[American Institute of Aeronautics and Astronautics 37th Joint Propulsion Conference and Exhibit -...

(c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

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A01-34529

AIAA 2001-3899EVALUATION OF PULSED PLASMATHRUSTER ELECTRICAL COMPONENTSEric J. PencilNASA Glenn Research CenterCleveland, OH

Lynn A. ArringtonQSS Group, Inc.Cleveland, OH

Travis J. CarterUnison Industries, Inc.Jacksonville, FL

37th Joint Propulsion Conference & Exhibit8-11 July 2001

Salt Lake City, UtahFor permission to copy or to republish, contact the copyright owner named on the first page.

For AIAA-held copyright, write to AIAA Permissions Department1801 Alexander Bell Drive, Suite 500, Reston, VA, 20191-4344.


AIAA 2001-3899

EVALUATION OF PULSED PLASMA THRUSTER ELECTRICAL COMPONENTS

Eric J. Pencil*NASA Glenn Research Center

Cleveland, Ohio

Lynn A. Arrington*QSS Group, Inc.Cleveland, Ohio

Travis J. CarterUnison Industries, Inc.Jacksonville, Florida

Abstract

The pulsed plasma thruster electrical components evaluated under this test program include abreadboard power processor unit (PPU), energy storage units (ESUs) or capacitors, discharge initiation(DI) electronics and sparkplugs. The main objectives of the breadboard test program are to characterizethe efficiency of the PPU main charge circuit and the ESU, to validate the ability of a single PPU tooperate multiple ESUs with multiple thruster sites, and to validate the effectiveness of the 'micropulsing'concept as a way to provide throttling and a reduction in system mass. Under an on-going componentevaluation program, tests with breadboard electrical components have been completed to demonstratebenchtop functionality, to quantify electrical component performance, to demonstrate vacuum functionality,and to measure thrust throttling capabilities. During the benchtop functionality tests, the PPU wassuccessfully operated over a range of input voltages and output ESU charge voltages. The maximum peakinput current to the PPU was less than 9.8 amperes for all conditions tested. The PPU conversionefficiencies ranged from 83 to 87 percent, while the PPU charge efficiencies ranged from 80.1 to 83.7percent for a nominal input voltage of 28 volts. The equivalent series resistance (ESR) for all ESUs(mica-foil capacitors) ranged from 5.0 to 11.4 milliohms with an average ESR of 8.75 milliohms. The PPU, ESU,and DI were successfully operated with two PPT thruster configurations over a nominal range of ESUenergies and pulse repetition rates. The constant thrust-to-power ratios, demonstrated for all PPTconfigurations, were ideal for micropulsing over a range of operating conditions.

Introduction

Pulsed plasma thrusters (PPTs) are lowpower electric propulsion devices, which havethe unique features of low impulse bit operationcombined with a high specific impulse. Thesolid fuel Teflon propellant eliminates thecomplexities of leaky valves or pressurizedvessels inherent with gas-fed systems. Thepropellant feed system is inherently simple andreliable due to its single moving mechanical part,a spring. The inert propellant simplifies storageand handling requirements during integrationwith the spacecraft. The technology has aheritage from applications stemming back to the1960's [Ref. 1]. The applications ranged fromcontrol propulsion for larger satellites to primarypropulsion for small satellites. PPT featurescombine to offer a favorable solution to NASA'snear-term and long-term deep spaceinterferometry propulsion requirements [Ref. 2,3].

Interest in pulsed plasma thrustertechnology was rekindled in the mid-1990s with

the development of the Earth Observing-1 (EO-1) PPT flight experiment [Ref. 4]. The projectwas a focused development effort, whichincorporated advances in power electronics andmaterials into existing thruster system designs.The EO-1 PPT was designed to replace thefunction of the pitch-axis control system.However, the potential use of PPTs forformation-flying applications require longer-lifecomponents and lower system mass to meetmission needs. In anticipation of future missionrequirements, NASA Glenn initiated acomponent development program to investigatelonger-life, lower-mass components, which canbe incorporated into multi-axis thrusterconfigurations.

Unison Industries, in collaboration withC.U. Aerospace and General Dynamics(formerly Primex Aerospace) designed,fabricated, and delivered breadboard electricalcomponents under contract to NASA Glenn.The electrical components delivered under thisprogram include breadboard PPUs, ESUs(capacitors), DI electronics, and spark plugs.

