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Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.

AIAA 97-3058Thruster Options for Microspacecraft:

A Review and Evaluation of Existing Hardwareand Emerging Technologies

Juergen Mueller*

Jet Propulsion LaboratoryCalifornia Institute of Technology

Pasadena, CA 91109

State-of-the-art thruster technologies are reviewed and evaluated in view of potentialmicrospacecraft applications. Microspacecraft are defined in this study as spacecraft withmasses between 1 and 20 kg. Based on this review of existing technologies, future developmentneeds for micropropulsion systems are defined and advanced new micropropulsion conceptsespecially designed with microspacecraft applications in mind are introduced. Of the state-of-the-art technologies, hydrazine thrusters and small solid rocket motors appear applicable t osome microspacecraft. Cold gas systems may provide near-term solutions to attitude control, atthe expense of leakage concerns and large and heavy propellant tankage. New thruster concepts,heavily relying on advanced microfabrication technologies have been designed and built at JPL,addressing some of the microspacecraft design challenges, and are introduced.

I. INTRODUCTION

Background and Significance

Within the National Aeronautics and SpaceAdministration (NASA), a research and developmentinitiative is currently underway to investigate the feasibilityof microspacecraft in the 1-20 kg class1. The motivationbehind this development is the desire to reduce launchmasses in order to reduce mission costs and greatly increaselaunch rates. Launch costs for a typical interplanetarymission may be as high as 30% of the overall mission cost,and these costs may be reduced significantly as a result ofsubstantially reduced spacecraft masses.

In addition, microspacecraft mission scenarios maybe envisioned where, rather than launching a single largespacecraft, the mission is accomplished by a fleet of severalsmaller microspacecraft, with the scientific payloaddistributed among the micro-craft to reduce mission risk.Loss of one microspacecraft would not eliminate the entiremission. A fleet of several small microspacecraft, possiblyin connection with a larger "mother"-spacecraft, could alsoincrease mission flexibility. For example, the smallermicrospacecraft could be placed on different trajectoriesaround the target planet and provide an almost instantaneous,global survey of the target. A "mother-craft" could also

Member of Technical Staff,Technology Group, Member AIAA

Advanced Propulsion

release smaller micro-craft to perform riskier portions of amission. For example, a close-up investigation of Saturn'sring objects may be envisioned2, with a swarm ofmicrospacecraft decending into Saturn's rings while the"mother-craft", providing high-data rate communication toearth via a large high-gain antenna, may cruise in a safedistance from the rings.

Building microspacecraft in the 1-20 kg class,however, will necessitate the miniaturization of everysubsystem in order to maintain the high degree of onboardcapability required to ensure an acceptable scientific returnfor the mission. One of the sub-systems that will beincluded in such a reduction in weight and size ispropulsion. Although in the past many very small spacecrafthave lacked propulsion systems altogether, future micro-spacecraft will likely require significant propulsioncapability in order to provide a high degree ofmaneuverability and capability3. In particular, interplanetarymission scenarios will require propulsion capability onmicrospacecraft for course corrections as well as attitudecontrol to accurately point the spacecraft for observation orcommunication3. Attitude control in low Earth orbit ispossibly achievable via other means, such as magnetictorquers, however propulsive capability is required in higherorbits, interplanetary space or around some other planet,either to directly control the spacecraft's attitude or to off-load momentum wheels3. In addition, very small spacecraftare often launched in a "piggy-back" configuration togetherwith larger spacecraft to save launch costs. Propulsive

Copyright © 1997 by the American Institute of Aeronautics and Astronautics, Inc. The U.S. Government has a royalty-free license toexercise all rights under the copyright claimed herein for governmental purposes. All other rights are reserved by the copyrightowner


capability may be required onboard the microspacecraft toadjust its trajectory according to the desired missionobjective3.

In order to meet microspacecraft propulsionrequirements, the use of lightweight, small sized, low-thrust and small impulse bit (I-bit) systems will be needed.It is the purpose of this paper to review and evaluateexisting propulsion hardware and emerging micropropulsiontechnologies with respect to their applicability tomicrospacecraft. This paper is an extension of an earlierstudy performed by the author at the Jet PropulsionLaboratory (JPL), distributed JPL-internally in late 19954.The JPL study contained company-discreet information andwas therefore not available for public release. In the paperpresented here, company-sensitive information has beeneliminated and new information gained in the emerging fieldof micropropulsion since the earlier study has been added

Definitions

There exists a wide variety of opinions regardingthe appropriate definition of what a microspacecraft is.Table 1 below gives a definition with respect to mass, sizeand power of the type of microspacecraft that are consideredin this study. To simplify this discussion, threemicrospacecraft classes, Class I through III, have beendefined. These microspacecraft classes distinguishthemselves from each other by their mass, power and sizeranges. These should be interpreted as approximate values .

Class I spacecraft, ranging in mass from about 5 -20 kg, are characterized by the fact that they may still beable to use the smallest propulsion hardware available today,or currently under substantial development. This hardwarewill be conventionally integrated by interconnecting it viaconventional feed lines. Development of some newpropulsion hardware, taking miniaturization to new extremesand possibly incorporating advanced microfabricationtechniques, such as MEMS (Micro-electromechanicalS_ystems) technologies, may also be required.

Figure 1 shows an example of a Class I-typespacecraft, designed and built at JPL and referred to as theJPL 2nd Generation Microspacecraft1'5. This craft, in itscurrent design iteration, has a mass of around 7-8 kg. The2nd Generation Microspacecraft is not designed for flight,but, rather, it is an evolutionary functional model of such acraft, with subsystem hardware constantly being upgraded tomore "flight-like" versions5. Based on the 2nd GenerationMicrospacecraft design regime, power densities formicrospacecraft have been estimated at IW/kg, resulting in10 W and 20W onboard power for the two Class I typemicrospacecraft (10 kg and 20 kg) listed in Table 1. Low-and high-mass versions of a Class I microspacecraft will beconsidered.

Microspacecraft with masses between 1 and 5 kghave been categorized here as Class II microspacecraft. In thecase of Class II microspacecraft, development of new,extremely miniaturized propulsion components, both fordelta-v maneuvers and in particular for attitude control,

Table 1: Definition and Classifications of Microspacecraft for the Purpose of the Study

Micro S/CClass

S/C Mass(kg)

S/C Power (W) S/CDimension

(m)

Comments

I 20

10

20

10

0.4 Use conventional components,possibly MEMS. Conventionalintegration (feed lines).

0.3 Same as above.II 0.1 MEMS components, high level of

integration between components of eachsubsystem and possibly between

______subsystems._______________III «1 «1 0.03 All MEMS. Very high level of

integration between all subsystems andwithin sub-systems required. Strongfeasibility issues. Not considered in this

______study._____________ __


Fig. 1: JPL Second Generation Microspacecraft

where multiple clusters of small thrusters are used, will berequired. These devices almost certainly will employ MEMStechnologies in some fashion. Also, because of the severevolume constraints on such a spacecraft, a high level ofintegration will be required between different propulsioncomponents, and between propulsion and other spacecraftsubsystems. For example, in the case of MEMS basedtechnologies, several propulsion components, such asthrusters and valves, plus the required control electronics,may be integrated onto a single chip, or a 3D-stack of chips.Integration approaches of this kind, not limited topropulsion however, are currently being addressed by JPL'sCenter for Integrated Space Microsystems (CISM). Thisrequirement for an increased level of integration, in additionto an even more pronounced degree of miniaturization overthe smallest availbale state-of-the-art propulsion hardwaremakes this Class II category of microspacecraft different inits design requirements from the Class I microspacecraft.

Even smaller microspacecraft with total masses ofsignificantly less than 1 kg, categorized as Class mmicrospacecraft, have recently been studied at JPL6, thesehowever, will not be considered in this study since thedesign concepts discussed in Ref. 5 did not requirepropulsion. If propulsion needs should arise for Class fflspacecraft, they will certainly require MEMS-basedtechnologies6. These propulsion systems would likely bebased on significantly scaled down versions of MEMS-basedClass II systems.

Scope of this Study

The goal of this study is to review currentpropulsion technology in view of its applicability to Class Iand II microspacecraft, identify future technology needs andto outline potential future thruster technology currentlyemerging, aimed at meeting these needs. Microspacecraftmission scenarios may involve a variety of propulsivemaneuvers, such as attitude control, course correction, delta-v maneuvers, orbit insertion or even landing and take-offfrom a distant planet. Depending on the maneuver and delta-v requirement, different propulsion technologies will beneeded . This paper will focus on relatively low-thrustpropulsion systems that can be integrated with amicrospacecraft bus of either Class I or II type for thepurposes of attitude control and delta-v maneuvers.

If take-off and landing operations are considered formicrospacecraft, a class of propulsion devices very differentfrom those to be considered here will be required. Given thelarge delta-v requirements associated with landing and take-off operations, even for relatively small payloads, fairlylarge chemical stages well exceeding the mass limitsconsidered here may result. As a consequence, high-thrust,and thus high-flow, devices will be required. The need tosustain high propellant flow rates will not allow forminiaturization of propulsion components used in theseapplications significantly beyond sizes already availabletoday (although significant research will have to be devotedto such areas as further component mass reduction andalternate propellant usage). These devices are not consideredmicropropulsion devices in the context of this study and aretherefore not included in the following discussion.

Micropropulsion subsystems will not only consistof thrusters, but will also require miniature feed systemcomponents, such as valves, tanks, and pressure regulators,etc. In some of these areas considerable design challengesarise during miniaturization. In particular the moving partsin valves make miniaturization difficult. However, alreadythe thruster material to be reviewed is so vast, thatsurveying miniature components in addition to thrustertechnologies could not be accommodated in this study.Evaluation of miniature components will, thus, not beincluded in this review.


H. REPRESENTATIVE MISSIONREQUIREMENTS

Requirements for microspacecraft missions aredifficult to predict accurately at this early stage of theirdevelopment and will vary greatly given the multitude ofconceivable missions. In this section, an attempt is made topresent a class of representative mission requirements, basedon our current understanding of these small spacecraftrequirements, to serve as a basis for micropropulsiontechnology evaluation. Both attitude control and delta-vrequirements (landing and take-off excluded, see Section I)have been considered. These mission requirements are basedon estimates provided in a recent workshop onmicropropulsion, held at NASA's Jet PropulsionLaboratory (JPL)7. As concrete future microspacecraftmissions develop, this set of mission requirements will nodoubt have to be modified.

Delta-v Requirements

Obviously, delta-v requirements do not depend onthe size of the spacecraft and therefore requirements for fourrepresentative small spacecraft missions, currently underinvestigation at JPL, have been listed below in Tables 2through 5 to serve as a reference. These missions include asmall body (asteroid) rendezvous, an outer planet (Europa)orbiter, a spacecraft formation flight (DS-3) and an earth-observing cluster.

Inspecting Tables 2 and 3, the large delta-vrequirements for deep space missions becomes immediatelyapparent. Electric propulsion applications result in evenlarger delta-v-requirements due to burn losses8, which mustbe offset by the higher specific impulses and more fuel-efficient operation of electric engines. Use of electricpropulsion may either lead to shorter trip times, or reducedspacecraft masses, or both. The benefit of using electricpropulsion in regard to spacecraft mass reduction will likelybe even more important for mass limited microspacecraftmissions.

An additional requirement for chemical primarypropulsion is the need to maintain large enough thrust-to-spacecraft weight ratios. Values around 0.1 - 0.3 are typical.Too small a thrust-to-weight ratio will again lead to burnlosses8 and increase the required delta-v. Since in the case ofchemical engines this increased delta-v requirement cannot beeasily offset by a sufficiently large specific impulse, too lowa thrust for chemical primary propulsion maneuvers mustbe avoided. Too large a thrust value, on the other hand, maygenerate accelerations too large to be tolerated by thespacecraft structure, in particular at times well into the

mission, when portions of the spacecraft structures may bedeployed.

Attitude Control Requirements

In order to estimate attitude control requirementsthe following assumptions were made7: (1) fine pointingrequirements are assumed, defined by the desire to staywithin a 0.2 - 2 mrad deadband and ACS firings occuring nomore frequently than one couple firing every 20 - 60 sec; (2)Slew rates of 180°/minute required with one couple ofthrusters firing. The spacecraft was assumed to be cubical inshape with the side of the cube being equal in length to thedimension listed in the forth column of Table 1. Theresulting minimum impulse bit and minimum thrustrequirements are listed in Table 6 for microspacecraft massesof 1, 10 and 20 kg. Very small impulse bit requirements canbe noted. It should be pointed out in this context, that thefine pointing requirements given above are not extreme fortoday's spacecraft.

HI. REVIEW OF THRUSTER TECHNOLOGIES

In this section, state-of-the-art thruster hardware,either available "off-the-shelf or under significantdevelopment, will be reviewed and evaluated in view ofmicrospacecraft applications. This section is structured intotwo main parts , focusing on primary and attitude controlapplications, respectively. Both primary and attitude controlsections have been further sub-divided into chemical andelectric thruster sections.

Primary Thrusters - Chemical

Bi-Propellant Engines

Bi-propellant engines are most commonly used forprimary propulsion applications of conventional spacecrafttoday due to its relatively high specific impulseperformances and considerable flight heritage. Advantages ofbi-propellant engines over other chemical systems, such asmono-propellant thrusters, are their higher specific impulseperformance, leading to lower fuel weights. Disadvantagesare their relative complexity, both with respect to enginetechnology and the feed system. Since separate feed systemsfor fuel and oxidizer, as well as pressurants are required,component part count is high, leading to large propulsionsystem dry masses. Therefore, bi-propellnat systems areusually used on mission requiring large delta-v's (> 1,000m/s) and large spacecraft.


