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SPACECRAFT PROPULSION

A BRIEF INTRODUCTION

Peter Erichsen

March 2005

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Table of Contents

Standard Notations ....................................................................................................... 2

1 INTRODUCTION........................................................................................................ 3

2 NEED FOR PROPULSION......................................................................................... 4

3 PROPULSION FUNDAMENTALS............................................................................ 6 3.1 Basic Propulsion Equations ................................................................................ 6 3.2 Propulsion Performance...................................................................................... 7

3.2.1 Thruster Performance Factor ................................................................ 7 3.2.2 System Performance Factor .................................................................. 8 3.2.3 Evaluation of Mass of Propulsion Systems .......................................... 8

4 SURVEY OF SPACECRAFT PROPULSION SYSTEMS......................................... 9 4.1 Spacecraft Propulsion System Options............................................................... 9 4.2 Chemical Propulsion......................................................................................... 10

4.2.1 Cold Gas ............................................................................................. 11 4.2.2 Hot Gas (survey)................................................................................. 12 4.2.3 Monopropellant................................................................................... 13 4.2.4 Bipropellant ........................................................................................ 15 4.2.5 Future Developments in Liquid Propellant Technology .................... 17 4.2.6 Solid Propellant .................................................................................. 18

4.3 Electric Propulsion............................................................................................ 19 4.3.1 Electric Propulsion Concepts.............................................................. 19 4.3.2 Electric Propulsion System Design .................................................... 20 4.3.3 Electric Propulsion System Performance ........................................... 21

5 PROPULSION SYSTEMS SELECTION CRITERIA .............................................. 23

6 TESTS OF PROPULSION SYSTEMS (SUB- AND INTEGRATED SYSTEM LEVEL) ...................................................................................................................... 27

7 SURVEILLANCE OF PROPULSION SYSTEMS DURING SPACECRAFT MISSIONS ................................................................................................................. 28

8 LITERATURE/REFERENCES ................................................................................. 29

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Standard Notations

Standard notation used throughout this paper is given below.

E energy [Ws]

F force [N] g acceleration of gravity, standard, 9.81 [m/s2] I impulse [Ns] m mass [kg] o

m propellant mass flow rate [kg/s] P power [W] v velocity [m/s] ∆v velocity-increment [m/s] η overall power conversion efficiency (Pjet/P) Subscripts e exhaust (effective) e opt exhaust (optimal) El electric (system) H/W hardware f final jet thruster nozzle exhaust 0 initial P propellant PS propulsion system PSS propellant storage system (propellant and corresponding tankage) S/C spacecraft sp specific ssp system-specific tot total

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1 INTRODUCTION This paper summarises key features and performance characteristics of existing and planned (near future) propulsion systems for use on spacecraft such as satellites, space probes, etc.

For a better understanding of spacecraft propulsion, the physical background of propulsion is discussed and basic propulsion mathematical equations are presented.

An overview of basic propulsion system designs is presented together with tables and graphs which should facilitate a preliminary selection of propulsion systems (chemical, electric) for spacecraft flight missions of given impulse and velocity-increment requirement.

However, for more detailed information about spacecraft propulsion see Ref. [1] to [5] noted in Section 8.

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2 NEED FOR PROPULSION Propulsion is needed:

- to place payloads into orbit: launch propulsion;

- to send payloads to the moon or to the planets: space propulsion;

- to position, adjust and maintain orbits of spacecrafts by orbit control: auxiliary propulsion;

- to orient spacecraft by attitude control: auxiliary propulsion also called reaction-control systems.

There are the following types of reaction-control systems:

- reaction jets (propulsion): which produce a control force by the expenditure of mass;

- solar sails, magnetic torquers (magnetic coils): which produce a control force by interaction with the environmental field;

- momentum-transfer devices (reaction-, flywheels): which produce no net control force, but simply transfer angular momentum to or from the spacecraft.

In this paper, only propulsion systems will be dealt with. These systems are based on jet propulsion devices that produce thrust by ejecting stored matter, called the propellant.

The main features of jet propulsion are:

a) LAUNCH PROPULSION for launching rockets with the following characteristics:

- high velocity increment capability (7 - 11 km/s)

- very high thrust levels (ratio thrust/launch vehicle weight: ≥ 1.3)

- low fraction of take-off mass of launch vehicle for payload (1 - 5%)

- powerful chemical rockets

b) SPACECRAFT PROPULSION is characterised in general by its complete integration within the spacecraft. Its function is to provide forces and torques to:

- transfer the spacecraft: orbit transfer incl. interplanetary travel

- position the spacecraft: orbit control

- orient the spacecraft: attitude control

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While jet propulsion systems for launching rockets are also called primary propulsion systems, spacecraft, e.g. satellites, are operated by secondary propulsion systems.

