Prepared as AIAA Paper 98-3631 for the 34th Joint Propulsion Conference, July 13 -15,1998,Cleveland, Ohio, U.S.A.

Low Bus Voltage Hydrazine Arcjet System for GeostationarySatellites

Dieter M. Zube*, Dan Fye**PRIMEX Aerospace Company, PO Box 97009, Redmond, WA 98073-9709, U.S.A.

Ideo Masudat

National Space Development Agency of Japan, World Trade Center Bldg. 27F,2-4-1, Hamamatsu-cho, Minatoku, Tokyo 105-60, Japan

Yoshifumi Gotoh*Mitsubishi Electric Corporation, 325, Kamimachiya Kamakura, Kanagawa 247, Japan

AbstractThe capabilities of arcjet propulsion systems were recently extended to accommodate operationon the NASD A Data Relay Test Satellite (DRTS) providing a power bus voltage between 31 and51.5 VDC. This paper summarizes the newly attained qualification status of the MR 509-A/Barcjet system demonstrating the flexibility of the current design. A redesign of the PowerProcessing Unit (PPU) became necessary as well as a delta-qualification of the thruster tovalidate spacecraft integration, and to provide compliance with the DRTS satelliteenvironmental requirements. Two types of thrusters with different thrust levels were madeavailable to meet mission requirements. The delta-qualification included a pyro-shock test,vibration tests to a higher level than previously tested, and performance mapping beyond theoriginal range. Included in the paper is an assessment of the PPU performance characteristicsas well as the discussion of the system operation and system telemetry.

IntroductionThe second generation MR 509 1.8 kWhydrazine arcjet system has been based on theMR 508 system, the first flight qualified1.8 kW arcjet system, by incorporating lessonslearned from the first flight missions. Thedesign upgrade successfully extended thelifetime to a more demanding 13 years on orbitrife. Its qualification history and flightexperience have been described by Smith etal.m. Consequently, the MR 510 2.2 kWhydrazine arcjet has been developed from theMR 509 system. The increase in power andincorporation of new technology enabled thearcjet to operate at higher performance. Formore details on the MR 510 development andqualification history see [2] by Smith et al.

The new 1.8kW MR 509-A/B system wasdeveloped to provide greater versatility andincreased environmental capabilities with aninput voltage range of 31 to 51.5 VDC for useon the Japanese NASDA Data Relay TestSatellite (DRTS)[3J. This paper describes thenew system, and the power processing unit(PPU) as well as thruster related changes thatbecame necessary for this application andproved the maturity and flexibility of arcjetsystems.

Since the thruster design had alreadydemonstrated adequate lifetime andperformance for the DRTS mission, a newcomplete qualification of the thrustercomponents was not necessary. A limiteddelta-qualification effort was required to show

*Principal Development Engineer, arcjet thrusters, member AIAA; **Senior Principal Electrical Engineer,TAssistant Senior Engineer, DRTS Project Team ^Senior Propulsion EngineerCopyright © 1998 by PRIMEX Aerospace Company, NASDA, and MELCO, published by AIAA with permission.

compliance with the environmental conditionsduring integration, launch, and on-orbit. Sincethe thruster integration on the satellite requiredthrusters with two different thrust levels (seechapter "Thruster Performance"), the twothrusters are now designated MR 509-A and-B respectively. A complete arcjet propulsionsystem for one satellite consists of four PPUs,four power cables, two MR 509-A and twoMR 509-B thrusters.

System RequirementsSystem redundancy is established by providingtwo thrusters for the eastern and westernsatellite side panels, each thruster beingconnected to its individual power supply. Areliability of 0.976 over the seven year missionlife is estimated for a single system (onethruster, one PPU) through the use of highreliability parts and already space qualified andproven components. The two redundantthrusters on each satellite side increase theoverall reliability of the arcjet stationkeepingsystem on the DRTS satellite to 0.9994.

