NASA/TMw97-206305 IEPC-97-127

A Performance Comparison of Pulsed Plasma

Thruster Electrode Configurations

Lynn A. Arrington

NYMA, Inc., Brook Park, Ohio

Tom W. Haag and Eric J. Pencil

Lewis Research Center, Cleveland, Ohio

Nicole J. Meckel

Primex Aerospace, Inc., Seattle, Washington

Prepared for the

25th International Electric Propulsion Conference

sponsored by the Electric Rocket Propulsion Society

Cleveland, Ohio, August 24-28, 1997

National Aeronautics and

Space Administration

Lewis Research Center

December 1997

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A PerformanceComparisonofPulsedPlasmaThrusterElectrodeConfigurations

Lynn A. Arrington

NYMA, Inc.

Tom W. Haag and Eric J. Pencil

NASA Lewis Research Center

Nicole J. Meckel

Primex Aerospace, Inc.

Abstract

Pulsed plasma thrusters are currently planned on

two small satellite missions and proposed for a

third. In these missions, the pulsed plasma

thruster's unique characteristics will be used

variously to provide propulsive attitude control,

orbit raising, translation, and precision positioning.

Pulsed plasma thrusters are attractive for small

satellite applications because they are essentially

stand alone devices which eliminate the need for

toxic and/or distributed propellant systems. Pulsed

plasma thrusters also operate at low power and

over a wide power range without loss of

performance. As part of the technical development

required for the noted missions, an experimental

program to optimize performance with respect to

electrode configuration was undertaken. One of

the planned missions will use pulsed plasma

thrusters for orbit raising requiring relatively high

thrust and previously tested configurations did not

provide this. Also, higher capacitor energies were

tested than previously tried for this mission.

Multiple configurations were tested and a final

configuration was selected for flight hardware

development. This paper describes the results of

the electrode optimization in detail.

Introduction

Pulsed plasma thrusters (PPTs) are low thrust

electric propulsion devices which can operate at

high specific impulse across a broad power range

(1-200W). PPTs typically use a fluorocarbon

polymer as a solid propellant, and applications for

these thrusters on satellites range from precision

positioning to orbit raising. In fact, PPTs are

scheduled on three upcoming space missions. In

1999, a PPT will fly on Earth Observer 1 (EO-I) to

demonstrate propulsive attitude control in NASA's

New Millennium program. In 2000, a PPT will be

used to demonstrate small satellite orbit raising on

the Air Force Mightysat ILl spacecraft.' Finally,

PPT technology has been baselined for both

spacecraft translation and precision positioning

(formation flying) in the proposed New Millennium

Deep Space 3 mission.

Compared to conventional propulsion systems, the PPT

is attractive in that this technology eliminates the need

for distributed and/or toxic propellant systems. PPTs

also operate at low power levels and its pulsed nature

permits operation over a relatively broad power range

without loss of performance. First developed during the

1970's and flown early into the 1980's, 2'3'4interest in

the PPT waned until NASA's On-Board Propulsion

(OBP) program began reevaluating the technology for

small satellite applications approximately three years

ago. Early technical and market assessments led to

renewed interest for small satellites, and new

generations of PPTs for both near and far term missions

are planned.

Most PPTs are ablative devices which accelerate

propellant through electromagnetic forces. Simply, the

heart of the PPT consists of a pair of electrodes

between which a bar of solid propellant is fed. Though

a variety of propellants can be used, a fluorocarbon

polymer is most typically employed. The electrodes are

connected to opposing sides of a high voltage

capacitor. Discharge is initiated via a spark plug

located in the cathode electrode. As the discharge

occurs, propellant is ablated, ionized, and then

accelerated via the Lorenz forces created by the

discharge. The PPT is typically charged up over a

period of milliseconds by a low voltage supply and the

energy is released in 1-10 microseconds. Except for

the spring which feeds the fuel bar, a typical PPT has

no moving parts.

