SPACECRAFT PROPULSION - KTH

-iv -Standard NotationsStandard notation used throughout this booklet is given below.C tank filling ratio (V p /V T )E energy [Ws]F force [N]g acceleration of gravity, standard, 9.81 [m/s 2 ]I impulse [Ns]K tank performance factor (P op /m T V T ) [m 2 /s 2 ]M molecular mass [kg/kmol]m mass [kg]m propellant mass flow rate [kg/s]P power [W]p pressure [N/m 2 ]R gas constant 8.314 [kJ/°K/kmol]S/C spacecraftT temperature [°K]x non-impulse dependent system mass (%)v velocity [m/s]z gas compressibility factorv velocity-increment m/s specific power [W/kg] overall power conversion efficiency (P jet /P)κ specific heat ratio specific mass of propellant [kg/m 3 ] thrust time sec]

-v -Subscriptsc motor chambercase motor casee exhaust (effective)e-opt exhaust (optimal)El electric (system)H/W hardwaref finaljet thruster nozzle exhaust0 initialop operatingP propellantPS propulsion systemPSS propellant storage system (tank with propellant)S/C spacecraftsp specificssp system-specificT tanktot total

- 1 -1 INTRODUCTIONThis booklet summarises key features and performancecharacteristics of existing and planned (near future) rocketpropulsion for use on spacecraft such as satellites, space probes, etc.In the frame of “Lecture Notes”, this booklet, presents a summary ofthe “Rocket Propulsion Course” contained in the CompEduHTPplatformwith focus on spacecraft propulsion systems.For a better understanding of spacecraft propulsion, the physicalbackground of propulsion is discussed and basic propulsionmathematical equations are presented.The aim of this presentation is to bring about the basics of spacepropulsion on system level including propulsion system performanceevaluation. As a supplement to rocket propulsion theory, the‘System-specific Impulse’ I ssp is introduced. The I ssp allows a moreaccurate determination of the propulsive performance than thecommonly used ‘Specific-Impulse’ I sp which is only related topropellant and thrust engine performances. The I ssp has the advantagein defining those parameters, which have a most significant impacton propulsion system impulse performance. This allowsunderstanding the significance of the various system performanceparameters, which means also a better understanding of systemdesign concepts with related performance in general. Relatedexercises are noted under ‘Spacecraft Propulsion’ (S1B8C4) in theCompEduHTP-platform: www.energy.kth.se/compedu.An overview of basic common propulsion system designs ispresented together with tables and graphs which should allow thevaluation and facilitate a preliminary selection of propulsion systems(chemical, electric) for spacecraft flight missions of given impulseand velocity-increment requirement.The literature noted in Chapter 9 is recommended for further readingabout spacecraft propulsion technology and its application.

- 2 -2 NEED FOR PROPULSIONPropulsion is needed:- to place payloads into orbit: launch propulsion;- to send payloads to the moon or to the planets: spacepropulsion;- to position, adjust and maintain orbits of spacecrafts by orbitcontrol: auxiliary propulsion;- to orient spacecraft by attitude control: auxiliary propulsionalso called reaction-control systems.There are the following types of reaction-control systems:- reaction jets (propulsion): which produce a control force bythe expenditure of mass;- solar sails, magnetic torquers (magnetic coils): which producea control force by interaction with the environmental field;- momentum-transfer devices (reaction-, flywheels): whichproduce no net control force, but simply transfer angularmomentum to or from the spacecraft.In this booklet, only propulsion systems will be dealt with which arebased on jet propulsion devices that produce thrust by ejecting storedmatter, called the propellant.

- 3 -The main features of jet propulsion are:a) LAUNCH PROPULSION for launching rockets with thefollowing characteristics:- high velocity increment capability (7 - 11 km/s)- very high thrust levels (ratio thrust/launch vehicle weight: 1.3)- low fraction of launch vehicle take-off mass for payload (1 - 5%)- powerful chemical rocketsb) SPACECRAFT PROPULSION is characterised in general by itscomplete integration within the spacecraft. Its function is toprovide forces and torques to:- transfer the spacecraft: orbit transfer incl. interplanetary travel- position the spacecraft: orbit control- orient the spacecraft: attitude controlWhile jet propulsion systems for launching rockets are also calledprimary propulsion systems, spacecraft, e.g. satellites, are operatedby secondary propulsion systems.In order to fulfil attitude and orbit operational requirements,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 exhaust plume impingement effect.

- 4 -3 PROPULSION FUNDAMENTALS3.1 Basic Propulsion EquationsThe essence of space propulsion is to modify the velocity vector of aspacecraft either in magnitude or in direction so as to modify theorbit or attitude. However, an isolated body, like a spacecraft, canmodify its momentum only if external forces act on it, since allinternal forces cancel each other in action-reaction pairs. This isexpressed by Newton's law of motion: d(mv)dv dmF m v , (1)dt dt dtUnfortunately, there are no such external forces in space (except veryweak perturbation forces which have to be compensated by thespacecraft onboard propulsion system) and Eq. (1) becomes:dv dmm vdt dt 0(2)Therefore, the only and obvious way out is, that the spacecraft mustbe split up such, that a part of the spacecraft can modify its velocitythrough the effect of action-reaction forces.In fact, this is realised by the ejection of mass in form of propellantfrom the spacecraft. If we assume a spacecraft with a mass m, ejecting propellant with a rate of dm/dt at constant velocity v veatnozzle outlet, Eq. (2) can be written with the action and reactionforces in balance:dv dmm ve(3)dt dt

- 5 -This expresses that the spacecraft experiences acceleration in theopposite direction to v e, or that the external force acting on thespacecraft is by definition the force of thrust. That is, constantexhaust propellant velocity v v at nozzle outlet ( ev v is theerelative velocity between spacecraft and exhaust propellant) gives thebasic equation for force of thrust:F dmdt v mv e e N, (4)withdmdt mkg s for the propellant mass flow rate. (5)Strictly speaking, veand F are vectors. They are here taken to becollinear, so no vector notation is needed.Eq. (3) can be integrated to get the accumulated velocity incrementv of a spacecraft:Integrated:vvv0dv vmfemmv velnmodmmfo. (6)m s , (7)where m 0 is the initial mass of the spacecraft at the beginning and m fis the final mass of the spacecraft at the end of its mission.

- 8 -An effective exhaust velocity of the jet is introduced, which isdetermined by test:vFe mm s . (15)From its definition as the thrust per unit rate of mass flow ofpropellant, it follows that v e is numerical the same as the I sp asdefined above with SI units of m/s. Note: v e hereafter is always theeffective exhaust velocity, although called simply ‘exhaust velocity’,if not stated otherwise. Further, in all related calculations with v e , theeffective velocity has to be applied.Propulsive performance is commonly associated with the specificImpulse I sp , (v e ). According to the ‘Rocket Equation’ (8), a high valueof I sp will result in a mission final high spacecraft mass, which meanshigh payload mass, because of lower propellant mass consumptionduring the spacecraft mission.Specific impulses are sometimes quoted in units of seconds,corresponding to a modification of the above definition to that of theimpulse delivered per unit weight of propellant. Such values inseconds then follow from those in Ns/kg by division with thegravitational acceleration standard, g (= 9.8 m/s 2 ).3.2.2 System Performance FactorWith regard to the evaluation of propulsion performance on systemlevel, propellant storage systems, and especially for electricpropulsion, electric power supply and power processing systems (seeChapter 4.3.2) may form a major ‘dead’ dry mass of the overallpropulsion system mass. Therefore, the choice and sizing ofpropulsion systems is not always clear on the basis of I sp alone.

- 10 -3.2.3 Evaluation of Mass of Propulsion SystemsThe mass of propulsion systems can be derived from the propulsionsystem mass fraction. The dependence of the propulsion system massfraction on mission velocity increment v is derived from the‘Rocket Equation’ in combination with the definition of the systemspecificimpulse I ssp .The first Eq. (17) below is obtained from Eqs. (9) and (10). Thesecond Eq. (18) is just the definition of I ssp , and the final expressionEq. (19), follows from the first two:IItotv v m v m (1 e e)(17)eIPeS / Cvtotssp Itot IsspmPS(18)mPSmmPSS / Cvvv Ieve sp v (1 e ) (1 e e) (19)IIsspwhere m S/C m 0 is the (initial) mass of the spacecraft and with theunderstanding of:m Ns ve Isp is numerical equal. (20) s kgsspWith the help Eq. (19), for given values of v e and I ssp , the propulsionsystem mass fraction can be plotted as a function of velocityincrement (v), as presented in Chapter 5 and realised by thecomputerised ‘Issp-Program’, see [4]. By this, the I ssp and the massfraction of propulsion systems can be evaluated for given missionimpulse and v requirements.

- 11 -4 SURVEY OF SPACECRAFT PROPULSIONSYSTEMS4.1 Spacecraft Propulsion System OptionsSpacecraft propulsion systems can be classified according to the typeof energy source. Both, space propulsion and auxiliary propulsion areperformed by the following two main on-board spacecraft propulsionsystem types:A) Propulsion Systems with self-contained energy in propellants,comprising cold gas and hot gas systems. The energy toproduce thrust is stored in the propellant, which is releasedmainly by chemical reactions (this is why these systems aremostly referred to as chemical propulsion systems) and thepropellant is then accelerated to a high velocity by expanding itin form of gas through a nozzle. These systems contain:- Storage and feed system that stores (tank) and feeds thepropellant to the thrusters to generate thrust.- Valves, piping which connects the propellant storage systemwith the thruster.- Electric control unit to operate electrically the valves andthrusters.PropellantStorage andFeed SystemThrusters,Valves,Piping, etc.ThrustExhaustElectricalControl UnitFigure 2: Schematic of Chemical Propulsion Systems

- 12 -With regard to the system-specific impulse, I ssp = I tot /m PS (seeEq. (16) above), its practical application, especially for systemperformance analysis, requires a very clear definition of what isincluded in the total mass of propulsion system m PS .Therefore, the I ssp can be further detailed according to Fig. 2, andwith Eq. (16) the I ssp for chemical propulsion systems can bewritten:ItotIssp, (21)m mH / WPSSwith m H/W , the propulsion hardware mass, such as thrusters,valves, piping, etc., which is independent of propulsion impulse,and m PSS , the mass of propellant storage system (propellant +tank), which is proportional to propulsion impulse; see colouredbox of Fig. 2.The following types of propulsion systems are part of systemswith self-contained energy in propellants:- Cold gas systems, comprising inert gases (e.g. nitrogen: N 2 )and high vapour pressure hydrocarbons (e.g. ammonia:NH 3and propane: C 3 H 8 ).- Monopropellant hydrazine systems (N 2 H 4 ).- Storable bipropellant systems (e.g. nitrogen tetroxide (NTO:N 2 O 4 ) oxidiser with anhydrous hydrazine (N 2 H 4 ) ormonomethyl hydrazine (MMH: CH 3 N 2 H 3 ) fuels).- Solid propellant motors (composite propellants: e.g.aluminium powder with hydroxyl-terminated polybutadiene(HTPB) binder and an oxidiser like ammonium perchlorate).More details about chemical propulsion are presented in Chapters 4.2below.

