system. Although it is a simpler system requiring less

hardware mass, the increase in fuel consumption due to its reduced performance more

then offset the component mass saved. Overall, for the same flight range, the

monopropellant system and propellant was heavier then that needed for the bipropellant

rocket system.

Two electric propulsion system concepts were studied. The battery powered system

weight was comparable to the bipropellant rocket system. This was mainly due to the

inability to take advantage of some of the high specific power batteries that are available

because of their low rates of discharge. This low discharge rate meant that to meet the

continuous power demand of the aircraft, the battery capacity had to be much larger then

NASA/CR—2003-212350 71
what was actually needed to meet the mission duration requirement. The battery system

also had significant issues with eliminating heat generated during operation as well as

volume and packaging constraints within the aircraft. Because of theses concerns the

battery system was not selected as the propulsion system.

The fuel cell system was lighter then the bipropellant rocket system. However the relative

newness of the technology and the complete lack of experience in utilizing a fuel cell as

the main power system in an aircraft made it a very high risk item. Because of the

relatively low TRL level (4) for a fuel cell in this type of application, it was not

considered a viable option for this mission.

Both of the types of combustion engines considered were lighter then the bipropellant

rocket system. The main concerns with these systems is operating them on either a

bipropellant or monopropellant fuel system. In order to accomplish this, a development

program would need to be implement to produce an engine that can operate within the

Mars environment. For near term missions or missions with limited development

funding, this would not be feasible. Heat transfer issues with operating these engines

within the Mars environment are also a concern.

In addition to the items listed above, both the electric and combustion engine systems

would require a propeller to generate thrust. There are significant concerns with operating

a propeller within the Mars environment. To generate the thrust required, the propeller

will need to operate near the sonic limit at its tip. This places the operation of the blade in

a low Reynolds number (~15,000) high subsonic Mach number environment. The

aerodynamics of operating a propeller within this regime is out of the experience of

conventional propeller design. A development program would need to be undertaken to

produce a propeller that can confidently operate under these conditions. In addition to its

aerodynamic operation, the size of the propeller requires it to be stowed and then deploy

from the aeroshell with the aircraft. Due to the aerodynamics and required deployment,

the propeller represents a considerable risk item for the operation of the aircraft.

Eliminating the need for a propeller significantly reduces the overall operational risk of

the aircraft and is a significant factor in selecting the bipropellant rocket system as the

baseline propulsion system for a near-term Mars aircraft.

NASA/CR—2003-212350 72
Appendix A.—TRL Level Definition

Level 1 Basic Principles Observed and Reported

Level 2 Technology Concept and/or Application Formulated

Level 3 Analytical and Experimental Critical Function and/or Characteristic Proofof Concept

Level 4 Component and/or Breadboard Validation in Laboratory Environment

Level 5 Component and/or Breadboard Validation in Relevant Environment

Level 6 System or Subsystem Model or Prototype Demonstration in a Relevant

Environment (Ground or Space)

Level 7 System Prototype Demonstration in a Space Environment

Level 8 Actual System Completed and Flight Qualified Through Test and

Demonstration (Ground or Space)

Level 9 Actual System Flight Proven Through Successful Mission Operations

NASA/CR—2003-212350 73

References

1. Colozza, A.J., “High Altitude Propeller Design and Analysis Overview,”

NASA/CR—1998-208520, October 1998.

2. Selig, M.S., Donovan, J.F., and Fraser, D.B., “Airfoils at Low Speeds,” Soartech 8,

H.A. Stokely publisher, 1989.

3. Colozza, A.J., “Hydrogen Storage for Aircraft Applications Overview,”

NASA/CR—2002-211867, September 2002.

4. Matherson Tir-Gas, Online Product Catalog, http://www.matheson-trigas.com, June

2002.

5. All Sensors, Online Product Catalog, http://www.allsensors.com, June 2002.

6. Swagelok, Online Product Catalog, http://www.swagelok.com, June 2002.

7. Ultralife Batteries Inc., Online Product Specifications, http://www.ulbi.com, June

2002.