The effect of varying propellant and electrode

geometry has been studied in the past. Vondra and

Thomassen investigated the variation in impulse bit

and mass loss per pulse as a function of fuel face

configuration and electrode flaring. _ Yaun-Zhu showed

that increasing the electrode gap increased

performance, particularly at higher energy levels, to

some limit. 6 A similar study was performed by

Polumbo and Guman in which the authors varied both

the gap between the electrodes and the included angle

NASA/TM--97-206305 1
betweenthem.7The PolumboandGumanstudy

alsoshowedthatangleandgapcanbeoptimized

to maximizeperformance.

Morerecently,theOBPprogramhasbeenworking

towardthedevelopment of a newgenerationof

PPTs. 8 As part of this effort, a breadboard PPT

was developed by the Primex Aerospace Company

(PAC) and this has been tested extensively both at

PAC and at the Lewis Research Center (LeRC).

As part of this testing, an extensive series of

experiments was performed at PAC to optimize

performance. In that experimental study, 9 different

electrode gaps, lengths, flares, and capacitor

energies were tested. The results were to be used

to determine configurations for the two near term

missions mentioned earlier. Of the two, the

planned orbital insertion is the most technically

challenging and that study focused on optimizing

a configuration for this application. The initial

capacitor energy levels studied were 22 and 43 J.

Various combinations of electrode spaces of 2.54,

3.81, and 4.08 cm, lengths of 2.54, 3.81, and 4.08

cm, and flares 0° and 20° were tested. It was

found that the configuration with 2.54 cm

electrodes, a 3.81 cm electrode gap, and no flare

provided the best performance with respect to

Mightysat mission requirements, i.e. the highest

impulse bit and moderately high specific impulse.

These results, however, were marginal for the

Mightysat insertion orbit at solar maximum and

further tests were deemed necessary to to finalize

the Mightysat configuration. The breadboard PPT

was moved to LeRC where previously untried

configurations with potential for performance

improvement, along with the most promising

configuration from the previous study, were tested.

In addition to testing at 43 J, testing at a higher

capacitor energy level was included. One

configuration was also tested at lower energy

levels to demonstrate the PPT could function over

a range of energies. This capability was critical

for precision positioning of the EO-1 mission. The

results of these tests are summarized in this paper.

In addition to the parametric performance tests, a

series of experiments were performed to assure that

short term test results were not biased due to

thermal effects. These effects had been observed

recently t° with other hardware where longer PPT

operation resulted in higher propellant

temperatures and thus greater mass loss impacting

performance. The results of the current study are

described and indicate that thermal impacts did

not affect the validity of short term tests with the

PPT used in this study.

Apparatus

Breadboard PPT

The Breadboard (BB) PPT tested in this study was

designed and built by Primex Aerospace Company as

part of the On-Board Propulsion PPT development

effort. The BB PPT has a modular design (see Figure

1) which allows easy exchange of components for

parametric study. As shown in the figure, the BB PPT

has two sets of electrode pairs which are parallel to

each other and fire in the same direction. The pair of

cathode electrodes are attached to a single stripline, as

are the anode electrodes. By removing the electrode

pairs, and replacing them with electrode pairs of

different dimensions, the gap between the electrodes,

the electrode lengths, and electrode flare can be

changed. The electrodes pairs are surrounded by a horn

assembly, and a spark plug is located in each cathode

electrode. Energy storage is provided by a 33 _tF

capacitor located below the horn assembly. The

electronics for the spark ignition and the capacitor

charging circuit are located behind the capacitor. Fuel

bars of fluorocarbon propellant are fed between the

electrode pairs by springs held in position by a

retaining shoulders built into the electrodes. The fuel

bars are removable so that they can be weighed to

determine mass loss.

Thrust Measurement

The transient nature of a PPT and the low thrust to

weight ratio required the use of a special torsional

thrust stand to determine the thrust and impulse bit of a

PPT. A precision balance designed and fabricated

under the OBP program was used for all thrust

measurements taken in the course of this testing. A

detailed description of the thrust stand can be found in

references 11 and 12, and a photograph of the thrust

stand with the PPT installed is shown in Figure 3.

Briefly, the stand resembles a swinging arm that

rotates around a vertical axis. The thruster is mounted

at a fixed radial distance from the axis on the end of

the arm with the thrust axis is tangent to the arc formed

by motion of the arm. A torsional restoring force is

used to resist the motion in the direction of the thrust.

