Pulsed Plasma Thruster

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Pergamon

rm srronautica Vol. 35. No. 9-1 I pp . 585-590. 1995Ekvier Science Ltd. Printed in Great Britain

0094-5766(95)ooo25-9

PULSED PLASMA THRUSTER OF THE EROSION TYPEFOR A GEOSTATIONARY ARTIFICIAL EARTH

SATELLITETS

A. I. RUDIKOV, N. N. ANTROPOV and G. A. POPOVResearch Institute of Applied Mechanics and Electrodynamics of the Moscow Aviation Institute,

Moscow, Russia

Received I Jul y 1994; received for publication 6 February 1995)

Abstract-Analysis is made for the possibility of erosion pulsed plasma thruster (PPT) application while

solving the task of maintaining the point sustaining a long-operating geostationary artificial Earth satellite.The concept of an erosion PPT with a pulse energy of 200-300 J, designed for holding the attitude of asatellite of 500 kg mass over 10 years is presented. The thruster, with a lifetime of 2-3 x 10 pulses,produces a total pulse of 2.5 x 10s Ns, consuming up to 13 kg of propellant (Teflon). Estimations haveshown that the thruster total mass will not exceed S&60 kg for the flying PPT. Even in the case of atwo-fold redundancy the thruster set mass should comprise no more than 0.20.25 of the satellite mass.The prospect of a rail design for the PPT discharge chamber with lateral propellant feeding isexperimentally examined

1. INTRODUCTION

The application of artificial Earth satellites in geo-stationary orbit for global communication systemcreation and various scientific purposes has given riseto huge interest in the last few years. For effectiveattitude control of the satellite in orbit it is necessaryto compensate the disturbing forces and momentsacting upon it which are being caused by the gravityof the Sun and Moon, solar wind pressure, Earthsgravitational field and non-centrality and by anumber of other factors. The task of the parameterssustaining geostationary orbit during a day requiresthe creation of thrust pulses, the calculated daily

values of which are shown in Fig. 1 as a function ofsatellite mass. Average thrust values for continuousthruster operation are also presented. The accuracyfor sustaining the satellite center of mass and its axesdirections depends upon the value of unit pulse beingproduced by the jet controls.

Gas-jet and catalytic (mono- and bi-propellant)engines are currently broadly used as the satellitecontrols[l]. For the satellite in low orbit for a rela-tively small (up to one year) period of active oper-ation, these engines have no competitors from amongthe other controls in the case of low requirement asregards the afteraction pulse and lifetime. They havelow mass and overall dimensions and are simple andconvenient to operate.

tPaper IAF-93-5.5.187 presented at the 44th I nternationalAstronau tical Federati on Congress, Graz, Austria, 1622October 1993.

$Due to circumstances beyond the Publishers control, thispaper appears in print without author corrections.

@ee Nomenclature at the end of this paper.

With the increase in the active operation period(from one to ten years) of the satellite and require-ment for spacecraft stabilization accuracy it is advis-able to use electric rocket propulsion in the controlsystem. Among the stationary plasma thrusters thethrusters with a closed drift of electrons or with amagnetic layer (TML) are optimized to the highestextent and tested in space[2]. Xenon is ordinarilyused as a propellant in these thrusters which have ahigh thrust efficiency. A lifetime of IO3 h of con-tinuous operation and of 10 switch-ons at pulsedoperation were confirmed experimentally. TML dis-advantages include the necessity of time losses forpreparation for operation (cathode-compensator

heating), the complexity of producing small unitpulses and afteraction pulse and the presence of aself-induced gas environment, caused by thrusteroutgassing. The mentioned disadvantages could beeliminated by application of an erosion plasmathruster with a solid dielectric as propellant[3].

RIAME MAI has vast experience in the develop-ment of pulsed plasma sources for flight and spaceexperiments under Earth ionosphere and magneto-sphere investigation. Twelve units of pulsed plasmasources with a pulse energy in the region of

100-1000 J were designed and successfully testedunder space vacuum conditions from 1975 to 1989[4].