* Senior Member, AIAACopyright © by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title17, U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein forGovernmental Purposes. All other rights are reserved by the copyright owner.


The power processing architecture was designedwith a centralized PPU with redundantconverters and high voltage switching to drivemultiple thruster modules with multiple thrustaxes at each module. The ESUs selected werethe mica-foil capacitors, which offer a low-risk,long-life technology, which can be readilyintegrated in a system design. The DI electronicswere designed to be distributed on the spacecraftnear a thruster location. The spark plugs havebeen designed to offer a lower-mass, longer-lifesolution as compared to the state-of-the-art sparkplug designs used previously. Details of thetrade studies and designs of these componentscan be found in Ref. 5.

An on-going test program was initiatedto evaluate the electrical components for theirapplication to pulsed plasma thruster designs.Individual tests were to verify functionality ofthe components at both atmospheric and vacuumconditions. Atmospheric tests quantified powerprocessor unit and energy storage unitefficiencies and characterized equivalent seriesresistance. Vacuum tests were conducted tovalidate operational functionality with multiplePPTs and to measure PPT thrust throttlingcapabilities. Future tests will evaluateelectromagnetic interference, conduct thermalcycling, quantify thermal parameters, anddemonstrate component life. At the time ofpublication, the electromagnetic interference,thermal cycling, thermal parameter analysis andlife tests were not completed and will be reportedat a later time.

Breadboard Electrical Component Hardware

A brief description of the advancedelectrical components is provided in thefollowing section. Detailed informationregarding the design requirements and theresulting component designs can be found inRef. 5.

Power Processor Units (PPUs)The PPU configuration used in

atmospheric tests was operated with a personnelcomputer, which simulated the spacecraftcontrols. It is configured by the microprocessorcontrol scheme and high-voltage switchingmatrix to operate two separate ESUs and DIelectronics modules. The input power is 28 +/- 6volts. The PPU operated in a 'micropulsing'mode, in which both the main discharge energyand the pulse repetition rate were varied toobtain a 2 to 50 watt throttle range for a single

ESU/PPT combination. The main dischargeenergy could be set from 2 to 10 joules, whilethe pulse repetition rate could be set from 1 to 5hertz. This feature allowed for the wideoperation range while minimizing requiredcapacitor size. As a result of these features, thedual channel unit is able to operate two dischargeinitiation units on two separate pulsed plasmathrusters and was used to demonstrate high-voltage switching concepts.

A single-channel PPU was designed tocontain one converter, a single energy storageunit output channel, and a single dischargeinitiation circuit. The unit was designed to bevacuum compatible and was principally used forthrust performance tests.

Energy Storage Unit (ESU)The energy storage unit design was a

2.6 microfarad, mica-foil capacitor. Eachcapacitor was assembled from four 0.65 mica-foil capacitor sections, which were joined in asingle block. All five capacitors were rated for3375 volts (but only operated to 2770 volts),which enabled a range of main dischargeenergies from 2 to 10 joules. With the solidmaterials design, these capacitors may allow fora wider range of thermal operations than state-of-the-art oil-filled capacitor designs.

Discharge Initiation ElectronicsEach DI electronics unit was connected

by coaxial cables to spark plugs at threeindividual thruster sites. The units are located inclose proximity to the thruster inside the vacuumchamber to minimize the potential ofelectromagnetic interference from the initiationof the spark. Each module contained a 33microfarad capacitor which allowed forindependent control of the spark plug voltage.

Spark PlugsThe spark plugs were a coaxial,

semiconductor design. The spark plugs wereelectrically isolated from the PPT electrodes viaa mechanical gap or insulator materials in orderto increase the longevity of the spark plugelectrodes. The return lead of the DI electronicswas electrically coupled to the return lead of theESU via a resistor, which will be referred to as acoupling component.

Coaxial PPT ConfigurationA six-nozzle coaxial thruster

configuration was provided by C.U. Aerospaceunder a Small Business Innovative Research

American Institute of Aeronautics and Astronautics2


Phase I contract. This type of thruster provides ahigher impedance discharge characteristic, whichresults in an electrothermally-dominated thrusteroperation. The design electrically connected sixthruster nozzles to one ESU. In this particularcoaxial design, the PPT has a center anode foreach nozzle and a common cathode at the end ofeach nozzle. The spark plug is imbedded in theTeflon propellant nozzle wall.