- 13 -B) Propulsion Systems with externally supplied energy topropellant, comprising e.g. electric propulsion. The energy toproduce thrust is not stored in the propellant but has to besupplied from outside by an extra power source, e.g. nuclear,solar radiation receivers (solar cells) or batteries. These systemscontain:- Storage and feed system that stores and feeds the propellantto the thrusters to generate thrust.- Valves, piping which connects the propellant storage systemwith the thruster.- Electric control unit to operate electrically the valves andthrusters.- Electric power generator and power processing system.ElectricalPowerGeneratorPowerProcessingSystemControl Unit,Harness,Piping, etc.ElectricalThrusterAssemblyPlasma/Ion JetPropellantStorage andFeed SystemFigure 3: Schematic of Electrical Propulsion Systems

- 14 -According to Fig. 3 and with Eq. (16) the I ssp for electricpropulsion systems can be written:IsspmH / WItot mPSS mEl(22)with m El , to be added to the system mass with regard to chemicalpropulsion. The m El comprise the mass the electrical powergenerator, the power processing system and the electricalthrusters assembly which are proportional to the power to behandled by electric propulsion systems; see coloured box ofFig. 3.The following types of propulsion systems are part of systems withexternally supplied energy to propellant, i.e. electric propulsion:- Electrothermal systems (resistojet and arc-jets):Here thrust is produced by expansion of hot gas (which is heatedby electric current) in a nozzle.- Electromagnetic systems (magnetoplasmadynamic: MPD).- Electrostatic systems (ion engines: Kaufman, radio-frequency,field emission, stationary plasma):Here thrust is produced by acceleration of charged particles inelectric or magnetic fields to high expulsion velocities.More details about electric propulsion are presented in Chapters 4.3below.Eqs. (21) and (22) for I ssp of the various propulsion systems have incommon the same numerator, representing the total impulse

- 15 -delivered by the propellant contained in the propellant tank, which iswith Eq. (17):I m v(23)totPewhile the denominator in Eqs. (21) and (22), with regard to theimpulse related system mass (m PSS , m El ), varies with the kind anddesign of propulsion systems. In this respect, a concise descriptionof common spacecraft propulsion systems is presented below.Details of derived mathematical formulas of I ssp for chemicalpropulsion systems are presented in Chapter 4.2 and those forelectrical in Chapter 4.3 below.An overview of actual Classification propulsion of system Propulsion options systemsaccording to theirsource of energy is shown in Fig. 4.PROPULSION SYSTEMSCHEMICALELECTRICALCOLD GASHOT GASELECTROTHERMAL(Resistojet; Arcjet)COMPRESSED GAS(Nitrogen)VAPORISING LIQUID(Propane)SOLID PROPELLANTMONO-PROPELLANT(Hydrazine)BI-PROPELLANT(MMH/N 2O)ELECTROMAGNETIC( MPD-Thruster)ELECTROSTATIC(RIT; Field emission)Figure 4: Classification of Spacecraft Propulsion Systems

- 16 -Historically, chemical propulsion, comprising cold gas and hot gassystems, was the first one available for space propulsion which isnow followed by the development of electric propulsion systems.Presentations of spacecraft propulsion systems in Chapter 4 belowwill be concentrated on today’s most commonly used systemdesigns, which include traditional chemical propulsion with evolvingenvironmental benign, so-called ‘green propellants’ as well aselectric propulsion still under development.Finally, Chapter 6 will present an outline of the potential futureevolution of spacecraft propulsion.

- 17 -4.2 Chemical PropulsionChemical propulsion is based on the principle of converting chemicalenergy (or pressure) contained in the propellant to kinetic energy ofthrust engine exhaust gases.Currently available chemical propulsion systems can be categorisedas either hot gas, or as a border-line case, cold gas.Propulsion operating with cold gas, represents the simplest form of apropulsion system, It comprise compressed (inert) gas which isstored at high pressures in a tank, and vaporising liquids (highvapour pressure hydrocarbons), which are pressurised by their ownequilibrium vapour pressure. Expelling these gases through a nozzlecreates a thrust force.Propulsion systems operating with hot gas comprise systemscontaining liquid and solid propellants. The energy from anexothermal combustion reaction of the propellant chemicals in athruster results in high temperature reaction product gases, which areexpelled through a nozzle. The maximum exhaust velocity will beachieved when all enthalpy contained in the gas at the inlet of thenozzle is transferred into kinetic energy by its expansion in thenozzle. This is described by the equation of ‘Saint-Venant’ for anideal nozzle with a complete expansion of the gas at the outlet of thenozzle:v 2 RTe max( 1) M, => in general: v e T , (24)Mwith the assumption that besides R (gas constant) also κ (specificheat ratio) is constant. Therefore, for high values of v e , high gastemperatures T and low molecular mass M are required.However, it has to be noted that nozzles are not of infinite length sothat gases are not expanded down to absolute vacuum and therefore

- 18 -gases leave the outlet of the nozzle with residual enthalpy. Inaddition, exhaust velocities v e will be limited by nozzles wall frictionloss, jet divergences, condensation of gas if temperatures becomelow enough.Typical values of v e /v e max for chemical propulsion systems are about:0.85 ÷ 0,95 for cold gas systems with no thermal losses andvery high area ratio nozzles thus higher nozzle expansionratios.0.6 ÷ 0.8 for hot gas systems because of heat losses.Finally, exhaust velocities v e will be limited by the available energyrelease per unit of mass of propellant which is according to Eq. (11):ve2E (25)mPOne of the most energetic chemical reactions release energies such asfor O 2 + H 2 is about 13.410 6 Joules/kg and v e is then 5200 m/stheoretically, while real values are being around v e = 4200 ÷ 4500m/s. Therefore, for chemical propulsion, maximum jet exhaustvelocities are limited to

- 19 -4.2.1 Cold GasCold gas systems operate with compressed inert gas (e.g. nitrogen:N 2 ) or high vapour pressure hydrocarbons (e.g. ammonia, NH 3 ); seeTable 3 below.Cold gas systems are shown schematically in Fig.5. The typicalsystem consists of a propellant tank, fill valve, filter, pressureregulator, line pressure transducers, control valves, and nozzles. Thepressure regulator provides propellant at constant pressure as the tankpressure drops. A relief valve is incorporated downstream of thepressure regulator to prevent system rupture in the case of a regulatorfailure. With regard to compressed gas systems, the cold gas is storedat high pressures (200 - 300 bar) in a tank.Figure 5: Basic Flow Scheme of Cold Gas Propulsion Systems

- 20 -The vaporising liquid system is characterised by a liquid propellantpressurised by its own equilibrium vapour pressure and the expulsionof this vapour through a nozzle. In order to provide completelyvaporised gas, a vaporiser is included in liquid cold gas systems.A typical cold gas thruster configuration is shown schematically inFig. 6 below.Figure 6: Cold Gas Thruster ConfigurationA cold gas thruster consists of a solenoid valve with mounted nozzle.The thruster is operated by opening the solenoid valve with help ofan electric current. Typical thrust range is 0.02 to 10N for spacecraftattitude and orbit control.System-specific Impulse, I sspWith regard to the derivation of the I ssp for cold gas systems, thedenominator of Eq. (21) has to be further evaluated.Starting with compressed cold gas systems, usually the cold gasused is stored at high pressures in a tank. Therefore, for calculatingthe gas mass content in the tank, the gas law applies as follows:RpopVT mPz T(26)M

- 21 -For calculating the tank mass, the so-called Tank- PerformanceFactor usually defined as:popVTK is to be used. (27)mTTable 1 below presents typical values of tank K-factors for differentbuilt tank designs and different tank materials for compressed gas,[5]. According to Eq. (27), the higher the K-factor, the lower will bethe mass of the tank. Consequently, the tank material shall be a hightensile strength material, such as Titanium Alloy or even better, afibre composite material, like Kevlar, see Table 1 below. With regardto the tank safety factor, see Chapter 4.2.5.Table 1: Ranges of typical Propellant Tank Performance Factors, K,for High Pressure TanksType of Tank Average K*(10 4 m 2 /s 2 )Range(+/- 1sigma)(10 4 m 2 /s 2 )Remarks*Tank SafetyFactor: S=2Titanium Alloy: Ti 6Al4V 6.43 5.87-6.99 High PressureComposite Over-wrappedPressure Vessel12.20 8.29-16.11 TanksWith Eq. (26) and (27) the combined mass of the tank and propellantis: zRT mPSS mP mT mP1 (28) KM

- 23 -Further, for vaporising liquids, the mass of liquids in a tank is:mP V P(31)In order to allow certain ullage, the volume of propellant is a certainfraction (0.5 to 0.9) of the available tank volume. Hence, with V P =CV T and Eqs. (27) and (31) for the combined mass of tank andpropellant we get:mPSS mT mP mP pop1 (32) CKAgain, with regard to the non-impulse dependent system mass, m HW ,as for compressed gas systems, it can be included as a mass fractionx (%) of the impulse dependent system mass m PSS . Hence, Eq. (32)can be expanded to include m HW :mPSS pop mHW mT mP mHW mP 11 CKx(33)With Eqs. (21), (23) and (33) the system-specific impulse forVAPORISING LIQUIDS becomes:IsspmPI mtotT mHWve pop1 1 CKx(34)From Eqs. (30) and (34) it is obvious, that the thruster exhaustvelocity, v e ≡ I sp , the non-impulse dependent system mass x as wellas the type of propellant and the tank performance factor K