Using the principle of angular momentum, the thrust

and impulse bit can be determined as a function of the

thrust stand deflection, spring stiffness, and natural

frequency. The natural frequency is determined by

disturbing the thrust stand with the damping circuit

deactivated. Since the thrust stand rotates on a

frictionless flexure pivot, the disturbed thrust stand sets

up a harmonic oscillation from which the natural

frequency can be measured. The spring stiffness of the

thrust stand can be calculated from the resultant

displacement of a known force applied to the stand.

When the PPT is operated in single pulse mode the

impulse bit can be determined by measuring the

displacement of the thrust stand. A single pulse starts

NASA/TM--97-206305 2
thethruststandfromrestanddeflects it until the

restoring force brings it to rest again and swings it

in the other direction, until it overshoots the

neutral position. The displacement is measured as

half of the peak-to-peak difference of the thrust

stand deflection. When the PPT is operated in

repetitive mode the thrust can be determined by

measuring the average deflection of the critically

damped thrust stand over a series of pulses. Since

the operating frequency of the PPT is many times

greater than the frequency response of the thrust

stand it is possible to do this. The restoring force

of the stand increases linearly with the

displacement and eventually balances with the

average thrust of the PPT. The average

displacement is compared to the displacements

generated by calibrated free hanging weights.

From these measurements, the thrust can be

calculated.

Vacuum Chambers

Most PPT testing described herein was performed

in a medium sized vacuum facility in which

ambient pressures were generally in the low 10 -5

torr range. The PPT was positioned horizontally in

the tank to fire along the long axis of the tank.

The PPT was fired at a frequency of less than 1 Hz

during thrust measurements to eliminate the need

for capacitor cooling. The lower energy level

tests were conducted at 2 Hz, since overheating

the capacitor was unlikely. This also decreased

the uncertainties of the thrust measurements at the

lower energies. The remaining mass loss testing

was performed in a different vacuum facility due to

availability. That chamber was a vertical facility

that is maintains a vacuum on the low 10 .5 torr

range via eryopumping. Thrust measurements

were not feasible in this upright facility. The tests

conducted in the upright facility were performed at

less than 1 Hz to again prevent capacitor

overheating.

Procedure

Experiment

The major objective of this study was to identify a

configuration which would provide the performance

required by the Mightsat mission, which required

performance the testing of different fuel bar face

shapes, electrode flare configurations, energy

levels and electrode lengths. The different

configurations are summarized in Table 1.

Previous tests indicated that a 3.81 cm electrode

gap was optimal, so this parameter was held

constant in this study. All electrodes were 2.54 cm

wide. The two fuel face configurations tested are

shown in Figure 2. One fuel bar had a flat face

perpendicular to the electrode faces. The other

had a notched face with a 120 ° inclusive angle. The

apex of the notch was perpendicular to the electrode

faces. The perpendicular distance from the center of

the spark plug to each face of the notch was the same

as the distance of the flat face fuel bar to the center of

the spark plug. Both types of fuel bars were

approximately 7.62 cm in length. Another condition

which was varied was the electrode flare configuration.

In one configuration, the electrode faces were parallel

to each other, called a 0° or no flare angle. In the other

a 20 degree inclusive flare angle was set between the

faces. In the main body of the testing pulse energy

levels of 43 and 54 J were used. The 43 J testing in

one case repeated previous data to provide

experimental confidence. The 54 J level was chosen

based on spacecraft considerations and the need for

higher thrust levels. Other tests at lower energy levels

were taken to expand the knowledge base and are

included here for completeness. Lower energy levels

tested were 5, 10 and 15 J, for one electrode and fuel

face configuration, that measured thrust and impulse

bit only. These energy levels where tested to

demonstrate the PPT could provide thrust at low energy

levels. Finally different electrode lengths were tested,

where the length is defined as the distance from fuel

face to electrode tip. Previous tests at Primex has

shown that 3.81 cm electrode gap with the 2.54 cm

long electrodes, and no flare showed the most promise

for performance increase, so that configuration was

repeated.

An additional test was performed to determine if the

mass loss per pulse varied with period of operation.

For each electrode configuration tested for performance

at a specific capacitor energy level, the BB PPT was

pulsed a total of approximately 2000 times. Thus, one

electrode configuration was chosen to run two tests of

different periods of operation at an energy level of 43 J.