No major re-design changes were required forthe thruster, since its previously demonstratedperformance was adequate for the intendedpurpose. Changes were only required to thetemperature sensors to meet the telemetryprovisions of this application.

The thruster arrangement on the DRTSsatellite demanded thrusters with differentthrust levels for the east and the west satelliteside panel. This could easily be achieved byproviding different propellant flow rates to thethrusters. These are established through theuse of differently sized fluid resistors, whichreduce the propellant flow to the thrusters tothe desired flow levels over the feed pressurerange available on the spacecraft.

The PPU for this new bus voltage levelrequired a major re-design. However, theheritage of the proven MR 509 power supplyarchitecture was maintained for theMR 509-A/B PPU.

System Operation and TelemetryA typical arcjet maneuver would be precededby heating the catalyst bed of the hydrazine gasgenerator for about 30 minutes prior to theintended start of the arcjet. This raises the gasgenerator temperature sufficiently above itsminimum operating temperature. Both valveand gas generator heaters are redundant toincrease system reliability.

The actual arcjet firing is initiated by openingthe propellant valve. At the same time thePPU is activated, and the ignition inductors arecharged. After establishing sufficientpropellant flow, the arc is ignited by applying ahigh voltage pulse to the arcjet electrodes.The PPU detects whether the arc is sustainedand then controls the arc at the output level ofnominally 1630 W (see Table 1 and 2 foradditional system parameters).

The thruster had earlier been qualified formore than 1050 hours of operation with morethan 1170 individual starts, most of thesecycles consisted of a 1 hour on periodfollowed by a 30 minute cool down phase.Thus, normally recommended burns can last upto 60 minutes. The thruster reaches steadystate thrust and specific impulse levels within2 minutes. The longer the burn period, thehigher the average impulse and the overall fuelefficiency. At the end of the arcjet maneuver,the PPU receives a "STOP" command and thearc is extinguished. The valve is then closed.

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PropellentFeed Pressure (nominal)

demonstratedThrust (nominal)

-A thruster-B thruster

Spec. Impulse, mission average-A-B

Total Impulse per ThrusterPPU Input VoltagePPU Input Power (nominal)Arcjet Input PowerPPU Efficiency (nominal)Qual. Firing TimeQual. StartsMass:

PPUArcjetPower Cable

Hydrazine14.2 - 17.6 bar13.1-20.7 bar

225 - 246 mN262 - 286 mN

>502 s>470 s

866,500 Ns31-51.5 VDC

1745 W1630 W93.4 %

> 1050 hours>1170

6.20 kg1.03 kg0.40 kg

Table 1: MR 509-A/B System Performancefor the DRTS Satellite

During thruster off-times, only the valveheaters have to be activated thermostatically toprevent propellant freezing in the fuel linesnear the valve, where the lines are exposed toopen space and its low temperatures. Theremainder of the propellant lines are routedinside the spacecraft at sufficient temperatures,or are equipped with line heaters to preventfreezing.

Both analytical and experimental plume studiesof this thruster type show that plume -spacecraft interaction is benign and does notpose a problem for the typical arcjetinstallation on a spacecraft. Information fromall arcjet equipped satellites on orbit confirmthat no arcjet induced system degradation isdetectable.

PPU- auxiliary power

0.5 W stby.10 W aux.

enabled- main power

Arcjetsvalvesvalve heaters

gas generator heaters

total

Non Operating

4 in stby. mode

2.0 W0

0

04 active, 50% dutycycle

22.0 W

024.0 W

Pre-Firing, typically30 minutes long

2 in stby. mode,2 aux. enabled

LOW20.0 W

0

04 active, 50% dutycycle

22.0 Wactive on the 2 thrustersto be fired

8.4 W51.4W

firing(2 thrusters at a time)

2 in stby.