2. JUSTIFlCATiON OF PPT PARAMETERS CHOICE@

At the first stage the task of PPT optimization isreduced to the definition of the thruster set minimummass at given values of total thrust pulse, unit pulse,average operation frequency and lifetime.

585


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A. I. Rudikov er al.86

10-5

IO-4

N

10-j

I I I2 4 103

kg

Fig. 1. Thrust characteristics of the satellite engine.

The PPT propulsion module mass could be pre-sented as a sum:

M,=M,,,+M,+M,+M,. 1)

Taking into account the expressions for the PPT masscomponents

(M, = mN, M, = ye IV M, = Y\Y Wfll?w, Mk = Y,M,),

the empirical relation for the thrust efficiencyr~ = K./m,)/* and dependencies C, = W/P, TJ =

P2/2m W, eqn (1), after simple transformation, couldbe written in the following form:

M,IM = AaC,(m, + YlN)ln(l - ok), (2)

where

C, = l/ 4K,m,)4, 3)

Y = Yc + YJh~ 4)

The value of the specific mass in the pulse, at whichMp minimum is achieved, is:

m, = y/3N. (5)

Accounting for eqns (3), (5) and (2) the expression forthe definition of the minimum thruster module masshas the following form:

M pm,n= (1.24Au/n(l - yk)K:j4)

x Y/N)~. 6)

The thruster unit comprises n simultaneously operat-ing modules, which ensure the velocity increment Aufor the satellite in the case of lifetime utilization.

The working body constant K. is 4 x lo- kg/J forTeflon, thus the minimum PPT mass in the case ofoperation with Teflon could be presented as:

Mpmin = [490Au/n(l - yt)](y/N)34. (7)

Making the same transformations, the expressionfor the definition of the minimum mass of a station-ary plasma thruster with a magnetic layer, operatingwith the gaseous propellant, could be obtained:

Mpmin= 2Aa( 1 + yy /no, 8)

IO3

102

N 24h

10

where

a = [2V9Vl(l + Y)l(Y, + Yr)l*. (9)

Let us discuss the advisability of PPT use forsolving the most energy-consumable task-attitude

hold in the geostationary orbit for a satellite of 500 kgin mass in the North-South direction over 10 years.For the total pulse of 2.5 x 1ONs the satellitecharacteristic velocity change during the 10 yearscomprises a value of N 500 m/s (see Fig. 1).

Currently the following specific parameter valuesof the PPT power supply unit are achievable:

yc = 3-5 x lo-* kg/J (for the foil capacitors)

yw = 3-5 x lo-* kg/W, rlw N 0.8

for the lifetime of its elements of l-3 x 10 pulsePI

In this case the specific parameters dependenceupon the lifetime could be described by theexpressions:

yc = 3 x 10-2(N/107)3,

Yw 3 x 10-2(N/107)3

assuming the relative mass of the constructiveelements for the PPT flight variant to be equal to0.2.

Figures 2 and 3 show the mass, energy and specificparameters of the PPT module as functions of life-time. The unit comprises two thruster modules, withoppositely directed discharge chambers, the axes ofwhich go through the satellite center of mass. Theaverage frequency for each module operation is0.5 Hz. As follows from the above-mentioned curves,the PPT lifetime increase from 10 up to 3 x 10 andcauses a reduction of the thruster set mass by twotimes. The minimum calculated value of the thrusterunit mass comprises 50 kg. Even in the case oftwo-fold modular redundancy the PPT mass would

not exceed 100 kg, which is 0.2 times that of thesatellite mass.

kg

I I I I1 2x 10

PulsesFig. 2. Mass and energy parameters of the PPT.


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PPT of the erosion type 587__---c --.

2x 109..

y,< I 1x 105WNTN 2L I I1 2x 107

PulsesFig. 3. PPT specific parameters.

Let us examine the thruster with a magnetic layerfor comparison. For estimation of the TML unitmass, assume the following values of specific par-ameters, which have been confirmed experimentally:

y, = 1, yw = 2 x lo-* kg/W,

yp = 1.5 x lo-* kg/W, q = 0.5, q\y = 0.8,

7 = 3.6 x 106s.