Table 2: Delta-v Requirements for a Vesta Rendezvous Mission7

Delta-v Requirements (m/s)Primary _____________ACS'"

Mission Duration (yrs)

Chemical3400*

Electric

7000"2.5-252.5-25

Chemical4.3

Electrical

2.7Assumes Mars Gravity Assist (MGA) trajectoryAssumes direct trajectoryAssumes 250 slews (180°/min ) and 3-axis stabilization in the range of 1 - 20 kg S/C mass, thrust 0.5 - 10 mN.

Table 3: Delta-v Requirements for a Europa Orbiter Mission7

Delta-v Requirements (m/s)Primary ACS'"

Chemical2500'945

Electric

5500"5-505-50


Chemical4.8

Electrical

5.8

Assumes Solar Electric Venus Venus Gravity Assist (Se-VVG) and 345 ms/s chemical delta-v required for Jupiter Orbit Insertion(JOI) and 600 m/s for Europa Orbit Insertion (EOI).Assume 500 180° /min slews and 3-axis stabilization (S/C masses ranging between 1 - 20 kg and thrusts between 0.5 - 10 mN).ACS requirements may be higher for low thrust trajectories.

Table 4: Delta-v Requirements for the Deep Space Interferometer Mission (DS-3)7

Delta-v Requirements (m/s)Primary________________ACS


100-300 0.5 - 1.0

Table 5: Delta-v requirements for an Earth Observing Cluster7

Delta-v Requirements (m/s)Primary'_______________ACS"


Chemical500

Electric

5505-505-50

Chemical5.0

Electrical

5.0Primary delta-v assumes 200 m/s for non-Keplerian orbit, 250 m/s for NSSK, and 50 - 100 m/s for phasing.ACS assumes 500 slews of 180° /min and 3-axis stabilization (S/C masses ranging between 1 - 20 kg and thrusts rangingbetween 0.5-10 mN).

Table 6: Representative Attitude Control Requirements for Microspacecraft7

S/C Mass (kg)20101

S/C Dimension* (m)0.40.30.1

Ibit (mNs)0.0130.0050.0002

Tmln (mN)4.651.750.06

* Assume cubical spacecraft shape


Consequently, most bi-propellant engines that havebeen built today provide fairly large thrust levels. Somesmaller engines in the 5 - 22 N (1 -5 Ibf) thrust range havebeen built or are under development. Up to this point,applications for these engines have been envisioned in theuse of attitude control purposes of larger spacecraft in orderto simplify the overall propulsion sytem, eliminatingseparate attitude control propellant tanks. Given these goals,considerable effort was devoted to fast thruster responsetimes and short impulse bits9"13.

Challenges encountered when building bi-propellantengines of such a small size, include combustion efficiencylosses due to the potential of reduced mixing andvaporization, thermal control issues of chambers, nozzlethroats and injector heads and related material issues, injectordesign issues and related accurate mixture ratio controlissues, and possibly, spacecraft contamination issues due thepotential of incomplete mixing and vaporization inside thethrust chamber.

Mixing and vaporization losses can occur in smallengines due to the small chamber size. In general, goodvaporization is obtained in longer chambers, at higherchamber pressures and for smaller injector orifice sizes14,while better mixing is achieved in engines having highchamber-length-to~diameter ratios and a larger number ofinjector inlets14. In addition, engine size also plays a role inthe mixing of the propellants14. Smaller engines have lowerchamber flow Reynolds numbers and thus lead to lessturbulent chambers, reducing mixing. The limitationsimposed on chamber length and diameter has an immediateimpact on the degree of miniaturization of a bi-propellantengine.

Thermal control of small bi-propellant engines isanother key design issue. Film cooling, or boundary layercooling (BLC), is often employed in bi-propellant enginesto keep the chamber wall temperature within its thermal andstructural design limits. Here, a fuel is injected close to thechamber wall. Since the propellant mixture is fuel rich, itdoes not burn completely and will shield the chamber wallsfrom the heat output of the combustion reactions occurringcloser to the center of the chamber. However, at the sametime combustion efficiencies are reduced due to incompletecombustion. While for more conventionally sized enginesabout 15 - 30 % of the fuel is commonly used for filmcooling, these values may reach up to 30 - 40% for smallerengines in the 22-N class, causing performance losses12 andpossibly resulting in spacecraft contamination concerns dueto the possibility that liquid fuel droplets may attachthemselves to sensitive surfaces (optical lenses, solar cells,etc.).

Elimination of film cooling was done in the smallbi-propellant attitude control engines developed by Rockwellfor the Kinetic Energy Anti-Satellite (KE ASAT)program.10'11 This increases combustion efficiency anddecreases injector head complexity since no separate BLCholes are required, and should facilitate miniaturization.However, in order to survive the punishing thermalenvironment, high temperature chamber materials have to beemployed. In the case of the KE ASAT technology10-11, acarbon-silicon carbide chamber was used. Despite use of thishigh temperature material, engine burn durations in a singleburn where limited to only a little over 20 seconds. Anotherchamber material under significant investigation is rhenium-iridium composite material. Rhenium is uses as thesubstrate material because of its high melting point (3453K)12 and coated with a iridium layer for chemical inertness.Iridium has a coefficient of thermal expansion (CTE) closelymatching that of rhenium and a high melting point of 2727K12. Using this chamber material, specific impulses inexcess of 300 sec have been obtained in a 22 N thrustchamber over burn durations of a maximum of 350seconds12.

Injector design also requires careful attention insmall bi-propellant rocket engine development. Due to thesmall flow cross sections encountered in small engines, flowrate control, and thus mixture rate control15, as well asmisalignment of impinging propellant jets12 could lead topoor engine performance repeatability or engine reliabilityproblems. In addition, thermal management of the injectorhead is important to ensure that heat diffusion from the hotchamber material to the injector head is minimized in orderto prevent vaporization of propellants in the injector and notto exceed thermal limits of the injector material, which maybe different from the high-temperature chamber material formachining reasons. Unlike-doublet injector types arefavoured14' 16 because of better mixing results and reducedheat load to the injector head by displacing the flame frontaway from the injector wall surfaces16. As mentioned above,more injector elements will lead to better mixing, however,limited engine size may limit the number of injectorelements. In the case of the Rockwell engine discussedabove, only a single unlike-doublet injector element isused10'11 (combustion efficiencies are maintained at highlevels due to the aforementioned elimination of the BLClayer).

Table 7 lists the smallest bi-propellant enginetechnology available today. Caution has to be exercisedwhen referring to the data presented in this table. Not allengines listed are space-qualified at this point and some maystill be experiencing problems in their respectivedevelopment programs. Schwende et al.13 point out in their1993 paper that the 4 N engine experienced "anomalies" due


Table 7: Comparison of Small Bi-Propellant Engines

Thrust(N)

4

4.45

10

10

10

22

22

2222

22

30

156

Manu-facturer

DASA

Marquardt

DASA

Marquardt

RoyalOrdnanceMarquardt

AtlanticResearchAerojetAerojet

RoyalOrdnanceRockwell

Marquardt

Type Fuel/ Isp (s) WeighOxidizer t (kg)

MMH/ 285 0.27MON-1

R-2/ MMH/NTO 280 0.43R-2B

MMH/ 290 0.3MON-1

R-52 MMH/NTO 295

LIT MMH/NTO

R-6C/ MMH or 289 0.67R-6D N2H4/NTO

A0809 MMH/NTO 290 0.55

SSD MMH/NTO 280 0.59MMH/NTO 313

Leros N2H4/MON 285 0.8520H

MMH/NTO 287 0.1

Divert N2H4/NTO - 0.1

Size Comments( length*Max.Dia)

(cm)Reportedtemperatureanomalies

26.1 x <9 No knownapplicationsRegenerativecooled throatin previousversion. 34flight unitsbuilt.No knownapplications

-

25x<13 FlightApplicationsfor R-6C, noknownapplication R-6D

21.7x5.4 FlightApplications

18.5x6.9 -Rh/Ir chamber.Indevelopment

20.9x6.6 UnderdevelopmentMax. 26 sec insingle burn.Max.accumulativeburn: 77 sec.1.25 mixtureratio.Developed forBMDO20 sec singleburndemonstratedDeveloped forLEAP

Ref

13

17

13

17

18

16

19

2012

18

10,11

9

to too high chamber wall temperatures. Also, although single duration burns which is too short for interplanetarycommonly referred to as examples for the degree of delta-v maneuvers. As mentioned above, the KE ASATminiaturization achieved for bi-propellant engines, the KE developments, as well as others, have focused on attitudeASAT engines have been tested only up to 26 seconds in control applications, rather than primary propulsion


applications. As a result, pulsing performances rather thanlong duration burns were emphasized and led to thecurrently exhibited design performances. Almost all enginesuse nitrogen tetroxide (NTO) and monomethylhydrazine(MMH) as oxidizer and fuels, respectively, due to storagereasons, acceptable performance values and relatively benignmixture ratio sensitivities. Using a O/F mixture ratio of 1.6results in equal propellant volumes for fuel and oxidizer, sothat identical tanks can be used (reducing development costand time) and the spacecraft will experience no center-of-gravity (e.g.) shifts during burns.

Given the engine data presented in Table 7, andusing the representative mission requirements of Section n,propellant mass fractions of 0.7 and 0.58 can be computedfor a spacecraft with delta-v requirements of 3400 m/s and2500 m/s, respectively, at a specific impulse of 290 s (seeTable 7). Although these values are not atypical for largerspacecraft, for a microspacecraft they may be too high giventhat component-mass-to-spacecraft-mass ratios are larger.For example, in the 3400 m/s case for a 20 kg spacecraft,merely 6 kg of available mass remains. This will have toinclude the entire dry mass of the propulsion system,structure plus all other subsystems. According to Table 7,one bi-propellant engine alone may take up about 5% of thatmass, even using the lightest engines available.

Smaller delta-v requirements around 1000 m/s, onthe other hand, would result in propellant mass fractions of0.3 for a 290 s bi-propellant system. However, a hydrazinemono-propellant system with a specific impulse of 220 s(see below) would result in a propellant mass fraction of0.37. In the case of a 10 or 20 kg Class I spacecraft, thisdifference would be a mere 0.7 or 1.4 kg in propellant mass,respectively. Given the lower component part count of amono-propellant feed system, this higher propellant fractioncan easily be offset by a simpler mono-propellant system.

It is, therefore, concluded that bi-propellant systemsare not suitable for either high or low delta-v requirementsonboard a microspacecraft due to too high dry weight of thesystem. Further, aggressive miniaturization may help,however, there exists considerable doubt that significantlysmaller, yet reliable and space-qualifyable bi-propellantengine technology can be developed in view of the designchallenges for small bi-propellent engines given above.

However, separate chemical stages for large delta-vmaneuvers may make use of bi-propellant engines. Anexample of such a stage is given in Ref. 21, describing aHydrazine (N2H4)/ Chlorine Pentafluoride (C1F5) chemicalupper stage, developed for the Lightweight Exo-Atmospheric Projectile (LEAP) program. Thrust levelsprovided by the LEAP stage are somewhat high (2056 N) for

microspacecraft applications and there are concerns regardingthe corrosivity and toxicity of C1F5. Stages like these, orsimilar ones using more conventional propellants, however,may be required for orbit insertion maneuvers around distantplanetary bodies, in particular when these bodies are lackingan atmosphere (no aerobraking possible) or when they arelocated too far from the sun (solar power levels too low touse solar electric propulsion). In addition, landing and take-off operations will require bi-propellant technology.However, those mission applications may require thrustlevels well exceeding those obtainable with the engineslisted in Table 7 due to high stage masses and large requiredvehicle accelrations to overcome the gravity of the respectiveplanetary body.

Mono-Propellant Hydrazine Engines

Hydrazine mono-propellant thrusters combineengine technology substantially simpler than that of bi-propellant engines, relatively simple and low part-count feedsystems, and high reliability with intermediate performance(specific impulses around 220 s for state-of-the art hydrazinethruster technology). In a hydrazine thruster, the propellantis passed through a catalyst bed and decomposed. Thedecomposition products are nitrogen, hydrogen andammonia. The reaction takes place in two stages: hydrazinedecomposes first through an exothermic reaction intoammonia and nitrogen. The ammonia then decomposesfurther through an endothermic reaction into hydrogen andnitrogen, however, leaving the overall reaction exothermic.The degree of ammonia decomposition depends on manyfactors, among them feed pressure, catalyst type andgeometry. Shell 405 is the standard catalyst used in the US,consisting of 1.5 - 3 mm dia. alumina pellets coated withiridium. The catalyst pellets are contained within a meshconstruction in a so called catalyst bed. Upon contact withthe iridium surfaces, the hydrazine decomposition reaction isinitiated.

Hydrazine thrusters have been used extensively onconventional spacecraft for attitude control as well asprimary propulsion sources for intermediate to low delta-vrequirements (about 1000 m/s or less). Of interest here arethe smallest available hydrazine thrusters, in the 0.9 - 4.45N range. These engines are being manufactured in the US by- Primex (formerly Olin Aerospace/Rocket Research),Kaiser-Marquardt and TRW companies, and abroad byDaimler Benz Aerospace in Germany. Typical enginecharacteristics are listed in Table 8.

