AIAA-2002-4208

Copyright 2002 by DLR-SART. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.1

American Institute of Aeronautics and Astronautics

EVALUATION OF PROPULSION SYSTEMS FOR SATELLITE

END-OF-L IFE DE-ORBITING

Holger Burkhardt, Martin SippelSpace Launcher Systems Analysis (SART), DLR, Cologne, Germany

Gerhard KrülleDLR Consultant, Aidlingen, Germany

Rolf Janovsky, Martin Kassebom, Hendrik Lübberstedt, Oliver RombergOHB System AG, Bremen, Germany

Bent FritscheHTG, Katlenburg-Lindau, Germany

Active post-mission disposal of space structures will have an increased importance in the future in orderto keep the Earth orbit in an acceptable condition for the safe operation of manned and unmannedmissions. In a European Space Agency study jointly performed by industry and the German AerospaceCenter, propulsive end-of-life de-orbiting concepts are elaborated. Two options, a direct, controlled de-orbit as well as a transfer to a limited lifetime orbit are considered. The suitability of various propulsionsystem options is assessed based on a set of evaluation criteria. These parameters are not limited to a merecomparison of the propulsion system but cover the impact on the entire satellite system as a whole. Theassessment strategy is demonstrated by the example of reference satellites.

Nomenclature

ACS Attitude Control SystemAOCS Attitude and Orbit Control SystemEoL End of LifeE.P. Electric PropulsionGEO Geostationary Earth OrbitH2O2 Hydrogen PeroxideIsp Specific ImpulseLEO Low Earth OrbitLOX Liquid OxygenMMH Monomethyl HydrazineN2O4 Nitrogen TetroxideOCS Orbit Control SystemPCU Power Control UnitS/C SpacecraftS/S SubsystemTC TelecommandTM Telemetry

1. INTRODUCTION

1.1. Background

The historical practice of abandoning spacecraft andupper stages at the end of mission life has allowedroughly 2 million kg of debris to accumulate in orbit.According to [1], the uncontrolled growth of thespace debris population has to be avoided in order toenable safe operations in space for the future. Spacesystem operators need to take measures now and infuture to conserve a space debris environment withtolerable risk levels, particularly in Low Earth Orbit(LEO) altitude regions.

Recent studies and publications show an increasingprobability of collisions between intact spacecraftand debris [1], [3]. If no countermeasures are taken,the number of debris particles will grow with agrowth rate in the order of 5% per year. Due to the

very high relative velocities in the order of km/s,even very small particles in the millimeter size rangecan destroy spacecraft subsystems and thuseventually lead to the loss of the complete spacecraft[2].

In principle two categories of measures to reducedebris growth can be distinguished:

• Debris avoidance• Debris removal

Simulations in [2] have shown, that a real reductionof the debris population can only be achieved withvery far-reaching measures. The only effective wayto limit the growth of the orbiting debris is tosystematically remove satellites and rocket upperstages at the end of their mission life from the nearEarth space.

The natural cleaning of the Earth environment by thetrajectory disturbances (substantially air drag) willno longer be sufficient in the future to keep the Earthenvironment in an acceptable status for the operationof manned and unmanned spacecraft.

1.2. Objectives

The work described in this paper is restricted tospacecraft passing through the LEO region. Onlypropulsive de-orbiting methods are considered. Non-propulsive methods like drag-increase e.g. withinflatable structures are not investigated within thisstudy.

The objective of this study was to provide anoverview and assessment of propulsion–relatedmethods to de-orbit spacecraft in LEO, or spacecraftwhich pass through LEO, to assess their applicabilityto different spacecraft-mission combinations and toestablish a know-how basis on end-of-life de-orbit

38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit7-10 July 2002, Indianapolis, Indiana

AIAA 2002-4208

Copyright © 2002 by the author(s). Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.

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strategies. A major task was the identification of themost suitable propulsion system to perform the de-orbit maneuver for the different classes of spacecraftwith mass ranges from below 10 kg to more than2000 kg. A set of evaluation criteria was establishedfor this purpose.

For spacecraft expected to be completely destroyedduring atmospheric re-entry with a negligible risk forthe ground population, an uncontrolled de-orbitmaneuver is permissible. In these cases it is possibleto perform a braking maneuver, resulting in a newspacecraft orbit with a limited lifetime, which wasselected in this study to 15-40 years. This remaininglifetime is assessed to be sufficient to provide theabove mentioned atmospheric cleaning effect. In thecase where the atmospheric destruction process isexpected to be incomplete, a controlled re-entry withprescribed re-entry location has to be carried out.

2. PROPULSION SYSTEMS

2.1. Performance Overview

Propulsion systems investigated in this study arechemical propulsion systems (including cold gas) aswell as solar-electric propulsion systems.

The following table gives a short overview overprincipal characteristics of spacecraft propulsionsystems. A more detailed description is given in therespective chapters.