LOW0

2 enabled3544 W

1.5 W (2 open)2 active on the non firingthrusters, 50% duty cycle

ll.OW

03557.5 W

Table 2: Average MR 509-A/B System Power Consumption

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Fig. 1: MR 509-A/B Arcjet Thruster

To determine the system status, the followingtelemetry signals are available from the arcjetsystem:- PPU input voltage- PPU output voltage- PPU output current- 2 PPU temperatures- PPU error flag- gas generator temperature- valve temperatureThese are either transmitted to the groundcontrol station or are used by the on-boardcontrol computer to monitor the arcjet systemhealth status, and to initiate commands duringnormal operation or anomalies.

Hydrazine propellant systems tend to generateand accumulate gas entrapments in the liquidfuel while on orbit. The passage of such gasbubbles briefly changes the fuel flow rate tothe arcjet thruster, causing temporary arcinstabilities. To cope with these problems, thePPUs of the MR 509 and MR 510 series havethe capability to handle such disturbances andto maintain safe arcjet operation during suchevents[1]. Experience with over 40 arcjetsystems has proven the validity of this

approach. The same approach has beenincorporated into the MR 509-A/B PPU andhas successfully been tested during the PPUqualification effort.

Thruster PerformanceThe arcjet system is designed for a range ofpropellant feed pressures. For the MR 509A/B(Fig. 1) application, the nominal feed pressurerange is 17.6 to 14.2 bar. This is within thepreviously established operating range.However, to comply with contingencyconditions, thruster operation had to bedemonstrated for a range from 20.7 to13.1 bar. This was successfully done duringthe arcjet and PPU qualification tests.

A performance model for feed pressures from13.1 to 18.6 bar (190 to 270 psi) was derivedfrom previous test results (Fig. 2), and wasvalidated with the test results from theMR 509-A/B qualification effort.

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referred to as the MR 509-A (standard thrust65

is 60

(A OU(An£ 45

MF 5O9-R

MR

12 14 16 18 20

300

280zE 260tfS 240

220

200

MR539-B

12 14 16 18 20

14 16 18

feed pressure, bar20

Fig. 2: Performance of the MR 509-A and -BThrusters (25°C Propellant Temperature)

The arcjet installation on the DRTS spacecraftrequired thrusters with two different thrustlevels for the thrusters on the east and westside panels. Under nominal conditions (fuelpressure 17.3 bar - 250 psi and fueltemperature 25°C), a 15% higher thrust wasrequested. This demand could be met byvarying the mass flow through the arcjetthrusters. Subsequently, the thrusters are

level)units.

and MR509-B (higher thrust level)

In the MR 509/510 arcjet family, the mass flowis controlled through the use of a fluid resistorwhich reduces the feed pressure in the gasgenerator and at the same time limits the massflow to the desired range.

To accommodate the higher thrust level, afluid resistor with a lower resistance, thusallowing a larger mass flow, was selected. Therequired mass flow was first estimated basedon the existing performance model, and the"equivalent feed pressure" for the MR 509-Athruster was calculated. During theMR509-A/B qualification tests, the -Athruster was performance mapped at feedpressures up to 31.1 bar, which isrepresentative for the 20.7 bar feed pressure ofthe -B thruster. These tests established apreliminary performance model for the -Bconfiguration as well as selection of theappropriate fluid resistor size for thesethrusters.

ParameterNominal FeedPressure (bar)

Mass Flow(mg/s)

Thrust (mN)

SpecificImpulse (sec)

MR 509-A

14.2 - 17.6

44-50

225 - 246

522-497

MR 509-B

14.2 - 17.6

54-62

262 - 286

492 - 462

Table 3: Performance Comparison ofMR 509-A and -B Thrusters at 25°C Fuel

Temperature

Parallel to the thruster test at the new thrustlevel, the PPU was tested with a thruster at thehigher mass flow levels to demonstrate thePPU capability to reliably start and operate the

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thruster under these conditions. A comparisonof the thruster performance for the -A and -Bthruster is given in Table 3 and Fig. 2.