The engine sizes, weights and thrust levels shouldallow for relatively easy integration on a Class I


Table 8: State-of-the-Art US Hydrazine Thrusters

Thrust Manufacturer Type Isp(N) (s)

Weight Size(kg) (Lengthx

Max. Dia)_________(cm)

Comments Ref .

0.9 Primex MR- 210- 0.33103 220

14.8x3.4 C/E, D and G models.Isp and thrust feedpressure dependent (340- <100 psia).

________Considerable flight use.

22

0.9 Marquardt KMHS 226 0.3310

14.6x3.2 Isp and thrust feedpressure dependent.Flight use________

17

2.2 Primex MR- 213- 0.3311 IE 224

16.9x3.8 Isp and thrust feedpressure dependent (370- 60 psia). Considerable

________flight use._______

22

4.45 Primex MR- 226- 0.33111C 229

16.9x3.8 Isp and thrust feedpressure dependent(400-80 psia).

________Considerable flight use.

22

4.45 Marquardt KMHS 230 0.3817

20.3x3.2 Isp and thrust feedpressure dependent.Flight use._______

17

TRW MRE-1 220 0.8215.2xN/A Mass is for dual thrustermodule. Isp and thrustfeed pressure dependent.Considerable flight use.

23

18 TRW MRE-4 230 0.41 20.3xunknown

Isp and thrust feedpressure dependent.Considerable flight use.

23

microspacecraft bus, mounted along the axis of the bus forprimary propulsion applications. All thrusters listed haveseen considerable flight use and potentailly would requireonly minimal re-development for use as Class I mainengines. Class II microspacecraft are too small to takeadvantage of these existing technologies.

One area of improvement in the use of state-of-the-art hydrazine thrusters as Class I main engines may be foundin the valve area. Currently, a considerable weight fractionof a small hydrazine thruster is taken up by the thrustervalve (greater than 50% for the small engines consideredhere). This fact may open an opportunity for further weightreductions. Since the smallest hydrazine thrusters have beenused mainly for attitude control purposes where fast valveaction is essential ( on the order of 15 ms on/off), thesevalves could possibly be replaced by slower valves, sinceprimary propulsion applications of these thrusters, asenvisioned here for microspacecraft, seldom would require

engine pulses that short. Slower valves, depending ondesign, may require less electromagnetic force action to openthe valve, which might reduce electromagnet masses.

A disadvantage of hydrazine propellant is itstoxicity and flammability and resulting ground handling andpropellant loading procedures. These procedures areobviously well established due to extensive hydrazinethruster use on conventional spacecraft, but willsignificantly contribute to the cost of small spacecraft. Inaddition, as was pointed out in the preceding section, ahydrazine propulsion system onboard a microspacecraft isonly practical if small or intermediate delta-v maneuvers arerequired (i.e. < 1000 m/s). In these cases, mono-propellantsystems will have an advantage over bi-propellant systemsdue to reduced system complexity, smaller component partcount and, thus, smaller volume requirements. If higherdelta-v's are required (see Section II), mono-propellantsystems become increasingly heavy due to large propellant


requirements. In these cases, bi-propellant engines, likely tobe mounted on a separate kick-stage, or electric thrusteroptions (see below) should be preferred. Thus, limited ClassI microspacecraft use of existing hydrazine thrustertechnology for primary propulsion applications appearsreasonable.

HAN-based Mono-Propellant Thrusters

Recently, so called HAN/TEAN thrusters4'24 havereceived attention. This thruster is of the mono-propellanttype, using mixtures of HAN (Hydroxylammonium Nitrate -NH3OH+NO3.), TEAN (Triethanolammonium Nitrate -(HOCH2)3HNOH+NO3.) and water as a propellant.HAN/DEHAN mixtures have also been studied, consistingof HAN, water and DEHAN (DiethylhydroxylammoniumNitrate - (CH3CH2)HNOH+NO3.)24. HAN is an oxygen richcomponent and TEAN or DEHAN are fuel rich components.Due to the water additive, both components can coexist in amixture without detonation, as long as the water content ismaintained at 10% or above25. Exposing the mixture to acatalyst causes a chemical reaction and exothermicdecomposition of the components into CO2, N2 and H2O4>24.

Mixtures of these propellants have been studied foruse as liquid gun propellants by the US Army and have beencategorized according to their composition. LP1846, forexample, consists of 60.8% HAN, 19.2% TEAN and 20%water, while LP1845 consists of 63.2% HAN, 20% TEANand 16.8% water24, i.e. has a lower water content thanLP1846. The amount of water in the mixture greatlyinfluences the decomposition temperature and, thus, theavailable specific impulse. Increasing the water content willlower the flame temperature. Jankowski24 quotes flametemperatures, based on numerical calculations, of 2022 K forLP1846 and 2125 K for LP1845, having the lower watercontent. These values result into theoretical specificimpulses, assuming a specific impulse efficiency of 92%, of233 and 239 sec24. Tests performed with HAN-basedpropellants with different, not specified additives haveresulted in specific impulses of 270 sec at a flametemperature of 2500 K24.

The flame temperatures of HAN/TEANcombinations are very high and approach values found insmall bi-propellant engines. Thus, many of the thermaldesign challenges found in the construction of small bi-propellant chambers would have to be overcome when usinghigh-performing HAN/TEAN mixtures with low watercontent. Engine lifetime restrictions may thus result if highperformance is required.

However, as mentioned above, flame temperaturesmay be lowered at the expense of specific impulseperformance if the water content is raised. Even though notproviding a significant performance advantage over existinghydrazine mono-propellant thrusters in those cases anymore,HAN/TEAN thrusters still offer advantages due to the lowtoxicity of both the NAN/TEAN propellant as well as itsreaction products, high storage densities (about 40% higherdensities than that of hydrazine) and lower environmentaltemperature handling capabilities. While hydrazine freezes atabout 0 C, HAN/TEAN mixtures may be used attemperatures as low as about -33 C24. At this point, theviscosity of HAN/TEAN mixtures increases and propellantfeeding will no longer be possible using conventional feedsystem technologies24.

Higher storage densities and lower environmentaltemperature handling capabilities of HAN/TEAN propellantsare beneficial for microspacecraft, since they allow forsmaller and lighter storage tanks and the elimination of, orreduction in power for, tank and line heaters, reducingoverall power requirements for the spacecraft. Thus,HAN/TEAN thrusters may find Class I microspacecraftapplications for those reasons. Reduction in engine size tomeet Class n requirements may not be possible due to thehigh heat loads to be expected in a HAN/TEANdecomposition chamber. Considerable development workwill be required to bring current HAN/TEAN thrusterconcepts to flight status, with reaction chamber thermaldesign issues being one of the most challenging steps in thedevelopment.

Other Mono-Propellant Thrusters

Hydrogen peroxide (H2O2) thrusters are consideredfrom time to time as an alternative to more conventionalmono-propellant systems25'26. Hydrogen peroxide, whensubjected to a suitable catalyst, decomposes into water andoxygen in an exothermic reaction. Although hydrogenperoxide has been used in flight applications as a mono-propellant in the past8'26'29, it is no longer in use due topropellant storability issues8'26"29. Hydrogen peroxide slowlydecomposes when heated or exposed to a catalyst. Almostany organic substance can serve as such a catalyst29. If slowdecomposition occurs in propellant tanks, as has beenobserved in the past8, tank pressure increases result overtime and propellant is lost due to its slow conversion intoits reaction products inside the propellant tank.

10


Solid Rocket Motors

Solid rocket motors are frequently used in kick-stages for orbit raising or orbit insertion of spacecraft,beginning with the Explorer 1 spacecraft8 and leading to themore recent Pioneer-Venus30, Magellan8'30 and Galileomissions, as well as numerous commercial missions (orbitraising). In solid motors, fuel (typically aluminum powder),oxidizer (typically ammonium perchlorate - NH4C1O4) andan organic binder (typically Hydroxyl-terminatedPolybutadiene - HTPB) are combined into a composite toform the solid propellant8'30. The advantages of solid rocketmotors are their compact size combined with a relativelyhigh specific impulse performance - less than that of bi-propellant systems but higher than that of mono-propellantsystems29. For obvious reasons, solid motors also do notsuffer from propellant leakage concerns. Propellantsublimation by exposure to space through an open nozzlewas a concern for the use of solid motors for deep-spaceapplications, but has been found to have no impact onmotor performance after 10 - 15 months of in-spacestorage8-30. In the case of the Magellan mission, a ThiokolSTAR 48B motor was fired for Venus orbit insertion after462 days in space30. A Thiokol STAR 24 motor was firedafter 6.5 months in space for the Venus orbit insertion ofthe Pioneer probe.30

Disadvantages of solid motors are that they aregenerally not restartable, and therefore do not allow for orbittrimming. If several delta-v burns are required it is necessaryto stack multiple stages leading to system complexities andhigher propulsion system dry masses. The issue of orbittrimming is of particular importance for solid motors sinceexact prediction of delivered total impulse is difficult toestimate due to uncertainties in the expected graintemperature, the exact propellant composition, and theamount of inert material consumed8. Thus, a separate smallliquid system may have to provide for the orbit trimmingmaneuvers30. A separate liquid system may also be requiredfor despin of the satellite. Solid motors for spaceapplications are usually not equipped with thrust vectoringcapability. Although some larger motors have been testedwith such a capability32, these nozzle gimbal systems maybe too heavy and complex for very small motors, such asthose required for microspacecraft applications. Thus,spacecraft generally are spin-stabilized before motor firings,and despin may be required after separation from the stagedepending on the mission,.

Table 9 shows some of the smallest solid motorsavailable today30"33. In addition to the companies listed inTable 9, Pacific Scientific, Inc. is also building small rocketengines for missile divert purposes34. As can be seen byinspecting Table 9, envelopes and masses of the smallest

available motors fit within the Class I category ofmicrospacecraft and specific impulse performances are quitegood. However, thrust levels are much higher than desiredand burn times are relatively short, which would lead to verylarge microspacecraft accelerations. For example, assuminga 20 kg overall spacecraft weight (incl. motor), using aThiokol STAR 6B motor results in accelerations at thebeginning of the burn of about 13 g's and at the end of theburn around 18 g's. The delta-v that can be achieved withthis motor for a 20 kg spacecraft would be 963 m/s. Asimilar calculation for a 10 kg spacecraft equipped with theSTAR 5A motor would lead to an initial acceleration of 1.7g's and an acceleration just prior to burn out of about 2.2g's, and a delta-v of 641 m/s. Achievable delta-v's arelimited by the reduced propellant mass fractions foundtypical for smaller motors.

The high thrust forces and short burn times are aresult of the intended design applications for most of thesesmall motors, i.e. stage separation or use as missile divertengines (missile attitude control). In both cases it isessential to provide a relatively large thrust in a shortamount of time. In the case of microspacecraft applications,this could lead to limitations of solid motor use due to therequirement of being able to fire only in a stowed vehicleconfiguration (no deployments) and possibly costly re-qualification of spacecraft components to account for thesehigh accelerations.

An exception to the fast burning, high-thrust smallsolid motors shown in Table 9 is the STAR 5A. Eventhough accelerations in the example given above are stillquite high, those values may be much more tolerable. Thelonger burn time and smaller thrust of this motor wereachieved by using an end burner propellant grain. This graintype was extensively used in the past in the so called JATO8

(Jet Assisted take-Off) units used in the 1940s to assist inthe take-off of aircraft from short runways or assistingaircraft requiring additional thrust for heavy-lift capability.This grain type may be the grain of choice formicrospacecraft applications. Even longer, lower-thrustburns could be accomplished if the length-to-diameter ratioof the motor case could be increased.

Solid rocket motors may thus present an interestingalternative to more complex liquid systems where missionprofiles are simple, require only single burns andintermediate delta-v values (< 1000 m/s), a small liquidsystem is onboard the microspacecraft for orbit trimming orif the required accuracy of the actually delivered delta-v is nottoo high. For example, in the Saturn-ring explorer scenariodiscussed in Section I2, quite sizable delta-v changes may beachieved for the microspacecraft probes using small solidmotors. Since multiple probes are being used anyway to

11


Table 9: State-of-the-Art Small Solid Rocket Motors

Manufacturer

Thiokol32

Thiokol32

Thiokol"Thiokol32

AtlanticResearch33

AtlanticResearch33

AtlanticResearch33

Type

STAR 5A*"STAR 5C

STAR 5CBSTAR6B

-

-

-

LoadedWeight

(kg)4.74.54.510.30.4

0.5

1.6

PropellantWeight

(kg)2.32.12.16.1

-

-

-

Size(LxD)(cm)

22.5x1334x1234x12

40x18.6-

-

-

BurnTime'

322.8

2.675.9

1

1

-

Isp(s)

250266270273

-

-

-

Thrust"(N)

169195320412513222

311

952

Burn time, 10% thrust at ignition, 90% thrust at shut-downBurn time averagedEnd Burner

account for the potential loss of some, total impulse anddelivered delta-v uncertainties for individual probes may beacceptable.

The absence of leakage concerns and the ability tocompactly package solid motors will be attractive formicrospacecraft applications Existing motor hardwareappears to fit the envelope of Class I spacecraft, althoughlonger burn times, lower thrust values and, thus, lowervehicle accelerations should be aimed for. The benefits ofusing small solid rocket motors would be even morepronounced for smaller microspacecraft, such as types fallinginto the Class II-category. Here, compactness plays an evengreater role than for Class I craft. Class II application ofsolid motors would require further miniaturization of solidmotor technology and dedicated full development programsto achieve the desired reductions in size, weight and thrust.