In order to fulfil attitude and orbit operational requirements of spacecraft, spacecraft propulsion systems are characterised in particular by:

- low thrust levels (1 mN to 500N) with low acceleration levels

- continuous operation mode for orbit control

- pulsed operation mode for attitude control

- predictable, accurate and repeatable performance (impulse bits)

- reliable, leak-free long time operation (storable propellants)

- minimum and predictable plume impingement effects

Candidate spacecraft propulsion systems are:

- chemical propulsion

- electric propulsion

An overview of spacecraft propulsion system options is shown in Fig.1

Fig. 1: Classification of Spacecraft Propulsion Systems

PROPULSION SYSTEMS

CHEMICAL

COLD GAS HOT GASELECTROTHERMAL(Resistojet; Arcjet)

ELECTROMAGNETIC ( MPD-Thruster)

ELECTROSTATIC(RIT; Field emission)

VAPORISING LIQUID

COMPRESSED GAS

(Nitrogen)

(Propane)

SOLID PROPELLANT

MONO-PROPELLANT(Hydrazine)

BI-PROPELLANT(MMH/N2O)

ELECTRICAL

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3 PROPULSION FUNDAMENTALS

3.1 Basic Propulsion Equations Newton's law of motion:

dtdmv

dtvdm

dtvmdF r

rrr+==

)( , (1)

Constant exhaust propellant velocity evv rr= at nozzle outlet ( evv rr

= is the relative velocity between spacecraft and exhaust propellant) gives the basic equation for force of thrust:

F dmdt

v m ve e= =o

[N], (2)

with dmdt

m=o

skg for the propellant mass flow rate. (3)

Strictly speaking, ev and F are vectors. They are here taken to be collinear, so no vector notation is needed.

From the law of conservation of momentum follows:

∆v m = - ∆m ve , (4)

which is change of momentum of spacecraft ≡ momentum of the expelled propellant. From this follows the total velocity increment ∆v of spacecraft:

dv v dmmv

v v

em

m

o

f

=

=

∫ ∫= −0

∆

; (5)

integrated:

∆v vmme

f

o

= − ln

sm , (6)

where mo is the initial mass and mf is the final mass of the spacecraft. This formula can be also written in the form of the basic 'Rocket Equation':

mm

efv

ve

0

=−

∆

(Tsiolkovski-Equation). (7)

The propellant quantity required for a spacecraft velocity change ∆v is with

mf = m0 - mP :

)1(0evv

P emm∆

−

−= [ ]kg . (8)

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The total impulse delivered by a certain quantity of propellant is calculated by:

Pe

mt

etot mvdmvFdtIP

=== ∫∫00

[ ]Ns . (9)

The kinetic energy of ejected matter is:

2

21

ePjet vmE = [ ]Ws . (10)

And the power of the jet is calculated by:

PdE

dtm v F

vjet

jete

e= = =12 2

2o

[ ]W . (11)

Therefore, the power input for an electric thruster will be:

PP

mv

Fvjet e e= = =

η η ηo 2

2 2 [ ]W , (12)

where η is the power conversion efficiency.

3.2 Propulsion Performance

3.2.1 Thruster Performance Factor The most useful parameter for determining thrust engine (or thruster) performance is thruster-specific impulse:

I F

msp = o

kgNs , (13)

which is defined as the impulse delivered per unit mass of propellant and which can be easily obtained by test.

An effective thruster exhaust velocity of the jet is sometimes introduced:

v F

me = o

sm . (14)

From its definition as the thrust per unit rate of mass flow of propellant, it follows that ve numerical the same as the Isp as defined above with SI units of m/s.

Specific impulses are sometimes quoted in units of seconds, corresponding to a modification of the above definition to that of the impulse delivered per unit weight of propellant. Such values in seconds then follow from those in Ns/kg by division with the gravitational acceleration standard, g (= 9.8 m/s2).

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3.2.2 System Performance Factor To describe the performance of the spacecraft propulsion system, a reference number is introduced, which defines the total impulse, Itot, delivered by the system, divided by the system mass, mPS:

I Imssp

tot

PS

=

kgNs . (15)

Because of the resulting dimension, - delivered impulse per kilogram of system mass mPS, this number is called system-specific impulse.

The Issp has to be further defined for the two different kind of systems applied commonly for spacecraft propulsion according to the configurations as shown in Fig. 2 and Fig. 3 in Section 4.1 below:

- Chemical Propulsion Systems, comprising cold gas and hot gas systems:

PSSWH

totssp mm

II

+=

/

, (16)

with mH/W, the propulsion hardware mass, such as thrusters, valves, piping, etc., which is independent of propulsion impulse, and mPSS, the mass of propellant and corresponding tankage (propellant storage system), which is proportional to propulsion impulse.

- Electric Propulsion Systems, where the electric power supply and control system, mEl, has to be added to the mass of the propulsion system:

ElPSSWH

totssp mmm

II

++=

/

. (17)

3.2.3 Evaluation of Mass of Propulsion Systems The mass of propulsion systems can be derived from the ‘propulsion system mass fraction’. With Equations (8), (9), (13, 14) and (15) the overall spacecraft propulsion system mass fraction is:

−=

∆−

evv

ssp

sp

CS

PS eII

mm

1/

, (18)

where mS/C ≡ m0 is the (initial) mass of the spacecraft.

For small values of evv∆ (e.g. ≤ 0.2) the propulsion system mass fraction is

approximately (≤ 10% different from (18)):

sspCS

PS

Iv

mm ∆

=/

. (19)

For details see 'Issp-Program' at SSC's website: http://www.ssc.se/ssd (Quick-Look software for evaluating spacecraft propulsion systems)

For more detailed information about propulsion fundamentals see Ref. [1] to [3] noted in Section 8.