Advantages DisadvantagesCold Gas • Simple

• Low system cost• Reliable• Safe

• Extremely low Isp• Moderate impulse

capability• Low density• High pressure

MonoPropellant

• Wide thrustrange

• Modulable• Proven

• Low Isp• (mostly) toxic fuels

Bi-Propellant(storable)

• Wide thrustrange

• Modulable• Proven

• Complex• Costly• Heavy• Toxic

SolidPropulsion

• Simple• Reliable• Low cost• High density• Low structural

index

• One thruster perburn

• Total Impulse fix• Currently not

qualified for long-term spaceapplication

HybridPropulsion

• Simple• Modulable• Low cost• Reliable

• Not qualified• Lack of suitable

oxidizer for long-term mission

ElectricalPropulsion

• Very high Isp • Low thrust• Complex• Long maneuver

time• Power

consumption

The following figures show the specific Impulse andthe thrust range of the different concepts. It has to benoted that the thrust varies over 10 orders of

magnitude, while the specific impulse varies over 2orders of magnitude.

Thrust Range

SolidBipropellant

MonopropellantCold Gas

ResistojetSPT / HCT

IonPPT

FEEP

Arcjet

0.000001 0.0001 0.01 1 100 10000

Thrust [N]

Figure 1: Thrust range of different spacecraftpropulsion systems

Isp Range

Cold Gas

Mono-propellant

Bipropellant

Solid

0 50 100 150 200 250 300 350

Isp [s]

Figure 2: Specific impulse range of chemicalpropulsion systems

Isp Range

SPT / HCT

Ion

PPT

FEEP

Resistojet

Arcjet

0 1000 2000 3000 4000 5000

Isp [s]

10.000

Figure 3: Specific impulse range of electricalpropulsion systems

Cold gas thrusters represent the smallest rocketengine technology available today. Cold gas systemsare valued for their low system complexity and theuse of benign propellants. They are howeverseverely limited in Isp, which results in increasedpropellant mass.

Vaporizing liquids (e.g. Propane) have a slightlylower specific Impulse than nitrogen cold gassystems. Due to the much lower tank pressurehowever, those systems have usually a considerablylower propulsion system dry mass and therefore alsoa slightly lower wet mass. Continuos operation ishowever problematic as important evaporation heatlosses reduce the propellant temperature and inconsequence also tank pressure and thrust [4].

Cold gas thrusters have a high degree of reliabilitydue to widespread practical experience and lowsystem complexity. There is however a greater riskof valve leakage problems, which are more mission

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critical with high pressure gas systems as comparedto mono- or bipropellant systems with viscousliquids.

Monopropellant thrusters use a single propellantwhich decomposes exothermically using a catalyst.The reaction products are expanded through a nozzleand generate thrust. The typical state of the artsystem for satellite attitude control is a hydrazineblow-down system. It has a simple architecture andmanufacturers can rely on a long design heritage.

Pressure regulated systems are an alternative. Thefeed pressure of the thruster is kept constant, thusavoiding the decline of specific impulse and thrustover consumed propellant. The more complexsystem architecture is however penalising, making ita very rare choice.

Hydrogen peroxide is the only alternative propellantin use. Its is however solely used for launcher rollcontrol systems, as it is the case for Soyuz, due tostability problems of the propellant for extendedstorage times.

The denotationbipropellant thruster is an umbrellaterm for a class of thrusters. Their common feature isthe use of a liquid fuel and a liquid oxidizer toproduce thrust. They are further classified intostorable, semi-storable (LOX + hydrocarbons) andcryogenic, in function of the propellant properties.Solely storable bipropellant systems are used insatellites AOCS for obvious reasons. MMH andN2O4 are the state of the art combination forbipropellant satellite thrusters.

The feed mode of the thrusters depends mainly onmass and performance constraints. Bipropellantthrusters on satellites are usually supplied with fuelby the means of a constant pressure system.

Extensive use of bipropellant thrusters for countlessmissions have proven a high reliability of thistechnology. There is however always an inherent riskwith growing system complexity that can not beeliminated completely. The reliability of bipropellantsystems as compared to a monopropellant systemwith equal redundancy level is therefore lower.

In solid motors, fuel, oxidizer and an organic binderare combined into a composite to form the solidpropellant.

The advantages of solid rocket motors are theircompact size combined with a relatively highspecific impulse performance. Propellant leakagecannot occur for obvious reasons, whereas propellantoutgasing during in-space storage does not seem tobe an insurmountable problem even for very longduration [5]. Solid propellant motors have a veryhigh reliability, mainly due to their simplearchitecture. It has however to be taken into accountthat the achievable accuracy is limited, so reliability

with respect to a constrained objective is onlymoderate.

Hybrid propulsion provides the potential forthrusters that are lower in cost, more reliable, saferand environmentally compliant. Most hybrid systemsunder study use a combination of solid, non-explosive fuel and a liquid oxidizer. Unlike solidrockets currently in use, the hybrid system can bethrottled, shut off and restarted during flight.

Hybrid propulsion technology has not yet found itsway to application. One problem is the lack ofsuitable oxidizers for long mission duration. NeitherLOX nor H2O2 are storable over long periods oftime. Nitrogen tetroxide requires a separate ignitionsource while chlorine-fluorides (ClF3/ClF5) arehighly toxic and corrosive.

The use of hybrid engines for EoL applicationswould require a full development program, startingat the point of propellant selection. If successful,hybrid engine technology could fill a useful gapbetween high performing, yet complex bipropellantengines and compact and simple, yet relativelyinflexible solid motor technology [6].