Hybrid Rocket Motors

In a hybrid rocket motor a solid fuel is combinedwith a liquid or gaseous oxidizer, which is stored in aseparate propellant tank and fed into the motor case27'35'36. Asa result of this separation between solid fuel and liquidoxidizer, hybrid rockets exhibit some interesting properties.Hybrid rockets are restartable, relatively safe when comparedwith solid motors, offer appreciable specific impulseperformances up to around 300 s when using storablepropellants, and still offer a higher degree of compactnessthan bi-propellant systems. Hybrid rockets, at first glance,may thus seem to be an attractive cross between highperforming bi-propellant engines and compact solid motortechnology.

One of the disadvantages of hybrid motors, inparticular when viewed in terms of space applications(microspacecraft or otherwise) with long mission durations,is the limited choice of suitable storable propellantcombinations available today. Typically, Hydroxyl-terminated Polybutadiene (HTPB) is being used as fuel andliquid oxygen (LO2) or hydrogen peroxide (H2O2) are used asoxidizers27'28'35"37. Although both oxidizers may be suitablefor launch applications, they are not storable over longperiods of time because they are either cryogenic (LO2) ormay slowly decompose over time in the propellant tankbefore use (H2O2 - see above). Among storable oxidizeroptions, nitrogen tetroxide (NTO) was used, but required aseparate ignition source27. Chlorine-fluorides, such as C1F3and C1F5, have also been used as oxidants. These substances,however, are highly toxic and corrosive.

Work on hybrid rocket motors started in the 1930'sto 40's in both Germany and the US, with some of the earlywork in Germany performed by Hermann Oberth35-36.Development has continued on an on-and-off basis over theyears. Focus was placed mainly on launch applications,leading to the development of the H-500 (312,000 N thrust)and the H-250F (1,000,000 N thrust) engines developed bythe recently failed AMROC company35. Both of the lattertwo motors used HTPB and LO2 as propellants. Extensiveresearch on hybrid rockets has also been performed at variousuniversity research laboratories around the world28. Thiswork was performed on smaller test devices. The smallestquoted thrust value for a hybrid rocket engine is found inSellers et al.28 at ION.

Given that attention in hybrid rocket enginedevelopment was focused mostly on large launch motors,

12


where oxidizer storage issues play a lesser role than for in-space applications, it is uncertain whether hybrid rocketsmay find space applications for interplanetary missions . Inparticular microspacecraft applications, requiring smallmotor technology comparable in size to solid motortechnology given in Table 9, remain uncertain. The use ofhybrid engines for these applications would require a fulldevelopment program, starting at the point of propellantselection. If successful, hybrid engine technology could fill auseful gap between high performing, yet complex bi-propellant engines and compact and simple, yet relativelyinflexible solid motor technology.

Primary Propulsion - Electric

Ion Engines

In an ion engine, the propellant (typically xenon) isionized in a plasma discharge. Ions are extracted from theplasma via electrostatic forces and accelerated across anelectric potential difference of about IkV. In the process,xenon ions achieve a velocity of about 30,000 m/s,corresponding to a specific impulse of about 3,000 sec.

An ion propulsion subsystem consists of severalcomponents that all will have to be miniaturized formicrospacecraft applications. These are the thruster itself,the power conditioning unit providing the required voltagesto the engine, and the feed system. Within the thrusterassembly, critical components include the cathode forcertain engine types, the accelerator grid system and theneutralizes ( used to neutralize the ion beam to avoidcharging the spacecraft). Different types of ion engines arebeing developed . DC electron bombardment types use anelectron current emitted from a hollow cathode inside theengine body to ionize the propellant gas by causingcollisions between the electrons and the propellant gasatoms. RF (radio-frequency) electron bombardment engines

Table 10: State-of-the-Art Small Ion Engines

use electrons accelerated in an inductive coupled RF field tocause propellant ionization.

Advantages of ion engines are their large specificimpulses which translate into significant propellant andspacecraft mass reductions. This fact is of particularimportance for mass constrained microspacecraft, especiallyfor interplanetary missions which have large delta-vrequirements. Using ion engine technology may lead tolighter overall spacecraft masses and shorter mission triptimes when compared with chemical bi-propellant systems.In addition, xenon propellant, when stored at about 2,000psia pressure, takes on a supercritical state with a densityabout twice that of water. Reduced propellant requirementsdue to the high specific impulse of the engine and highdensity will allow for compact propellant storage.

On the other hand, propellant mass reductions dueto higher specific impulses will have to be traded off withelectric power requirements. Power requirements drive powerconditioning unit and power supply masses. In order toreduce overall system wet masses, an optimum operatingpoint must be selected for the engine, allowing for bothsignificant propellant savings and low power systemmasses. For typical interplanetary mission requirements andcurrent power system technology, this optimum is typicallyfound around 3,000 s specific impulse.

Table 10 lists the smallest available ion enginetechnology today38"41. As can be seen, all current ion enginesystems are too large for use on a microspacecraft, both withrespect to mass (incl. PPU) and power requirements. Theseengines may, however, be used on separate electrical stagesjust like the bi-propellant thruster technology discussedabove, and provide large delta-v changes for Class I - typemicrospacecraft in that configuration. In the case of electricstages, a dedicated power supply will have to be provided.

ManufacturerDischarge TypeThrust (mN)Isp (s)Power (W)Thruster Mass (kg)PPU Mass (kg)

Beam Diameter (cm)

Hughes40

DC17.825854395.06.8

13

DASA38'39

RF5-153000240 (5 mN) - 600 (15 mN)1.68.0 (PPU)1.3 (RF generator)10

JPL41

DC21-312500 - 3900500-9002.5-

15

13


In order to integrate ion engine technology onboarda microspacecraft bus within the mass margins provided inSection II, new technologies will have to bedeveloped.Thruster sizes will have to be reduced. Challengesto be overcome here will be to maintain plasmas in smalldischarge chambers, where electron wall losses may be high.Cathodes and neutralizers will have to be miniaturized andcold cathode technology may be explored. Micromachinedgrid systems, allowing for electrostatic beam steering,eliminating heavy engine gimbals (see Section V) andminiaturized power processing units will be needed.Miniature ion engine technology may possibly be based onhollow cathode technology, currently used for conventionalengine cathodes and neutralizers. Thus, miniature ionengines have to be considered as very advancedmicropropulsion concepts that still have to overcome manyfeasibility concerns before they can be seriously consideredfor microspacecraft applications.

Hall Thrusters

Hall thrusters42'43 are electrostatic propulsiondevices which use xenon propellant. Plasma generationand ion beam accelration are different from those found inion engines and lead to a more compact thruster technology.In a Hall thruster42'43, electrons emitted from a hollowcathode external to the thruster are accelerated towards apositive anode located upstream and inside an annulardischarge chamber. On their way to the anode, the electronscross a radial magnetic field extending across the annularchamber. Due to Lorentz-force action, the electrons gyratearound the magnetic field lines, and drift azimuthallythrough the annular channel, colliding with propellant gasatoms (xenon) and ionizing them. The ions are acceleratedaway from the engine by the same electric field thatattracted the electrons. The ion beam is neutralized byadditional electrons streaming off the cathode.

Due to the high electron density in the magneticfield region, a dense ion beam can be formed, overcommingspace, charge limitation effects found in ion engines. Hallthrusters are thus more compact for the same delivered thrustlevel than ion engines. On the other hand, Hall thrustertypically deliver specific impulses around 1500 - 2000 s,making them more suitable for near-earth missions (orbittransfer, repositioning, etc.), rather than interplanetaryflights. A high-Isp Hall thruster would be an attractivealternative to ion engines.

Current Hall engine technology is far too heavy andpower consuming to be used within the microspacecraftdesign envelope. However, efforts are underway to

miniaturize the technology for small satellite applications43.Design goals are a power level of 50 W at a thrust of 5 mNand a specific impulse of 1600 s. The device is estimated tobe about 4 mm in diameter. Given the smaller channeldimensions, larger magnetic fields have to be provided toachieve smaller electron gyration radii. Required estimatedfield strengths are 0.5 T. For even smaller devices, magneticfield strengths would have to be increased further. Currently,samarium-cobalt permanent magnets are able to deliverabout 1 T at their surface, and thus further miniaturizationbeyond the engine size outlined above may be difficult toachieve.

Power levels even for the miniaturized Hall thrusterversion currently being studied at MIT43 are quite high forthe microspacecraft considered in this study, but may be fitwithin the upper Class I category if power requirements canbe relaxed and a dedicated power supply is provided for thethruster. However, as mentioned, a Hall thruster deviceappears unsuitable for interplanetary applications requiringlarge delta-v's due to their lower specific impulse capabilitywhen compared with ion engines. In addition, as for ionengine applications, substantial reductions in powerprocessing unit weight and size will have to be made.

Field Emission Thrusters

Field emission thrusters44"46, or Field EmissionElectric Propulsion (FEEP) devices as they are commonlybeing referred to, have traditionally been envisioned forspacecraft attitude control, providing ultra-fine control ordrag make-up to establish virtual-drag-free environments onscientific missions. However, FEEP devices, due to theiralready small size and low available thrust levels, may alsobe able to provide primary propulsion for microspacecraft.

In a FEEP device, thrust is generated throughelectrostatic forces as in ion engines or Hall thrusters, butthe ionization mechanism is different yet again. In a FEEPthruster, a liquid metal propellant (Cs) is fed by capillaryforces through a small channel. The channel ends formingsharp edges that are located opposite a negative electrode,separated by a small gap (about 1 mm) from the channel tip.The channel structure itself carries a positive potential. Anelectric field develops between the two electrodes and the freesurface of the liquid CS metal column near the tip of thechannel deforms, forming cusps which protrude from thesurface of the liquid. As the liquid forms ever sharper conestructures due to the action of the electric field, the localelectric field strength near these cusps increases. Once a localelectric field strength of about 107 V/cm is reached, electronsare ripped off the Cs metal atoms. The electrons are collected

14


through the liquid metal column and the channel walls, andthe positive ions are accelerated away from the liquidcolumns through a gap in the negative electrode by thesame electric field that created them.

Accelerating voltages between the two electrodestypically reach values of around 10,000 V, resulting inspecific impulse values up to 8,000 s48. Rectangular, slit-shaped channel geometries are most frequently investigatedin the laboratory . Depending on slit width and voltagesapplied, thrust levels between 10"6 and 10"3 N have beenachieved. Required power levels for the thruster itself as lowas 60W/mN have been achieved. In addition to thrusterpower, power has to be provided to a neutralizer. Theneutralizer is required to prevent spacecraft charging whileemitting the positive ion beam. Different neutralizerconcepts have been tested, among them a hollow cathodetype and a tungsten filament configuration47. In the hollowcathode type, which requires less power than the tungstenfilament type, a Cs compound is heated, creating a Cs vaporthat is ionized in a hollow cathode discharge, providing theelectrons required for beam neutralization. This neutralizertype delivers 0.1 mA per Watt of electric power, includingheater power.

Advantages of FEEP devices for microspacecraftprimary propulsion applications are its high-specificimpulse and compact size and propellant storage (liquidmetal). Disadvantages are potential contamination issues dueto the use of Cs propellant, and high voltage and powerrequirements, requiring dedicated power processing unitsadding to the propulsion system weight. FEEP thrustersalso require a fairly narrow operating temperature range toavoid solidification of the Cs propellant.

Table 11 lists two FEEP thrusters underdevelopment at Centrospazio in Italy, currently the onlyprovider of FEEP thrusters. As can be noted, thrust specificpower requirements are about 90W/mN for the smallermodel and 75 W/mN for the larger thruster. Given that theneutralizer requires 1 W per 0.1 mA, and , according toPetagna et al.49, an emitter current of 7mA/mN is required,the power required per mN of thrust is (90W/mN +10W/rnA*7mA/mN), i.e. 160W/mN or 145 W/mN in caseof the larger thruster. These power requirements will limitachievable thrust values with a FEEP thruster on board a 20W Class I spacecraft to about 0.125 - 0.14 mN. At thrustvalues this low, thrust-to-spacecraft mass ratios will bearound 0.006 mN/kg. The Europa orbiter mission currentlybeing studied at JPL uses solar elelctric propulsion (SEP)and has thrust-to-weight ratios of around 0.23 mN/kg at thebeginning of the mission, dropping to about 0.03 mN/kg atthe end of the mission50. These values are significantly

Table 11: State-of-the-Art FEEP Thrusters

ManufacturerPropellantThrust (uN)Isp (s)Power (W)Thruster Mass (kg)PPU Mass (kg)Thruster Size (cm)PPU Size (cm)Total Impulse (Ns)

CentrospazioCs100800090.452.91.2x1.2x0.81.2x1.0x0.6160

CentrospazioCs8008000603.54.12.5x2.5x1.52.5x1.3x1.560,000

higher than those obtainable with a state-of-the art FEEPsystem onboard a microspacecraft.

Using FEEP thrusters for microspacecraft primaryapplications will therefore require higher power levels thanthose assumed in this study in Section II. In addition, thepotential Cs contamination issues will have to be studiedand evaluated for each mission. Total impulse capabilitiesappear low for the smaller of the two thrusters listed inTable 11 (corresponding to about 450 hrs of run timeassuming steady-state nominal thrust values), however, arequite high for the larger thruster (corresponding to over20,000 hrs run time assuming nominal steady-state thrustconditions). The latter would be sufficient for electricprimary propulsion applications. PPU masses appearcompatible with the microspacecraft mass allocations.