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4 SURVEY OF SPACECRAFT PROPULSION SYSTEMS

4.1 Spacecraft Propulsion System Options Space propulsion and auxiliary propulsion are performed by the following on-board spacecraft propulsion system types:

a) Chemical Propulsion Systems:

- cold gas systems, comprising inert gases (e.g. nitrogen: N2) and high vapour pressure hydrocarbons (e.g. propane: C3H8);

- monopropellant hydrazine systems (N2H4);

- storable bipropellant systems (e.g. nitrogen tetroxide (NTO: N2O4) oxidiser with anhydrous hydrazine (N2H4) or monomethyl hydrazine (MMH: CH3N2H3) fuels);

- solid propellant motors (composite propellants, e.g. hydroxy-terminated polybutadiene (HTPB)).

Fig. 2: Schematic of Chemical Propulsion Systems

b) Electric Propulsion Systems:

- electrothermal systems (resistojet and arc-jets);

- electromagnetic systems (magnetoplasmadynamic: MPD);

- electrostatic systems (ion engines: Kaufman, radio-frequency, field emission, stationary plasma).

Propellant Storageand Feed System

Thrusters, Valves,Piping, etc

Electrical Control Unit

ThrustExhaust

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Fig. 3: Schematic of Electrical Propulsion Systems

Historically, chemical propulsion, comprising cold gas and hot gas systems, was the first one available for space propulsion, which is now followed by the development of electric propulsion systems. Since chemical propulsion is most commonly applied to today’s spacecraft, the description of spacecraft propulsion systems will be concentrated mainly on this type of space propulsion, while electric propulsion systems will be looked at in general only.

4.2 Chemical Propulsion Chemical propulsion is based on the principle of converting chemical energy to kinetic energy of the exhaust gases in a nozzle.

Currently available chemical propulsion systems can be categorised as either cold gas or hot gas.

Propulsion operating with cold gas, comprise compressed (inert) gas which is stored at high pressures in a tank, and vaporising liquids (high vapour pressure hydrocarbons), which are pressurised by their own equilibrium vapour pressure. Expelling these gases through a nozzle creates a thrust force.

Propulsion systems operating with hot gas comprise systems containing liquid and solid propellants. The energy from an exothermal combustion reaction of propellant chemicals in a thruster results in high temperature reaction product gases, which are expelled through a nozzle. In contrast to cold gas, liquid propellants need to be pressurised in their storage tank to feed the thrusters with propellant. For the reason of reliability, propellant pump-fed systems are not applied for (long) spacecraft flight missions.

Electrical PowerGenerator

Propellant Storageand Feed System

Power ProcessingSystem

Control Unit Plasma/Ion-Jet

ElectricThrusterAssembly

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4.2.1 Cold Gas Cold gas systems operate with compressed inert gas (e.g. nitrogen: N2) or high vapour pressure hydrocarbons (e.g. propane: C3H8). Cold gas systems are shown schematically in Fig. 4. The typical system consists of a propellant tank, fill valve, filter, pressure regulator, line pressure transducers, control valves, and nozzles. The pressure regulator provides propellant at constant pressure as the tank pressure drops. A relief valve is incorporated downstream of the pressure regulator to prevent system rupture in the case of a regulator failure. With regard to compressed gas systems, the cold gas is stored at high pressures in a tank. The vaporising liquid system is characterised by a liquid propellant pressurised by its own equilibrium vapour pressure and the expulsion of this vapour through a nozzle. In order to provide completely vaporised gas, a vaporiser is included in liquid cold gas systems. Although of moderate impulse capability, cold gas systems, in particular systems operating with compressed cold gas, are still of interest in view of their simplicity, high reliability and repeatability of impulse bit. Main performance and operating characteristics are presented in Tables 2 and 3.

Fig. 4: Basic Flow Scheme of Cold Gas Propulsion Systems

ADVANTAGES:

- simplicity and reliability - lowest cost propulsion system - very low thrust (≥ 0.02 N) and impulse bit (≥ 10-4 Ns) capability - low plume contamination

DISADVANTAGES:

- low Isp (≤ 700 Ns/kg) ≡ low Issp (≤ 500 Ns/kg) with resulting high system mass for e.g. gaseous nitrogen.

N2

T

P

P

Compressed Gas System

Press.

Relief Valve

C H3

T

P

P

Vaporising Liquid System

vvvvvvvvvvvvvvvvvvvvv Vaporizer

Temperature SensorTank

Fill Valve

Pressure Transducer

Filter

Pressure Regulator

Pressure Transducer

Thruster

8

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4.2.2 Hot Gas (survey) For increasing absolute levels of thrust and impulse requirements for spacecraft propulsion (e.g. attitude and orbit control), cold gas systems are inadequate and more energetic propellants generating hot gas for mass expulsion are required.

Hot gas systems are the most common type of propulsion systems for space applications. They can be divided into three basic categories:

- liquid, comprising monopropellant hydrazine (N2H4) and storable bipropellant (MMH/N2O4);

- solid, with composite propellants;

- hybrid (so far not used for spacecraft propulsion).

The terminology refers to the physical state of the stored propellants.