- 24 -influences the I ssp . However, high values of I ssp are mainly dictatedby high values of v e and low values of x, while all other parametersare of secondary importance.With regard to vaporising liquids, while no great improvement overinert gas thrusters exhaust velocity v e can be obtained, considerablesavings in propellant storage mass result from the propellant’s highdensity and low pressure. This is illustrated by values of I sp and I sspas in indicated by examples for actual cold gas propulsion systems,presented in Table 3 below. Note, that calculated values of I ssp showa good agreement with actual values of I ssp . This confirms that theI ssp -analytical tool describes very well those design parameters whichcharacterise the system’s propulsive performances.Table 3: Actual Cold Gas Propulsion Systems Performances(Listed data are examples and therefore only indicative)PROPELLANTTHRUSTERSPEC.-IMPULSEI sp (missionaverage)(Ns/kg)TOTALIMPULSEI tot(Ns)PROP.SYSTEMMASSm PS(kg)SYSTEMSPEC.-IMPULSEI ssp(Ns/kg)Nitrogen (N 2 ) 706 845 4.4 193 Vela IIIMol.Mass (M):28 kg/kmolz = 1.13, at tankpressure 250 bar706 6780 24 283 COS-B; 8REMARKSActual Propulsion SystemsI ssp values derived fromRef. 7 if not notedotherwise706 - - 273 Calculated: tank material:Ti 6Al 4V; K= 6.8·10 4 m 2 /s 2x = 5.6% (small S/C)Argon (A) 510 3900 16.8 232 OGO A,B,CMol.Mass (M):39.9 kg/kmolz = 1.02, at tankpressure 250 bar510 5500 24.3 226 TD-1A 9510 - - 248 Calculated: tank material:Ti 6Al 4V; K= 6.8·10 4 m 2 /s 2x = 5.6% (small S/C)Ammonia (NH 3 ) 800 4450 6.8 654 NRL Explorer 30 (1965)=0.62 (kg/m 3 )·10 3Max. op. pressureat 30C: 12 bar800 - - 663 Calculated: Al tank withvaporizer: K=1.5·10 4 m 2 /s 2C=0.9; x = 5.6% (small S/C)

- 25 -ConclusionAlthough of moderate impulse capability, cold gas systems, inparticular systems operating with compressed cold gas, as withNitrogen, N 2 , are still of interest in view of their simplicity, highreliability and repeatability of impulse bit.ADVANTAGES:- simplicity and reliability;- lowest cost propulsion system;- very low thrust ( 10 N) and impulse bit ( 10 -4 Ns)capability;- low plume contamination.DISADVANTAGES:- low I sp ( 950 Ns/kg) low I ssp ( 650 Ns/kg) withresulting high system mass.4.2.2 Hot Gas (survey)For increasing absolute levels of thrust and impulse requirements forspacecraft propulsion (e.g. orbit transfer and orbit control), cold gassystems are inadequate and more energetic propellants generating hotgas for mass expulsion are required, see Eq. (24).Hot gas systems are the most common type of propulsion systems forspace applications. They can be divided into three basic categories:- liquid, comprising monopropellant hydrazine (N 2 H 4 ) andstorable bipropellant (MMH/N 2 O 4 );- solid, with composite propellants;- hybrid (so far not used for spacecraft propulsion).

- 26 -The terminology refers to the physical state of the stored propellantsas illustrated in Fig.7 below. Note, only systems containing liquid inform of monopropellant hydrazine or bi-propellants and solidpropellants are used for spacecraft propulsion, while hybridpropulsion systems are used for launch propulsion.OxidiserFuelLiquidOxidiser (liquid)Fuel (solid)HybridSolid PropellantSolidFigure 7: Schematic of Hot Gas Propulsion SystemsIn contrast to compressed gas and vaporising liquids, liquidpropellants need to be pressurised in the tank to feed the thrusterswith propellant. Note that due to long spaceflight mission durations,only pressure-fed systems are used because of their inherentsimplicity compared with pump-fed systems, which are usedcommonly for launch propulsion. Therefore, hot gas propulsionsystems for spacecrafts in the gravity-free environment needpropellant tanks equipped with propellant management devices inorder to separate liquids from the pressurising gas; - details seeChapters 4.2.3 and 4.2.4 below.

G- 27 -4.2.3 MonopropellantMonopropellant systems use a single (Mono) propellant to producethrust. The most commonly used monopropellant is anhydroushydrazine (N 2 H 4 ), as noted in Table 5 below. The hydrazinepropellant is decomposed in a thruster by a catalyst and the resultinghot gas is expelled through a nozzle, thus generating thrust force onthe spacecraft. A typical monopropellant system, as shownschematically in Fig.8, uses generally nitrogen gas to expel thepropellant from a diaphragm tank into the chamber catalyst beds ofthe thrusters. The typical system contains fill and drain valves for thepressurant gas and for the monopropellant hydrazine.TFill Valve (Nitrogen)Temperature SensorN 2 Propellant TankN 2N 2 H 4(diaphragm)N 2 H 4TTFill Valve (Hydrazine)TPPressure TransducerPLatch-ValveFilterPvvvvvvvSDThrusterXHydrazine System with Catalytic ThrustersFigure 8: Basic Flow Scheme of Monopropellant HydrazinePropulsion SystemsHydrazine Gas Genera

- 28 -In addition the system contains latch valves and line pressuretransducers. Filters are provided upstream of line valves to preventdamage of the valve seats or clogging the injectors of thrusters byentrained foreign material.Since the pressurant gas is stored (at a pre-selected but relatively lowpressure, e.g. 22 bar) in the propellant tank, the propellant pressurevaries with propellant usage. A typical selection of the ullage volumeof 25% filled with pressurant gas (thus containing 75% propellant)will result in a propellant feed pressure decay, and thus in a thrustdecay of 4:1. This mode of operation is also referred to as the blowdownmode, in contrast to the pressure constant mode, whichrequires the storage of a high-pressure gas in a tank external to thepropellant tank (see ‘Bipropellant systems’).A typical monopropellant hydrazine thruster configuration is shownschematically in Fig.9 below. Thrust is produced by decompositionof hydrazine into hot gas in the presence of a catalyst such as iridiummetal supported by high-surface-area aluminum oxide granulates.The catalyst causes the hydrazine to decompose into ammonia,nitrogen gas and hydrogen gas at high temperature up to 1100 o C.This results in a fairly high specific impulse of up to I sp = 2300 Ns/kg.Typical thrust range is 0.5 to 22N for spacecraft attitude and orbitcontrol maneuvers.Propellant Inlet with FilterSolenoid Valve Decomposition Chamber(not shown)with CatalystHeat BarrierInjectorInjector HeadThermalInsolationNozzleFigure 9: Monopropellant Hydrazine Thruster Configuration

- 29 -System-specific Impulse, I sspWith regard to the derivation of the I ssp for systems operating withliquid propellants, the denominator of Eq. (21) has to be furtherdetermined. In contrast to compressed gas and vaporising liquids,liquid propellants need to be pressurised to feed the thrusters withpropellant. Therefore, the mass of the pressurising gas m pr and, ifnecessary, an extra tank with m Tpr for the pressurising gas has to betaken into account.No extra tank for the pressurising gas is needed for the blow-downmode, which is the most widely used means of tank pressurisationfor monopropellant hydrazine. As already mentioned above, at thebeginning of a mission the volume of the propellant is a certainfraction C (mostly 0.75 for a blow-down ratio of 4:1) of the internaltank volume. Consequently, the volume of pressurising gas in thepropellant tank will be V pr = (1-C)V T . Therefore, the mass of the gascan be derived easily from the gas law and will be with Eqs. (26) and(27):mprKmT( 1C)M (35)zRTWith Eq. (33) and (35) the combined mass of tank with propellantincl. pressurising gas and non-impulse dependent system mass m HWis given by:mPSS m m m m m =HWTPprHWp CKK(1 C)M zRT op= m 1 11 xP(36)

- 30 -With (21), (23) and (36) we obtain the final expression for the I ssp ofsystems operating with stored liquid and with contained pressurisinggas in the propellant tank, representing the ”BLOW-DOWN MODE”:I totsspmT mP mpr m=HWIvepop K(1 C)M 111CK zRT x(37)Eq. (37) shows, that both the type of propellant (represented by v e , ,- a high v e and a high are desirable) and the propellant storageconditions (propellant-storage pressure p op , tank-filling ratio C, typeof pressurising gas, tank performance factor K) influence the systemspec.impulse. However, as already mentioned for cold gas, it has tobe noted from Eq. (37), that high values of I ssp are mainly dictated bymaximum values of thrusters exhaust velocity v e (I sp ) and low valuesof impulse independent system mass x, while all other parametersnoted above will have only a secondary impact on values of I ssp .With regard to tank K-factors, propellant tanks, with liquidpropellants in contrast to cold gas systems, need propellantmanagement devices. In the case of monopropellant hydrazine, tanksare equipped with positive expulsion devices, which are diaphragms,and which mechanically separate the pressurizing gas from the liquidpropellant in the tank during the gravity-free condition of spaceflightmissions. Diaphragms are made typically of Buthyl-and EthylenePropylene rubber materials. For more aggressive bi-propellantliquids (see Chapter 4.2.4. below), only surface tension devices madeof stainless steel screens can be used. They work by using surfacetension forces between the propellant liquid and the metal screen toseparate liquid from the pressurising gas.