Future work in field emission thruster technologyis focusing on the use of microfabricated emitter arrays45

consisting of a series of micro-"volcano" structures on awafer. The significance of these arrays is that ions can beproduced at much lower voltages, reducing powerrequirements. Since extractable currents from each micro-emitter are much lower than those obtained withconventionally machined emitters, arrays of many of theseemitters will be required to operate in parallel.

Colloid Thrusters

Colloid thrusters were studied extensively duringthe late 1960s and early 1970s for spacecraft attitude controland drag-makeup, but due to their failure to produce highenough thrusts at reasonable power levels fell out of favor.With the advent of microspacecraft designs, a potentialapplication for microspacecraft primary propulsion may havearrived.

In a colloid thruster, thrust is produced byelectrostatically accelerating fine liquid droplets ejected from

15


a capillary51'52. A strong electric field applied between thesharp-edged exit of the capillary and an external electrodecauses charge separation inside the liquid propellant, whichin most cases is doped with an additive to increase itselectric conductivity. Through a combination ofhydrodynamic instabilities, causing jet break-up into smallliquid droplets, and the action of the applied field acting onthe conductive liquid, charged droplets are extracted from thecapillary at high velocities, producing thrust.

Depending on the propellant used, either positiveor negative liquid droplets can be produced. Mostapplications studied in the past used glycerol doped withsodium iodine to produce positive droplets and glyceroldoped with 2 - 10 % sulfuric acid to produce negativedroplets. The ability to produce both positive and negativedroplets was termed a "bi-polar" thruster by Perel et al.51. Itssignificance is that it can potentially be self-neutralizing,provided the same amount of current can be drawn from eachset of capillaries, eliminating the need for a separateneutralizer.

Several trade-offs have to be made in the design of asuccessful colloid thruster to optimize performance. In orderto obtain good performance, the specific charge, measured inCoulomb per droplet mass, has to be high in order to obtainhigh specific impulses at reasonable applied voltages. Thecolloid thruster is an electric thruster, and as such additionalpropulsion system masses associated with the power supplyor conditioning will have to be offset by sufficientpropellant mass savings that can only be obtained through ahigh enough specific impulse. Specific impulse in a colloidthruster is also determined by the so called specific chargeefficiency, which measures the distribution of specific chargein a droplet stream. A more "peaked" specific chargedistribution will lead to higher specific impulses and higherpropulsion system efficiencies.

Specific charge efficiency in turn depends on severalparameters, such as electric field strengths (higher electricfield strengths reduce the efficiency since they create highercharged particles in addition to lower charged ones), flow rate(lower mass flow rates result in higher specific chargeefficiencies), fluid conductivity (a higher conductivity leadsto lower specific charge efficiencies since it becomes easierto produce more charges in a more conductive stream) andcapillary tip design (which affects the specific chargeefficiency mostly through its effect on the local electric fieldstrength near the tip)51. Some of these design considerationswork against each other. A large potential drop caused by astrong electric field, for example, will accelerate the dropletsto a higher exhaust velocity (which would help to raise thespecific impulse), however, it will also decrease the specificcharge efficiency (which will lower the specific impulse

again). Low flow rates, aiding in obtaining higher specificcharge efficiencies and, thus, specific impulses, will alsoreduce thrust.

In addition to these performance considerations,careful propellant selection will need to be made to ensureproper thruster function and long lifetime. High solvationcapability (to take up dopants), low vapor pressure (to avoidcrystallization of dopants on capillary walls near tip,potentially clogging the system), low freezing point (toavoid clogging) and low corrosivity (to ensure long thrusterlifetime) are key parameters in the selection of thepropellant51. The glycerol propellants discussed above werefound to have a somehwat high vaporization pressure, butwere chosen for its superior ability to dissolve dopants51.Platinum capillaries have been used because of theirresistance to corrosion51.

Based on the design considerations above, colloidthrusters have been designed and operated with platinumcapillaries having ID's of about 200 microns, using sodiumiodine and sulfuric acid doped glycerol propellants, producingthrusts between 0.2 - 0.5 mN at power levels of about 4.4W/mN, requiring voltages of +4.4 kV and -5.8 kV,depending on droplet polarity51. Specific impulses wereestimated between 450 - 700 sec51. In other cases, specificimpulses of up to 1350 sec at thrust levels of 0.55 mN wereobtained (power levels were not reported in that case)52.

Of all micro-electric primary propulsion optionsreviewed so far, colloid thrusters are quite possibly the mostsuited for microspacecraft primary propulsion applications atthis stage of micro-electric propulsion development. Powerrequirements are much lower than those of FEEP devices andfit well within Class I constraints. Thrust levels of 0.5 mNcan easily be achieved using only about 2 W of power in aself-neutralizing, bi-polar array51. Thruster specific masses(excluding power supply and conditioning) have beenestimated at 0.2 - 0.5 kg/W50, and would also fit withinClass I microspacecraft envelopes. However, specificimpulse values are somehwat low for interplanetaryapplications. Thus, future work should focus on obtaininghigher specific impulse values.

Attitude Control - Chemical

Cold Gas Thrusters

Cold gas thrusters represent the smallest rocketengine technology available today. Cold gas systems arevalued for their low system complexity, small I-bit andthrust capability and the fact that, when using benign

16


propellants (e.g. N2), they present no spacecraftcontamination problems. Sometimes high reliability is alsoreferred to as one of the advantages. However, valve leakageproblems have resulted in repeated losses of spacecraft due topremature depletion of propellant through valve internalleaks.

This leakage problem is a result of a combinationof small amount of microscopic contaminants (to be foundin even the cleanest propulsion system) and high pressurepropellant storage. On the contamination side, the propellanttank is one of the major contaminant sources (microscopicmetal flakes left over from fabrication, etc.). Thesecontaminants may locate themselves on valve seats, carriedalong with the propellant flow, and subsequently prevent thevalve from sealing completely. Even though the remainingopening across the valve seat may be so small that for aliquid propellant application it would pose no problem dueto the higher liquid viscosity, in cold gas systems, where thepropellant is stored at very high pressures, propellant mayescape even through these microscopic openings.

Another disadvantage of cold gas systems are theirlow specific impulse performances, unless very light gases(H2, He) are used. Neither hydrogen or helium is commonlyused, however, since storage problems and large and heavytankage would result due to low gas densities, and additionalleakage concerns would have to be considered with theselight gases.

Table 12 gives a list of typical cold gas

performance values, based on data found in Refs. 8 and 26.Of the gases listed, nitrogen is by far the gas mostfrequently used as a cold gas, due to a combination ofreasonable propellant storage density, performance and lackof contamination concerns.

Table 13 lists some of the smallest cold gas thrustersavailable today. The Moog 58x125 thruster is merely 4.3cm in length and 1.4 cm max. dia., including fitting. Thefitting accounts for roughly half of the size of the totalenvelope. Size, mass and power requirements fit well withinthe Class I microspacecraft envelope. However, even a coldgas thruster this size may perform only marginally withrespect to the impulse bit requirements (compare data inTable 13 with comments in Section II).

Using the data found in Tables 12 and 13, requiredleak rates for microspacecraft can be estimated and currentcold gas thruster performances be evaluated in this regard.Using the information provided in Section II, a total "delta-v" requirement of 50 m/s is assumed. For a specificimpulse of 70 sec (N2) and an assumed microspacecraft massof 10 kg (Class I), the required attitude control propellantmass is 0.7 kg of nitrogen. At a storage density of 0.28g/cm3 for nitrogen at 3500 psia and 0 C, a tank volume ofabout 2500 cm3 is required. Taking into account thepossibility of propellant leakage, assume that 10% morepropellant is loaded onto the spacecraft, now requiring a tankvolume of 2750 cm3 at the same storage pressure. Assuminga spherical tank, this translates into an inner tank diameterof roughly 37 cm. This tank size is slightly larger than the

Table 12: Cold Gas Propellant Performances

Propellant

HydrogenHeliumNeon

NitrogenArgon

KryptonXenon

Freon 12Freon 14MethaneAmmonia

Nitrous OxideCarbon Dioxide

MolecularWeight

(Kg/Kmol)2.04.020.428.039.983.8131.31218816174444

Density(3500 psia 0 C)

(g/cm3)0.020.040.190.280.441.08

2.74'"-

0.960.19liquid

-liquid

Isp*(Theoretical)

(s)2961798280573931

46"5511410567*67

Isp'(Measured)

(s)27216575735237283745105966161

at 25 C. Assume exoansion to zero pressure in case of theoretical value.at 38 C (560 R) and area ratio of 100.Likely stored at lower pressure values (2000 psia) to maximize propellant to tank weight ratio.

17


Table 13: Small Cold Gas Thrusters

ManufacturerTypeThrust (N)Ibit (Ns)Isp (s)Pressure (kPa)OpenResponse (ms)Power(pull-in) (W)Weight (g)

Moog55

58x1250.0045io-4

65 (N2)34.50.94

2.4

7.34

Moog"58x1152.89--14603.5(spec)30

13

Marquardt7

-4.5--8840<1.1

-

5.4

envelope assumed for a 10-kg spacecraft, but is within theright range. The tank, however, will dominate the spacecraftdesign layout completely. Assuming further that all of theadditional 10% of the propellant may be lost over the courseof the mission (corresponding to 250 cm3 at 3500 psiastorage pressure or almost 59,000 sec, assuming zerocompressibility of nitrogen in this rough estimate),maximum allowable leak rates for a 2 year mission wouldbe 9 x IO'4 scc/s and for a 3 year mission 6 x IO"4 scc/s.These small leak rate requirements are a consequence of thesmall spacecraft size. Since smaller spacecraft carry smalleronboard propellant supplies for the same attitude controlrequirements, less propellant can be afforded to be lost due toleakage and valve leak rates for microspacecraft consequentlyhave to be even lower than for bigger spacecraft.

Recent leak tests with cold gas thrusters have shownleak rates lower than the ones calculated above53, however, atthruster operating pressures which where a mere 34.5 kPa,or about one third of an atmosphere. The low leak ratescalculated above will have to be maintained throughout theentire feed system, however, in particular across componentsthat are exposed to the full tank pressure. A cold gas thrustervalve recently developed by Marquardt7 is subjected topressures of about 8900 kPa (1300 psig) and maintains leakrates of <2.8 x IO"2 scc/s GHe7, corresponding to roughly 1x IO"2 scc/s GN2, assuming a ll^jM (M being themolecular weight) dependence for the leak rate. These valuesare far higher than the estimated leak rate requirements givenabove.

The leak rate requirements could be relaxed if morepropellant reserves were carried. Performing the calculationsabove for the same case of a 10 kg microspacecraft requiringa 50 m/s attitude control budget over the course of themission, but raising the propellant margin to 50% wouldyield maximum allowable leak rates of 4 x IO"3 scc/s in thecase of a 2 year mission and 2.5 x IO"3 scc/s for a 3 yearmission. While these leak rates approach obtainable values,

it has to be noted that these rates will have to be maintainedover the course of the entire mission, even after the valvehas been subjected to many cycles and substantial propellantflow, carrying contaminants through the valve. In addition,raising the propellant reserves from 10% to 50% leads to abigger required tank diameter of 41 cm ID compared to 37cm before. These tank designs will completely dominate thespacecraft design layout and consume a substantial portionof the overall spacecraft mass and volume.

An interesting alternative to conventional cold gaspropulsion using high pressure gas tanks is the use ofammonia as a propellant. As pointed out by Nakazono54,ammonia has a vapor pressure of 33 psia (224 kPa) at -18C. Thus, even without tank heaters, sufficient pressure couldbe provided to an ammonia cold gas thruster merely usingthe boil-off of the propellant. As can be seen from Table 12,specific impulses obtainable with ammonia are higher thanthose achievable with nitrogen. The ammonia system wouldallow for liquid storage, reducing tank size and mass andpropellant leakage concerns. Depending on availablevaporization rates (dependent on tank temperature),propellant flow rates may be limited39.

It may therefore be concluded that cold gas propulsionsystems using high-pressure gas supplies do not appear tobe a viable option for microspacecraft unless attitude controlrequirements can be substantially reduced, valve leak ratescan be lowered by at least one order of magnitude, and severemass and volume constraints on the remainder of themicrospacecraft, caused by large and heavy propellant tanks,can be tolerated. Cold gas systems based on liquid storage ofammonia, on the other hand, appear as a very attractiveoption for microspacecraft attitude control. In either case,obtainable input bits have to be reduced further, even forClass I microspacecraft applications. This requires thedevelopment of either faster valves or smaller nozzle throatareas. Fabricating nozzle throat diameters smaller than theones obtainable today may require the exploration of newtechnologies, such as MEMS (see Section V).

Warm Gas Thrusters

Cold gas technology discussed above could beadapted to warm gas thruster concepts. In this casehydrazine propellant is decomposed in a separated gasgenerator, consisting in essence of a Shell 405 catalyst bed,and gaseous hydrazine decomposition products are then fed toa plenum and finally through a cold gas thruster54. Such asystem would not require separate propellant tanks if aconventional hydrazine thruster was used as themicrospacecraft primary propulsion device. In addition,several cold gas thrusters in existence today already claim

18


compatibility with hydrazine decomposition products. Aseparate heater would be required to heat the catalyst bed inorder to allow for a sufficient number of starts, just as inconventional hydrazine thruster technology. By feeding thedecomposition products into a plenum, from which they canbe drawn to the various attitude control thruster clustersupon demand would lower the required number of catalyststarts54.