Thruster QualificationAs already stated, the previously achievedqualification status of the arcjet thruster^ was,with minor exceptions, adequate for the DRTSmission. The system requirements for totalpropellant throughput and the anticipatednumber of cycles will be met by the qualified1170 starts and the total impulse capability of866,500 Ns per thruster. This is equivalent to1050 hours of arcjet operation, or athroughput of roughly 175kg of hydrazinethrough one thruster of the MR509-Aconfiguration.

The MR 509-B thruster was not individuallylifetime qualified. Previous studies had shownthat arcjet thruster life is limited by the numberof starts as well as the constrictor closurephenomenon [2>4]. The latter is caused bythermal stresses in the constrictor area. Whilethe -B thruster will have the same number ofstarts than the -A thruster, the thermal loadson the constrictor region will be smaller sincethe -B thruster will have a higher mass flow.As a result, the thruster will experience lowerthermal stresses in the constrictor region.Consequently, total lifetime expectancy for the-B thruster will even be higher than for the -Athruster, and a new lifetime test was notdeemed necessary.

A qualification thruster was, however,subjected to an additional delta qualification todemonstrate its compliance with previously nottested environmental conditions. The launchenvironment will expose the thruster to highervibrational loads. Additionally, the systemapproach calls for a "wet"-launch, i.e.propellant will be present at the propellantcontrol valve mounted on the thruster. Thus,

the thruster was tested for leakage at loadlevels 3 dB above the anticipated vibrationalloads during launch. The test was completedsuccessfully, no leakage occurred at theselevels.

During the deployment of the solar arrays onthe spacecraft, the thruster will be exposed tothe sudden shock loads. To demonstratestructural integrity during this event, thethruster was subjected to a shock test.

This shock test, the first ever to be performedon an arcjet thruster, was conducted as apyroshock test, since this was considered to bethe most reliable duplication of the loadsimposed during the solar array deploymentTogether with the thruster, the power cablewas also tested to verify its capability towithstand the shock loads.

After the vibration and shock test toqualification levels, the thruster was subjectedto a standard acceptance test firing. This firingdemonstrated that the thruster had passed thetests without damage to the internal insulatorsand the electrodes, which could otherwise onlyhave been determined by a destructivedisassembly. Furthermore, the test firing wasalso used to extend the validity of the originalperformance model into the mass flow andthrust range of the -B thruster. Due to theidentical design of the -A and the -B thruster,only one thruster test was required.

Thermal AnalysisTo support system integration, an intensivethermal analysis was conducted. A detailedanalytical thermal model was developed topredict thruster temperatures within 15°C fortemperatures generally below 250°C and towithin 30°C for the hotter anode and gasgenerator parts. For all critical components(spacecraft interface areas, valve seats, gas

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generator temperatures), predictedtemperatures were conservative whencompared to test results, thus confirming theadequacy of the design.

This detailed thermal model was run for bothhot and cold biased environments to establishcompliance with the two extremes the thrusterwill experience on orbit. It was also used tocalculate the anticipated heat loads from thethruster to the DRTS spacecraft. The heatloads from a single operating 1.63kW arcjetthruster were found not to exceed 6.5 Wconductive heating and 78 W radiative heating.The temperature of the arcjet mountingstructure at the interface with the satellitevaries with the specific design and is below190 °C for the DRTS application.

The detailed thermal model was also used toassess soak back temperatures after firing onthe valve and the spacecraft structure. Notemperature limits were exceeded.

To allow a fast assessment ofinstallation environments, a simplified thermal

varying

model was created that allows heat-load andtemperature predictions in the critical areasduring both firing and non firing conditions. Itbreaks the thruster down into a 19 nodenetwork with the thermal conductancebetween the nodes adjusted to meet thetemperature predictions of the detailed modelfor steady state conditions.