- 31 -Table 4: Ranges of typical Propellant Tank Performance Factors-Kfor Liquid Propellant Tanks, [5]Type of Tank Average K*(10 4 m 2 /s 2 )Range(+/-1 sigma)(10 4 m 2 /s 2 )Remarks* Tank SafetyFactor: S = 2Diaphragm 2.25 1.53-2.97 Tank Material:Surface tension 3.32 2.28-4.36 Titanium AlloyNo propellant propellantmanagement device4.06 3.41-4.71 Ti 6Al4VBecause of the more energetic propellant hydrazine generating hotgas for mass expulsion, higher values of v e , and I ssp are achievedwhen compared with cold gas thrusters; see Table 5 below.Table 5: Performances of Actual Hydrazine Propulsion Systems(Listed data are examples and therefore only indicative)PROPELLANTTHRUSTERSPEC.-IMPULSEI sp (missionaverage)(Ns/kg)TOTALIMPULSEI tot(Ns)PROPUL-SIONSYSTEMMASSm PS(kg)SYSTEMSPEC.-IMPULSEI ssp(Ns/kg)REMARKSActual PropulsionSystemsI ssp values derivedfrom Ref. 10Monopropellant 2163 2.64 10 5 142 1859 ECSHydrazine: 2134 6.40 10 5 375 1707 ERS-1N 2 H 4 2110 9.50 10 4 66 1440 EXOSAT2110 6.41 10 4 38 1687 GEOSv e 2300 m/s; 2163 1.49 10 5 80.1 1860 GIOTTO = 1.0·10 3 kg/m 3 2163 7.25 10 4 48.8 1486 ULYSSES2168 2.36 10 5 130 1815 MARECS2060 8.24 10 4 53 1555 METEOSAT2168 3.04 10 5 168 1810 TELECOM-1N 2 pressurant gas,Max. op. pressure:p op =22 bar2150 − − 1781Calculated: I ssp fordiaphragm tank:C=0.75,K=2.310 4 m 2 /s 2 ,x = 6% (medium S/C)

- 32 -For comparison, a calculated I ssp -value for typical propellant storageand non-impulse system mass x, parameters are also presented inTable 5, showing an overall good agreement with I ssp -values of actualspacecraft propulsion systems.ConclusionMonopropellant hydrazine for spacecraft attitude and orbit control isone of the most widely used propellants. The primary reason for suchwide acceptance of monopropellant hydrazine propulsion systemslies in their inherent simplicity (reliability) while still providingadequate propulsive performance.ADVANTAGES:- simplicity and reliability (monopropellant);- lowest cost propulsion system (other than cold gas);- space storable for long periods (> 15 years demonstrated);- low thrust capability ( 0.5 N);- moderate thrust levels available ( 22 N).DISADVANTAGES:- moderate I sp ( 2300 Ns/kg) with moderate I ssp( 1900 Ns/kg) resulting in medium to high system mass;- limited life of catalyst.

- 33 -4.2.4 BipropellantBipropellant systems are characterised by the combustion of two(Bi) propellants, a fuel (e.g. monomethyl-hydrazine, CH 3 NHNH 2 )and an oxidiser (e.g. nitrogen tetroxide, N 2 O 4 ), to produce thrust.A typical bipropellant system is shown schematically in Fig.10.HeTemperature SensorTank for Pressurant GasPTeFill Valve (H )Pressure TransducerPyro-Valve(normally closed)FilterPressure RegulatorPressure Relief ValveCheck ValvePyro-Valve(normally open)Test PortMONMMHTemperature SensorPropellant Tank(surface tension)TPPTFill Valves (MON/MMH)Pressure TransducerPyro-Valve(normally closed)FilterThrusterFigure 10: Basic Flow Scheme of Bipropellant Systems

- 34 -The propellants are injected separately into the thruster combustionchamber where they react spontaneously (hypergolic propellant) toperform high-temperature, low molecular weight combustionproducts, which are then expelled through a nozzle. A typicalbipropellant thruster configuration is shown schematically in Fig.11below. Typical thrust range is 4 to 500N for spacecraft attitude andorbit control.Figure 11: Bipropellant Thruster ConfigurationThe Bipropellant system basically consists of a pressurising-gassystem, propellant tanks (with surface tension propellant

- 35 -management devices), propellant lines and thrusters. Unlikehydrazine thrusters, bipropellant thrusters accept only a limited rangeof propellant inlet pressure variation of < 2. Therefore, the highpressuregas, generally helium, contained in a separate high pressuretank, is regulated to the desired tank pressure, e.g. 17.5 bar. Thismode of operation is also referred to as the pressure constant mode.The system contains check valves upstream of the propellant tanks toprevent possible back-flow, mixing, and combustion of thepropellant vapours in the common pressurant gas line. Relief valvesare incorporated in the system upstream of the propellant tanks toprevent system rupture in the event of a pressure regulator failure.Filters are provided in the propellant lines upstream of the line valvesto prevent damage of the valve seats or clogging of injectors ofthrusters by contamination. Finally, the systems contains pyro- orlatch valves, line pressure transducers, fill and drain valves andvarious test ports for system check out.System-specific Impulse, I sspIn the case of the constant pressure mode, which is the commonmode of tank pressurisation for storable bipropellants, C is usuallyclose to 1 (e.g. 0.95) and the mass of the tank containing thepressurising gas has to be added to the tankage mass; see Fig.10above.To include the constant pressure mode in our calculations, Eq. (37)has to be modified to include the mass of the extra gas storage tank.For the mass of the pressurising gas plus the extra gas storage tankwe get with Eq. (28) - as already derived for compressed gases: zprRTmpr mTpr mpr 1 (38)KprM

- 36 -The pressurising gas will have to fill the propellant tank plus the gasstorage tank at the end of the spacecraft mission. Therefore, with thegas storage tank estimated to have a volume of about 10% of that ofthe propellant tank, the mass of the pressurising gas can be calculatedwith the help of the gas law (see Eq. (26)):mprpop1.1VTM (39)zRTWith help of Eqs. (38) and (39), Eq. (36) can be now expanded to:mPSS m m m m m m =HWTpprTprHW pop Kp M z RTP(1 C)M 1.1oppr mp 1 1 1 1 xCKP zRT zRTKprM (40)With (21), (23) and (40) we obtain the final expression for thesystem-spec. impulse of systems operating with liquid propellants inthe ”CONSTANT PRESSURE MODE”:IsspmT mpI mprtot mTpr mHW pop1CKP K1P(1 C)MzRTve (41)1.1popM zprRT1 1 xzRTKprM

- 37 -From Eq. (41) it is obvious that, as in the case of the blow-downmode, both the type of propellant and propellant storage conditionsas well as the non-impulse dependent propulsion system mass x havea major effect on the I ssp . Because of the even more energetic bipropellantcombinations, when compared with monopropellanthydrazine, higher values of v e , and I ssp can be achieved; see Table 6.For comparison, a calculated I ssp -value for typical propellant storageand non-impulse system mass parameters are presented in Table 6,showing an overall good agreement with I ssp -values of actual builtspacecraft propulsion systems.Table 6: Actual Bipropellant Propulsion Systems Performances(Listed data are examples and therefore only indicative)PROPELLANT THRUSTERSPEC.-IMPULSEI sp (missionaverage)(Ns/kg)TOTALIMPULSEI tot(Ns)PROPUL-SIONSYSTEMMASSm PS(kg)SYSTEMSPEC.-IMPULSEI ssp(Ns/kg)REMARKSActual PropulsionSystemsI ssp values derivedfrom Ref. 10Bi-Propellant 2963 2.22 10 6 849 2615 DFS2900 2.89 10 6 1101 2625 EUROSTARFuel:Monomethyl- 2900 3.10 10 6 1170 2650 EUTELSAT-2hydrazine (MMH) 2900 2.70 10 6 1147 2354 GALILEOCH 3 N 2 H 3 2900 2.20 10 6 847 2597 INMARSAT-2Oxidiser: Nitrogen 2930 5.05 10 6 1839 2746 OLYMPUSTetroxide (NTO)N 2 O 42960 3.05 10 6 1147 2659 TVSAT/TDF1/TELE-Xv e 3120 m/s 2900 3.34 10 6 1253 2666 TELECOM-2r = 1.65; mix. ratio 1.15·10 3 kg/m 3Max. op. pressure:p op =17.5 barHe pressurant gas:K= 10 5 m 2 /s 2Kevlar tank2950 - - 2639 Calculated: I ssp forsurface tension tank:C=0.95,K=3.310 4 m 2 /s 2x = 4.5% (large S/C)

- 38 -ConclusionBipropellant systems are more complex and therefore moreexpensive than monopropellant hydrazine systems. However, theirpotential high system costs is compensated by their higher impulseperformance (high I ssp ) resulting in lower propulsion mass fractionallowing a higher payload mass. Therefore, bipropellant systems aremainly used for commercial spacecrafts with missions of highimpulse requirements. E.g. for geostationary communicationsatellites, they form a single unified propulsion system, givingmaximum flexibility in the shared use of the propellant between theorbit transfer operation, as well as the apogee and attitude controlfunctions.ADVANTAGES:- high I sp : ( 2900 Ns/kg) for F 25 N, I ssp ( 2800 Ns/kg)( 3110 Ns/kg) for F 500 N,- high thrust capability, - up to 45 000 N.DISADVANTAGES:- system complexity with added valves, regulators, etc.;- higher cost in comparison to monopropellant hydrazine systems.

- 39 -4.2.5 General System Design ConsiderationsIn order to ensure safety of personnel during spacecraft groundoperations, in general the following pressure ratings ofpressurized systems have to be followed:- The burst pressure (causing rupture) of theintegrated system shall be not less than fourtimes the maximum system operating pressure.Only the tank burst pressure in general is twotimes the maximum propellant storage pressurefor the reason of low tank mass.- The proof pressure (checking safety) of theintegrated system shall be not less than 1.5 timesthe maximum system operating pressure. Thesystem has to pass successfully the proofpressure before operating it for the first time.This applies also for the case of system repairwhere faulty equipment has to be replaced. Afterrepair, again the system has to pass successfullya proof pressure cycle.In general, propellant feed systems are an all-welded design in orderto minimize mass and ensure leak-tightness. Screw mountedconnections are used only for the connection of thrusters. This allowseasy mounting and even later replacement of this equipment ifrequired.All components, which are in contact with the propellants aredesigned for and have demonstrated their long term compatibility.Therefore, high strength titanium alloy 6AL4V and pure titaniumA40 are normally used for tanks and all other components includingtubing lines.