Electric propulsion is known for its high specificimpulse of 500 to over 3000 s which leads tosubstantial propellant mass savings compared toclassical propulsion. The considerable powerconsumption of typically 500 to 1500 W (perthruster), however, constitutes a serious usageconstraint to E.P. Yet, the applicability of electricpropulsion for de-orbiting will finally be decided bythe compatibility of the low thrust level with missionrequirements. For background literature, see [9]-[14].

The assessed E.P. types are:� Arcjets (Electrothermal Arc Thrusters)� Resistojets/EHT (Electrothermal [Hydrogen]

Resistance Thrusters)� SPT/HCT (Stationary Plasma Thrusters / Hall

Current Thrusters)� Ion Thrusters (Electrostatic Discharge or Radio

Frequency Thrusters� PPT (Pulsed Plasma Thrusters)� FEEP (Field Emission Electric Propulsion)

The magneto plasma dynamic thrusters (MPD),having a high thrust capability in the order of 1 Nand above, are not considered in this study becauseof their high power demand and the limiteddevelopment status.

The individual characteristics of electric propulsiontypes are presented in Table 1. In principle, numbersrefer to the range of near flight qualified specimensamong types. Smaller and, particularly, larger andmore powerful units are being developed. In cases ofadvanced developments, the extended range is givenin parenthesis.

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Arc Jet Resisto/EHT SPT/HCT Ion Thruster PPT FEEPThrust level [mN] 100-250 (100-) 300-400 (20-) 80 –100 5 -40 (-200) 0.1 – 2. 0.001 - 1. (-2.)Specific impulse [s] 480-700 300 (-1000[H2]) 1600-2000 2500-3800 (500-) > 1000 6000-10000Overall efficiency [%] Up to 40 80 – 90 (*) Up to 50 Up to 60 < 20 (*) Up to 50 (*)Total impulse [Ns] Up to 1.E+06 > 0.5E+06 Up to 2.5E +06 Up to 1.2E +06 Up to > 5000 ?Lifetime [h] >=1200 (ver.) > 400 (ver.) Up to 7000 Up to 15000 * Dep.on no.pulses ?Maximum no. of cycles 1000-2000 (v.) 5.E+5 off-pulses <= 6500 (ver.) Up to 3000 >50 M pulsesPropellant (alternatives) NH3,N2H4,

(MMH, H2) etc.N2H4, H2O, (Xe,H2)

Xe (Kr) Xe Teflon Cs or similar

Propellant mass flow [mg/s] 20-50 100. (2.-) 5. – 10. 0.5 (- 5.) ? per pulse ?Thruster Mass [kg] 0.5-1.5 0.3 – 0.8 (1.-) 3. - 5. 2. (- 7.) ∼ 1. (*) 0.5 (*)Propulsion S/S Mass (per unit)[kg]

5.0-10.0 (*) 2. – 5. 10. – 25. (*) 15. –20. (- 30.) 2.5 – 5. ∼ 3. (*)

Thrust vector misalignment [°] < 1 < 1 < 2 (*) < 1 < 10 (*) < 10 (*)Thrust homogeneity (level) [%] ± 2 (*) ± 2 ± 3 (*) ± 2 ± 30 (pulse) (*) ± 5 (*)Exhaust beam diverg. [halfangle,°]

< 5 (*) < 5 > 30 (90%) (*) 12 (90%), 25(95 %) (ver.)

> 30 ∼ 20

Reliability (of unit S/S tocomplete de-orbiting mission)

98 (*) > 98 (*) 95 (*) 95 (*) 98 (*) 90 (*)

Subsystem emission Broadband noise (arc jet; SPT, Ion [neutralizer]), thermal radiation (arc jet), slight cathode[neutralizer]+anode erosion (arc jet, SPT, Ion); in operation: no serious problem verified !

El. Power demand (system)[W]

750-2000 (50-) 500-600 (300-) 1300-1500

500-800(-5000)

10-100 (-200)(contin.)

?, low

Environmental conditions Normal S/C equipment requirements, non op. temp. limits (thruster)∼ -50/+100°CS/S integration Must be compliant with S/C configuration and mission requirements (thrust vector)S/S control requirements Internal control by S/S PCU including start-up and shut-down sequence;

on/off, preheating (?*) commands by central electronic or remotelyS/S ground handling Normally no pb.; performance tests executed on S/S level; possible pb with FEEPNumbers marked by stars (*) are estimated.

Table 1: Electric Propulsion Characteristics

It has to be mentioned as an advantage that arcjetsand resistojets (EHT), when fuelled with hydrazine,can work together with a monopropellant AOCSmaking use of the same tank and fuel managementsystem.

2.2. Requirements

The integration of a de-orbit propulsion functionalityhas a considerable impact on satellite design anddetermines a set of additional requirements. Theconsequences are more profound for obvious reasonsif the satellite does not need an OCS for the nominalmission profile.

Requirements include suitable volume allocatedonboard the spacecraft for thrusters and relatedhardware, a sufficient power supply, compatibilitywith the ACS, thermal control in certain cases andthe integration with the on-board computer andsatellite control. In order to achieve the desiredobjective of reducing space debris, a minimumreliability Rde-orbit function>90% of the de-orbit functionwas required for the purpose of this study.