Then the mass of a thruster unit consisting of two

modules would be 80 kg at a velocity of the plasmajet efflux v = 12 x lo3 m/s. Thus, the PPT and TMLunits are comparable in mass.

3 C HOICE OF PPT DESIGN

The reactive thrust in the PPT is produced due tothe dielectric erosion products efflux out of the dis-charge chamber as a result of a high current pulseddischarge between the electrodes along the dielectricsurface. Figures 4 and 5 show some of the types ofPPT discharge chambers with solid dielectric feedingto the discharge area. The discharge chamber of theerosion PPT comprises: (1) cathode; (2) anode; (3)working body grains; (4) igniter. The propellant

-

L

Fig. 4. Discharge chambers with longitudinal grain feeding.

Fig. 5. Discharge chamber with lateral grain feeding.

feeding system transports the grains in the directionindicated by the arrows, while their exhaust PPTelectrodes are connected to the capacitive battery.

Development of a discharge chamber with a highlifetime requires, in particular, the assurance of:

-high lifetime and reliability for the igniter, gener-ating plasma igniting the main discharge,

--constancy of the discharge chamber geometricdimensions for ensuring thrust pulse stability[6].

A discharge chamber of the coaxial type withlongitudinal grain feeding (see Fig. 4) might ensurethe calculated value of the specific mass in a pulse of0.8-1.6 x 10e9 kg/J for the plane working surface ofthe grain. However, as the lifetime test showed, theinitial shape of the dielectric working surface isvarying substantially during PPT operation andacquires a parabolic shape after _ 10 pulses. In thiscase the blob specific mass increases up to

2-3 x 10m9 kg/J, while the thrust efficiency decreases.Another substantial disadvantage of this design is thelarge length of the grain, which is more than 1 m inour case. The inductance of conducting bushes growsproportionally as the grain length increases, whichleads to thrust efficiency reduction. In view of theabove, the first design, presented in Fig. 4, should beused in cases where the operation lifetime of the PPTis not high with a small volume of dielectric being fed.

The rail discharge chamber with longitudinal pro-pellant grain feeding is shown in Fig. 4. In the caseof this design the geometric dimensions of the dis-charge gap practically do not change during oper-ation. Inductance and resistance of conductingbushes could be made minimum and high thrustefficiency values could be ensured due to correctcommutation. But the specific mass output from thedischarge is not high here and comprises less than0.8 x 1O-9 kg/J. Thrust pulse is less than the calcu-lated value too. For the task solution this wouldrequire a lifetime increase up to 3 x 10 pulses.


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A. I. Rudikov ef al

4 PPT DISCHARGE CHAMBER LABORATORY TEST

Two- and three-electrode high voltage spark plugsof the surface breakthrough were mainly used asigniters in the flight variants of the pulsed accelera-tors, designed in RIAME MAI. Teflon was used asthe material dividing the electrodes. In the case of anenergy of N 1 J, such igniters initiate discharge witha high reliability at a distance between the electrodesof up to 10 cm. However, the Teflon consumption insuch igniters does not allow them to be used in PPTwith a high lifetime. To substantially increase theigniter lifetime the dielectric separating the ignitingelectrodes should be made of a ceramic (aluminumoxide, for example). As carbon is in the compositionof Teflon, the working surface of the ceramic shouldbe covered by a carbonic film during the discharge

process. The igniting plasma blob will be formed asa result of a high voltage breakthrough in the film,protecting the surface of the ceramic and electrodes.As regards making the correct choice of the igniterlocation inside the discharge chamber and of theinitiation of energy discharge, the constancy of thecarbon film thickness should be ensured during thePPT operation and erosion of the igniter workingelements should be eliminated.

Constructive design for the PPT discharge chamberlaboratory model with lateral propellant feeding is

presented in Fig. 7. The chamber comprises: (1)cathode; (2) anode; (3) propellant grains; (4) igniter;(5) end insulator. Teflon grains are fed to the dis-charge zone as their consumption takes place with thehelp of a spring (6) and pushers (7).