- 40 -4.2.6 Solid PropellantThe solid propellant rocket motor consists of a motor case,containing a propellant grain, a nozzle and an igniter. The schematicis shown in Fig.12.There are two principal types of propellants:- homogeneous propellants, which are composed of fuels thatcontain enough chemically bonded oxygen to sustain thepropellant burning process,- composite propellants, which are a mixture of powderedmetal (fuel), crystalline oxidiser and a polymer binder.Most common is the use of composite propellants, usually based onsolid aluminium powder held in a hydroxyl terminated polybutadiene(HTPB) synthetic rubber binder and stable solid oxidiser likeammonium perchlorate (AP). The propellant is premixed and batchloaded into lightweight SCHEMATIC simple OF SOLID motors. PROPELLANT MOTORIgniter Motor Case NozzleHot GasExhaustPropellant GrainFigure 12: Schematic of Solid Propellant Motor

- 41 -System-specific Impulse, I sspWith regard to solid propellant rocket motors propulsion, Eq. (34)of vaporising liquids (see Chapter 4.2.1 above) may be applied. In asolid motor (see Fig.12), the propellant tank and the combustionchamber are contained in the motor case. The motor case m case m Tis filled with propellant m P according to the volumetric loadingfraction C case ( 90%), and during motor operation, the motorchamber pressure will be: p c p op .Therefore, with Eq. (34) the System-spec. Impulse becomes forSOLID PROPELLANT ROCKET MOTORS:IsspmPI mtotcase mHW1 CcasevepcKcase1Table 7: Performances of Actual Solid Propellant Motor Systems(Listed data are examples and therefore only indicative)PROPELLANT THRUSTERSPEC.-IMPULSEI sp (missionaverage)(Ns/kg)Solid Propellant(HTPB)v e ≤ 3000 m/sTOTALIMPULSEI tot(Ns)PROPUL-SIONSYSTEMMASSm PS(kg)SYSTEMSPEC.-IMPULSEI ssp(Ns/kg)xREMARKS(42)Actual PropulsionSystems2842 7.73 10 6 2960 2611 Orbus-6 Inert.Upp.Stage Motor [1]2852 1.17 10 6 447 2617 MAGE 1S ApogeeKick Motor [11]2880 1.41 10 6 528 2670 MAGE 2 ApogeeKick Motor [11]2858 4.58 10 6 1729 2649 Solid End-burningMotor [12]2842 − − 2620Calculated: I ssp forP c = 5.8 10 6 N/m 2 ,C=0.92, x=0%K=4.2·10 4 m 2 /s 2 ,ρ=1.76·10 3 kg/m 3

- 42 -From Eq. (42) it is obvious, that the motor exhaust velocity v e ≡ I sp ,as well as the type of propellant (ρ), the motor case performancefactor K case , and the volumetric loading fraction C case influence theI ssp . As to be seen from Eq. (42), for high values of I ssp , above all thethrust exhaust velocity v e shall be high while the hardware mass xshall be low. All other parameters are of secondary importance forthe value of the I ssp . Table 7 shows a good overall agreement ofcalculated with actual values of I ssp .ConclusionIn general, solid propulsion motors can only deliver their totalimpulse potential in one firing, because off-modulation is notpossible. Therefore the usage of solid propulsion is restricted to:- orbit change (e.g. apogee or perigee manoeuvre);- impart acceleration (e.g. liquid reorientation maneuvers,separation maneuvers).ADVANTAGES:- relatively simple operation;- very high mass fraction, excellent bulk density andpackaging 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 I sp performance (2800 - 3000 Ns/kg);- limited redundancy with associated reliability and safetyissues.

- 43 -4.3 Electric Propulsion4.3.1 Propulsion ConceptsIn order to increase propulsion system impulse performance, e.g. forinterplanetary missions, the jet exhaust velocity has to be increasedbeyond the v e ≤ 5000 m/s, which is best available from chemicalpropulsion. This can be achieved by electrical propulsion that relieson externally provided electric power to accelerate the working fluid(propellant) to produce useful thrust. There are three main methodsby which the electrical energy may be converted into the kineticenergy of thrust:- Electrothermal Systems, where the propellant (gas) isheated 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-dynamicexpansion in a nozzle. Typical applications of this principleare the monopropellant hydrazine operated PowerAugmented Catalytic Thruster (PACT) and the Hydrazine-Arcjet.Heater+_Hot GasExhaustPower Supply Gas Inlet Heat ExchangerFigure 13: Schematic of Resistojet Thruster

- 44 -- Electromagnetic Systems, where a gas is heated in an arcdischarge to such a high temperature, that it is converted toneutral plasma (plasma thruster). The plasma is thenexpelled at high velocity by the interaction of the dischargecurrent with the magnetic field (Lorentz force). A typicalapplication of this principle is the Magneto-Plasma-Dynamic (MPD) type of thruster.+PlasmaExhaustArc-DischargePower SupplyGas Inlet Cathode AnodeFigure 14: Schematic of Magnetic Arcjet Thruster- Electrostatic Systems, where usually a high molecularpropellant, such as xenon gas, is ionised (ion thruster) bye.g. electron bombardment (Kaufman thrusters), or in a highfrequency electro-magnetic field (radio-frequency thrusters)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 highvelocity (30 to 100 km/s) by a strong electric field.Electrons are injected into the exhaust ion beam from anelectron emitter in order to keep it electrically neutral, thuspreventing an electric charge build-up of the spacecraft.

- 45 -To the above described category of ion thrusters, theStationary Plasma Thruster (SPT), which belongs to thecategory of Hall-effect Thrusters, uses an applied magneticfield to control electrons in a quasi-neutral plasmadischarge.+_Power SupplyElectron EmitterPositiv Ion BeamPropellant SupplyIoniserIoniser Exit GridAccelerator GridFigure 15: Schematic of Ion ThrusterFinally, as an example of liquid propellants for electrostaticelectric propulsion, Caesium (Cs) is the propellant of choicewith a melting point of 29°C. Caesium, as a liquid metal, isalso desirable because it has a high atomic mass andeffectively wets metal surfaces. It is used for field emissionthrusters, or Field Emission Electric Propulsion (FEEP)devices.In a FEEP thruster, a strong electric field is established atthe tip (tailored cones) of a pair of closely spaced electrodes,which even form a capillary propellant feed system for theliquid caesium. See also the schematic of the FEEP thrusterwhich is shown below.

Field Emission Electric Propulsion (FEEP• V• F• C• E• Sr• N- 46 -Fig. 16: FEEP Thruster SchematicWhen the field reaches a threshold value, which is in theorder of 106 V/mm (for caesium), atoms on the surface of thetip of the electrodes are ionised and eventually removed.They are then accelerated to a high velocity in between thepositive emitter (tailored cones) and the negative acceleratorelectrode. Expelled ions are replenished by the flow of liquidpropellant in the capillary feed system. A separate neutraliseris required to maintain charge neutrality of the system.FEEP-thrusters can achieve very high exhaust velocities up to10 5 m/s at the expense of low thrust levels. This is due to thelimit of power available from the spacecraft. E.g., if we takeP = 1kW available, assuming an overall power conversionefficiency of η=0.5, then with Eq. (13) we will get for the20.5thrust force: F 1000 0. 0105N.10Conclusion: Thrust levels of electric propulsion are

- 47 -4.3.2 Propulsion System Design and PerformanceAn electric propulsion system consists of a power generator (solar ornuclear), power processing system (unit), electric control unit,thruster assembly, propellant storage and propellant feed system (seealso Fig. 3, Chapter 4.1 above).Solar ArraySolar ArrayS/CInterfaceControlUnitPowerProcessingUnitThrusterGimbalPlasma/Ion-JetPropellantTankPropellantFeedSystemFigure 17: Electric Propulsion System Block Diagram- Power GeneratorElectric power can be obtained from either sunlight or from anuclear reactor. In the case of solar electric propulsion, solar

- 48 -photons are converted into electricity by solar cells. In nuclearelectric propulsion, thermal energy from the nuclear reactor isconverted into electricity by either a static or dynamic thermal-toelectricpower conversion system. Static, thermoelectric systemshave the advantage of no moving parts for high reliability, but theyhave low efficiency. Dynamic systems have moving parts (e.g.,turbines, generators, etc.) and do not scale well for small systems,but they do have a higher efficiency.- Power Processing SystemPower processing systems are required to convert the voltage fromthe 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.) tokilovolt levels for use in an ion thruster. The power processingsystem is often referred to as the Power-Processing Unit (PPU).- Electric Control Unit to operate electrically valves and thrusters.- Propellant Storage & Feed SystemsVarious combinations of propellant and thruster are possible,depending on the specific application. In general, liquid or gaseouspropellants are stored and fed to the thruster assemblies as inchemical propulsion. Details see also [1].System-specific Impulse, I sspElectric propulsion relies on externally provided electric power tocreate or augment the kinetic energy of the exhaust jet. Therefore, forthe evaluation of the system-spec. impulse, the mass of the electricpower (supply- and processing) system, as depicted in Fig.3 and

- 49 -Fig.17, has to be considered in addition to the propellant storagesystem as already dealt with for chemical propulsion systems.To describe the performance of electric propulsion systems, thedenominator of Eq. (22) has to be further determined. Both, the massof the propellant storage system and the mass of the electric powersupply system have to be considered.For systems operation with gaseous propellants, e.g. xenon, thecombined mass of tank and propellant is calculated according to theEq. (28) as derived for cold gas systems above. The mass of theelectric power supply and processing system is calculated with thesystem specific power W/kg:where:Pm El (43)v eP F(44)2is the system input of electrical energy and = P jet /P is the overallenergy conversion efficiency.With Eqs. (29), (43) and (44) the combined mass of the propellantstorage system and the electric power system is calculated forsystems operating with gaseous propellants:mPSS mEl mHW mP mT mEl mHW zRT Fve mP1 1 KM mP2x(45)

- 50 -The system-specific impulse becomes with Eqs. (22) and (45):IsspmP mTItot mEl mHW1zRTKMveFve 1mP2x(46)And for:mPFt F (47)v veewith t =, which is the thruster operating time (s), the system-spec.impulse for ELECTRIC PROPULSION SYSTEMS, operating withgaseous propellants, becomes finally:Issp1zRTKMve2v e 12x(48)Eq. (48) shows that the I ssp for electric propulsion systems dependson the parameters of propellant storage as well as on the energysupply and processing systems. According to Eq. (48), for highvalues of I ssp , high values of v e and low values of x are required.However, with regard to the impact of the energy supply andprocessing system on the I ssp , low values of v e and high values of γand as well as long thrust operation times τ are preferred. Withregard to the controversial requirement for v e , there must be optimalvalues of v e-opt , which will result in maximum values of I ssp .