Furthermore, additional ground handling duringintegration and during launch preparations comealong as well.

3. DE-ORBIT CONCEPTS

3.1. De-Orbit Options

The principal disposal regions for spacecraft areindicated in Figure 4.

In general, the post-mission disposal options are:

Option 1: maneuver to an orbit for whichatmospheric drag will remove the structure within agiven time frame, e.g. 25 years, (uncontrolled de-orbiting)

Option 2: direct retrieval and de-orbiting,(controlled de-orbiting)

Figure 4: Disposal regions and storage orbitoptions for post-mission disposal

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Option 3: maneuver to one of a set ofdisposal regions in which the structures will notinterfere with future space operations.

Uncontrolled De-orbiting (Option 1)The de-orbiting is initiated by one or moredeceleration maneuvers, either with relatively shortpulses of a “high”-thrust propulsion system or withextended low-thrust maneuvers of electric propulsionsystems. The deceleration maneuvers reduce theperigee altitude of the orbit, thus increasing theaerodynamic deceleration and therefore reducing theorbital lifetime. The satellite will be transferred intoan orbit in which, using conservative projections forsolar activity and atmospheric drag, the lifetime willbe limited to no more than e.g. 25 years.

Controlled De-orbiting (Option 2)In case that the atmospheric destruction process isexpected to be incomplete, an uncontrolled re-entryis not permissible due to the risk for life on ground.In these cases, the re-entry maneuver has to becontrolled, which means that the point of re-entry –and thus also the point of impact on ground will haveto be predetermined. The controlled re-entry isinitiated by one or several propulsive retro-burnmaneuvers of sufficient∆V, in order to lower theperigee altitude to a level which ascertains animmediate atmospheric capture. Such a controlledre-entry under a relatively steep atmosphericincidence angle produces a confined ground impactarea of break-up fragments which have survived theaerothermal heating. The ground impact area mustbe selected such that a tolerable residual risk topersons on ground can be achieved.

The option 3, which might be attractive from anenergetic point of view for orbit altitudes above1500 km with a transfer to a disposal orbit of about2500 km altitude and which is common practice forGEO satellites EoL (‘graveyard’), will not beconsidered here.

3.2. De-Orbit Strategy Trade-Off

The main design driver for a de-orbit system is themagnitude of the braking maneuver to be performedat the end of mission life of the spacecraft. In cases,where an uncontrolled de-orbit is permissible, thespacecraft will be sent to a new orbit with a pre-defined remaining lifetime, after which thespacecraft enters Earth’s atmosphere for anuncontrolled destructive re-entry.

Figure 5 shows the remaining orbit life-time ofspacecrafts as a function of the Area-to-Mass ratioand the altitude. The main influencing factor for theorbital lifetime is the initial orbital altitude. In theshown range of Area-to-Mass-ratios, the lifetimevaries from a few days to a few months for an initialorbital altitude of 300 km to several hundred or eventhousand years at 900 km initial altitude. The Area-to-mass ratio has a relative strong impact for

compact satellites (Area-to-Mass≤≈0.005 m2/kg).For very large and lightweight spacecraft theimportance of this factor decreases. Three horizontallines limit the range with 15, 25 and 40 yearsmaximal orbit life. Spacecraft with an circular orbitof ≈600 km are in the vicinity of these boundaries.

0,01

0,1

1

10

100

1000

0 0,002 0,004 0,006 0,008 0,01 0,012 0,014

Area-to-Mass Ratio [m^2/kg]

Rem

aini

ngO

rbit

Tim

e[y

ears

]

Initial circular orbit 300 km

Initial circular orbit 600 km

Initial circular orbit 900 km

lifetime.graf

Initial circular orbit 400 km

Initial circular orbit 500 km

Safir-2

Abrixas

ERS-1&2

m=2400 kg-S/C

MuninIRS-1C

152540

Figure 5: Orbital lifetime of satellites

Figure 6 shows the required∆V as a function of theinitial orbital altitude for the transfer of twospacecraft to an orbit with 25 years remaininglifetime. The two reference spacecraft cover thepractical range of Area-to-Mass ratios and thus canbe considered as lower and upper boundaries. Twobraking strategies are compared. The first strategyuses a high-thrust impulsive Hohmann-typemaneuver, sending the spacecraft to an elliptic orbitwith 25 years remaining lifetime. The secondstrategy uses a low-thrust propulsion system withcontinuous operation, resulting in a spiral-typetrajectory and leaving the spacecraft at the end of themaneuver in a circular orbit with 25 years remaininglifetime.