Fig. 5: Schematic of Hot Gas Propulsion Systems

In contrast to compressed gas and vaporising liquids, liquid propellants need to be pressurised in the tank to feed the thrusters with propellant. Note that due to long spaceflight mission durations, only pressure-fed systems are used because of their inherent simplicity compared with pump-fed systems, which are used commonly for launch propulsion.

Typically, rockets using solid propellants are called motors and rockets using liquids are called engines or thrusters.

Oxidiser

Oxidiser (liquid)

Fuel

Fuel (solid)

Liquid

Hybrid

Solid

Solid Propellant

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4.2.3 Monopropellant Monopropellant systems use a single (Mono) propellant to produce thrust. The most commonly used monopropellant is anhydrous hydrazine (N2H4). The hydrazine propellant is decomposed in a thruster by a catalyst and the resulting hot gas is expelled through a nozzle, thus generating thrust force on the spacecraft. A typical monopropellant system shown schematically in Fig. 6 uses nitrogen or helium gas to expel the propellant from a diaphragm tank into the chamber catalyst beds of the thrusters. Since the pressurant gas is stored (at a pre-selected but relatively low pressure, e.g. 22 bar) in the propellant tank, the propellant pressure varies with propellant usage. A typical selection of the ullage volume of 25% filled with pressurant gas (thus containing 75% propellant) will result in a propellant feed pressure decay, and thus in a thrust decay of 4:1. This mode of operation is also referred to as the blow-down mode, in contrast to the pressure constant mode, which requires the storage of a high-pressure gas in a tank external to the propellant tank (see bipropellant systems).

In a hydrazine gas generator system, the hydrazine decomposition gases are exhausted into a gas storage tank for later gas expulsion.

The catalytic thruster and gas generator systems have identical propellant feed systems consisting typically of propellant tank(s) with diaphragm expulsion device(s), propellant and gas fill valves, eventually latch valves (start valves), line pressure transducers and filters.

Monopropellant hydrazine for spacecraft attitude and orbit control is one of the most widely used propellants. The primary reason for such wide acceptance of monopropellant hydrazine propulsion systems lies in their inherent simplicity (reliability) while still providing adequate propulsive performance. Main performance and operating characteristics are presented in Tables 2 and 3.

Fig. 6: Basic Flow Scheme of Monopropellant Hydrazine Propulsion Systems

N2

T

P

Hydrazine System with Catalytic Thrusters

T

Hydrazine Gas Generator System

vvvvvvvvvvvvvvvvvvvvv

N H2 4

N2

N H2 4

T T

X

Fill Valve (Nitrogen)

Temperature Sensor

Propellant Tank(diaphragm)

Fill Valve (Hydrazine)

Pressure Transducer

Filter

Thruster

P

P

Latch-Valve

SolenoidValveHydrazineDecomposit.Chamber

Gas-StoreTank

Filter

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ADVANTAGES:

- simplicity and reliability (monopropellant) - lowest cost propulsion system (other than cold gas) - space storable for long periods (> 12 years demonstrated) - low thrust capability (≥ 0.5 N) - moderate thrust levels available (≤ 400 N) DISADVANTAGES:

- moderate Isp (≤ 2300 Ns/kg) with moderate Issp (≤ 1900 Ns/kg) resulting in medium/high system mass - limited life of catalyst - surface contamination (NH3), moderate

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4.2.4 Bipropellant Bipropellant systems are characterised by the combustion of two (Bi) propellants, a fuel (e.g. MMH) and an oxidiser (e.g. NTO), to produce thrust. The propellants are injected separately into the thruster combustion chamber where they react spontaneously (hypergolic propellant) to perform high-temperature, low molecular weight combustion products, which are then expelled through a nozzle.

A typical bipropellant system is shown schematically in Fig. 7. The system basically consists of a pressurising-gas system, propellant tanks (with surface tension propellant management devices), propellant lines and thrusters. Unlike hydrazine thrusters, bipropellant thrusters accept only a limited range of propellant inlet pressure variation of < 2. Therefore, the high-pressure gas, generally nitrogen or helium is regulated to the desired tank pressure, e.g. 17 bar. This mode of operation is also referred to as the pressure constant mode. The system contains check valves upstream of the propellant tanks to prevent possible back-flow, mixing, and combustion of the propellant vapours in the common pressurant gas line. Relieve valves are incorporated in the system upstream of the propellant tanks to prevent system rupture in the event of a pressure regulator failure. Filters are provided in the propellant lines directly upstream of the thruster valves to prevent clogging of the injector or damage of the valve seat by entrained foreign material. Finally, the systems contains pyro- or latch valves, line pressure transducers, fill and drain valves and various test ports for system check out. Main performance and operating characteristics are presented in Tables 2 and 3.

Fig. 7: Basic Flow Scheme of Bipropellant Systems

He

T

P

P P

MON MMH

TT

Temperature Sensor

Tank forPressurant Gas

Fill Valve (H )

Pressure Transducer

Pyro-Valve(normally closed)

Filter

Pressure Regulator

Pressure Relief Valve

Check Valve

Pyro-Valve(normally open)

Test Port

Temperature Sensor

Propellant Tank(surface tension)

Fill Valves (MON/MMH)Pressure Transducer

Pyro-Valve(normally closed)

Filter

Thruster

e

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ADVANTAGES: - high Isp: (≤ 2900 Ns/kg) for F ≤ 25 N (≤ 3110 Ns/kg) for F ≤ 410 N with high Issp (≤ 2800 Ns/kg) resulting in low system mass - high thrust capability, - up to 45 000 N DISADVANTAGES: - bipropellant system complexity with added valves, regulators, etc.