One test ran for approximately 2000 pulses and the

second for 10,000 pulses. The configuration used was

the 3.81 cm long electrodes, 3.81 cm gap, 20 ° flare,

and flat fuel face. The mass loss was measured after

each test.

Mass Loss Measurements

The typical fuel bar weighed approximately 200g. All

mass loss measurements were made on a precision

balance capable of weighing up to 1000g. The balance

is accurate to +/-0.001g. Prior to weighing each fuel

bar, the balance was checked against known calibrated

weights. All fuel bars were handled with gloves to

avoid contamination.

Performance Calculations

The performance equations for a PPT have been well

defined in previous references. _3'14 The impulse bit and

thrust are calculated from measurements made on the

thrust stand and are defined above. From these

NASA/TM--97-206305 3
parametersthe specificimpulse,efficiencyand

powerofthePPTcanbecalculated.Thespecific

impulse,I_p,is definedbytheequation:

mg

whereT is thethrust, m is the mass ablated per

pulse, f is the frequency at which the PPT

discharges, and g is the gravitational constant. The

thruster efficiency, 11, is then defined as:

where E is the capacitor discharge energy. The

power, P, of the PPT is defined as:

P=Ef.

Results and Discussion

As noted in Table 2, several electrode and fuel

face configurations were evaluated in a parametric

study designed to identify a final configuration

which would meet Mightysat requirements. A

range of capacitor energies were also evaluated.

The results of the electrode configurations showed

that increasing the length of the electrodes from

2.54 cm to 3.81 cm electrodes increased the

efficiency but not the thrust to power ratio.

However, when the electrodes were flared at 20 °,

both the efficiency and thrust to power ratio

improved. Notching the fuel face increased the

thrust to power ratio but decreased the efficiency.

Increasing the capacitor energy level was found to

increase the efficiencies in all cases. Also, the

data showed that the impulse bit increased nearly

linearly over the range of energies tested,

particularly for the 0° flare case. Results are

shown in Figure 4, with the highest impulse bits at

the 54 J energy level. While the overall results of

this study did not show one configuration which is

clearly superior in all performance categories (e.g.

efficiency, impulse bit, thrust to power, etc.), once

plotted with respect to a mission analysis, the

selection of a configuration became more apparent.

The impact of the various configurations on the

Mightysat mission are shown in Figure 5. In this

figure, plots of total specific impulse for a given

amount of fuel and an average orbit thrust at

various insertion altitudes (nautical miles), are

reproduced from Reference 1. The figure also

shows various orbit raising trip times, with trip

times decreasing to the fight. The projected

performance of the various configurations

examined in this study are located in the upper left

hand comer of the figure. Though it is difficult to view

on the figure the 3.81 cm long electrodes with the 20 °

flare at 43 J coincides with the notch fuel face at 54 J.

From this it can be seen that the notched fuel face at

43 J would provide the shortest orbit trip time. The

notched face feed system, however also required the

most rigorous design and no long term testing has been

performed to demonstrate unequivocally that the

notched face will maintain its shaped (and

performance) over the millions of pulses required for

small satellite insertion. Thus it was desirable to avoid

this configuration until these concerns could be

addressed. The 2.81 cm electrodes with the 20° flare

produced only a slightly longer trip time and almost as

high an average thrust at both 43 and 54 J. This

configuration was then deemed the "best". On the

component level, higher energies put a greater strain

on the capacitor and thus are expected to reduce the

maximum mission life. Since both energy levels for

the 20 ° flare configuration met the total impulse

requirement in theory, in deference to the life of the

capacitor, the lower energy of 43 J was selected for the

mission. The final selection for Mightysat was the 3.81

cm long electrodes, 3.81cm gap, and 20 ° flare with a

flat fuel face at 43 J. In an effort to maintain

uniformity between missions, the same electrode

configuration would have been chosen for EO1.

However, dimensional restrictions placed on the PPT

by the spacecraft required the use of electrodes with no

flare though the same length and gap.