A hydrazine warm gas system is a very attractiveoption for microspacecraft, in particular when a hydrazinepropellant supply is already onboard for primary propulsionpurposes. Relatively high performance, comparable to theammonia cold gas system described above, can be combinedwith compact propellant storage and the relatively near termavailability of the required propulsion components, drawingupon cold gas heritage. As in the case of cold gas systems,either faster valves or smaller nozzle orifices are required tolower impulse bits to the requirements for both Class I andClass II microspacecraft applications.

Tri-propellant Thrusters

In a tri-propellant thruster, a propellant mixture ofhydrogen, oxygen and an inert gas, such as helium or, morecommonly for propellant storage reasons, nitrogen isused7'57. The propellants are stored fully mixed, no separatetanks are required. The addition of the inert gas to themixture renders the mixture non-combustible, until exposedto a suitable catalyst. Different catalysts are being studied,typically based on nobel metal compounds27'58, althoughdetails on catalyst compositions are mostly treated asproprietary information. Thruster performances rangebetween 70 s and 140 s of specific impulse57. A recentdesign by French7 is aimed at 125 sec Isp and a thrust ofabout 2N. However, this design was not specificallyintended for microspacecraft use.

Tri-propellant systems, even though able to deliverhigher performance than cold gas systems, suffer the samedisadvantage of high-pressure propellant storage and theassociated leakage problems as cold gas systems. In addition,even though required propellant masses can be reduced overconventional cold gas systems due to the higher specificimpulse performance, required high-pressure propellant tankswill likely continue to dominate the spacecraft design. Thus,the advantages gained with the use of tri-propellant systemsover cold gas systems onboard microspacecraft may belimited.

Hydrazine Mono-Propellant Thrusters

Hydrazine mono-propellant thrusters readilyavailable today (see Table 8) are far too large and heavy to beused in attitude control clusters around a microspacecraft.The development of new, miniature hydrazine thrusters isrequired. A research initiative to study the feasibility ofminiature hydrazine thrusters was recently begun at JPL,but it is too early to present results from this acitivity.

Bi-propellant Thrusters

Bi-propellant thrusters were not considered formicrospacecraft attitude control. Currently available enginetechnology is far too big and heavy for use inmicrospacecraft attitude control clusters. Bi-propellantengines have been considered for attitude control purposes onlarger spacecraft to allow for easier integration into primarypropulsion systems, eliminating separate attitude propellantsupplies. It was already concluded, however, that bi-propellant systems are probably not suitable for onboardmicrospacecraft primary propulsion. In addition, eventhough bi-propellant thrusters do offer higher performances,reducing propellant requirements, required propellant massesfor attitude control are usually small, not providing anopportunity for large spacecraft mass reductions. Even if asubstantial reduction of the typically small attitude controlpropellant mass could be achieved, it would likely be offsetby the higher dry mass of a bi-propellant system due toincreased component parts count when compared with muchless complex cold or warm gas, or even miniature hydrazinesystems. Finally, in view of the survey of state-of-the-artminiature bi-propellant technology given above, there existsconsiderable doubt whether further substantial reductions inthruster size can be made while still being able to providereliable, space-qualifyable engine technology.

Attitude Control - Electric

Pulsed Plasma Thrusters (PPTs)

In a pulsed plasma thruster, electrical power is usedto ablate, ionize and electromagnetically accelerate atoms andmolecules from a bar of solid Teflon50'59. The Teflon bar ispushed against a retaining lid between two electrodes bymeans of a negator spring. The electrodes are connected to acapacitor, which is unable to discharge because the vacuum

19


and solid Teflon bar between the electrodes do not provide aconductive path. A spark plug located near the solid Teflonsurface is fired, removing a portion of the Teflon andionizing it. In the process, the Teflon bar is pushed forwardtoward the lid and brought in position for the next pulse.The capacitor now discharges through the ionized Teflongas. The particles inside the Teflon discharge consist of avariety of molecular fluorocarbons, which are ionized in theprocess and give rise to a current flow between theelectrodes. This discharge current generates a strongmagnetic field surrounding it. Lorentz forces acting on thedischarge current as a result of this magnetic field push theTeflon plasma out of the thruster at high exhaust velocitiesof about 10 - 20 km/s. Non-ionized particles are expandedfrom the thruster due to Joule-heating.

Thrusts generated per pulse are on the order of 10sto 100s of micro-newtons. However, because the pulselength is on the order of milliseconds, the capacitor can becharged and discharged several times per second, thuscreating an accumulative thrust in the micronewton tomillinewton range. Pulsing frequencies between 1 Hz and 6Hz are common. The ability of a PPT to produce very smallthrust levels per pulse allows for the possibility to deliververy small impulse bits on the order of less than 1 mNs.Future designs are aimed at providing merely 10s of |J.Ns.PPT thrusters have been developed since the 1950s and haveflown on several US satellites. State-of-the art thruster areable to provide specific impulses in the range of 800 - 1500s, thrusts of 220 - 1100 \iN, and efficiencies between 5 - 15% at a wet mass of 5 kg.

Current PPT designs are therefore too large formicrospacecraft attitude control. Future designs are beingcontemplated and are predicted to have wet masses as low as0.5 kg. Providing 12 of these thrusters for attitude controlon a microspacecraft will lead to a total (wet) system massof 6 kg. These mass values are compatible with Class Imass guidelines. With minimum impulse bits for the future, 0.5 kg thrusters predicted at 10 - 100 |iNs, PPT are alsoable to provide the minimum impulse bit requirements inthe upper Class I microspacecraft category. However,minimum thrust requirements for attitude control of Class Ispacecraft, ranging between 1.75 - 4.5 mN for the largerspacecraft masses, can only be met at high power levels.According to Ref. 59, a 1.75 mN thrust requires a powerlevel of 120 W and pulsing frequencies between 3 - 6 Hz,depending on capacitor size. These values far exceed thepower levels that will be available for attitude control onany of the microspacecraft considered here. PPTs can thusonly be used on Class I craft if slew rate requirements can bereduced substantially. Current and predicted PPT

technologies are too large and heavy for Class II spacecraft.

Field Emission Thrusters

Field emission, or FEEP, thrusters, surveyedabove, are reviewed here in terms of their applicability asattitude control thrusters on microspacecraft. Large PPUmasses and high power requirements may prohibit their usein such a function. Following the example given in the coldgas thruster section above, for a typical Class Imicrospacecraft mission, 0.7 kg of nitrogen gas would berequired to meet a 50 m/s delta-v budget. If a FEEP systemwith a specific impulse of 8,000 s was used, the requiredpropellant mass would be reduced to 12 g. While this is asubstantial reduction, the required PPU mass is, according toTable 11, is 2.9 kg for the smaller of the two FEEPthrusters considered. Since Cs propellant can be stored in itsliquid state, whereas nitrogen gas would have to be stored inhigh-pressure tanks, tank weight reduction will benefit theFEEP system. According to recent tank data60, a 3,244 cm3

tank (roughly the size required for nitrogen storage in theexample above), capable of maintaining a maximumexpected operating pressure (MEOP) of 10,000 psia (farmore than required for a storage pressure of 3,500 psia asassumed), weighs around 1.8 kg (these data are based on acylindrical tank). Even neglecting tank masses for the FEEPsystem, the cold gas system will still be lighter than theFEEP system by about 0.4 kg (although volumerequirements will be higher due to the large tank volume). Inaddition, the FEEP system will not be able to provide theminimum thrust for the assumed slew rate requirementswithin the power constraints expected to be found on amicrospacecraft. FEEP systems may find applications asattitude control devices for microspacecraft if slew raterequirements can be relaxed significantly and very smallimpulse bits are required.

Colloid Thrusters

Much the same comments as made with respect tothe applicability of a field emission systems tomicrospacecraft attitude control can be made for colloidsystems. Minimum thrust requirements cannot be met , andadditional power processing unit masses may offsetpropellant reductions gained over cold gas systems throughthe use of the higher specific impulse colloid thruster. Sincecurrent colloid thrusters provide lower specific impulses andrequire less power, the comparison between colloid and coldgas systems may be shifted somewhat more in favor of thecolloid system than was the case for the FEEP system.

20


Resistojets

State-of-the art resistojet technology22 is far tooheavy, and requires far too much power (in excess of 350 W)to be useful for microspacecraft attitude control. However,some work on small water resistojets was performed61"63.Work focused on measuring small nozzle performancesusing water vapor. No heater power requirements werereported.

Water is not the most suitable propellant forresistojet use to its high heat of vaporization, even though itdoes simplify laboratory testing due to its lack of safety andtoxicity concerns. Table 14, based on data from Ref. 26, listseveral relevant properties of candidate propellants for aresistojet system. Of the propellants listed, ammonia andwater immediately stand out due to their low molecularweight, resulting in high specific impulse performance. Ofthese two propellants, ammonia requires less heat tovaporize and is thus the propellant of choice.

Since no small resistojet technology exists, it isnot clear how well this technology is suited formicrospacecraft use. Liquid storage of propellants will reducesystem weights over high-pressure cold gas storage systems. Comparing ammonia resistojets with the ammonia coldgas thruster discussed above, the resistojet should enablehigher duty cycles and longer burns, making resistojettechnology more versatile. The disadvantage is their needfor a separate power supply, adding to the system weight.Work is currently underway at JPL and the Aerospace

Corporation to develop micro-resistojet technology usingMEMS technologies (see Section IV).

IV. FUTURE DEVELOPMENT NEEDS ANDEMERGING TECHNOLOGIES

Based on the thruster survey performed in SectionIII, technology needs for microspacecraft propulsion can beidentified. Several emerging micropropulsion technologiesare being introduced. These emerging technologies are thefirst propulsion components designed specifically withmicrospacecraft applications in mind.

Identification of Technology Needs

The results of the survey conducted in Section fflare summarized in a matrix, shown in Table 15. Only state-of-the-art technologies, or those under significantdevelopment, are listed in this matrix. The technologieslisted were evaluated in view of their application to Class Iand II microspacecraft attitude control and primarypropulsion applications by grouping them into threecategories: those technologies that appear applicable to thetask ("yes" category), those that do not ("no" category), andthose that fall somewhat in between ("maybe"). The lattercategory serves to classify technologies that may fulfillsome mission requirements, but not others, or those thatcould be made to fulfill all requirements for the specific

Table 14: Properties of Candidate Resistojets Propellants

Propellant

AmmoniaPropane

Ethyl ChlorideButane

Freon 12Water

Hydrogen FluorideMethanol

Methyl ChlorideEthane

Ethyl MethylEther

Mono MethylAmine

Formula

NH3C3H8

C2H5C1C^HIO

CC12F2H2OHF

C2H3OHCH3OH

C2H8C2H5OCH3

CH3NH2

MolecularWeight

(kg/Kmol)17.044.164.558.1120.918.020.144.051.030.060.0

31.0

Liquid Density(g/cm3)

0.60.490.920.570.981.0

0.990.790.910.560.8

0.77

Heat ofVaporization

(kj/kg)1159.7339.3388.1360.2141.8

2442.51505.91099.3376.5313.7350.9

873.8

21


Table 15: Matrix of Applicability Status of State-of-the-Art Technologies, or Technologies under SignificantDevelopment, to Microspacecraft

Technologies(State-of-the-Art

Only)Bi-Propellant

HydrazineHAN/TEAN

PeroxideSolid

HybridCold Gas

Ammonia ColdGas

Warm GasTri-Propellant

IonHall

FEEPColloid

PPTResistojet

Class IPrimary

NoYes

MaybeNo

YesMaybe

NoNo

NoNoNoNo

MaybeMaybe

NoNo

ACS

NoNoNoNoNoNo

MaybeMaybe

MaybeMaybe

NoNo

MaybeMaybeMaybe

No

Class IIPrimary

NoNoNoNo

MaybeNoNoNo

NoNoNoNoNoNoNoNo

ACS

NoNoNoNoNoNoNoNo

NoNoNoNoNoNoNoNo

microspacecraft category and task with only minortechnology advances required.

Inspecting Table 15, the lack of suitable thrustertechnology for microspacecraft applications becomesobvious. This is not too surprising a result, given that mostof the technologies reviewed here were developed forspacecraft much larger than the spacecraft considered here.Some technologies stand out, however, and may appearapplicable to microspacecraft even in, or close to, theircurrent design stage. For primary Class I applications, thesmallest available hydrazine attitude control thrusters appearas a suitable and reliable thruster option if only small tointermediate delta-v's are required. Solid motors also mayprovide low to intermediate delta-v capability within theircurrent design limits, although thrust values may need to bereduced further. However, no thruster options providing largedelta-v and compatible with the microspacecraft assumptionsare currently available. Both bi-propellant chemical as wellas ion in their current manifestations would require separatepropulsion stages due to high system masses and volumesresulting in larger launch masses. Possible exceptions areFEEP and colloid thrusters. While FEEP thrusters stillrequire too much power, and colloid thrusters deliver fairlylow specific impulses, they may be marginally suitable atthis point of their development, but will likley requirefurther development. PPT thrusters do not have the total

impulse capability and also require too high power levels forthe minimum thrust levels needed.

For Class I attitude control applications, cold gasand warm gas systems currently offer the greatest near-termpotential. In both cases, however, delivered impulse bits aretoo high. Impulse bits maybe reduced through faster valvetechnologies or increased flow restriction through the use ofsmaller nozzle throats or separate flow restrictors. The lastoption may be the easiest to achieve, but fabricating smallorifices with low tolerances and protecting them fromclogging through contaminants may require the use of newtechnologies (MEMS).