0

100

200

300

400

500

600

700

600 800 1000 1200 1400 1600 1800 2000

25 years, aver. A/m, Delta-V (Hohmann) [m/s] (Safir-2)25 years, aver. A/m, Delta-V (Spiral) [m/s] (Pathfinder)

25 years, aver. A/m, Delta-V (Hohmann) [m/s] (Pathfinder)

25 years, aver. A/m, Delta-V (Spiral) [m/s] (Safir-2)

Del

ta-V

[m/s

]

Initial orbital Altitude [km]Ref_SC_Lifetime(Defined Lifetime)

∆∆∆∆V required for transfer to orbit

Comparison of impulsive and low-thrust maneuverSafir-2 A/m=0.0047 m2/kgPathfinder A/m=0.02 m²/kg Low-Thrust maneuver

Hohmann- maneuver

with limited lifetime 25 years

Figure 6: ∆∆∆∆V-requirement for uncontrolleddisposal to orbits with 25 years remaining lifetime

For low orbits below <≅ 615 km no maneuver isnecessary to limit the orbital lifetime to 25 years.Above≅ 615 km a∆V of up to 380 m/s for a circularorbit with 2000 km initial altitude is required for ade-orbit with an impulsive Hohmann-transfer. Alow-thrust-maneuver requires in the range of up to670 m/s for a very compact spacecraft (Safir-2) witha 2000 km initial LEO. This big difference in overall∆V between a continuous low-thrust maneuver and

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an impulsive maneuver, especially for high initialorbits, led to the conclusion that, even with low-thrust systems like electric propulsion, it is moreeffective to subdivide the overall braking maneuverinto a number of sub-maneuvers with braking only inthe Apogee-region of the orbit. This multi-burnstrategy limits the required overall∆V to a valueclose to that of a single-burn Hohmann-typemaneuver. Furthermore, the accumulated thrusteroperating time is reduced and operation from batterypower in eclipses is avoided for E.P. if periapsisdirection is chosen properly. As a disadvantage, thischoice leads to an increased mission time and therequirement of cycled thruster operation.

The differences in the required∆V between the twospacecraft, reflecting the range of compact and notcompact designs is in the order of 15-20%.

A controlled de-orbiting of a spacecraft resulting inan immediate atmospheric re-entry at a predefinedposition is only feasible with a high-thrustpropulsion system, at least for the final braking. Inorder to confine the impact area in its size to a fewhundred km2, it is necessary to have a sufficientlysteep flight path angle at the atmospheric re-entry.Simulations predicting the destruction of thespacecraft during re-entry and the impact area sizehave shown, that the Perigee of the transfer orbitleading to the re-entry has to be 60 km or lower. Thisrequires to have a flight path angle at theatmospheric re-entry window, defined here at 120km, in the range of -1.5°÷ -2.5°, depending on theinitial orbital altitude before the final brakingmaneuver. Figure 7 shows the required∆V toperform a controlled re-entry as a function of theflight path angle in 120 km and the initial orbitalaltitude.

100

150

200

250

300

350

400

450

500

-4 -3,5 -3 -2,5 -2 -1,5 -1 -0,5 0Flight Path An gle at 120 km [°]

Del

ta-V

[m/s

]

Hohmann-Transfer:Initial circular orbit

500 km 650 km

850 km

1000 km

2000 km

re-entry conditions

Figure 7: Effect of initial altitude (circular orbits)and re-entry flight path angle on ∆∆∆∆V

From Figure 7 it is obvious, that the∆V-requirementvaries strongly with the initial orbital altitude.Whereas a satellite with an initial altitude in theorder of 500 km needs only about∆V500 km=150 m/s,a satellite in 2000 km requires more than∆V2000

km=450 m/s.

3.3. Influence of Atmospheric Drag

A controlled de-orbit maneuver requires a stablesatellite trajectory and attitude until completion ofthe last de-orbit boost. In the lower atmosphere thestrongest disturbing forces and torques are theaerodynamic actions. Aerodynamic force and torquedepend on the satellite geometry and on theatmospheric conditions. During re-entry they vary bymany orders of magnitude. This variation is mainlydue to the linear dependence on the atmosphericdensity. The forces and torques also scale with thesize of the satellite.

During re-entry the aerodynamic drag equals thegravitational attraction along the trajectory usuallytwo times: the first time, when the satellite leaves itshigh altitude orbit and reaches the lower atmosphere;the second time, when the satellite approachesground. The first case occurs typically at an altitudeof about 95 km. This means, that the drag can beneglected in the considerations about the de-orbitcontrol.

The relation between the aerodynamic torque and thegravity-gradient torque depends strongly on theaerodynamic moment coefficient of the satellite.Typically both torques become comparable between250 km and 300 km altitude. The aerodynamictorque decreases with the satellite size, but thenecessary thrust to control the attitude increases withsize. The aerodynamic torque depends strongly onthe satellite shape and on the location of the center ofmass. This implies a lower limit for the altitude, atwhich a given satellite (shape, size) can be de-orbited in a controlled way with its on-board attitudecontrol capabilities. This is especially true for ACSwith electrical propulsion thrusters.

If the required attitude cannot be maintained atperigee, the satellite will leave the aerodynamictorque dominated region with a superimposed (free)rotation, which may be difficult to compensate andmay inhibit any subsequent directed de-orbit thrustmaneuvers. This is even true for satellites whichhave an aerodynamically stable attitude and whichwill oscillate at perigee about this attitude.

From the aerodynamic point of view it may beconcluded, that a single-burn maneuver with the aimof an immediate re-entry should be performed above300 km altitude (This restriction can be relaxed forvery heavy or fast spinning satellites). Multi-burnmaneuvers or low-thrust spiraling re-entries have toconsider the torque to be provided by the onboardACS and corresponding propellant consumption tocontrol the attitude against the aerodynamic torquebelow 250 km altitude.