Parallel Plate PPT ConfigurationA two-channel parallel plate thruster

configuration was fabricated similar to an EO-1configuration for baseline thruster comparisons.This configuration provides a lower impedancedischarge characteristic. Both thruster nozzlesare electrically attached to one ESU. Eachthruster nozzle had a single spark plug that wasisolated from the main cathode electrode by avacuum gap.

Test Description

Benchtop Functionality and PerformanceEvaluation

These tests were conducted atatmospheric conditions for all components. Adual channel PPU was connected to two ESUsthat could be discharged into a triggerableresistive load. The PPU was also connected totwo DI modules with three spark plugs on eachmodule. The spark plug firing initiated thedischarge of the ESUs into the resistive loads asshown in Figure 1. The system was functionallychecked over all operating conditions to verifydesign capabilities. In conjunction with thesetests, measurements were made to characterizethe electrical efficiency of the PPU for ESUcharging, and energy transfer to the DI. ThePPU efficiency relates the operating performanceof a given ESU/thruster combination to the totalsystem input requirements. This information isessential for thermal design of thePPU/spacecraft interface. Current and voltagewaveforms were recorded with a digitaloscilloscope using current and voltage probes.Specifically, the PPU input voltage and current,the PPU output voltage to an ESU, PPU DIcircuit charge voltage, and DI voltage weremeasured to compute the PPU charge and PPUconversion efficiencies. Parameters weremeasured over a range of PPU input voltages,which varied from 22 to 34 volts.

The equivalent series resistance (ESR)or electrical efficiency of the ESU was evaluatedto determine energy sinks in evaluating system

performance and thermal loads. A frequencyresponse analyzer was used to measureimpedance as a function of ESU frequency. Thisis done by measuring the attenuation of asinusoidal signal injected to the ESU for a rangeof frequencies. The results showed a minima atthe resonant frequency of the ESU. At thisresonant point the capacitance and the equivalentseries inductance (ESL) cancel each otherleaving just the value for the ESR. Themeasurement was repeated several times toconfirm consistency.

Vacuum Functionality TestsComponents were installed in a 1.52-

meter diameter by 3.05-meter tall, cryopumpedvacuum facility, shown in Figure 2. The facilitybase pressure was typically 0.1 milliPascal (1microtorr). The vacuum functionality tests werethe first operation of the PPU with a plasmathruster. The PPU features verified in thisportion of the test include simultaneous firing oftwo thrusters, switching between two ESUs,switching between two converters, operationthrough the full range of micropulsing, andoperation of five separate nozzles via fiveseparate spark plugs. Three spark plugs werelocated in coaxial nozzles and the remaining twowere placed in both parallel plate nozzles.Operation was verified via discharge current andvoltage measurements and via observation of thePPT discharge through an optical port in thevacuum chamber.

Thrust MeasurementA torsional-type thrust stand, located in

a 1.5-meter diameter by 4.5-meter long oil-diffusion pumped facility, shown in Figure 3,was used to perform thrust measurements. Thefacility base pressure was typically 0.1milliPascal (1 microtorr). The thrust stand canbe used to determine impulse bit and thrust as afunction of thrust stand deflection, springstiffness, and natural frequency. In-situcalibration weights were used to apply a knownforce to determine the deflection of the thruststand. Both steady-state and single pulseoperation can be used to determine steady-statethrust and impulse bits. Additional details forthe thrust stand can be found in Ref. 6. Both thecoaxial and parallel plate PPT configurationswere tested. The thrusters were operated over arange of discharge energies and pulse repetitionrates to examine the variation in thrust to poweras a metric to evaluate the effectiveness of



micropulsing. The thrust levels determined wereused to calculate the impulse bit.