Discharge electrodes are made of copper. Theanode working surface is plane. As tests showed, theanode surface, salient into the chamber, leads tocarbon film formation at the grain edges borderingthe anode. Carbonic film presence at the Teflonsurface prevents its evaporation in the discharge. Aswas mentioned above (see Fig. 5) the grains are fedby the pusher up to the fixing device at the cathode.For reduction of the erosion of the fixator edges, thecathode is made in the form of a cylinder. Thiscathode form allows the channel width to be changedby variation of the distance between the electrodeswithout variation in the grain dimensions. In order to

Fig. 6. PPT arrangement.

Figure 5 shows the rail chamber design with lateralTeflon feeding to the discharge region. In this casegrains are made in the form of half-rings having arectangular radial cross-section. The feeding systemensures their transportation around the system axisup to the lock of the fixing device at the cathode.

Disposition of the grain working surface along theelectrodes increases the zone of discharge location,reducing electrode erosion. A ring form of grainsallows substantial decrease of the overall dimensionsof the feeding system in the case of a large propellantmass. Varying the discharge channel dimensions onecould obtain the calculated value of the specific massoutput at a level of low9 kg/J. All the above obser-vations of the latter design show prospect for design-ing a PPT for a geostationary satellite.

Figure 6 shows one of the variants for the PPT

module, designed for attitude hold of a geostationarysatellite of 500 kg mass over 10 years. Propellant(Teflon) load (5-6 kg) is calculated for a lifetime of2-3 x 10 pulses at a pulse energy of 220-320 J. Massand energy parameters of this PPT are presented inFig. 2. The module has two discharge chambers. Eachchamber is equipped with its own feeding systemwith a Teflon load and igniter with the dischargeinitiation unit. This design allows, firstly, thementioned elements to be made redundant and, sec-ondly, the reduction of the overall PPT moduledimensions. In the simplest case, the grain feedingsystem might be made in the form of a torsion spring(see Fig. 6).

The power unit, comprising the capacitive battery,voltage converter and discharge initiation units, islocated between two units of Teflon storage andfeeding made in the form of disks. The capacitanceof the battery is SO-70 pF. The maximum overalldimension of the module does not exceed 0.5 m andits mass is 25-30 kg.

h4- + ?6

Fig. 7. PPT laboratory model circuit.


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PPT of the erosion type

prevent carbon film formation at the grains edgesbordering the cathode, its diameter should be morethan three channel widths.

The end insulator is made of ceramic. Its workingsurface, turned to the channel, has a concave cylinder

form, the diameter of which is not less than thechannel width. The depth of the forechamber (8)formed by the insulator, is approximately 3/2 of thechannel width. The igniter should be mounted nearthe rear surface of the forechamber normal to thecathode axis. As experiments have shown, such aform of the end insulator working surface and igniterlocation ensure uniform Teflon evaporation at thebeginning of the discharge channel. In order toprevent electric self-breakthroughs at the surface ofthe carbon film from forming at the forechamberwalls, lateral grooves are made in them.

The igniter is made in the form of a ceramic rod ofaluminum oxide of 3 mm in diameter having longi-tudinal channels of 1 mm in diameter. Copperigniters are mounted inside the channels. The planeworking end of the igniter does not run off thecathode surface.

The discharge initiation unit (9) generates a highvoltage pulse of -20 kV. Pulse energy is _ 1 J. Theelectrodes of the main discharge are connected to thecapacitive battery (10) having a capacitance of 36 PFand a maximum voltage of 3 kV. Inductance and

resistance of the conducting bushes are correspond-ingly equal to 1.2 x IO- H and 2 x lo- ohms. Thefrequency mode of PPT operation is defined by thepulse generator 1 1), connected to the dischargeinitiation unit. The maximum pulse repetition fre-quency is 0.4Hz.

The PPT experimental test was conducted in avacuum chamber at a residual gas pressure of notmore than 10m4 orr.