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4. EVALUATION

4.1. Propulsion Evaluation Criteria

The choice of a propulsion system for the de-orbitfunction is a complex task as design choices affectmany parameters. Any potential system has to fulfil anumber of minimum requirements, which are givenby physical or operational requirements. Therelevant parameters are listed subsequently. Theimpact differs considerably depending if apropulsion system for nominal mission life is alreadyforeseen for the satellite and solely the performancehas to be adapted or if a propulsion system has to beadded.

Thrust level

The maximum thrust level is defined by themaximum tolerable acceleration level that thesatellite can sustain. In the case of low thrustpropulsion systems, it may be necessary to perform amulti-impulse Hohmann transfer. The minimumthrust level is hereby constraint by the upper limit ofone year for the time span of the active de-orbitmaneuver.

Duration of steady state firing

The duration of the de-orbit thruster firing has to becompatible with the chosen steady-state burn-time ofthe thrusters.

Lifetime / Total Impulse of Maneuver

Life time and total impulse limits of thrusters have tobe respected. This is especially to verify if the de-orbit propulsion thrusters are also used duringnominal mission life.

Electrical Power Demand

Electrical power demand of chemical thrusters islow. Compliance with available power at end ofsatellite mission life has to be checked, though canbe considered unproblematic in most cases.

The electrical power needed to operate electricalpropulsion subsystems is significant. This power isto be made and to be held available over the wholede-orbiting duration taking into account eclipses andbattery operation if applicable.

Environmental Requirements

During launch, satellites and its components sustainimportant shock, vibration and acceleration loads. Inorbit, changing thermal conditions prevail especiallyfor 3-axis stabilized satellites. The propulsion systemhas to be compatible with these loads.

Size / Volume

Mainly the significant volume for propellant tankshas to be accommodated in a suitable, preferablycentral location. Liquid propulsion systems have anadvantage as the different components can bedistributed within a given limit. This is not possiblewith solid propellant motors, making the placement

sometimes more difficult in spite of theircompactness. Electric propulsion systems comealong with a number of auxiliary equipment, e.g. thePCU’s which have to be mounted within the S/Cbody.

4.2. Quantification of Evaluation Criteria

The selection of the most suitable de-orbitingmethod is a task influenced by many parameters. Onthe one hand, the impact of the de-orbiting methodon the overall spacecraft budgets (e.g. cost, mass,volume) shall be as small as possible, on the otherhand, the maneuver has to be performed with asufficient probability of success, in order to get thedesired LEO-cleaning effect. For the assessment ofthe different de-orbiting methods, the following setof evaluation criteria was defined:

Cost

The additional cost for the capability of performing ade-orbit maneuver incl. additional AOCS andpotentially a complete de-orbit propulsion systemhas to be carried by the satellite operator. This canbe a critical subject for a planned mission.Depending on the de-orbit method additional de-orbit cost can consist of:

� cost for the development of the enablingtechnologies,

� cost for possible adaptation, qualification,manufacturing, testing and integration of de-orbit system hardware & software includingpossible S/C design changes,

� increased launch cost due to increased mass andcomplexity of satellite,

� de-orbit ground control station operation cost.

Reliability

The reliability of the de-orbit system, in particularthe de-orbit propulsion module, is an importantcriterion since the component shall be reliablyactivated at end of satellite’s mission, i. e. after atime period in the order of 5 – 10 years. Anotheraspect is the risk potential for the satellite due tomalfunction, i. e. in worst case explosion of the de-orbit device or its unintended early activation.Especially if a controlled de-orbit is necessary theaccuracy of meeting the target point plays anothervital role.

Mass / Volume

In most cases, system mass will be a considerablecost driver. High specific impulse thereby inherentlyleads to lower propellant mass. Nevertheless, thosesystems have not necessarily lowest system mass, assystem dry mass differs considerably depending onthe functional principle [7].

When looking at very small satellites, the importanceof the specific Impulse of the propellant mightbecome negligible. The smaller the spacecraft gets,the higher the dry mass fraction of the propulsion

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system tends to become. Furthermore, there are oftenflight opportunities as secondary payloads, where thespacecraft has to fit within a predefined volume,which is often more of a design driver than the mass[8].

Power Demand

Additional power, exceeding the nominal missionlife power, might be necessary for the de-orbitfunction due to the propulsion system itself,additional or extended ACS and TM/TC to ground.This can impact the power generation as well asenergy storage.

Duration

The time duration is important since, besides theoperation cost, it is directly related to the probabilityof failures, i. e. reliability, the requirements for thepropulsion system, the energy balance, the powerneed and maneuver accuracy.

Debris Casualty Area

Satellites with lower mass are normally completelydestroyed during re-entry, while wreckage parts anddebris of larger satellites can impact on ground. Alsosatellites consisting of or including special materialsmight not melt completely during re-entry.

Impacting parts mean risk for live, groundinfrastructure and environment. The parameter forevaluation is the debris casualty area, which varies infunction of a changing path angle at the entrywindow into the atmosphere.

Complexity

System complexity reflects impacts on interfaces,structure, geometry, performance, power supply andadditional sensors. It directly affects costs foracquisition, assembling and testing.

Operational Capability

The operational capability of a propulsion system ischaracterized by factors as controllability, re-startcapability, maximal operation time, throttle range,temperature range, thrust level oscillation, efficiencyand capability of autonomous operation.