Results and Discussion

Benchtop Functionality and PerformanceEvaluation

The PPU input current, PPU inputvoltage, and ESU charge voltage were collectedacross the range of ESU charge voltages. Figure4 shows the waveforms for an ESU energy levelof 2 joules. The input voltage is a nominal 28volts, but drops to 27 volts during ESU chargingdue to limitations of the laboratory input powersupply. The PPU input current oscillates around7.5 amperes, before returning to zero after 14milliseconds. The ESU voltage ramps up to thepeak 1200 volts, before discharging through atriggerable resistive load. Figure 4 displays thesuccessful functioning of the PPU at atmosphericconditions for 2 joules ESU operation.

Figure 5 shows the same waveforms foran ESU energy level of 10 joules. The inputvoltage is a nominal 28 volts, but drops to 27volts during the 80 millisecond charging cycle.The input current oscillates around 8.0 amperesfor 16 milliseconds, before dropping to 6.8amperes for 20 milliseconds and then drops to 5amperes for the remaining 40 milliseconds. Thecharge voltage ramps to a peak of 2770 volts,before discharging through the same triggerableresistive load. Figure 5 displays the successfulfunctioning of the PPU at atmospheric conditionsfor 10 joules ESU operation.

In addition, stand-by currents andcharge currents at 2 joules and 10 joules weremeasured for a range of PPU input voltages.These parameters are vital to determine if anyfuture spacecraft interface requirements areviolated. The results are tabulated below:

Stand-bycurrent

Chargecurrent(10J)

Chargecurrent(2J)

RMS

Max

RMS

Max

RMS

Max

Input22V

85mA-

4.28 A

9.76 A

2.83 A

9.76 A

Input28V

70mA

78mA

5.23 A

8.64 A

2.51 A

8.16 A

Input34V

62mA

70mA

4.32 A

6.96 A

2.12 A

6.88 A

The stand-by current decreases as the inputvoltage increases, partially due to the fact thatthe input power is nearly constant, so a higherinput voltage requires a lower input current tomaintain the same input power.

The PPU conversion efficiency wascalculated by computing the ratio of ESU and DIinput energy to the PPU input energy as shownin the following equation:

J__ 2 -0)

The PPU conversion efficiency should beinterpreted as a performance metric of the PPUtopology to deliver the required input energy toboth the ESU and DI. It should not beinterpreted as a component efficiency used torelate input energy from the spacecraft to thethruster's directed kinetic energy. The PPUcharge efficiency was calculated from the ratioof ESU input energy to PPU input energy asshown in the following equation:

I PPU Charge (2)

The PPU charge efficiency combined with thePPT thrust efficiency can be used in relatingavailable spacecraft energy to directed kineticenergy (thrust energy). It should not beinterpreted as a performance metric of the PPU.The PPU conversion efficiency and the PPUcharge efficiency were calculated for a range ofinput voltages at 10 joules and 2 joules. Theresults for both efficiencies are tabulated below:

InputVoltage(Volts)222834222834

ESUEnergy(Joules)101010222

PPUConversionEfficiency0.8170.8300.8510.8360.8700.871

PPUChargeEfficiency0.7890.8010.8220.8050.8370.840

The PPU conversion efficiency is about 3percent larger than the PPU charge efficiency,which represents the energy supplied to the DI.For a given ESU energy, both efficiencies



increase as the input voltage to the PPUincreases. For a given input voltage, bothefficiencies increase as the ESU energy (or ESUvoltage) decreases.

The ESR and resonance frequencieswere measured and tabulated in Table I. ESRranged from 5.0 to 11.4 milliohms for all fivecapacitors. The individual capacitors hadaverage ESRs of 5.51, 8.35, 8.61, 10.24, and11.02 milliohms, respectively. Themeasurement-to-measurement variation wassmall given that ESR varied by only 0.7milliohms, which represents a maximumvariation of 14 percent. The average ESR for theset of five capacitors is 8.75 +/-1.98 milliohms.

Vacuum Functionality TestsFigure 6 displays the ESU current and

ESU voltage (Channels 1 and 2, respectively) aswell as the DI current and DI voltage (Channels3 and 4, respectively) measured in order todemonstrate the functionality of the coaxialPPT/ESU configuration with the breadboardPPU. The DI and ESU are fully charged when atrigger is commanded to discharge the DI. Aspark is formed as indicated by the ramping DIcurrent and collapsing DI voltage. In this figure,the spark is formed about 8 microseconds fromthe display beginning. The DI proceeds tooperate via an oscillatory discharge for twoadditional microseconds. Once sufficient plasmais generated to fill the gap between the PPTelectrodes, the main PPT discharge occurs asevident by the ramping ESU current andcollapsing ESU voltage. Figure 6 shows that themain PPT discharge occurs approximately 1.5microseconds after DI spark initiation. The PPTproceeds to discharge through an oscillatorywaveform for an additional 6-8 microseconds.Voltage reversal through the ESU was about 33percent of the original charge voltage. Figure 6displays the successful operation of thebreadboard PPU, ESU, and DI with the coaxialPPT.