PPTs, FEEPs and colloid thrusters may beapplicable for Class I spacecraft attitude control only whenslew rate requirements can be lowered dramatically. Giventhat current microspacecraft mission requirements may stillundergo considerable evaluation and may vary from missionto mission, these three technologies have been listed aspossible candidate technologies. If slew rate requirementswill not change significantly, large reduction in powerrequirements and PPU specific masses will be needed.

For the Class II category, virtually no state-of-the-art propulsion technologies appear suitable. Once again, thereason for this is that current hardware was designed for

22


larger spacecraft and currently available components mayeven exceed the size of a typical Class II spacecraft. Theonly primary propulsion device that may be applicable toClass II spacecraft using state-of-the-art technology (yet stillrequiring a full development program) may be solid motors.For attitude control, no current propulsion hardware appearssuitable, either because the size of available components donot allow for the distributed mounting of a dozen or sothruster around a Class II bus, or because deliveredminimum impulse bits are too high.

Therefore, high-specific impulse primarypropulsion for both spacecraft categories, as well as low I-bit attitude control thruster technology for Class I, but inparticular Class II, are identified as major propulsiontechnology needs for the microspacecraft considered here.Class II attitude control applications will also requireextreme miniaturization.

Emerging Technologies

The increasing interest in microspacecraft for deep-space missions has led to the exploration of technologiessuitable for these spacecraft at JPL, NASA's lead center forrobotic interplanetary space exploration and one of the firstpotential users of these technologies. Several newmicropropulsion technologies were proposed in the course ofthese activities. Three of these emerging micropropulsiontechnologies are discussed here. Among them are twomicro-attitude control phase-change thrusters (vaporizingliquid or resistojet and subliming solid) and micro-ionengines for high delta-v primary applications. All threetechnologies rely heavily on MEMS-fabrication techniques,allowing for the potential of order-of-magnitude reductionsin component mass and size, possibly suitable for bothClass I as well as Class II applications. Associated witheach technology considered here, however, are significantfeasibility issues that will have to be thoroughlyinvestigated in coming years. Thus, the thrustertechnologies introduced here have to be viewed as veryadvanced micropropulsion options.

Vaporizing Liquid Micro-Thruster

The vaporizing liquid micro-thruster, or micro-resistojet, is a concept that was proposed by Leifer andMueller , as well as Janson at the Aerospace Corportaion, acouple of years ago7'64. In this thruster concept, described indetail in a companion paper65, a suitable liquid (ammonia orhydrazine) is heated via a thin-film heat exchanger, micro-fabricated onto a silicon substrate. A conceptual sketch andpicture of the assembled JPL-device are shown in Figs. 2

and 3. Two identical silicon wafers, featuring thin-filmheaters and micro-nozzles, are bonded to a Pyrex spacer that,in its final assembly, is sandwiched between the two siliconwafers. The liquid propellant, pressure-fed through one of theopenings machined into one of the silicon wafers, enters thethruster and is vaporized as it flows between the heaterelements. Propellant vapor is then exhausted through thesecond nozzle. A recess machined into the silicon underneaththe heater creates thermal chokes near the heater edges,reducing conductive losses to the structure. Silicon has avery high thermal conductivity, providing a bigger challengefor the thermal design of this device. However, no othermaterial can currently be micromachined to the degree andflexibility that silicon can, and silicon was thus chosen asthe substrate material.

A key aspect of this design is its simplicity. Itdoes not contain any complex moving parts (such asMEMS-pumps, turbines, etc.) which could decrease the

Intel

Thin-Film Heater(1of2)

Nozzle

Silicon Substrate

Fig. 2: Concept of the Vaporizing Liquid Thruster

Fig. 3: Vaporizing Liquid Micro-Thruster Chip

23


reliability of the device. Performance targets are 0.5 toseveral mN thrust (to meet minimum thrust requirementsfor microspacecraft attitude control), 50% efficiency andpower requirements of a few W or less. The latter condition,dictated by available assumed power levels onboard amicrospacecraft, will limit performance. Specific impulsevalues around 100 to 150 sec have been estimated dependingon propellant and heater temperature obtainable. However,given that delta-v budgets for attitude control are fairly low(see Section 31), high performance is not necessarily required.

Impulse bits will largely be determined by thevalve technology to be used. Presently, no suitable MEMSvalve technology exists to be interfaced with this thruster,but miniature conventional valves, approaching in both sizeand mass of fully packaged MEMS valve technology, maybe used initially. With open-to-close times on the order of afew ms, impulse bits may range as low as 10"5 Ns or less.Work on integrating MEMS technology with conventionalvalve technology is currently being explored at MarottaScientific Controls, Inc under an SBIR program.

Several prototype devices ( see Fig. 3) have beenassembled and are awaiting testing this summer. Focus ofthe initial work is to optimize heater designs, search fordesign options to reduce power losses to the structure,determine minimum heater lengths, test the reliability of theelectric contacting through thruster cycling and gain initialthrust performance data. Work on micro-resistojets is alsounderway at the Aerospace Corporation7. These devices aresimilar in function, but feature slightly different heaterdesigns7.

Subliming Solid Micro-Thruster

Subliming solid thruster concepts are not new andsubstantial development work was performed with thesethrusters in the 1960's. Main contributors to this field wereRocket Research66"68 (now Primex), Lockheed69'74 (nowLockheed-Martin), and NASA Goddard75'76. Some work wasalso performed at Aerospace Industries77, the LewisResearch Center78 and the Martin-Marietta company79 (nowLockheed-Martin also). In the subliming solid thrusterconcept, a solid propellant is chosen with a highsublimation pressure, such as ammonium hydrosulfide(NH4HS) or ammonium carbamate (NH4CO2NH2)- Uponheating, gas pressure builds up inside the propellant tank andthe vapor is vented through a valve and nozzle to producethrust.

The simplicity of this design, and the solidstorability of the propellant appear to easily lend themselves

to miniaturization. Based on the 1960's work in this area, asubliming solid micro-thruster concept was proposed by theauthor using MEMS technology80. A conceptual drawing ofthe device is shown in Fig. 4. The chip consists of twolayers, one made from silicon the other from Pyrex. Thesilicon portion of the chip was machined at Sandia Nat'1Labs. Propellant vapor enters the chip from the tankthrough the circular Pyrex hole and flows along a recessmachined into the silicon towards the nozzle orifice. On itsway, the propellant passes through a micromachined combfilter. This filter will prevent solid particles from driftingout of the tank and towards the nozzle in the zero-genvironment of space, preventing nozzle blockage.

The micro-nozzle has a throat area of about 50 x 50ftm and is shown in Fig. 5. The square nozzle shape is aresult of the anisotropic etch used in the fabrication of thisdevice, resulting in preferential etching of some crystalplanes over others. The wall surfaces of the nozzle arecomposed of {111} planes, etching slowest in the fabricationprocess. Using this technique, simple converging-divergingnozzle shapes can easily be fabricated. Nozzle shapes are notoptimized in this design, since performance optimization isnot a goal with the current, first generation of chips, andfuture design iterations may explore different contours.

A mock-up of the thruster concept is shown in Fig.6, consisting of a tank, a valve and the thruster chipassembly. The valve shown in Fig. 6 is a MEMS-typevalve based on thermopneumatic action, designed byRedwood Microsystems, Inc. This valve has leakage issuesand currently only serves as a place holder. It should bepointed out, however, that due to the solid propellant storageand relatively low vapor pressures inside the tank, leak raterequirements for valves may be relaxed significantly overthose required for cold gas systems.

Entrance Hole'.*> Propellant Met

(from Tank)Pyrox Cover

Nozzle Entrance{Square Nozzle Contour)

PropeBanlOulet(Through Nozzle In Si-

Substrate

Fig. 4: Subliming Solid Thruster Concept

24


Fig. 5: Subliming Solid Micro-Thruster Nozzle and Filter

Fig. 6: Mock-up of a Subliming Solid Micro-thrusterSystem

Propellant condensation may occur as propellantvapors are exposed to the cold surfaces. In the event thatpropellant condensation leads to clogging of the filter, aheater was thin-film deposited just underneath the filter onthe opposite side of the silicon substrate (facing up in Fig.6). Similar heaters may be required around the valve toprevent recondensation on the valve seat, preventing sealing.Valve heating will necessitate the selection of a differentvalve design than the one shown in Fig. 6 (with no thermal

activation as in the case of the Redwood valve shown inFig. 6 allowed).

Advantages of the subliming solid micro-thrusterconcepts are its compact size and propellant storage, lightweight tankage due to low pressure requirements, relativeimmunity to leak rate concerns and suitable projectedperformances. Thrust values of about 0.5 mN up to severalmN are targeted with this design. If the chip can beinterfaced with fast acting valves very low impulse bitswould result. Disadvantages of the system are its lowperformances (50- 75 s Isp estimated) and toxicity issuesassociated with the propellants, requiring special handlingduring testing and propellant loading.

Micro-Ion Engines

High specific impulse micropropulsion devices willbe required in order to achieve high-delta-v capability formicrospacecraft. MEMS-based devices have been studied81'82,however, were found to have problems associated withelectron wall losses. In addition, for the thrust and powerlevels required, conventionally machined thrusters may beused7. MEMS technology, however, does hold promise foruse in various ion engine components. MEMS acceleratorgrid system technology is currently being studied at JPL.Using this technology, ion beams may be steeredelectrostatically by designing a special third, or deceleratorgrid, of a three-grid system, allowing for the applicationdifferent electric potentials to different grid section (see Fig7).

In order to fabricate MEMS-scale grids, electricbreakdown voltages of MEMS materials, such as siliconoxide, will have to be tested. Insulator materials, such as

DecelGridAccelGrid

^ScreeGrid

n

Fig. 7: Concept of Electrostatic Gimbaling

25


silicon oxide, will be needed as grid insulator materials sincefree standing grids may not be practical at these dimensions.Figure 8 shows a test chip recently completed at JPL to testsilicon oxide breakdown strengths. The oxide layer isdeposited between two electrodes (a polysilicon layerunderneath the oxide layer, accessible through a via etchedthrough the oxide, and an aluminum layer on top of theoxide). Applying voltage to the two contact pads shown inFig. 8 will allow silicon oxide breakdown voltages to betested. An integral heater allows for breakdown tests atvarious temperatures.

Other areas in the development of miniature ionengine technology that require attention in the future atemicro power processing units and micro flow control andfeed systems. In the latter area, micro-ion engine systems(and conventional systems due to their low flow raterequirements) are likely to benefit from the integration ofMEMS technologies. Another SBIR contract performed byMarotta Scientific Controls, Inc addresses the issues ofmicroflow control on a chip in a device termed a "micro gasrheostat."

V. CONCLUSIONS

Existing thruster technologies were reviewed in theview of potential applications for microspacecraft, defines asspacecraft with masses of 1 - 20 kg. Based on this review,technology needs were defined and several emerging,advanced micropropulsion concepts specifically designed formicrospacecraft applications were introduced.

Only a few of the currently existing thrustertechnologies appear applicable for spacecraft of the sizeconsidered here. For primary propulsion applications, smallhydrazine thrusters and solid motors may provide

Fig. 8: Grid Breakdown Test Chip

intermediate to low delta-v capability. Thrust values of solidmotors may have to be reduced further in order to avoidexcessive spacecraft accelerations. FEEP and colloid thrustermay possibly be used as primary propulsion devices ifpower requirements can be lowered in the case of FEEPdevices, and specific impulse can be raised for colloidthrusters.

For attitude control functions, currently availablecold gas systems approach in performance the requirementsimposed by microspacecraft designs with respect tominimum thrust and impulse bit values. Impulse bits,however, will have to be lowered even further beyond thevalues obtainable with todays smallest thrusters for Class nmicrospacecraft. In addition, significant leakage concernsexist for cold gas systems and the required high-pressurestorage tanks will completely dominate microspacecraftdesign with respect to both size and mass, even for relativelybenign attitude control requirements. Ammonia cold gasthrusters or hydrazine warm gas systems may provide fairlynear-term solutions to the propellant storage and leakageissues. For spacecraft with masses of less than 5 kg,virtually no suitable propulsion hardware exists, neither forprimary propulsion, nor attitude control. In many cases,existing thrusters are larger than the spacecraft in question.

Future technology needs for both primary andattitude control propulsion will be required, in particular forspacecraft below 5 kg in mass. However, even for largerspacecraft (5 - 20 kg), significant improvements for high-specific impulse primary propulsion will need to be made inorder to reduce propellant masses , thus aiding in keepingmicrospacecraft small. Also, further reductions with respectto I-bit performance are required.

New advanced micropropulsion technologiesintroduced include both attitude control thrusters of thevaporizing liquid (resistojet) and subliming solid thrustertypes, and micro-ion engines for primary propulsionapplications. All devices make heavy use of advancedmicrofabrication techniques, such as siliconmicromachining. Several devices were designed and builtover the course of the past year and represent the latesttechnology advances in the micropropulsion area,specifically aimed at addressing microspacecraft technologyneeds.

VI. ACKNOWLEDGMENTS

The author would like to thank Ross Jones fromJPL and the NASA Advanced Propulsion Concept programfor supporting this work. He also wishes to thank WilliamTang, Wen Li, Indrani Chakraborty, Lilac Muller, Thomas

26


George, Andrew Wallace, Russell Lawton, Dave Bame,Morgan Parker and John Blandino for their contributions tothe JPL micropropulsion activities, and John Brophy forreviewing this paper.

The work described in the paper was conducted atthe Jet Propulsion Laboratory under contract with theNational Aeronautics and Space Administration.