8. Wilson Greatbatch Technologies, Inc, Electrochem Commercial/Industrial Products,

Online Product Specifications, http://greatbatch.com, June 2002.

9. Tadiran Lithium Batteries, Online Product Specifications,

http://www.tadiranbat.com, June 2002.

10. Lithion Products, Inc., Online Product Specifications, http://www.yardney.com,

June 2002.

11. Eagle Pitcher Technologies, LLC, Online Product Specifications,

http://www.tech.epcorp.com, June 2002.

12. BST Systems, Inc., Online Product Specifications, http://www.bstsys.com, June

2002.

13. Bolz, R.E. and Tuve, G.L., editors, Handbook of tables for Applied Engineering

Science, 2nd Edition CRC press, 1973.

14. Schmidt, E.W., Hydrazine and its Derivatives Preparation, Properties, Applications,

John Wiley and Sons, Publisher, 1984.

15. Clark, C.C., Hydrazine, Mathieson Chemical Corporation, 1953.

16. Atlantic Research Corporation, “Mars Flyer Rocket Propulsion Risk Assessment:

ARC Testing,” NASA/CR—2001-210709, April 2001.

17. Kaiser Marquardt, Inc., “Mars Flyer Rocket Propulsion Risk Assessment,”

NASA/CR—2001-210710, April 2001.

18. Reed, R.D. (1978), High-Flying Mini-Sniffer RPV: Mars Bound, Astronautics and

Aeronautics, Vol. 16, No. 6, pp. 26–39.

19. Fuji Engines, Inc., Online Product Specifications, http://www.fujiengines.com, June

2002.

20. O.S. Engines, Inc., Online Product Specifications, http://www.osengines.com, June

2002.

21. Colozza, A.J., Miller C., Reed, B., Kohout, L., Loyselle, P., “Overview of

Propulsion Systems for a Mars Aircraft,” NASA report February, 2000.

22. Morgan, O.M., Cassady, J., and Christofferson, S., “Cost Effective Propulsion

Systems for Small Satellites Corners to Cut, Corners to Honor,” 14th AIAA/USA

Conference on Small Satellites, August 2000.

23. Loral Space Systems, GOES I-M Data Book, S-415-19, August 1996.

24. Vacco Aerospace Products, Product Data Sheet F1D10558-01, January 2001.

25. Vacco Aerospace Products, Product Data Sheet F1D10856-02, December 2000.

26. Vacco Aerospace Products, Product Data Sheet V1E10362, August 2000.

NASA/CR—2003-212350 75
27. Vacco Aerospace Products, Product Data Sheet 70440, June 2001.

28. Conax, Product Data Sheet, 1832-192, October 2002.

29. Conax, Product Data Sheet, 1801-090, October 2002.

30. Atlantic Research Corporation, Online Product Specifications,

http://www.atlanticresearchcorp.com, June 2002.

31. Moog, Inc., Online Product Specifications, http://www.moog.com, June 2002.

32. Hill, P.G. and Peterson, C.R., Mechanics and Thermodynamics of Propulsion,

Addison-Wesley Publishing Company, 1970.

NASA/CR—2003-212350 76
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NAS3–00145

84

Comparison of Mars Aircraft Propulsion Systems

Anthony J. Colozza

Aircraft engines; Microrocket engines; Mars exploration

Unclassified -Unlimited

Subject Categories: 07 and 34 Distribution: Nonstandard

Analex Corporation

3001 Aerospace Parkway

Brook Park, Ohio 44142

Project Manager, Lisa Kohout, Power and On-Board Propulsion Technology Division, NASA Glenn Research Center,

organization code 5420, 216–433–8004.

The propulsion system is a critical aspect of the performance and feasibility of a Mars aircraft. Propulsion system mass

and performance greatly influence the aircraft’s design and mission capabilities. Various propulsion systems were

analyzed to estimate the system mass necessary for producing 35N of thrust within the Mars environment. Three main

categories of propulsion systems were considered: electric systems, combustion engine systems and rocket systems.

Also, the system masses were compared for mission durations of 1, 2, and 4 h.