Figure 7 displays the ESU current andESU voltage (Channels 1 and 2, respectively) aswell as the DI voltage (Channel 4) measured inorder to demonstrate the functionality of theparallel plate PPT/ESU configuration with thebreadboard PPU. Channel 3 recorded a voltagepotential on a return lead. The DI and ESU arefully charged when a trigger is commanded todischarge the DI. A spark is formed as the DIvoltage collapses. In this figure, the spark isformed about 4 microseconds after the displaybeginning. The DI proceeds to operate via an

oscillatory discharge for two additionalmicroseconds. Once sufficient plasma isgenerated to fill the gap between the PPTelectrodes, the main PPT discharge occurs asevident by the ramping ESU current andcollapsing ESU voltage. Figure 7 shows themain PPT discharge occurs less than 0.2microseconds after DI spark initiation. The PPTproceeds to discharge through an oscillatorywaveform for an additional 14 microseconds.Voltage reversal through the ESU was about 80percent of the initial charge voltage. Figure 7displays the successful operation of thebreadboard PPU, ESU, and DI with the parallelplate PPT.

In order to demonstrate the feasibility ofreducing the magnitudes of the ESU voltagereversal and oscillations, 'clamping' diodes wereinstalled across the output of the PPU. Figure 8displays the ESU current and ESU voltage(Channels 1 and 2, respectively) as well as theDI voltage (Channel 4) measured in order todemonstrate functionality with the 'clamping'diodes. Channel 3 recorded a voltage potentialon a return lead. The DI and ESU are fullycharged when a trigger is commanded todischarge the DI. In this figure, the breakdowntiming was similar to previous tests. The PPTproceeds to discharge through a reducedoscillatory waveform for an additional 14microseconds. The clamping diodes reduced thevoltage reversal to 21 percent. Figure 8 displaysthe successful operation of the breadboard PPU,ESU, and DI with the parallel plate PPT withreduced oscillation magnitude.

Figure 9 displays the DI charge voltageand the ESU charge voltage for the coaxial PPT(Channels 1 and 2, respectively) and the DIcharge voltage and ESU charge voltage for theparallel plate PPT (Channels 3 and 4,respectively) measured in order to demonstratethe concurrent operation of two ESU/PPTconfigurations using one PPU. First the DI andESU were charged for the coaxial thruster. TheDI and ESU were fully charged after about 10milliseconds and about 80 milliseconds,respectively, when the command signal is sent toinitiate the discharge. Both the DI and ESUcharges are dissipated in the arc discharge in 12-14 microseconds, which is shown as a nearvertical line, given the differences in timescales.The cycle is repeated alternately with the parallelplate PPT and the coaxial PPT. Figure 9displays the successful operation of boththrusters at an ESU energy of 10 joules and apulse repetition rate of 5 hertz.



Thrust MeasurementThe thrust was measured for a range of

ESU energies and pulse repetition rates. Theaverage impulse bits are calculated from the ratioof average thrust to pulse repetition rate. Theresults for the coaxial PPT are listed in Table II.The results for the parallel plate PPT both withand without clamping diodes are listed in TablesIII and IV, respectively. For the range ofconditions tested, the thrust-to-power ratio forthe coaxial PPT was nearly constant at 29.7+7-0.7 microNewton/watt. For the range ofconditions tested the thrust-to-power for theparallel plate PPT was nearly constant at 9.6 +/-0.5 microNewton/watt when the clamping diodeswere installed. For the range of conditionstested, the thrust-to-power for the parallel platePPT was nearly constant at 10.4 +/- 0.4microNewton/watt when the clamping diodeswere removed. The constant thrust-to-powerratios, demonstrated for all PPT configurations,were ideal for micropulsing over a range ofoperating conditions.