- 51 -The term of an optimal exhaust velocity v e-opt can be elucidatedschematically by the following picture:Figure 18: Term of Optimal Exhaust VelocityWith increasing exhaust velocity v e, the combined mass of propellantand tank is decreasing while the mass of the power supply (withpower processing) system is increasing. The point of inter-section ofthe two curves determines the minimum of the system mass by v e-optresulting in a maximum value of I ssp .This diagram shows clearly, that with increasing thrusters exhaustvelocity v e , the mass of propellant becomes less important for themass of electric propulsion systems and therefore the system-spec.impulse I ssp , describes better the system performance than the usuallyused I sp , which is only propellant related.Maximum values of I ssp can be derived by observing the first andsecond derivates of Eq. (48) with regard to v e .

- 52 -The first derivate of Eq. (48) with regard to v e has to be set equal tozero: d d( ) veIssp 0(49)2dvv edvee1a b where:zRTa and b 2(50)KMExplicitly it follows with Eq. (49)2ve1a d( I ) 0 bsspdv2e v e1a b Hence maximum and minimum values of I ssp will be for:2(51)v e b( 1a)(52)For a maximum value of I ssp , the second derivate of Eq. (48) shallbe

- 53 -resulting in a maximum value for I ssp with the optimal thrusterexhaust velocity of: zRT e 2 1 (54) KM v optIn a similar way, the system-spec. impulse and the optimal thrusterexhaust velocity can be determined for electric propulsion systemsoperating with liquid propellants, by taking into account the relevantexpressions for mass of the propellant storage systems as derived forvaporising liquid gas- or hot gas systems above. For ELECTRICPROPULSION operating with e.g. vaporising liquid gas, the systemspec.impulse becomes:Issp pop1 CKve2v e 12x(55)The I ssp will be a maximum for the optimal thruster exhaust velocityof:veopt pop2 1(56) CKTherefore, for electric propulsion high impulse performance is notdictated by maximum exhaust velocity, like for chemical propulsion,but rather by optimum values of thrusters exhaust velocity v e-opt .Here, high values of I ssp , that is high values of v e-opt , will be achievedmainly for high values of overall specific power , overall powerconversion efficiency , and thrust operation time . The thrustoperation time will be mainly dictated by mission manoeuvreoperating times and/or max. life of thrusters.

- 54 -Parameters of the xenon gas storage system, like gas compressibilityfactor z, tank performance factor K, gas storage temperature T, andgas molecular mass M, will have only a secondary impact on valuesof I ssp .A precise quantitative determination of the I ssp of electric propulsionsystems is more difficult than for chemical propulsion systems. In thecase of electrical propulsion, the electrical power can be sharedpartly and/or temporarily with the payload of a spacecraft. Here, I sspis dependent on the operative conditions of a spacecraft. Therefore,in order to allow a more quantitative comparison of actual electricpropulsion systems, Table 8 lists examples without considering theirpower supply (solar array) systems. Listed examples are electrostaticsystems with Stationary Plasma Thrusters, SPT, and Kaufman-typeof ion thrusters.Table 8: Performances of Actual Electric Propulsion Systems(Listed data are examples and therefore only indicative)PROPELLANT THRUSTERSPEC.-IMPULSEI sp (missionaverage)(Ns/kg)TOTALIMPULSEI tot(Ns)PROPUL-SIONSYSTEMMASSm PS(kg)SYSTEMSPEC.-IMPULSEI ssp(Ns/kg)REMARKSActual PropulsionSystemsElectricPropulsionXenonXenon (Xe)Mol.Mass (M):14700 7.65· 10 5 128 5980 GALS [13];Stationary PlasmaThruster, SPT-10015107 1.210 6 111 10811 SMART-1; 14, 15;PPS 1350 (SPT)4 kg/kmol 21580 1.15·10 6 96 11980 ETS-VIII 16,Kaufman-type XenonIon Thrusterz = 0.3, at tankpressure 150 bar;K= 110 5 m 2 /s 2 ;15000 − − 11287Calculated:Xe-Propellant,γ=82 W/kg; =50%,=5000h, x=10%;SPT-100

- 55 -Table 8 notes the differences between the propulsion performancereference numbers I sp (v e ) and I ssp . Hence, the differences becomes ofparticular interest with respect to the calculation of the ‘propulsionsystem mass fraction’, m SP /m SC . Usually this is done by taking intoaccount the rocket equation and calculating the mass of propellant,Eq. (9). However, taking into account the entire propulsion systemmass, the ‘propulsion system mass fraction’ has to be calculated withEq. (19), which is related to the I ssp .The differences in calculating the ‘propulsion system mass fraction’can be illustrated by the SMART-1 project (the first Europeanspacecraft DELTA-V travelled PERFORMANCES to and orbit OF the S/C Moon PROPULSION [14] [15]) SYSTEMS as noted inTable Calculated 8. Results by are System-spec. elucidated Impulse in diagram Equation Fig. versus 19 below. Rocket Equationm PS/m S/C0,400,350,300,250,20Mission to the Moon:delta-v=3.7 km/s(constant low thrust space manoeuvres)(1) Calculated by System-spec. Impulse EquationI ssp= 10 811 Ns/kgv e= 15 107 m/s(1)(2)0,150,100,050,00(2) Calculated by Rocket EquationI sp= 15 107 Ns/kgv e= 15 107 m/s0 1000 2000 3000 4000delta-v (m/s)Fig. 19: Delta-v Performance of Spacecraft Propulsion SystemDiagram Fig.19 shows clearly that for electric propulsion, where theelectric subsystem mass, m EL , may form a major ‘dead’ dry mass ofthe overall propulsion system mass, the ‘rocket equation’ does notapply alone.

- 56 -Table 9 presents a summary of the comparison between electrical(ion) and chemical (bipropellant) propulsion.Table 9: Comparison of typical electrical vs. chemical figuresType ofThrusterSpec. -Impulse(Ns/kg)Thrust F(N)DC PowerRequired(W)Electrical(Ion thruster)Chemical(Bipropellant)Order ofmagnitude ofthe ratioION/Chemical 30 00010 -3 – 0.2 400 – 800≈ 3 000 4 – 5004 – 8(short term)10 2Conclusion:- While chemical propulsion is limited to specific impulse exhaustvelocity of

- 57 -5 PROPULSION SYSTEMS SELECTIONCRITERIAA detailed procedure for the selection of propulsion systems forgiven spacecraft mission requirements is beyond the scope of thisbooklet. The process for selecting and sizing the elements ofpropulsion systems is detailed in 17.However, in general, an important consideration for the selection of asuitable propulsion system will be the trade-off between its impulseor velocity increment (v) capability and the system mass.Consequently, when selecting a spacecraft propulsion system forgiven mission impulse demand, primarily the system will have tomeet the impulse or delta-v requirement with highest possiblespacecraft payload mass. Therefore, an important requirement of thespacecraft designer will be that the mass of the propulsion systemshall be a minimum or at least shall not exceed a certain percentageof the overall mass of the spacecraft. As already mentioned inChapter 3.2.2 above, the performance of propulsion systems cannotbe assessed only by the specific impulse I sp , but requires also takinginto account the system- specific impulse I ssp .To assess the suitability of spacecraft propulsion systems forspacecraft mission impulse requirements, value ranges of I sp and I ssp(see Eq. (16)) of various actual built systems can be derived frompublished data as summarised in Table 10. In addition, with Eq. (19)the dependence of the propulsion system mass fraction m PS /m S/C , onmission velocity increment v can be derived for any given value ofI sp (v e ) and I ssp . Curves of m PS /m S/C, plotted as a function of v fordifferent propulsion system designs with typical value ranges of I spand I ssp from Table 10, are shown in diagram Fig. 20 below.When suitable systems are selected, a refinement of the selection hasto be carried out. This process takes into consideration additionalparameters such as cost, operability, complexity and reliability, etc.

- 58 -Table 10: Comparison of typical Spacecraft Propulsion SystemsPerformancesPropellantThruster-spec.Impulse I sp(mission average)(Ns/kg)System-spec.Impulse, I ssp(Ns/kg)RemarksCold GasCompressed Gas(N 2 , A)510 – 706193 – 283Compressed Gas:Titanium TankVaporisingLiquid, (C 3 H 8 ,NH 3 )618 – 800486 – 654Liquid Gas:Al-Tank withHeat ExchangerLiquidMonopropellantHydrazine (N 2 H 4 )BipropellantMMH/ NTO,(CH 3 N 2 H 3 /N 2 O 4 )2100 – 2 300 1440 -1860 Hot Gas; Tankwith Diaphragm2900 – 3120 2354 - 2746 Hot Gas; Tankwith SurfaceTension DeviceSolid(Composites,HTPB)ElectricPropulsionElectrostatic:Stationary PlasmaThruster (SPT)/Kaufman-type Xe-Ion Thruster2800 – 3000 2611 -267014000 - 34000 5980 - 11287E.g. MAGE 1, 2Apogee KickMotorsXenon PropellantE.g. GALS/SMART-1/ETS-VIII ProjectsN.B.: Listed data, which can be derived from published data, are examplesand therefore only indicative.

- 59 -m PS/m S/C1,0I ssp=193 Ns/kg(Nitrogen)Cold GasI ssp=654 Ns/kg(Ammonia)I ssp=1440 Ns/kgHydrazineI ssp=1860 Ns/kg0,8I ssp=2356 Ns/kgBipropellantI ssp=2746 Ns/kg0,6I ssp=5980 Ns/kg0,4SPT/Ion ElectricPropulsion Systems0,2I ssp=11980 Ns/kg0,0500 1000 1500 2000 2500 3000NITROGEN (1PROPANEHYDRAZINEPACTdelta-v (m/s)Figure 20: Delta-V Performance Range of Spacecraft Propulsion System Concepts (Examples)

- 60 -The diagram in Fig. 20 gives the first and most important indicationfor the selection of propulsion systems. If we assume a system massratio of m PS /m S/C < 0.30, we can read directly from Fig. 20:- for low v 150 m/s, compressed cold gas and vaporising liquidpropulsion systems seem to be the best choice, because they meetthe requirement and have the lowest cost;- for 150 v 650 m/s, monopropellant hydrazine fed propulsionsystems are the best choice, because of their inherent simplicity(reliability) and potential low cost, while still meeting the Δvrequirement;- for high v 650 m/s, bipropellant systems, monopropellanthydrazine fed resistojet systems (power-augmented thrusters,arcjets), and electrostatic (electromagnetic) systems will satisfy thev-requirements best.Finally, for any given value of total impulse I tot , the mass of thepropulsion system m PS can be calculated of course directly fromvalues of I ssp .When a suitable spacecraft auxiliary propulsion system is selected,however, the cost, complexity, operability and reliability of thesystem also play an important role.With regard to low v-requirements, compressed cold gas systemsused for auxiliary propulsion of spacecraft’s (attitude and orbitcontrol), although of moderate impulse capability, are still of interestin view of their simplicity, high reliability, repeatability of impulsebit and low system costs.