The PPT was also tested at lower energy levels in the

3.81 cm x 3.81 x 0° flare configuration. The PPT was

successfully discharged at the 5, 10, and 15 J energy

levels. The resultant thrust to power and impulse bit

measurements are recorded in Table 2. However, mass

measurements were not made in the interests of time.

The purpose here was to prove discharge capability

across a wide range of capacitor energies, which was

the case.

Measurement uncertainties were determined for the

performance parameters using standard propagation

techniques. The calculated performance parameter of

power, specific impulse, impulse bit, and efficiency

were calculated from the uncertainties of the measured

quantities. The largest contributing factor was the

thrust measurement, followed by the uncertainty in the

mass loss measurements. The uncertainties are

presented in Table 3. Generally for the 43 and 54 J

cases, the specific impulse uncertainty was less than

5%, the power uncertainty 1.2%, the impulse bit

uncertainty less than 5%, and the efficiency

uncertainty less than 9.5%. At the lower energy, mass

loss measurement were not available, so only power

and impulse bit uncertainties were calculated which

were slight higher than at the higher energy levels.

NASA/TM--97-206305 4
For the massloss test conductedin vertical

facility,themasslossperpulsewas64.8ktgand

64.1p.g, for the 2000and 10,000pulsetests,

respectively.The differencebetweenthe two

measurements is approximately1.1%whichis less

thanthemeasurement uncertaintyof theweighing

devicealone(1.5%for the shorterpulsetest).

Therefore, nomasslossperpulsedifferencewas

discernibleforthisconfigurationandenergylevel.

This testwasperformedbecausetherewasa

concernexpressedthattheheatingof thefuelbar

couldcauseanincreasein fuel consumptionthat

wouldeffect the validity of the performance

parameters. Thelongertestwasconductedto see

if additionalheatingof thefuelbarincreasedthe

masslossperpulse.

Summary

PPT testing at LeRC showed that a configuration

with 3.81 cm long electrodes, a 3.81 cm gap and a

20 ° flare had the highest efficiencies at the

respective energy levels. The same gap and length

without a flare and a notched fuel face had the

highest thrust to power ratio. Also, the PPT was

successfully discharged at lower energy levels,

showing the PPT as viable across a wide range of

energy levels.

The 3.81 cm x 3.81 cm x 20 ° flare at configuration

43 J was chosen for the Mightysat II.1 mission

because of its high efficiency and good thrust to

power ratio, and the ability to best approach the

maximum mission requirements for the allotted

fuel mass. The lower energy level was chosen to

increase the life expectancy of the capacitor. The

notched fuel face had better thrust to power ratios,

but this particular configuration had the most

difficult fuel feed system design. Also, the PPT

was not operated over millions of pulses to

determine if the fuel face would retain its original

notched shape over its life. The same electrode

configuration but without the flare was chosen for

the EO-I mission because of the high performance,

but the flare was forgone because of dimensional

restrictions place on the PPT by the spacecraft.

References

i LeDuc, J.R., Bromaghim, D.R., Peterson, T.,

Pencil, E., Arrington, L., Hoskins, W.A., Meckel,

N.J., and Cassady, R.J., "Mission Planning,

Hardware Development, and Ground Testing for

the Pulsed Plasma Thruster(PPT) Flight

Demonstration on Mighty Sat II.1", AIAA 97-2779,

July 1997.

2 Guman, W.J., "Pulsed Plasma Microthruster

Propulsion system for Synchronous Orbit Satellite",

Journal of Spacecraft and Rockets, Vol. 7, No. 4, April

1970.

3 Guman, W.J., "Pulsed Plasma Solid Propellant

Microthruster for the Synchronous Meteorological

Satellite", NAS5-11494, August 1972.

4 Brill, Y., Eisner, A., and Osborn, L., "The Flight

Application of a Pulsed Plasma Microthruster: The

NOVA Satellite", AIAA-82-1956, November 1982.

5 Vondra, R.J. and Thomassen, K.I., "Performance

Improvements in Solid Fuel Microthrusters", Journal of

Spacecraft and Rockets, Vol. 9, No. 10, October 1972,

pp 738-742.

6 Yuan-Zhu, K., "Effects of Propellant Geometry on

PPT Performance", IEPC 84-94, 1984.