Conclusions

Under an on-going componentevaluation program, tests with breadboardelectrical components have been completed todemonstrate benchtop functionality, to quantifyelectrical component performance, todemonstrate vacuum functionality, and tomeasure thrust. During the benchtopfunctionality tests, the PPU was successfullyoperated over a range of input voltages and ESUoutput charge voltages. The maximum peakinput current to the PPU was less than 9.8amperes for all conditions tested. The PPUconversion efficiencies ranged from 83 to 87percent, while the PPU charge efficienciesranged from 80.1 to 83.7 percent for a nominalinput voltage of 28 volts. The ESR for all ESUsranged from 5.0 to 11.4 milliohms with anaverage ESR of 8.75 milliohms. The PPU, ESU,and DI were successfully operated with bothPPT thruster configurations over a nominal rangeof ESU energies and pulse repetition rates. Theconstant thrust-to-power ratios, demonstrated forall PPT configurations, were ideal formicropulsing over a range of operatingconditions.

Future work will include tests tocharacterize the electromagnetic interference

levels of the PPU, ESU, and DI as well as theireffects on the EMI produced by the thruster, ascompared to previous data. Future work willinclude the operation of the ESU over a range oftemperature conditions expected in orbit as wellas characterize thermal parameters for futurethermal analysis. Finally the breadboardcomponent evaluation program will culminate ina life demonstration test of the PPU, ESU, andDI.

Acknowledgements

The authors wish to acknowledge thekey team members involved in this project: LuisPinero, Charles Sarmiento, and Thomas Haag atNASA GRC, John Frus and Dawn Thomas atUnison Industries, W. Andrew Hoskins atGeneral Dynamics, and Rodney Burton, DavidCarroll, Darren King, and Julia Laystrom at C.U.Aerospace. The authors wish to acknowledge allof the dedicated members of NASA GRC's TestInstallations Division, who have contributedgreatly to PPT testing activities.

References

(1) Burton, R.L. and Turchi, P.J., "Pulsed PlasmaThruster," Journal of Propulsion and Power, Vol.14, Num. 5, pp 716-35, Sept.-Oct, 1998.

(2) Deininger, W.D., Weiss, M.A., Wiemer, D.J.,Hoffman, C.N., Cleven, G.C., Patel, K.C.,Linfield, R.P., Livesay, L.L., "Description of theStarLight Mission and Spacecraft Concept,"Paper ID 277, 2001 IEEE AerospaceConference, March 10-17, 2001.

(3) Blandino, J.J., Cassady, R.J., Sankovic, J.M.,"Propulsion Requirements and Options for theNew Millennium Interferometer (DS-3) Mission,AIAA-98-3331, July 1998.

(4) Benson, S.W., Arrington, L.A., Hoskins,W.A, and Meckel, N.J., "Development of a PPTfor the EO-1 Spacecraft," AIAA-99-2276, June1999.

(5) Benson, S, and Frus, J. "Advanced PulsedPlasma Thruster Electrical Components," AIAA-01-3984, July 2001.

(6) Haag, T.W., "PPT Thrust Stand," AIAA-95-2917, also NASA TM-107066, July 1995.



Figure 1 - Breadboard electrical component in atmospheric functionality tests.

Figure 2 - Breadboard electrical components and both PPT configurations in vacuum functionality tests.

Figure 3 - Breadboard components and coaxial PPT in thrust measurement test.



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Ch2 y1 520mA

C1 RMS29.82V

€2 RMS2. SOS A

C2 Max8.16 A

C3 Max1.24kV

4 Dec 200016:48:20

Figure 4 - Representative charge curve for 2 Joule ESU charging performed at atmospheric conditions anddischarged into a resistive load.

[Ch 1: PPU input voltage (lOV/div), Ch 2: PPU input current (2A/div),Ch 3: ESU charge voltage (500V/div)]

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4 Dec 200016:01:27

Figure 5 - Representative charge curve for 10 Joule ESU charging performed at atmospheric conditionsand discharged into a resistive load.

[Ch 1: PPU input voltage (lOV/div), Ch 2: PPU input current (2A/div),Ch 3: ESU charge voltage (500V/div)]



idomv ^ ctii V.ookv M2.oojjis Ch2 SI V.ookv10.0V Ch4 1.00KV

Figure 6 - Representative discharge curve performed at vacuum conditions and discharged into the coaxialPPT.