- 61 -In considering the merits of the various compressed cold gas andvaporising-liquid systems, the following major points must beconsidered carefully:- Additional heat may be necessary for vaporising, e.g.propane for use in gas jets. For high thrust levels andlong thrust duration’s, this can give rise to thermalproblems in the propane system. As the latent heat ofammonia is about three times that of propane, additionaltechnical problems may occur here.- For zero-g conditions (non-spinning satellites), thestorage of liquefied propellants is more complex than thatof pressurised, inert gases, as tanks with bellows, surfacetension devices etc. have to be provided to separate liquidand vapour. In addition, propellant gauging is muchmore complex (with resulting higher costs) for liquefiedpropellants under zero-g conditions than for compressedgases. Moreover, fuel slosh of liquefied propellants maycause extra problems for the dynamic behaviour ofspacecrafts.Therefore, for low v-requirements, compressed cold gas systemsutilising N 2 are the most commonly used.For higher v-requirements, in the trade-off between monopropellanthydrazine systems, bipropellant systems and electric propulsionsystems, the following major points have to be considered carefully:- Because of their inherent simplicity, hydrazinemonopropellant systems still represent the lowestpossib1e cost technology in the field of liquid propulsion.Such a technology is therefore of interest whenever amoderate velocity increment is required or where mass isnot a critical design driver.

- 62 -- For delivering low impulse-bits or impulses at lowspacecraft torques or acceleration forces, hydrazinethrusters of potential low thrust levels (down to 0.5N)will have to be used.- High v-requirements for spacecraft in-orbit transfer andattitude and orbit-control can only be met by bipropellantsystems (e.g. unified propulsion system) and electricpropulsion.- For the selection of a hydrazine resistojet system (e.g.power-augmented thrusters, arc-jet), thrust level and dutycycle requirements have to be considered. The fact thatthe limited power available and heat capacitance of theelectrothermal thruster impose a limit on thrust- andduty-cycle levels may give rise to technical problems.- With regard to electric propulsion like electromagneticand electrostatic systems, this technology, although stillunder development, has proven to achieve thrustersexhaust velocities v e an order of magnitude higher thanthe best performing chemical propulsion systems; seeTable 9. Therefore, electric propulsion is essential forfurther reduction of system (propellant) mass, enablinghigher payload mass and coping best with future highenergy mission requirements. However, depending onthrust levels, electric propulsion can impose severepower requirements on the spacecraft power supply andpower processing system.When spacecraft propulsion systems are to be selected, the abovementioned points have to be assessed properly with reference to theflight mission and design requirements of the spacecraft itself.Table 11 below presents an overall comparison summary ofcandidate spacecraft propulsion systems.

- 63 -Table 11: Survey: Typical Candidate Spacecraft Propulsion SystemsTypeThrustLevelRange(N)ThrusterExhaustVelocity(m/s)AdvantagesDisadvantagesCold Gas(N 2 , A, NH 3 ,C 3 H 8 )0.02 – 10 500 – 800Extremelysimple, reliable,very low costVery lowperformance,Solid Motor (e.g.Apogee KickMotor)Liquid:MonopropellantHydrazine(N 2 H 4 )28 000 –47 0002 800 –30000.5 – 22 2 100 –2 300Simple, reliable,relatively lowcostSimple, reliable,low-costLimitedperformance, higherthrustModerateperformanceBipropellant(CH 3 N 2 H 3 /N 2 O 4 )ElectricPropulsionElectrothermal:Resistojet(NH 3 , N 2 H 4 , H 2 )Arcjet (NH 3 ,N 2 H 4 , H 2 , N 2 )Electromagnetic:Pulsed plasma,(Teflon)Electrostatic:Stat. PlasmaThruster (SPT)(Xenon: Xe)4 – 500 2 900 –3 1205∙10 -3 –0.51 300 –5 0005∙10 -2 – 5 4 000 –15 000HighperformanceHighperformanceHighperformance,5∙10 -6 – 15 000 High5∙10 -3 performance10 -2 – 0.5 15 000 –25 000HighperformanceIon (Hg, A, Xe) 10 -3 – 0.2 30 000 Very highperformanceMore complicatedsystem thanmonopropellantLow thrustHigh power,complicatedinterfacesHigh power, lowthrust, complicatedHigh power, lowthrust, complicatedVery high powerN.B.: Listed data, which can be derived from published data, are examplesand therefore only indicative.

- 64 -6 OUTLINE OF POTENTIAL FUTURESPACE PROPULSIONSo far, chemical propulsion has given access to space and has eventaken spacecraft through the solar system. Electric propulsion, stillunder development, offers a further vast increase in propulsionsystem mass efficiency.The prevailing goal of future propulsion in form of advancedspacecraft propulsion systems is to enable cost efficient spacemissions and extended exploration of the solar system up tointerstellar missions.In order to achieve efficient mission costs, an important applicationof advanced spacecraft propulsion is to reduce cost by:- reduction of the total mass that must be launched fromEarth,- reduction of propulsion system mass fraction, allowingfor higher payload mass,- increase of mission impulse performance, allowing forsatellite extended orbit maintenance and attitude control.A second goal of advanced spacecraft propulsion is to performextended (manned) exploration of the solar system and previously‘impossible’ missions, like interstellar travel.Consequently, the evolution of advanced spacecraft propulsionsystems will mainly focus on increased performance that is highvalues of I ssp .In a first instance, advanced propulsion systems can be derived fromexisting systems, by increasing the performance of chemical andelectric propulsion with regard to their mission impulse and velocityincrement,Δv, capabilities.

- 65 -6.1 Potential Improvement of ChemicalPropulsionFor chemical propulsion high performance, i.e. high values of I ssp ,resulting in low values of ‘propulsion system mass fraction’, isprimarily dictated by maximum values of I sp (v e ). See to this Eq. (21)with the mass of propellant storage system (propellant + tank), m PSS ,which is proportional to the system impulse capability and sized byI sp (v e ); see Eq. (10). However, the performance of state-of the-artspacecraft engines operating with cold and hot gas can be considerednear to the theoretical limit for actual space storable propellantcombinations.But the I ssp can be still improved by also taking into account the nonimpulsesystem mass m H/W , - see Eq. (21). The emerging class ofmicro-and nanospacecrafts requires miniaturization of thepropulsion system with help of ‘Microelectromechanical System’(MEMS) technology for acceptable values of I ssp , in order to achievea low value of m H/W .Fig.21: MICRO PROPULSION: Laser Induced Etched Nozzle withthrust force F= 0.5 ÷ 10 mN (Courtesy of Ångström SpaceTechnology Centre, Uppsala/Sweden)

- 66 -For further reading about MEMS technology in particular withreference to its space applications [18] is recommended.In addition, with increasing interest in environmental and safetyissues, non-toxic monopropellant systems are under development.Current satellite users and manufacturers are looking for moreenvironmentally friendly, safer propellants. Safer propellants canreduce costs by eliminating the need for self-contained atmosphericprotective ensemble (SCAPE) suits that are needed for toxicpropellants by personnel for propellant filling and drainingoperations. Also, extensive and prohibitive propellant safetyprecautions and isolation of the space vehicle from parallel activitiesduring propellant loading operations can be minimised or eliminated.Therefore, if environmentally safe and low toxic propellants areused, the costs for operating satellites on ground can be lowered, insome cases even dramatically.A new family of environmentally friendly monopropellants has beenidentified as an alternative to hydrazine. These new propellants arebased on blends of e.g. hydroxyl ammonium nitrate (HAN),ammonium dinitramide (ADN), hydrazinium nitroformate (HNF),nitrous oxide (N 2 O), and hydrogen peroxide (H 2 O 2 ). When comparedto hydrazine, e.g. HAN and ADN blends have a range of specificimpulse (I sp ) which can exceed that of hydrazine [19]. Testing ofHAN and ADN based propellants has begun to show promise andcould soon be adopted for spacecraft on-board propulsion systemsuse.To summarize, actual designs of chemical spacecraft propulsionsystems are well developed, but are being mainly complemented byminiaturised cold/hot gas-, as well as by low-toxic monopropellantsystems.

- 67 -6.2 Potential Improvement of ElectricPropulsionMost promising for further increase of propulsive performancecapabilities is the use of electric propulsion.For electric propulsion, high values of I ssp will be achieved mainlyfor high values of v e-opt , which requires in particular high values ofoverall specific power γ (watt per unit mass), combined with highoverall power conversion efficiency η, resulting in low values of m El .This can be illustrated by considering planetary missions to the edgeof the solar system, - mission to rendezvous with Pluto or othermembers of the Kuiper belt. For constant low thrust maneuvers, atotal Δv ≥ 37 km/s has been assumed to cover escape from theEarth’s gravitational field as well as escape from the solar systemwith no gravity assist maneuvers, thus giving wide launch windows[20]. A rough propulsion system analysis will show the needs forfurther performance improvements.Parametric investigations have been performed by altering overallsystem specific power and thrust time , considering an overallsystem power efficiency of = 0.67. In order to demonstrate theimpact of these parameters on I ssp with resulting values of‘propulsion system mass fraction’, parameters have been combinedfor extreme cases of and as follows. Overall system values of containing electric power generators, power processing systems andthrusters have been assumed in the range of = 30 ÷ 130 W/kg witha main emphasis on Radioisotope Thermoelectric Generators (RTG),needed for deep space missions. Performances of RTG’s have beenassumed for = 33 W/kg at 100 kW to = 625 W/kg for potentialfuture RTG designs at 10 MW [2]. Thruster operation times havebeen assumed for max. life of thrusters, ranging from 7000 h to20000 h [20]. In addition, it is assumed, that systems are operatingwith xenon-gas propellant.