4.3. Comparative Assessment for ReferenceSatellites

The integration of a de-orbit function can be realizedin two ways. Either the spacecraft OCS is adapted tothe extended requirements, comprising the nominalmission and the de-orbit task (‘Dual-use’), or adedicated de-orbit OCS is installed in addition(‘Add-on’). However the first option typically is notavailable for small spacecrafts as they oftentimes canperform their nominal mission without any OCS.

When assessing add-on systems, thrusters, PCU’s ifapplicable, a propellant management system, tubing,cabling, (delta-) tank, and possiblly additional powergenerating equipment is added to the S/C. In

contrast, when assessing dual-use systems, many ofthe above listed items are readily available on thespacecraft and have solely to be adapted to theaugmented performance requirements. In this case,only the adaptations have to be assessed.

Subsequently, only add-on systems are presented andassessed. The additional S/C equipment, which isneeded for the de-orbit function, is accounted foraccording to the criteria of section 4.2. Using astandardized quantitative evaluation scheme, anabstract value, the so called System Cost Figure(SCF) is calculated. A lower SCF correspondsthereby to a better performance.

The SCF shows the relative influence of theindividual assessment criteria. However it should notbe mistaken to have a direct monetary equivalent, i.e.that the cost of the de-orbit function is directlyrelated to the SCF figure.

In this study, cost, reliability, mass and volume wereselected to have the highest priority and thus thehighest importance in the selection of the de-orbitingconcepts.

For the comparative assessment of the de-orbitingconcepts, two maneuvers generating a total∆V of100m/s respectively 200m/s were defined. This∆Vmust be generated by the propulsion systems eitherin one single-burn maneuver or in a number of finiteburns close to the apogee of the initial orbit. A∆V=100 m/s is for example sufficient to perform acontrolled de-orbit from an initial 400 km orbit. It isalso sufficient to send a spacecraft from an initialaltitude of up to 850 km to an orbit with a remaininglifetime of 25 years. The results of the comparativeassessment for two of the selected referencespacecrafts are presented in the followingparagraphs.

Pathfinder

Figure 8: Pathfinder

The Pathfinder-S/C with a nominal mass of 1.6 kg isthe smallest spacecraft considered in this study.Pathfinder is presently in a conceptual design phaseand intended to investigate the magnetic field of theEarth. For this purpose, a number of 100-200satellites shall be flown. Pathfinder has a cold gassystem used for a spin-up to stabilize the spacecraft.Due to its small size, the implementation of a de-orbit function has a tremendous impact on the S/C-design, regardless what type of components areselected. This is mainly caused by the fact, that not

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all necessary components for a de-orbit can beminiaturized and adapted to the small size of theS/C. Thus the additional mass and volume requiredfor the implementation of the de-orbit function are ofthe same order of magnitude like the original S/C.Also the cost linked with the implementation of apropulsive capability of 100 m/s are roughly of thesame order of magnitude as the original S/C withoutde-orbit function, having thus a significant impact onthe overall SCF.

The original spacecraft is designed for an averageelectric power of 1 W with a peak power of 5 W.This is by far too low for the operation of anyelectric propulsion system to be used for the de-orbiting. Thus only a cold gas system, a solidpropellant system, a mono-propellant system and abi-propellant system were included in theassessment. The impact of the implementation of ade-orbit function on the Pathfinder is shownsubsequently, showing individual contributions tothe SCF.

0

500

1000

1500

2000

2500

Cold Gas SolidPropellant

Mono-Propellant

Bi-Propellant

Capability

Complexity

Cas. Area

Man. Time

Power

Volume

Mass

Reliability

Cost

Sys

tem

Cos

tFig

ure

[-] Pathfinder System Cost Figure

De-Orbit Tas k ∆∆∆∆V=100 m/s,Single Burn

Figure 9: SCF for Pathfinder De-Orbit options

The solid propulsion system is by far the best choicewith regard to the overall SCF. The numericaladvantage in the overall SCF is in the order of 40%,compared to the second choice, which is the mono-propellant system. This result is linked to theadditional cost of the de-orbiting system, whichranges from about 0.7 to 1.5 Mÿ. The solidpropellant system represents thereby the cheapestsolution. Also with regard to additional mass andvolume, the solid propellant system is the bestchoice, followed by the mono-propellant system andthe cold gas system. Mass and volume increase onlyby less than 1 kg respectively 2.5 ltrs, whereas it ismore than +2.7kg respectively 4ltrs for the mono-propellant system as second best choice.

It is important to note that the assumptions for asolid propulsion system are based on published dataof small solid motor prototypes. Currentcommercially available solid motors have too highthrust levels and too short burn times, leading toexcessive acceleration levels.

IRS-1C

Figure 10: IRS-1C

The IRS-1C S/C belongs to the medium-sizedspacecraft and it is designed for Earth observationpurposes. It has a nominal mass of 1250 kg withoutde-orbit device. The spacecraft is 3-axes stabilizedand uses a hydrazine OCS. The multi-burn strategyallows to perform a de-orbit maneuver with a thrustlevel of less than 10-50 Newton for the “chemical”propulsion systems. Here electric propulsion is alsoan option, but the available power level is still low.The maneuver has to be subdivided into some 1000burns in the apogee-region, resulting in overallmaneuver duration of 20 days or more. The impactof the implementation of a de-orbit function on theIRS-1C is shown in Figure 11.