- higher cost in comparison to monopropellant hydrazine systems

Dual-Mode systems use hydrazine (N2H4) both as fuel for a bipropellant (N2H4/N2O4) Liquid Apogee Engine (LAE) and as monopropellant for on-orbit Attitude and Orbit Control Systems (AOCS) from a common fuel tank. The propulsion system layout is shown in Fig. 8 with the propellant feed system design similar to that of the bipropellant system, as described in Section 4.2.4 above.

Fig. 8: Basic Flow Scheme of Dual-Mode Systems

He

T

P

P P

MON N H

TT

Temperature Sensor

Tank forPressurant Gas

Fill Valve

Pressure Transducer

Pyro-Valve(normally closed)

Filter

Pressure Regulator

Pressure Relief Valve

Check Valve

Pyro-Valve(normally open)

Test Port

Temperature Sensor

Propellant Tank(surface tension)

Fill Valves (MON/N H )Pressure Transducer

Pyro-Valve(normally closed)

Filter

ThrusterHydrazine Bi-propellant(e.g. LAE)

2 4

2 4

X

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ADVANTAGES: - high LAE Isp (3110 Ns/kg) with potential weight savings over existing

Unified Propulsion System (UPS) (≈ 3-5%)

- nearly zero propellant residuals (oxidiser depletion)

- common AOCS and LAE fuel tank

- can use higher performance station-keeping thrusters (Power Augmented Catalytic Thrusters (PACT) at Isp=3000 Ns/kg versus Isp=2900 Ns/kg for F = 10 - 22 N bipropellant thrusters) if required

DISADVANTAGES: - Dual-Mode system complexity with added valves, regulators, etc. - higher cost in comparison to monopropellant hydrazine systems

4.2.5 Future Developments in Liquid Propellant Technology Current spacecraft and satellite users and manufacturers are looking for more environmentally friendly, safer propellants. Safer propellants can reduce costs by eliminating the need for self-contained atmospheric protective ensemble (SCAPE) suits that are needed for toxic propellants. Also, extensive and prohibitive propellant safety precautions, and isolation of the space vehicle from parallel activities during propellant loading operations can be minimised or eliminated. If used on these satellites, the costs for operating the vehicles will be lowered, in some cases dramatically.

A new family of environmentally friendly monopropellants has been identified as an alternative to hydrazine. These new propellants are based on blends of hydroxyl ammonium nitrate (HAN: [N+H3OH]NO3). When compared to hydrazine, HAN blends have a range of specific impulse (Isp) which can exceed that of hydrazine. Testing of HAN based propellants has begun to show promise and could soon be adopted for on-board propulsion systems on communications satellites and Low Earth Orbit (LEO) satellites and constellations.

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4.2.6 Solid Propellant The solid propellant rocket motor consists of a motor case, containing a propellant grain, a nozzle and an igniter. The schematic is shown in Fig. 9.

There are two principal types of propellants:

- homogeneous propellants, which are composed of fuels that contain enough chemically bonded oxygen to sustain the propellant burning process,

- composite propellants, which are composed of organic fuel binders and oxidisers.

Most common is the use of composite propellants, usually rubber based, polymer-type fuel (CTPB ) and stable solid oxidiser (ammonium perchlorate or nitro-cellulose, -double base). The propellant is premixed and batch loaded into lightweight simple motors. Main performance and operating characteristics are presented in Tables 2 and 3. In general, solid propulsion motors can only deliver their total impulse potential in one firing, because off-modulation is not possible. Therefore the usage of solid propulsion is restricted to:

- orbit change (e.g. apogee or perigee manoeuvre)

- impart acceleration (e.g. liquid reorientation manoeuvres, separation manoeuvres).

Fig. 9: Schematic of Solid Propellant Motor ADVANTAGES: - relatively simple operation - very high mass fraction, excellent bulk density and packaging characteristics - good, long-term storage characteristics DISADVANTAGES: - not readily tested and checked-out prior to flight - very difficult to stop and restart, throttle, pulse, etc. (hybrid) - limited Isp performance (2800 - 3100 Ns/kg) - limited redundancy with associated reliability and safety issues

Igniter Motor Case Nozzle

Propellant Grain

Hot GasExhaust

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4.3 Electric Propulsion

4.3.1 Electric Propulsion Concepts Electrical propulsion relies on externally provided electric power to accelerate the working fluid (propellant) to produce useful thrust. There are three main methods by which the electrical energy may be converted into the kinetic energy of the useful thrust:

- Electrothermal Systems, where the propellant (gas) is heated by passing it over an electric heated solid surface (resistojet), or by passing it through an arc discharge (arcjet). The heated gas is then accelerated by gas-dynamic expansion in a nozzle. Typical applications of this principle are the monopropellant hydrazine operated Power Augmented Catalytic Thruster (PACT) and the Hydrazine-Arcjet.