7 Palumbo, D.J., and Guman, W.J., "Effects of

Propellant and Electrode Geometry on Pulsed Ablative

Plasma Thruster Performance", Journal of Spacecraft

and Rockets, Vol. 13, No. 3, March 1976, pp. 163-167.

8 Curran, F.M., Peterson, T.T., and Pencil, E., "Pulsed

Plasma Thruster Technology Directions", AIAA 97-

2926, July t997.

9 Meckel, N. J., Hoskins, W.A., Cassady, R.J., Myers,

R.M., Olsen, S.R., McGuire, M.L., "Improve Pulsed

Plasma Thruster System for Satellite Propulsion",

AIAA 96-2735, July 1996.

10 Spanjers, G.G., Malak, J.B., Leiweke, R.J., and

Spores, R.A., "The Effect of Propellant Temperature on

Efficiency in a Pulsed Plasma Thruster", AIAA 97-

2920, July 1997.

11Haag, T.W., "PPT Thrust Stand", AIAA Paper 95-

2917, July 1995.

12 Haag, T.W., "Thrust Stand for Pulsed Plasma

Thrusters", Rev. Sci. Instrum., Vol. 68. No. 5, May

1997, pp. 2060-2067.

13Solbes, A. and Vondra, R.J., "Performance Study of a

Solid Fuel-Pulsed Electric Microthruster", Journal of

Spacecraft, Vol. 10, No. 6., June 1973.

t4 Guman, W.J., "Designing Solid Propellant Pulsed

Plasma Thrusters", AIAA Paper 75-410, March 1975.

NASA/TM--97-206305 5
Table1:ElectrodeConfigurations

Configuration ElectrodeLength,

cm

Flare,

degrees

0

Fuel Face

1 3.81 Flat

2 3.81 20 Flat 54,43

3 3.81 0 Notch 54,43

4 2.54 0 Flat 43

Energy Levels,

joules

54,43,15,10,5

Table 2: Summary of Primex Breadboard PPT Testing at LeRC

Electrode

Length,

cm

Fuel

Face

Flare

Angle,

degrees

0

Capacitor

Energy,

.ioules

43

Ibil,

gN-sec

Isp,

sec

Thrust/

Power,

_tN-sec/J

Efficiency

%

3.81 Flat 694 1200 16.1 9.5

3.81 FI at 0 54 875 1351 16.2 10.7

3.81 Flat 20 43 734 1228 17.1 10.3

3.81 Flat 20 54 914 1331 16.9 11.0

3.81 Notch 0 43 775 1059 18.0 9.4

3.81 Notch 0 54 950 1168 17.6 10.1

2.54 Flat 0 43 726 1121 16.9 9

3.81 Flat 0 5.2 96 - 18.5 -

3.81 Flat 0 10 97 - 9.4 -

3.81 Flat 0 15 172 - 11.5 -

Table 3: Performance Parameter Uncertainties

Configuration

3.81 x 3.81 x 0°, 54J

U_rmeifie imnlll_a_, %

4.8

3.81 x 3.81 x 0% 43J 4.4 4.5 1.2

3.81 x 3.81 x 20°, 54J 3.1 3.2 1.2

3.81 x 2.4 2.5 1.2

3.81 x 3.81 x 0% 15J 6.2

3.81 x 20°, 43J

3.81 x 3.81 x notch, 54J 2.4 2.5

3.81 x 3.81 x notch, 43J 2.1 2.2

2.54 x 3.81 x 0°, 43J 3.6 3.6

3.81 x 3.81 x 0°, 5.2J 4.4 -

3.81 x 3.81 x 0°, 10J 5.7 -

1.2

1.2

1.2

1.8

3.8

2.2

% U_mcl ..... %

9.5

8.9

6.3

4.9

4.9

4.3

7.3

NASA/TM--97-206305 6
Figure1: PhotographofPrimexBreadboardPPT.

gap / spark plug

/"-" _ _ /

teflon fuel bar teflon fuel bar

flat face notch face

Figure 2: Schematic of Fuel Face and Electrode Configurations for a PPT.

Figure 3: Photograph of Thrust Stand with Primex Breadboard PPT.

NASA/TM--97-206305 7
IO0o

900

800

,_ 70o

z_z

_6O0

_, 50o

300

200

16O

0

• •

I0 20 30

Energy,Joules

!