[Ch 1: ESU current (~10kA/div), Ch 2: ESU voltage (1 kV/div),Ch 3: DI current (lOOA/div), Ch 4: DI voltage (IkV/div)]

Chi iddmv ^ Chi i.obkv1 M 2.dt)iis Ch2 \. 980 vCh3 10.0V HJJH I.OOkV

Figure 7 - Representative discharge curve performed at vacuum conditions and discharged into the parallelplate PPT without clamping diodes.

[Ch 1: ESU current (~10kA/div), Ch 2: ESU voltage (1 kV/div),Ch 3: return lead current (lOV/div), Ch 4: DI voltage (IkV/div)]



Chi icidmv W Ch2 i.ookv M i.'objis 612 "V 980 vCK3 10.0V SB3 LOOkV

Figure 8 - Representative discharge curve performed at vacuum conditions and discharged into the parallelplate PPT with clamping diodes.

[Ch 1: ESU current (~10kA/div), Ch 2: ESU voltage (1 kV/div),Ch 3: return lead current (lOV/div), Ch 4: DI voltage (IkV/div)]

a i.obkv cih2 Y.ookv' M ibbms Ch2 x V.ookvI.QOkV Ch4 KOOkV

Figure 9 - Successful concurrent operation of two separate PPTs at 5 hertz each.[Ch 1: coaxial PPT DI voltage (IkV/div), Ch 2, coaxial PPT ESU voltage (IkV/div),

Ch 3: parallel plate PPT DI voltage (IkV/div), Ch 4: parallel plate PPT ESU voltage (IkV/div)]



Table I - Equivalent Series Resistances and resonance frequencies for ESUs.

ESUS/N 1239

Freq(kHz)459.1

459.1

459.6

ESR(mohm)

5.81

5.656

5.063

ESUS/N 1240

Freq(kHz)479.9

483

482.5

ESR(mohm)8.343

8.49

8.225

ESUS/N 1241

Freq(kHz)379.6

379.6

379.5

ESR(mohm)

8.48

8.483

8.877

ESUS/N 1242

Freq(kHz)379.6

381.5

381.9

ESR(mohm)9.985

10.21

10.52

ESUS/N 1243

Freq(kHz)379.6

380.5

381.2

ESR(mohm)

10.84

10.88

11.35

Table II - Coaxial PPT performance.

Power

(watts)

5040405040303030305040302010403224168

Energy

O'oules)

108101010107.5610101010101088888

Frequency

(hertz)

5545434535432154321

Average Thrust

(micronewtons)

15001190119014161190876885882879152612128886002881224984732468228

AverageImpulse Bit

(micronewtons-seconds)

300238298283298292221176293305303296300288245246244234228

Thrust/Power

(micronewtons/watt)30.029.829.828.329.829.229.529.429.330.530.329.630.028.830.630.830.529.228.5



Table III - Parallel plate PPT performance (with clamping diodes).

Power

(watts)

504040303030202020101010 "55

Energy

(joules)

10108107.5610

6.675105

3.335

2.5

Frequency

(hertz)

54534523 -412312

Average Thrust

(micronewtons)

50340940230529628120419518394.691.585.446.343.8

AverageImpulse Bit

(micronewton-seconds)

10110280.410274.056.210265.045.894.645.828.546.321.9

Thrust/Power

(micronewton/watt)10.110.210.210.29.99.410.29.89.29.59.28.59.38.8

Table IV - Parallel plate PPT performance (without clamping diodes).

Power

(watts)

50404030303020202020101010105

Energy

(joules)

10108107.5610

6.6754105

3.332.55

Frequency

(hertz)

545345234512341

Average Thrust

(micronewtons)

53942844031331930721120819020410810510296.054.0

AverageImpulse Bit

(micronewtons-seconds)

10810788.010479.861.410669.347.540.810852.534.024.054.0

Thrust/Power

(micronewtons/watt)10.810.711.010.410.610.210.610.49.510.210.810.510.29.610.8


Date post:	11-Dec-2016
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[American Institute of Aeronautics and Astronautics 37th Joint Propulsion Conference and Exhibit -...

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