The SCFs of the de-orbit options for the IRS-1Cspacecraft are considerably lower than the SCFs forthe Pathfinder spacecraft. It can be deduced fromthis fact that the relative impact of the addition of ade-orbit capability is more profound on thePathfinder spacecraft. However, conclusions that theaddition of the de-orbit function is e.g. more costlyor heavier in absolute terms on the Pathfinderspacecraft as compared to the IRS-1C spacecraft arenot permissible and not valid.

0

20

40

60

80

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120

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Cold Gas SolidProp.

Mono-P. Bi-Prop. Arcjet Ion

CapabilityComplexity

Cas. AreaMan. Time

PowerVolume

MassReliability

Cost

Sys

tem

Cos

tFig

ure

[-] IRS-1C System Cost Figure

De-Orbit Task ∆∆∆∆V=100 m/s,Multi Burn

Figure 11: SCF for IRS-1C De-Orbit options

For IRS-1C the solid propellant propulsion system aswell as the mono-propellant system are performingcomparable with regard to the overall SCF. The solidpropellant system has some advantages with regardto cost, mass and volume, the mono-propellantsystem with regard to operational flexibility. The bi-propellant system is only a few percent behind. Thetwo E.P-options are ranked on 4th and 5th place withsome distance. The cold gas system is clearly on thelast rank due to its mass- and volume-penalty, whichalso has an impact on the overall cost (increasinglaunch cost).

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The implementation of a de-orbit function into theIRS-1C S/C implies an additional cost of about 2 to10 Mÿ, with the solid propulsion system being themost inexpensive solution. This is in the order of2%-3% of the overall S/C-cost without de-orbitfunction. The mass impact for a braking maneuverwith a total∆V=100m/s is, with the exception of thecold gas option, in the order of 60 to 80 kg or 5% ofinitial spacecraft mass.

Due to the fact that the available power level is to beconsidered moderate, time needed for de-orbitingusing E.P. becomes relatively long. Fuelconsumption is relatively low, yet, E. P. hardwaremasses including additional solar power mass ifnecessary kill a considerable fraction of the fuelsaving advantage. E.P. turns out not to be thesolution of choice in this configuration. If in contrastsolar power is sufficient to operate E. P., or, goingfurther, E.P. is already installed on the S/C for thenominal mission and there is no problem to use theS/S for de-orbiting as well (requirement for thrustvector directions while respecting compatibility withsolar generator orientation and performance), E.P.will perform better in the comparative ranking.

5. CONCLUSION

Propulsive methods for end of life de-orbiting ofspacecraft passing through LEO regions have beenassessed and classified. Satellites having an orbitabove≅ 615 km altitude have a natural lifetime ofmore than 25 years and need an active de-orbitmaneuver. The transfer to an elliptical orbit with alimited lifetime of 15 to 40 years has emerged as theoptimal solution for satellites where an uncontrolledreentry is admissible. This strategy leads to aminimal ∆V requirement. In cases that a controlledreentry is necessary, a flight path angle between-1.5° and -2.5° has to be achieved at the reentrywindow at 120 km altitude to limit the ground-impact footprint.

The addition of a de-orbit function on a spacecrafthas in most cases a significant effect on satellitedesign due to the fact that the∆V requirements for ade-orbit can rise up to 450 m/s and above. Theimpact on small spacecrafts is thereby even greateras the nominal mission usually does not require anOCS and propellant hardware dry mass has aminimum value. It is found that the spacecraft massof micro- and nanosatellites can double due to theaddition of the de-orbit function. Miniaturizationefforts and ongoing developments in this domaincould improve the situation.

Solid propulsion appears as an attractive solution inmany cases due to its compactness and a relativesmall mass increase. However it has to be mentionedthat adequate solid motors with respect to thrustlevel and burn duration exist solely as prototypesnowadays.

Monopropellant thrusters seem to be a goodalternative in most cases, having the advantage ofbeing readily available. Bipropellant thrusters seemto be competitive only on heavy satellites.

Cold gas systems are tremendously impaired by theirlow Isp. A use for a de-orbit function seemstherefore only advantageous for small satellites andlow orbits with small∆V requirements.

Electric propulsion was not identified as an optimalsolution for any of the chosen reference satellites.This is due to the fact that none of the spacecraftswas equipped with E.P. for nominal mission life andthat available electric power was low. It can beexpected that the outcome is considerably differentfor sufficiently heavy satellites with a sufficientlylarge power/mass ratio and a relative high initialorbit.

This study was intended to highlight advantages anddrawbacks of different propulsion options and togive a general indication on suitability for de-orbittasks. However the selection of a propulsion systemwill have to be made in a case by case study, takinginto account the total∆V requirement of the satellite,both for the nominal mission and the de-orbitfunction, synergetic effects and possibly other (e.g.non-technical) boundary conditions.

6. ACKNOWLEDGEMENTS

The work presented in this paper was performed byOHB System, DLR and HTG under ESA ContractNo. 5316/NL/CK.

7. REFERENCES

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