- 68 -Parametric investigations performed according to Eqs. (55) with (56)are depicted in Fig. 22 for optimal thruster exhaust velocity v e-opt .DELTA-V PERFORMANCES OF S/C ELECTRIC PROPULSION SYSTEMSm PS/m S/C1,00,90,80,70,6(1) I ssp= 15 489 Ns/kg,v e-opt= 32 703 m/s,gamma = 30 W/kg,tau = 7 000h(2) I ssp= 28 278 Ns/kg,(1) v (2)e-opt= 59 707 m/s,gamma = 100 W/kg, (3) I ssp= 47 799 Ns/kg,tau = 7000 hv e-opt= 100 923 m/s,gamma = 100 W/kg,tau = 20 000 h(3)(4)0,50,40,30,20,1Assumption: Constant low thrust maneuversIon Propulsion, Xenon propellant(4) I ssp= 54 499 Ns/kgv e-opt= 115070 m/s,gamma = 130 W/kg,tau = 20 000 h0,00 10 000 20 000 30 000 40 000delta-v (m/s)Fig. 22: Delta-v Performance of Potential Future Electric PropulsionThe parametric investigation performed by altering and shows theimportance of the large range of specific power mainly caused bythe electric power generators which have a major impact on theoverall value of . Here mainly power supply systems with highspecific power γ need to be further developed to achieve high valuesof I ssp , while for thrusters, the “Dual-Stage 4-Grid” type of griddedion thruster with a capability of v e ≤ 186 000 m/s is already underdevelopment [20]. In addition, for high values of I ssp , thrust operationtimes should be always a maximum within the frame of missionmaneuver time in order to minimise power consumption, thus lowermass of power supply system. This could imply multiple (cluster)thruster configurations with thrusters operating serial in time in orderto obtain extra long thrust operation time, if required.

- 69 -6.3 New ApproachesNew approaches are studied or are under development, withexamples like:- Solar- thermal Propulsion:reflectorFUELHeat exchangerthrustchamberFig. 23: Schematic of Solar-Thermal Propulsion System(Courtesy by SNECMA)Propulsion system is based on using solar radiation energy to heatliquid hydrogen propellant which is passed through a heatexchanger, reaching very high temperatures up to 2500 o K, beforeexpanding through a nozzle. The advantage would be: higher thrust levels than achieved for electric propulsion, e.g.F = 5 to 10 N continuous for 70 kW (solar), higher exhaust velocities than achieved for chemicalpropulsion, e.g. v e =8000 m/s.Status: Several concepts for solar-thermal propulsion systemshave been proposed, however, so far none have been realised.

- 70 -- Nuclear-thermal rockets:There are two main different categories of nuclear technologyfor space power and propulsion: Radioisotope thermoelectric generators (RTG) and closecycle[Type a quote from(e.g.the documentSterlingor the summarytechnology)of an interesting point.forYou cannuclearposition the textelectricbox anywherepower,in the document. Use the Text Box Tools tab to change the formatting of the pull quote text box.]NEP, to power electric propulsion. Flight heritage of RTG’swith power level < 10 kWe while future NEP’s aim at>10 kWe to MWe’s for electric propulsion: v e = 20 000m/s toSCHEMATIC OF A NUCLEAR ROCKET ENGINE≤ 186 000 m/s [20]. Open-cycle nuclear thermal reactors, NTR, which heat e.g.liquid hydrogen propellant directly to produce rocket thrust.Liquid hydrogen propellant absorbs heat from the core of afission reactor, before expanding through a nozzle: v e = 8000to 9000 m/s, F = 20kN to 70 kNTurbinH2 PropellantNuclearReactorReactorFuel ElementNozzle CoolantJacketFig.24: Schematic of NTR

- 71 -Extensive research has been performed into nuclear-thermalrockets in USA in 1960 as part of the NERVA program.Status: Environmental and political concern about safe groundtest and launch of fueled reactor has reduced research in nucleartechnology.- Exotic propulsion methods, such as:Exotic Propulsion Systems are those “far out” ideas still understudy and are pure speculation so far. They will be required forthe ultimate dream of space exploration to travel to other starsystems, as depicted in TV shows like ‘Star Trek’. Two examplesof such exotic propulsion systems are outlined below. Antimatter Propulsion: Matter- antimatter annihilation offersthe highest possible physical energy density of any knownreaction substance. Since matter and antimatter annihilate eachother completely, it is an incredibly compact way of storingenergy. E.g. a round trip to Mars with a 100-ton payload mightrequire only 30 gram of antimatter. However, sufficientproduction and storage of antimatter (with potential complexand high storage system mass) is still very much in the future. Photon Propulsion: The generation of usable thrust by ejectionof photons is still very hypothetic. The generation of photonsby e.g. laser technology and their subsequent decay in space,involves the mass-energy transfer expressed by Einstein’sequation, E = mc 2 . Consequently, very large quantities ofenergy will be required even for nominal levels of thrust.Possibly, matter-antimatter annihilation can be harnessed forphoton propulsion in the futureFor a further reading about advanced propulsion systems, [21] isrecommended.

- 72 -7 GROUND TESTING OF PROPULSIONSYSTEMSThe following parameters can be measured during system groundoperations:- pressure;- temperature;- electric current and voltage on e.g. solenoid valves.The following tests can be performed:- functional tests of electric actuated valves by measuringelectric current and voltages response time of valves;- measurement of leak tightness of components, parts ofsystems and overall systems by measurements oftemperatures and pressures at various points of time;- measurement of leak tightness of the integrated propulsionsystem with the help of a gas spectrometer, either by“sniffing” or during test operations of the spacecraft in avacuum chamber;- measurement of leak tightness of valves (thrusters) withthe help of 'glass pipettes'.

- 73 -8 MISSION SURVEILLANCE OFPROPULSION SYSTEMSDuring spacecraft mission operations the following propulsionsystem 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 differentparts of systems at various points of time;- propellant consumption and remaining mass of propellant by: evaluation of pressure and temperature data (gas law) inpropellant tanks, related to initial mass of propellant atbeginning of life (BOL), also called ‘P.V.T’ (Pressure,Volume, Temperature) method; 'book keeping' of thruster operations;- thrust and spec. impulse of individual thrusters by: evaluation of propellant consumption during in-orbitthruster operations and comparison of planned andachieved in orbit spacecraft movements.

- 74 -9 LITERATURE/REFERENCES1G. P. Sutton, Oscar Biblarz, 2001, “Rocket PropulsionElements”, 7th Edition, John-Wiley & Sons, Ltd.,ISBN: 0-471-32642-9,2 P. Hill, R. Peterson, "Mechanics and Thermodynamics ofPropulsion". Addison-Wesley Publishing Company, Inc., USA.ISBN 0-201-14659-2,3456P. Erichsen, “Performance Evaluation of Spacecraft PropulsionSystems in Relation to Mission Impulse Requirements”, ESA,1997, SP-398, Proceedings of the Second European SpacecraftPropulsion Conference, 27-29 May, 1997,P. Erichsen, “A Quick-Look Analysis Tool for the ImpulsePerformance of Spacecraft Propulsion Systems”, EUCASS,Europe, 2007, EUCASS Paper 01-05-03,D.R. Trotsenburg, “A Design Tool for Low Thrust RocketPropulsion Systems”, MSc. Thesis in Aerospace EngineeringTU-Delft, August 2004,Adib Najib, „Ermittlung der impulsunabhängigen Masse imVerhältnis zu den impulsabhängigen Massen für Raumfahrtantriebssysteme“,Study work at the Institute for Aerospace,Technical University Berlin, Jan. 2004,[7 Lee B. Holcomb; “Satellite Auxiliary-Propulsion SelectionTechniques”, NASA Technical Report 32-1505, November 1,1970,[8 COS-B Project, AOCS Design Specification, Ref. D-310.0200,dated 15 March 1975,

- 75 -[9 W. Inden; “Development Results of the ESRO TD SatellitePneumatic System”, Paper Reprint from Lecture Series No. 45on Attitude Stabilization of Satellites in Orbit, AGARD,10 D. Gale, et. al., “Bibliography of Liquid Propellant PropulsionSystems (LPPS) of European Spacecraft”, Study Note:SN/ESA-P/001/89/BAe Issue 1, January 1990,11 P. Erichsen, “Catalogue of Propulsion Motors for Spacecraft”,ESTEC Working Paper No. 1348, September 1982,[12] H.F.R. Schöyer, “Some New European Developments inChemical Propulsion”, ESA Bulletin No. 66, 1991,[13] A. Bober et al., “Development and Qualification Test of a SPTElectric Propulsion System for ‘GALS’ Spacecraft”; IEPC-93-008, IEPC-93-008, 23 rd Int. Electric Prop. Conf. 1993,[14] D. Estublier et. al., “Electric Propulsion on SMART-1”, ESABulletin 129, February 2007,[15] J. Kugelberg, “Accommodating Electric Propulsion on a SmallSpacecraft”, IAF-00.S.4.09, 2000,[16] T. Ozaki et. al., “In orbit Operation of 20mN Class Xenon IonEngine for ETS-VIII”, IEPC-2007-084, 30 th Int. Electric Prop.Conf. 2007,17 W. J. Larsson and J. R. Wertz, “Space Mission Analysis andDesign”, Chapter 17: Space Propulsion Systems, Microcosm,Inc. Torrance, California and Kluwer Academic PublishersDodrecht/Boston/London,

- 76 -18 J. Köhler, “Bringing Silicon Microsystems to Space”, ActaUniversitatis Upsaliensis, Uppsala 2001,19] M. Sc Niklas Boman, Dr M. Ford, ”Reduced Hazard Propellant– Propulsion System Impact”, Proc. ’2 nd Int. Conference onGreen Propellants for Space Propulsion’, Cagliari, Sardinia,Italy, 7-8 June 2004 (ESA SP-557, October 2004),20] D.G. Fearn, R. Walker, “Interstellar Precursor Missions usingAdvanced Dual-Stage Ion Propulsion Systems”, InternationalWorkshop on Innovative System Concepts. ESTEC, 21February 2006,21 M. Tajmar, “Advanced Space Propulsion Systems”, Springer-Verlag/Wien, 2003 (ISBN 3-211-83862-7).

Notes