PPU DesignThe PPU design baseline was the MR 509PPU. It was, however, updated, and modifiedfor a power bus voltage range of 31 to51.5VDC. PPU output characteristicsremained unchanged. The MR 509 arcjetthruster requires high voltage arc startingpulses, controlled power start-upcharacteristics, and constant power. Thethruster voltage can range from -105 V to-140 V while the power is maintained constantat 1630 ±30 W. Arcjet operation outside thisenvelope, especially at lower voltages, ispossible, however with losses in efficiency andcontrol stability.

OUTPUT SIGNALS

INPUT SIGNALSCONSTANTPOWERCONTROLLER

OUTPUT RETN

INPUT POWER

POWER RETN

Fig. 3: MR 509-A/B PPU Block Diagram

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The PPU power train (Fig. 3) consists of anEMI filter, and a DC to DC Switch ModeConverter power control. In addition there isan auxiliary housekeeping supply, a commandinterface, and telemetry signal generation, thetwo latter tailored to the DRTS requirements.

The DC to DC converter is a classic push-pull,transformer coupled pulse width modulatedpower supply. The input power is specified at1886W maximum at an input voltage of31 VDC for all environmental conditions. Thisresults in a maximum input current of 61 A.Consequently, the design objective became oneof minimizing resistance losses from the PPUinput terminals to the transformer output. Acareful voltage stress analysis for everycomponent confirmed the design for all inputvoltages over the large input range.

The EMI conducted emissions requirementsare stringent and require filtration. Theaverage input current, and the resulting RMSswitching current affects nearly all componentsof the power train. The filter characteristicsdetermined the bulk capacitor value. Parallelcapacitors sum to the needed value. Theindividual capacitors are assembled into aseparate module.

The original MR 509 PPU uses powerMOSFETs packaged in custom power hybridcircuits. The new MR 509-A/B PPU utilizespower MOSFETs connected in parallel. Theirhigh count effectively reduces the ONresistance to a low value. Discretecomponents are chosen for economy,availability, and distributed current and heatprofiles. The MOSFETs mount on heat sinks,conducting heat directly to the PPU thermalplate and mounting base.

The transformer primary and secondarywindings contain sufficient copper tape forcurrent density below 107 A/m, with nearly

equal current density in each winding. Heatdeveloped within the transformer conducts tothe thermal plate by paths designed for lowthermal resistance.

Inductors, developed for this application, havelow resistance and high DC current capacitybefore saturating. A latching relay is seriesconnected between the EMI filter input andoutput. This allows the user to apply a powersource with the relay open preventing highinrush currents from charging the bulkcapacitor. Spice analysis of the filter madepossible the design iterations necessary forsolving conflicting requirements. Predictedsmall signal response was confirmed onbreadboard hardware using a networkanalyzer. The measured responsecorresponded very well with the predictions.An engineering development PPU was testedby an independent laboratory to MIL-STD-461; conducted emissions measured below thespecified limits with margin.

The user command interface, while similar tothe heritage MR 509 PPU, was redesigned tocomply with user requirements. Analog anddigital telemetry output signals provided keyoperational parameters. Rather than calibrateeach analog signal to a precision transferfunction, the function is measured accuratelyduring acceptance testing and reported to theuser.

Repackaging the PPU accomplished the goalsof modularization for ease of assembly,distributing heat loads uniformly over thethermal baseplate, and weight optimization.This resulted in a PPU with superiormechanical integrity. The highly modulardesign allows testing on subsystem levels,which is verified by functional testing of thecompleted units.

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PPU PerformanceEfficiency is higher than 93.5% under typicaloperating conditions. The design alsoachieved the goal of 1886W maximum inputpower at the lowest input voltage. Constantpower regulation, at 1630 ±30 W, for linevoltage, output voltage; and temperature worstcase extremes met requirements. Usercommand interfaces and telemetry signalscomplied with customer requirements andfunctioned flawlessly at the system level. PPUtesting with the AJT demonstrated systemcompatibility under all operating conditions.

Reliability levels are extremely high, achievedby selection of NASA approved components.Component operating stresses are withinconservative (MIL-STD-975) deratingguidelines. Radiation hardness is assured byselecting radiation hard components.