Fig. 10: Schematic of Resistojet Thruster

- Electromagnetic Systems, where a gas is heated in an arc discharge to such a high temperature, that it is converted to a neutral plasma (plasma thruster). The plasma is then expelled at high velocity by the interaction of the discharge current with the magnetic field (Lorentz force). A typical application of this principle is the Magneto-Plasma-Dynamic (MPD) type of thruster.

Fig. 11: Schematic of Magnetic Arcjet Thruster

+

Gas Inlet Cathode AnodePower Supply

PlasmaExhaust

Arc-Discharge

Heater

∫∫∫∫∫∫∫∫

Power Supply Gas Inlet Heat Exchanger

_

+Hot GasExhaust

- 20 -

- Electrostatic Systems, where usually a high molecular propellant, such as xenon gas, is ionised (ion thruster) by e.g. electron bombardment (Kaufman), in a high frequency electro-magnetic field (radio-frequency) or by extracting ions from the surface of a liquid metal (caesium) under the effect of a strong electrostatic field (field emission). The ions are then accelerated to high velocity (30 to 60 km/s) by a strong electric field. Electrons are injected into the ion beam from an electron emitter in order to keep it electrically neutral, thus preventing an electric charge build-up of the spacecraft.

In addition to the above described category of ion thrusters, the Stationary Plasma Thruster (SPT), which belongs to the category of ‘Hall-effect Thrusters’, uses an applied magnetic field to control electrons in a quasi-neutral plasma discharge.

Fig. 12: Schematic of Ion Thruster

4.3.2 Electric Propulsion System Design An electric propulsion system consists of a power generator (e.g. solar or nuclear), power processing system, electric control unit, thruster assembly, propellant storage and propellant feed system (see also Fig. 3 above).

Fig. 13: Electric Propulsion System Block Diagram

Ioniser Exit GridIoniserPropellant Supply

Positiv Ion Beam

Electron Emitter

Power Supply

AcceleratorGrid

+ _

Plasma/Ion-JetPPUControl

Unit

Solar Array Solar Array

PropellantTank

PropellantFeedSystem

S/CInterface

Gimbal

Thruster

- 21 -

- Power Generator Electric power can be obtained from either sunlight or from a nuclear reactor. In the case of solar electric propulsion, solar photons are converted into electricity by solar cells. In nuclear electric propulsion, thermal energy from the nuclear reactor is converted into electricity by either a static or dynamic thermal-to-electric power conversion system. Static systems have the advantage of no moving parts for high reliability, but they have low efficiency; dynamic systems have moving parts (e.g., turbines, generators, etc.) and do not scale well for small systems, but they do have higher efficiency.

- Power Processing System Power processing systems are required to convert the voltage from the power generator to the form required by the electric thruster. For example, a solar array produces low-voltage DC (typically ~ 100 V); this would need to be converted (via transformers, etc.) to kilovolt levels for use in an ion thruster. The power processing system is often referred to as the power-processing unit (PPU).

- Electric Control Unit to operate electrically valves and thrusters.

- Propellant Storage & Feed Systems Various combinations of propellant and thruster are possible, depending on the specific application. In general, liquid or gaseous propellants are stored and fed to the thruster assemblies as in chemical propulsion. Details see also Ref. [1]

4.3.3 Electric Propulsion System Performance For electric propulsion high impulse performance is not dictated by maximum exhaust velocity, like for chemical propulsion, but rather by optimum values of thruster exhaust velocity, ve opt.

The term of optimal exhaust velocity ve opt can be elucidated schematically by the following picture:

Fig. 14: Term of Optimal Exhaust Velocity

Mass of Propellant + Tank

Mass of Power Supply

Mass of Propellant/Tank + Power Supply

Minimum System Mass

Exhaust Velocity (ve)

Mass

ve opt

- 22 -

With increasing exhaust velocity, ve, the combined mass of propellant and tank is decreasing while the mass of the power supply is increasing. The point of inter-section of the two curves determines the minimum of the system mass by ve opt resulting in a maximum value of Issp.

For details see 'Issp-Program' at SSC’s website: http://www.ssc.se/ssd/ (“Quick-Look software for evaluating spacecraft prop. systems”, - under products and services, Satellite Subsystems).

Table 1 presents a summary of the comparison between electrical (ion) and chemical (bipropellant) propulsion.

Type of Thruster Spec. Impulse (Ns/kg)

Thrust F (N)

DC Power Required(W)

Electrical (Ion thruster) Chemical (Bipropellant)

Order of magnitude of the ratio ION/Chemical

≥ 30 000

≈ 3 000

101

10-3 – 0.2

4 – 500

10-4

400 – 800

4 – 8 (short term)

102

Table 1: Thruster Comparison: typical electrical vs. chemical figures

Main performance and operating characteristics are presented in Tables 2 and 3.

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5 PROPULSION SYSTEMS SELECTION CRITERIA A detailed procedure for the selection of propulsion systems for given spacecraft mission requirements is beyond the scope of this paper. The process for selecting and sizing the elements of propulsion systems is detailed in Ref.[4].