!

n • 3.81 ¢rnxo °

dl, 3.81 crux 20 _

• 381cm x Notch

2-_4 cm x 0 °

40 50 6O

Figure 4: Impulse Bit versus Energy Level for Various Electrode Configurations.

I4

12

10 A

2

o

500

i s6

4

Required Total Impulse and Thrust to Raise MightySat rl.1

to a One-Year-Life Orbit

t 5-0-54 I • 1 5-20-54 PPT_ Initial Altitude

1 00 kgfuel

67 Ww:g pwr I250

1 5-0-43 _2_,"PPU eft .... 265

1..5-20-43 • rotch-54 ....... 275

t 0-0-43 •

• notch-43

"_ i00daw

_/ _ FI0.7 - 241

i_ FI0.7 = 2110

/_, 0.7 = 240

sm

/ .................................. no.7-240

- --_ F10.7 - 160

":_:T""_'.";"_'.'::":2=:_==__'_':7'"'; .....;....;-._o.r,= 2_ , ,

I000 1500 2000 2500

Average Orbit Thrust (ttN)

Figure 5: PPT Total Impulse Vs. Average Orbit Thrust Available reproduced from Reference 1.

NASA/TM--97-206305 8

REPORT DOCUMENTATION PAGE Forr,,A_orov_

OMB No. 07040188

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1. AGENCY USE ONLY (Leave b_ 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

December 1997 Technical Memorandum

,4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

A Performance Comparison of Pulsed Plasma Thruster Electrode Configurations

6. AUTHOR(S)

Lynn A. Ardngton, Tom W. Haag, Eric J. Pencil, and Nicole J. Meckel

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135-3191

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

WU-632-1B-1B-O0

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-II002

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TMB97-206305

IEPC-97-127

111. SUPPLEMENTARY NOTES

Prepared for the 25th International Electric Propulsion Conference sponsored by the Electric Rocket Propulsion Society,

Cleveland, Ohio, August 24--28, 1997. Lynn A. Arrington, NYMA, Inc., 2001 Aerospace Parkway, Brook Park,Ohio

44142 (work funded by NASA Contract NAS3-27] 86); Tom W. Haag and Eric J. Pencil, NASA Lewis Research Center,

Nicole J. Meckel, Primex Aerospace, Inc., Seattle, Washington. Responsible person, Lynn A. Arrington, organization code

5430, (216) 977-7486.

|21. DISTRIBUTI_FI_AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Categories: 20, 15, and 13 Distribution: Nonstandard

This publication is available from the NASA Cent_" for AeroSpace Information, (301) 621--0390.

121=. DISTRIBUTION CODE

l& AB Sl"P.ACT (Ma,_dmum 200 tmrds)

Pulsed plasma thrusters are currendy planned on two small satellite missions and proposed for a third, in these missions, the pulsed plasma thruster's unique characteristics will be used variously to provide propulsive attitude control,

orbit raising, translation, and precision positioning. Pulsed plasma thrusters are attractive for small satellite applications

because they are essentially stand alone devices which eliminate the need for toxic and/or distributed propellant

systems. Pulsed plasma thrusters also operate at low power and over a wide power range without loss of performance.

As part of the technical development required for the noted missions, an experimental program to optimize performance

with respect to electrode configuration was undertaken. One of the planned missions will use pulsed plasma thrusters

for orbit raising requiting relatively high thrust and previously tested configurations did not provide this. Also, higher

capacitor energies were tested than previously tried for this mission. Multiple configurations were tested and a final

configuration was selected for flight hardware development. This paper describes the results of the electrode optimization in detail.

14. SUBJECT TERMS

Pulsed plasma thrusters; Electrodes; Performance; Electric propulsion

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURn'Y CLASSIFICATION

OF REPORT OF THIS PAGE OF ABSTRACT

Unclassified Unclassified Unclassified

NSN 7540-01-280-5500

15. NUMBER OF PAGES

_4

16. PRICE CODE

A03

20. UMITATION OF ABSTRACT

Standard Form 298 (Rev. 2-89)

Pm_:dl)ed by ANSI Std. Z39-18

298-102