Thrust pulse indirect measurements were made bya dynamic thrust meter. Grains were weighted beforeand after a series of 103-lo4 pulses for definition ofthe propellant mass per pulse.

Experimental tests of constructive elements showed:

4 10-7

kg

20 40

1 mm

Fig. 8. PPT integral parameters.

10

km-s

5 -It--Y

I I20 40

1 mm

Fig. 9. PPT specific parameters.

589

0.1

0.05

-absence of electrode erosion effects inside thechannel (electrodes are covered by a densecarbon film) and weak cathode erosion at thechannel exit;

-equality of the Teflon mass, coming to thedischarge from each grain;

-good uniformity of Teflon evaporation fromthe grain working surfaces;

-stoppage of carbon film growth at the fore-chamber walls after lo4 pulses;

-erosion absence at the igniter working endcovered by the carbon film.

Figures 8 and 9 show the PPT performance for apulse energy of 160 J and current amplitude of 35 kA.The grain length I varied in the range of 15-55 mmin the case of the channel width h = 8 mm and heightH = 35 mm. It is obvious from Fig. 8, that the massper pulse is approximately proportional to the grainlength. Such a dependence substantially simplifies thegrain length choice for ensuring the given value fatthe blob specific mass.

Refinement of the discharge chamber and of theelectric circuit was made to ensure the calculated PPTparameters for a lifetime of 3 x 10 pulses (see Fig. 3).The width and height of the modified channel were 20and 50 mm, respectively. The capacitance of thebattery was increased up to 50 p F, while inductanceand resistance of the conducting bushes were de-creased to 5 x IO- H and 5 x 1O-4 ohm, respectively.Tests of the new PPT laboratory mock-up modifi-cation confirmed the correctness of the engineeringforecast. The calculated PPT parameters were

achieved at a grain length I = 20 mm.Thus, the results obtained confirm the correctnessof the choice of PPT design, discharge chamber andelement geometry for the development of an engineset with a long lifetime.

Acknowledgemenrs-The authors would like to express theirthanks to S. Yu. Shibanov, D. V. Khorkov, I. G.Krivonosov and G. V. Soganova for help in the experimentsand execution of this paper.


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REFERENCES

H. D. Schmitz, Techni cal A spects on the Development ofLow Thrust H ydrazi ne Propulsion Systems. ERNO,Germany (1971).N. V. Belan, V. Kim, A. I. Oransky and V. B. Tikhonov,Stationary plasma thrusters. USSR State Committee on

National Education, KhAI (1989) (in Russian).W. J. Guman and D. M. Nathanson, Pulsed plasmamicrothruster propulsion system for synchronous orbitsatellite. J. Spacecraff Rock et s 7, 409 1970).S. I. Avdyushin, I. M. Podgorny, G. A. Popov and A.A. Porotnikov, Plasma accelerators application for thestudy of physical processes in space. In Plasma A cceler-ators and I on I njectors, pp. 232-239. Nauka, MOSCOW1984) (in Russian).

L. Golkomb, Electric rocket propulsion sets for satel-lites. J. Vopr. raker. t ekh. 10, 39-66 (1972) (in Russian).D. J. Palumbo and W. J. Guman, Effect of electrodegeometry and propellant on pulsed ablative thrusterperformance. AIAA Paper 75-409, March (1975).

APPENDIX

Nomenclature

Au = satellite velocity variationv = effective velocity for the plasma plume effluxw = energy in the dischargey = specific mass for the power supply unitye = capacitive battery specific massy, = energy converter specific massyk = relative mass of the constructionyv = relative mass for tanks with propellanty,, = thruster specific massq = thrust efficiency

1, = voltage converter efficiency5 = lifetime

C, = energetic price of thrust I, H, h = length, height and width of the discharge chan-f = PPT operation frequency nel in Fig. 7.

K. = working body constantm = plasma blob mass

m, = blob specific massM = satellite mass

M, = PPT module massM,,, = working body massM, = capacitive battery massM, = voltage converter massMk = construction mass

n = number of thruster modules operating simul-taneously

N = total pulse numberP = unit pulse

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Pulsed Plasma Thruster