However, in general, it can be said that an important consideration for the selection of a suitable propulsion system will be the trade-off between its velocity increment capability and system mass. An important requirement of the spacecraft designer will be that the mass of the propulsion system shall not exceed a certain percentage of the overall mass of the spacecraft (propulsion system mass fraction). To assess the suitability of spacecraft propulsion systems with spacecraft mission impulse requirements, the velocity increment (∆v) capability of the various systems must be estimated on the typical propulsion performance figures presented in Section 4 and summarised in Table 2. Propellant Thruster-spec. Impulse

Isp (mission average)

(Ns/kg)

System-spec. Impulse, Issp

(Ns/kg)

Remarks

Nitrogen (N2) 700 291 Cold Gas; Titanium Tank

Ammonia (NH3) 950 700 Cold Gas; Al-Tank with Heat Exchanger

Monopropellant Hydrazine (N2H4)

2150 1860 Hot Gas; Tank with Diaphragm

Bipropellant MMH/ NTO, (CH3N2H3/N2O4)

2950 2730 Hot Gas; Tank with Surface Tension Device

Solid propellant (Composites, CTPB)

2880 2670 MAGE 2 Apogee Kick Motor

Electrothermal Thruster (N2H4)

3000

5000

2780

4540

Power Augment. Catalytic Thruster (PACT);

ARC-JET;

Stationary Plasma Thruster (SPT) (Xenon: Xe)

14700 10800 SMART-1

Table 2: Comparison of typical Spacecraft Propulsion Systems Performances

From Equation (18) the dependence of the propulsion system mass fraction, mPS/mS/C , on mission velocity increment ∆v can be derived for any given value of Isp (ve) and Issp. Curves of mPS/mS/C plotted as a function of ∆v for different propulsion system designs with typical values of Isp and Issp, which are summarised in Table 2, are shown in Figure 13.

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0 500 1000 1500 2000 2500 3000

0,0

0,1

0,2

0,3

0,4

0,5

0,6

0,7

0,8

0,9

1,0

(8) Issp

= 10801 Ns/kg

(7) Issp

= 4540 Ns/kg

(6) Issp = 2780 Ns/kg(5) I

ssp = 2730 Ns/kg

(4) Issp = 2670 Ns/kg

(3) Issp

= 1860 Ns/kg

(2) Issp = 700 Ns/kg(1) Issp = 291 Ns/kg

mPS

/mS/C

delta-v (m/s)

(1) NITROGEN (2) AMMONIA (3) HYDRAZINE(4) SOLID PROPELLANT(5) BI-PROPELLANT(6) PACT(7) ARC-JET(8) ELECTRIC (SPT-100)

Fig. 15: Typical delta-v Performance of Spacecraft Propulsion Systems The diagram in Figure 15 gives the first and most important indication for the selection of propulsion systems. If we assume mPS/mS/C < 0.30, we can read directly from Figure 15:

- for low ∆v ≤ 150 m/s, compressed cold gas and vaporising liquid propulsion systems seem to be the best choice, because they meet the requirement and have the lowest cost;

- for 150 < ∆v ≤ 650 m/s, monopropellant hydrazine fed propulsion systems are the best choice, because of their inherent simplicity (reliability) and potential low cost, while still meeting the requirement;

- for high ∆v > 650 m/s, bipropellant systems, monopropellant hydrazine fed resistojet systems (power-augmented thrusters, arcjets), and electrostatic (electromagnetic) systems will satisfy the ∆v-requirements best.

Finally, for any given value of total impulse, Itot, the mass of the propulsion system, mPS, can of course be calculated directly from values of Issp. When a suitable spacecraft auxiliary propulsion system is selected, however, the cost, complexity, operability and reliability of the system also play an important role.

With regard to low ∆v-requirements, compressed cold gas systems used for auxiliary propulsion of spacecraft’s (attitude and orbit control), although of moderate impulse capability, are still of interest in view of their simplicity, high reliability, repeatability of impulse bit and low system costs.

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In considering the merits of the various compressed cold gas and vaporising-liquid systems, the following major points must be considered carefully:

- Additional heat may be necessary for vaporising, e.g. propane for use in gas jets. For high thrust levels and long thrust duration’s, this can give rise to thermal problems in the propane system. As the latent heat of ammonia is about three times that of propane, additional technical problems may occur here.

- For zero-g conditions (non-spinning satellites), the storage of liquefied propellants is more complex than that of pressurised gases, as tanks with bellows, surface tension devices etc. have to be provided to separate liquid and vapour. In addition, propellant gauging is much more complex (with resulting higher costs) for liquefied propellants under zero-g conditions than for compressed gases. Moreover, fuel slosh of liquefied propellants may cause extra problems for the dynamic behaviour of spacecrafts.

Therefore, for low ∆v requirements, compressed cold gas systems utilising N2 are the most commonly used.

For higher ∆v-requirements, in the trade-off between monopropellant hydrazine systems, bipropellant systems and electric propulsion systems, the following major points have to be considered carefully:

- Because of their inherent simplicity, hydrazine monopropellant systems still represent the lowest possib1e cost technology in the field of liquid propulsion. Such a technology is therefore of interest whenever a moderate velocity increment is required or where mass is not a critical design driver.

- For delivering low impulse-bits or impulses at low spacecraft torques or acceleration forces, hydrazine thrusters of potential low thrust levels (down to 0.5N) will have to be used.

- High ∆v requirements for spacecraft in-orbit transfer and attitude and orbit-control can only be met by bipropellant systems (e.g.: unified propulsion system) and electric propulsion because of system mass limitations.