VH. REFERENCES

'Collins, D., Kukkonen, C, and Venneri, S., "Miniature,Low-Cost, Highly Autonomous Spacecraft - A Focus for the NewMillennium", IAF Paper 95 - U.2.06, Oslo, Norway, Oct. 1995.

2West, J., Pers. Comm., JPL, Fall 1995.'Fleeter, R., "Microspacecraft". The Edge City Press, Reston,

VA, 1995."Mueller, J., "Propulsion Options for Microspacecraft", JPL

Internal Document, JPL-D-13444, Dec. 4, 1995.'Jones, R., "JPL Microspacecraft Technology Development

(MTD) Program", JPL Internal Document, May 31, 1996.'West, J., "Microelectromechanical Systems

(MEMS)/Nanotechnology Studies", JPL Internal Document,JPL-D-13302, January 12, 1996.

'Proceedings, JPL Micropropulsion Workshop, April 7-9,1997, to be published.

8Brown, C.D., "Spacecraft Propulsion". AIAA EducationSeries, 1995.

'Ruttle, D. and Fitzsimmons, M., "Development of Miniature35-lbf Fast Response Bipropellant Divert Thruster", AIAAPaper 93-2585, Monterey, CA, June 1993.

'"Hodge, K.F., Alien, K.A., and Hemmings, B.,"Development and Test of the ASAT Bipropellant AttitudeControl System (ACS) Engine", AIAA Paper 93-2587,Monterey, CA, June 1993.

"Craddock, J. and Janeski, B., "Design and Development ofthe Army KE ASAT ACS Thruster, AIAA Paper 93-1959,Monterey, CA, June 1993.

!2Rosenberg, S.D. and Schoenmann, L., "New Generation ofHigh-Performance Engines for Spacecraft Propulsion", JPP,Vol.10, No.l, Jan-Feb. 1994. pp.-40-46.

13Schwende, M.A., Schulte, G., Dargies, E., Gotzig, U. andScharli-Weinert, E., "New Generation of Low-Thrust Bi-Propellant Engines in Qualification Process", AIAA Paper 93-2120, Monterey, CA, June 1993.

14Chirivella, J., "Analysis of the Transient Behavior of the 4-Newton Attitude Control Thruster of the Comet RendezvousAsteroid Flyby (CRAF) Spacecraft", JPL Internal Document, JPLP.O. 958214, Nov. 25, 1988.

"Sutherland, G.S. and Maes, M.E., "A Review of MicrorocketTechnology: 10'6 to 1 Ibf Thrust", J.Spacecraft and Rockets,Vol. 3, No. 8, August 1966, pp. 1153 - 1165.

"Wichmann, H. and Fitzsimmons, M., "Miniature HighPerformance Delta V Engine", AIAA Paper 93-2582, Monterey,CA, June 1993.

"Kaiser-Marquardt, Company Information.

18Royal Ordnance, Company Information."Atlantic Research, Company Information.20Aerojet, Company Information.2'Bryant, K., Knight, C., and Hurtz, R., "Planetary Lander

Vehicles Utilizing LEAP Technology", AIAA Paper 94-2748,Indianapolis, IN, June 1994.

22Primex, Company Information.23TRW, Company Information.24Jankowsky, R.S., HAN-Based Monopropellant Assessment

for Spacecraft", AIAA Paper 96-2863, Lake Buena Vista, FL,July, 1996.

"deGroot, W.A. and Oleson, S.R., "Chemical MicrothrusterOptions", AIAA Paper 96-2868, Lake Buena Vista, FL, July1996.

26Grossmann, I., Jones, I.R., and Lee, D.H., "AuxiliaryPropulsion Survey, Part III: Survey of Secondary Propulsion andPassive Attitude Control Systems for Spacecraft", AFAPL-TR-68-67, Part HI.

"Sutton, G.P., "Rocket Propulsion Elements". 6th ed.,Wiley, 1992.

"Sellers, J.J., Meerman, M., Paul, M., and Sweeting, M., "ALow-Cost Propulsion Option for Small Satellites", Journal ofthe British Interplanetary Society, Vol. 48, pp. 129-138, 1995.

29Heister, S., "Solid Rocket Motors", in "Space PropulsionAnalysis and Design", edited by Humble, R.W., Henry, G.N.,and Larson, W.J., Space Technology Series, McGraw-Hill,1995, pp. 295.

30McGrath, D., "The History of Thiokol STAR™ MotorMissions", AIAA Paper 95-3129, San Diego, CA, July 1995.

"McGrath, D., "Motors with Movable Nozzles", AIAA 95-3020, San Diego, CA, July 1995.

32Thiokol Space Motors Catalogue."Atlantic Research Solid Motor Catalogue."Todd, M., Personal Comm., Pacific Research, Oct. 13,

1995."Altman, D. and Humble, R., "Hybrid Rocket Propulsion

Systems", in "Space Propulsion Analysis and Design", editedby Humble, R.W., Henry, G.N., and Larson, W.J., SpaceTechnology Series, McGraw-Hill, 1995, pp. 365.

36Helmy, A.M., "Chronicle Review of the Hybrid RocketCombustion", AIAA Paper 94-2881, Indianapolis, IN, June1994.

"Wernimont, E.J. and Meyer, S.E., "Hydrogen PeroxideHybrid Rocket Engine Performance Investigation", AIAA 94-3147, Indianapolis, IN, June 1994.

38"RITA Ion Thruster Assembly", DASA Company Brochure."Bassner, H., Berg, H.P. and Kukis, R., "The Design of RITA

Electric Propulsion System for Sat 2 (Artemis)", AIAA Paper 90-2539, Orlando, FL, July 1990.

40Beattie, J.R., Williams, J.D., and Robson, R.R., "FlightQualification of an 18-mN Xenon Ion Thruster", IEPC Paper 93-106, Seattle, WA, Sept. 1993.

"'Brophy, J.R, Pless, L.C., Mueller, J., and Anderson, J.R.,"Operating Characteristics of a 15-cm-dia. Ion Engine for SmallPlanetary Spacecraft", IEPC Paper 93-110, Seattle, WA, Sept.1993.

27


42Brophy, J.R, "Stationary Plasma Thruster Evaluation inRussia", JPL Internal Document, JPL Publication 92-4, March15, 1992.

43Khayms, V. and Martinez-Sanchez, M, "Design of aMiniaturized Hall Thruster for Microsatellites", AIAA Paper 96-3291, Lake Buena Vista, FL, July 1996.

"Marcuccio, S., Ciucci, A., Oest, H., Genovese, A., andAndrenucci, M., "Flight Demonstration Opportunities forFEEP", AIAA Paper 96-2724, Lake Buena Vista, FL, July 1996.

45Mitterauer, J., "Miniaturized Liquid Metal Ion Sources(MILMS)", IEEE Transactions on Plasma Science, Vol. 19,No.5, Oct. 1991, pp. 790-799.

46Mitterauer, J., "Prospects of Liquid Metal Ion Thrusters forElectric Propulsion", IEPC Paper 91-105, Viareggio, Italy, Oct.1991.

47Genovese, A., Marcuccio, S., Petracchi, P., and Andrenucci,M., "Neutralization Tests of a mN FEEP Thruster", AIAA Paper96-2725, Lake Buena Vista, FL, July 1996.

"""Micro Thrust FEEP Systems", Centrospazio CompanyBrochure.

49Petagna, C., von Rhoden, H., Bartoli, C., Valentian, D.,"Field Emission Electric Propulsion (FEEP): ExperimentalInvestigation on Continuous and Pulsed Modes of Operation",IEPC Paper 88-127, 1988.

50Blandino, J.J., Pers. Comm., June 26, 1997.51Perel, J., Bates, T., Mahoney, J., Moore, R.D., and Yahiku,

A.Y., "Research on Charged Particle Bipolar Thrustor", AIAAPaper 67-728, Colorado Springs, CO, Sept. 1967.

52Hubermann, M.N. and Rosen, S.G., "Advanced High-ThrustColloid Sources", J.Spacecraft, Vol. 11, No. 7, July 1974, pp.475 - 480.

"Strand, L., Toews, H., Schwartz, K., and Milewski, R.,"Extended Duty Cycle Testing of Spacecraft PropulsionMiniaturized Components", AIAA Paper 95-2810, San Diego,CA, July 1995.

"Nakazono, B., "Second-Generation MicrospacecraftPropulsion Analysis and Alternatives", JPL Internal Document,JPL-D-14038, Nov. 8, 1996.

"Morash, D.H. and Strand, L., "Miniature PropulsionComponents for the Pluto Fast Flyby Spacecraft", AIAA Paper94-3374, Indianapolis, IN, June 1994.

56Bzibziak, R., "Miniature Cold Gas Thrusters, AIAA Paper92-3256, Nashville, TN, July 1992.

"Barber, H.E., Falkenstein, G.L., Buell, C.A., and Gurnitz,R.N., "Microthrusters Employing Catalytically Reacted N2-O2-H2 Gas Mixtures, Tridyne", J.Spacecraft and Rockets, Vol. 8,No. 2, February 1971, pp. I l l - 116.

58Mueller, J.M. and Mcfarlane, J.S., "Design of TridynePressurization Systems for Liquid Oxygen Polybutadiene HybridRocket Motors", AIAA Paper 91-2406, Sacramento, CA, June1991.

59 "Pulsed Plasma Thruster", Primex Company Information."Lincoln Composites Company Information.61Miksch, R.S. and Heller, K.G., "Design and Development of

a Vaporjet Attitude-Control System for Space Vehicles", ATL-D-648, Contract AF 33(616)-7044, Wright-Patterson AFB, Dec.1961.

62Tinling, B.E., "Measured Steady-State Performance of WaterVapor Jets for Use in Space Vehicle Attitude Control Systems",NASA TN-D-1302, NASA, Washington, May 1962.

63Kanning, G., "Measured Performance of Water Vapor Jetsfor Space Vehicle Attitude Control Systems", NASA TN D-3561,NASA, Washington, 1966.

"Mueller, J., Leifer, S.D., Muller, L, and George, T., "LiquidVaporizing Thrusters", NASA Tech Brief, Vol. 21, No. 5, 1997,pp. 73.

"Mueller, J., Tang, W., Wallace, A., Li, W., Bame, D.,Chakraborty, I., and Lawton, R., "Design, Analysis andFabrication of a Vaporizing Liquid Micro-Thruster, AIAA Paper97-3054, Seattle, WA, July 1997.

66"Subliming Solid Control Rocket", Final Report, NASACR-711, Contract NAS 5-3599, Rocket Research Corp., June1965.

""Development of the Subliming Solid Control Rocket PhaseII", NASA CR-712, Contract NAS 5-9070, Rocket ResearchCorp., Seattle WA, March 1967.

68"Application of the Valveless Subliming Control Rocket tothe NRL Gravity Gradient Satellite", RRC-66-R-59, ContractNonr-5123(00)(X), Naval Research Laboratories; May 4, 1966.

"Hardt, A.P., Foley, W.M., and Brandon, R.L., "TheChemistry of Subliming Solids for Micro Thrust Engines",Astronautica Acta, Vol. 11, No. 5, 1965, pp. 340 - 347.

70Owens, W.L., Jr., "Design Aspects of Subliming SolidReaction Control Systems", AIAA Paper 68-516, Atlantic City,NJ, June 1968.

"Kindsvater, H.M., "Simplified Space Mechanisms UsingSublimng Solids", Proceedings of the First AerospaceMechanisms Symposium, University of Santa Clara, SantaClara, CA, May 19-20, 1966.

"Owens, W.L., "An Experimental Study of SuperheatedSubliming Solid Thruster Performance", AIAA Paper 70-210,8th Aerospace Sciences Meeting, Jan. 1970.

"Hardt, A.P., Foley, W.M., and Brandon, R.L., "TheChemistry of Subliming Solids for Micro-Thrust Engines",AIAA Paper 65-595, Colorado Springs, CO, June 1965.

74Kindsvater, H.M., "Design Criteria for Subliming SolidApplications", ICPRG/AIAA Solid Propulsion Conference,Anaheim, CA, June 6-8, 1967.

"Forsythe, R.W., "Impulse and Thrust Stand Test of aSubliming Solid Micropropulsion System", NASA TN D-3245,NASA Goddard Space Flight Center, March 1966.

76Federline, F., "Development of the Subliming Solid ControlRocket - Contributions from the Explorer XXXV (AIMP-E)Satellite Program", Document X-723-68-227, NASA GoddardSpace Flight Center, June 1968.

77Greer, H. and Griep, D.J., "Dynamic Performance of aSubliming Solid Reaction Jet", Report No. TR-1001 (2230-33)-I, Aerospace Industries, Dec. 1966.

78Berkopic, F.D., "Performance of Two Subliming SolidPropellant Thruster Systems for Attitude Control of Spacecraft",NASA TN D-3841, Lewis Research Center, 1967.

79Maycock, N. and Pai Vemeker, V.R., "A PhotochemicalMicrorocket for Attitude Control", J.Spacecraft and Rockets,Vol. 6, No. 3, March 1969.

28


80Mueller, J, Muller, L., and George, T., "Subliming SolidMicro-Thruster for Microspacecraft", New Technology ReportNPO-19926/9525, JPL Technology Utilization Office, April22, 1996.

"'Brophy, J.R., Polk, I.E., Blandino, J., and Mueller, J.,"Micromachined Ion Accelerators", NASA Tech Briefs, Vol. 20,No. 2, February 1996, pp. 36-37.

"Hastings, D, Progress Report on the MIT MicrothrusterContract, Jan. 1997.

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