Thermal imaging of internal electroniccomponents during the testing of theengineering unit demonstrated uniform heatdistribution. All temperatures are well belowcomponent temperature ratings, thusincreasing reliability.

The unit weight is approximately 6.2 kg, anincrease over the MR 509 design. This can beexplained by the weight of added copper tomeet the efficiency goal for all input voltages.Mechanical design integrity was validated byvibration testing and finite element predictionsfor shock loads.

PPU QualificationThe general concept of the MR 509-A/B PPUhad been tested with an EDM-unit, which ledto the qualification model. This unit, which isrepresentative for the flight models, wassubjected to a comprehensive qualification test,which it passed successfully. The qualificationtest included (among others):

- corona tests,- EMI tests,- sinusoidal and random vibration tests,- thermal vacuum cycling,- a high temperature burn in,- firing into an arcjet thruster,and a- 240 hour burn in.These tests were separated by electricalfunctional tests, which verified the unit hadsuccessfully passed the previous test sequence.

The EMI test was performed to the MIL-STD-461 guidelines for:- conducted emissions,

methods CEO 1.CE03;- radiated emissions,

methods RE01.RE02;- conducted susceptibility,

methods CS01.CS02;- radiated susceptibility,

methods RS02, RS03;and- spike susceptibility,

method CS06.Sufficient margin in all these areas wasdemonstrated.

The firing test verified the MR 509-A/B PPUcapability to reliably operate both the -A aswell as the -B thruster. Included was ademonstration to successfully cope with gasingestion and the resulting temporary brief arcinstability.

SummaryThrough the development of the newMR 509-A/B arcjet system, the range ofsatellite buses for which arcjet thrusters arenow qualified, could be greatly extended. Thisrequired design and qualification of a powerprocessing unit for an input voltage of 31 to51.5VDC and additional environmental teststo the thruster design.

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The MR509-A/B system showed the greatflexibility of the arcjet design to accommodatedifferent thrust levels, and increasedenvironmental requirements. The change inthrust level could be met by changing the flowcontrolling fluid resistor. While the higherthrust unit has a lower specific impulse, it issufficiently high to make hydrazine arcjetsattractive for unique north-south stationkeeping applications. This example showedthat the current line of arcjets can easily betailored to meet mission requirements beyondthe original performance range.

[4] Lichon, P.O., Sankovic, J.M.: "Development andDemonstration of a 600 s Mission AverageArcjet", 23rd International Electric PropulsionConference Paper 93-087, Seattle, Washington,September 1993

AcknowledgmentsThe authors wish to acknowledge theexceptional efforts of Mr. N. Shikagawa, andMr. S. Yamazaki and their colleagues ofMitsubishi Electric Corporation, and of Mr. Y.Fujiwara and Mr. Y. Sudo from NASD A'sDRTS project team who provided engineeringand program support for the use of theMR509-A/B arcjet system on NASD A'sDRTS spacecraft.

References[1] Smith, R.D., Roberts, C.R., Aadland, R.S.,

Lichtin, D.A., Davies, K.: "Right Qualification ofthe 1.8kW MR-509 Hydrazine Arcjet System",25th International Electric Propulsion ConferencePaper IEPC 97-081, Cleveland, Ohio, August1997

[2] Smith, R.D., Aadland, R.S., Roberts, C.R.,Lichtin, D.A.: "Flight Qualification of the 2.2 kWMR-510 Hydrazine Arcjet System", 25thInternational Electric Propulsion ConferencePaper ffiPC 97-082, Cleveland, Ohio, August1997

[3] Hotta, N., Ichino, H., Toriyama, K., Kosugi, S.,Kanamori Y.: "Development of SystemEngineering Model for Data Relay Test Satellite"17th AIAA International CommunicationsSatellite Systems Conference Paper AIAA-98-1219, Yokohama, Japan, February 1998

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