- For the selection of a hydrazine resistojet system (e.g. power-augmented thrusters, arc-jet), thrust level and duty cycle requirements have to be considered. The fact that the limited power available and heat capacitance of the electrothermal thruster impose a limit on thrust- and duty-cycle levels may give rise to technical problems.

- With regard to electric propulsion like electromagnetic and electrostatic systems, this technology, although still under development, has proven to achieve thruster exhaust velocities ve an order of magnitude higher than the best performing chemical propulsion systems; see Table 1. Therefore, electric propulsion is essential for further reduction of system (propellant) mass, enabling higher payload mass and coping best with future high energy mission requirements. However, depending on thrust levels, electric propulsion can impose severe power requirements on the spacecraft power supply and power processing system.

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When spacecraft propulsion systems are to be selected, the above mentioned points have to be assessed properly with reference to the flight mission and design requirements of the spacecraft itself.

Table 3 below presents a comparison summary of candidate spacecraft propulsion systems.

Type Thrust Level

(N)

Thruster-Spec. Impulse (Ns/kg)

Advantages Disadvantages

Cold Gas (N2, CF4, NH3, C3H8) Solid Motor (e.g. Apogee Kick Motor) Liquid: Monopropellant Hydrazine (N2H4) Bipropellant (CH3N2H3/N2O4) Electric Propulsion Electrothermal:

Resistojet (NH3, N2H4, H2)

Arcjet (NH3, N2H4, H2, N2) Electromagnetic: Pulsed plasma,

(Teflon) Electrostatic:

Stationary Plasma Thruster (SPT) (Xenon: Xe)

Ion (Hg, A, Xe)

0.02 – 5 28 000 – 47 000

0.5 – 22

4 – 500

5·10-3 – 0.5

5·10-2 – 5

5·10-6 – 5·10-3

10-2 – 0.5

10-3 – 0.2

440 – 950

2 800 – 3100

2 100 – 2 300

2 850 – 3 110

1 300 – 5 000

4 000 – 15 000

15 000

15 000 - 25 000

≥ 30 000

Extremely simple, reliable, very low cost Simple, reliable, relatively low cost Simple, reliable, low-cost High performance

High performance, low power, simple feed system High performance, simple feed system

High performance

Very high performance

Very low performance, highest mass of all systems Limited performance, higher thrust Low performance, higher mass than bipropellant More complicated system than monopropellant

More complicated interfaces, more power than Chemical thrusters, low thrust High power, complicated interfaces (especially Thermal)

High power, low thrust, complicated

Very high power, low thrust, complicated

Table 3: Survey of Candidate Spacecraft Propulsion Systems

For more detailed information about spacecraft propulsion technology and its application, a list of literature and references is noted in Section 8 below.

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6 TESTS OF PROPULSION SYSTEMS (SUB- AND INTEGRATED SYSTEM LEVEL) The following parameters can be measured:

- pressure

- temperature

- electr. current and voltage on e.g. solenoid valves

The following tests can be performed:

- functional tests of electric actuated valves by measuring electric current and voltages ⇒ response time of valves;

- measurement of leak tightness of components, parts of systems and overall systems by measurements of temperatures and pressures at various points of time;

- measurement of leak tightness of the integrated propulsion system with the help of a gas spectrometer;

- measurement of leak tightness of valves (thrusters) with the help of 'glass pipettes'.

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7 SURVEILLANCE OF PROPULSION SYSTEMS DURING SPACECRAFT MISSIONS The following propulsion system related telemetry data are available:

- pressure

- temperature

- thruster operations

- system valves operations

The propulsion systems can be checked for the following items:

- leak detection of systems (parts and overall system) by:

• measurement of pressure and temperature of different parts of systems at various points of time;

- propellant consumption and remaining mass of propellant by:

• evaluation of pressure and temperature data (gas law) in propellant tanks, related to initial mass of propellant at beginning of life (BOL), also called ‘P.V.T’ method;

• 'book keeping' of thruster operations;

- thrust and spec. impulse of individual thrusters by:

• evaluation of propellant consumption during in-orbit thruster operations and comparison of planned and achieved in orbit spacecraft movements.

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8 LITERATURE/REFERENCES

[1] Rocket Propulsion Elements; An Introduction to the Engineering of Rockets, George P. Sutton, A Wiley-Interscience Publication, John Wiley & Sons.

[2] Handbook of Astronautical Engineering; Section 17: Propulsion Fundamentals; Heinz Hermann Koelle, Editor; McGraw-Hill Book Company, Inc. 1961

[3] Performance Evaluation of Spacecraft Propulsion Systems in Relation to Mission Impulse Requirements; Peter Erichsen, Proceedings of the Second European Spacecraft Propulsion Conference, 27-29 May, 1997 (ESA SP-398, Aug. 1997)

[4] Space Mission Analysis and Design, Chapter 17: Space Propulsion Systems; Wiley J. Larsson and James R. Wertz (editors), Published Jointly by Microcosm, Inc. Torrance, California and Kluwer Academic Publishers Dodrecht/Boston/London

[5] Space Propulsion, Charles D. Brown; AIAA Education Series, Published by American Institute of Aeronautics and Astronautics, Inc., 370 L’Enfant Promenade, SW, Washington, DC 20024-2518