[American Institute of Aeronautics and Astronautics 41st AIAA/ASME/SAE/ASEE Joint Propulsion...

1American Institute of Aeronautics and Astronautics

Intelsat 10 Plasma Propulsion System Initial FlightOperations

Howard L Gray*

EADS Astrium Ltd, Portsmouth, PO3 5PU, UK

Ulhas P Kamath†

Intelsat, Washington DC 20008

The latest generation of geostationary telecommunications satellites provide increases inpower, life and payload mass. For these platforms electric propulsion systems offerattractive overall mass savings, as the propellant gains are significantly higher than anyadditional hardware mass. The function of the Plasma Propulsion System (PPS) used on theEurostar 3000 platform is to provide inclination and eccentricity control for north-southstation keeping. It is based on the SPT-100 Stationary Plasma Thruster, and in the mainexploits hardware already developed and qualified for the French experimental Stentorsatellite. This paper presents an overview of the Eurostar 3000 platform PPS asimplemented on the Intelsat 10 satellite, along with the in-orbit testing and initial flightoperations results.

I. IntroductionElectric propulsion (EP) systems have been advocated for many years for various satellite applications. The

large fuel savings they offer compared to conventional chemical propulsion systems (CPS) (a factor in the order of 5for hall effect thrusters) offers very significant mass savings at satellite level; this technology thereby allows theintroduction of large high power satellites within existing launcher constraints. As a result, most satellite primecontractors now offer some form of electric propulsion for their latest large high power platforms, either as abaseline or as an option. It should be noted, however, that there is a penalty in terms of additional hardware mass,making EP particularly useful for satellites which otherwise would have high propellant requirements.

Geostationary telecommunications satellite propulsion systems are typically used for orbit raising andcircularisation, north-south stationkeeping, east-west stationkeeping, and attitude control (momentum dumping).The above list is in order of decreasing total impulse requirement. As orbit raising and circularisation normally hasto be completed quite quickly (to ensure the satellite can start earning revenue as soon as possible after launch), andas EP typically provides very low thrust levels (less than 1N), its usefulness for primary spacecraft propulsion islimited in this context. Consequently, it has been the north-south stationkeeping requirement which has been themain focus for EP activities, where a high propellant throughput is required but low thrust levels can beaccommodated.

EADS Astrium have developed a plasma propulsion system (PPS) based around the Russian SPT-100 thrusterfor use on their Eurostar 3000 (E3000) platform1,2, and the Intelsat 10 spacecraft is the first commercial exploitationof this. This thruster has extensive flight experience with no reported failures on a number of Russian satellites (inparticular, the GALS and Express programmes, and more recently Sesat), and so is considered to be mature with arelatively low residual risk. The Western ground qualification for the overall PPS and most of its equipments waslargely achieved through the French Stentor programme3, although the loss of that satellite due to a launcher failuremeans the corresponding in-orbit experience on a European satellite has not been previously available. Groundqualification not provided by Stentor has been achieved through a dedicated EADS Astrium test programme,including full end-to-end testing of all the active elements of the PPS4.

* Electric Propulsion Specialist Engineer, EADS Astrium Ltd, Anchorage Road, Portsmouth, PO3 5PU, UK† Propulsion Engineer, Intelsat, 3400 International Drive NW, Washington, DC 20008

41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit10 - 13 July 2005, Tucson, Arizona

AIAA 2005-3672

Copyright © 2005 by EADS Astrium Ltd. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.


FUFUFU

PPU APPU B

XST

FDVHPXE

XFC XFC XFC XFC

FDVTPXE

Orientation mechanism (PY) Orientation mechanism (MY)

HET2PYThruster

HET1PYThruster

HET2MYThruster

HET1MYThruster

FU

LPT (2)

HPT (2)XMRS

FDVXE

PV

XEF

PY TMA MY TMA

TSU TSU

PressureRegulators

Isolation valves

GLPT

FUFUFU

PPU APPU B

XST

FDVHPXE

XFC XFC XFC XFC

FDVTPXE

Orientation mechanism (PY) Orientation mechanism (MY)

HET2PYThruster

HET1PYThruster

HET2MYThruster

HET1MYThruster

FU

LPT (2)

HPT (2)XMRS

FDVXE

PV

XEF

PY TMA MY TMA

TSU TSU

PressureRegulators

Isolation valves

GLPT

Figure 1. PPS schematic

Table 1. PPS Equipment Product TreeTitle Nº Off

PPS Plasma Propulsion SubsystemXST Xenon Storage Tank 1PV Pyro Valve 1XEF Xenon Filter 1XMRS Xenon Mechanical Regulation System 1

HPT High Pressure Transducer 2 / XMRSREG Mechanical Regulator 2 / XMRSLV Latch Valve 2 / XMRSLPT Low Pressure Transducer 2 / XMRS

GLPT Ground Low Pressure Transducer 1FDV Fill and Drain Valve 3PPSPIP Pipework (HP and LP) 1 setTMA Thruster Module Assembly 2

HET Hall Effect Thruster 2 / TMAXFC Xenon Flow Controller 2 / TMATOM Thruster Orientation Mechanism 1 / TMAFU Filter Unit 2 / TMAN/A TMA Thermal Device (MLI) 1 set / TMABP Base Plate 1 / TMAN/A TMA Pipework 1 set / TMAN/A FU - HET Harness 1 set / TMAHIB Hot Interconnection Box 2 / TMA

PPU Power Processing Unit (with TSU) 2EMH PPU - TMA Harness 1 set

Acronym

II. PPS Description

A. PPS OverviewThe function of the PPS is to provide inclination and eccentricity control for North/South station keeping

(NSSK) of the satellite. The PPS (shown schematically in Fig. 1) uses Xenon as propellant and includes all devicesto store and supply Xenon to the plasma thrusters (four SPT-100), which are accommodated by pairs on two

orientation mechanisms (TOM) and which are powersupplied by two electronic units (PPU). The PPUs areconnected to the Spacecraft Computing Unit (SCU) viaa digital 1553B interface, and the TOM is driven by theMechanical Drive Electronics (MDE) function of theActuator Drive Equipment (ADE). The latch valvedrivers and pressure transducer power and signalacquisition are also managed by the ADE. The PPUtakes power from the main spacecraft regulated bus.

The PPS feed system comprises a Xenon storagetank (XST), a pyro valve (PV) with its associated Filter(XEF), a mechanical pressure regulation system(XMRS), and three fill and drain valves (FDVs); thePPS is then completed by two thruster moduleassemblies (TMAs) and two power processing units(PPUs), along with their associated pipework andharnesses. The PPS equipments are listed in Table 1.

Each TMA comprises two SPT-100 thrusters and theirassociated Xenon Flow Controllers (XFC), a thrusterorientation mechanism (TOM) bearing and canting theSPTs (with its associated thermal hardware), a set of MLI for thermal control (with a dedicated structure), two FilterUnits (FUs, one per thruster), and the associated pipework and harnesses. The TOM and the XFCs are mounted ontoan interface base plate (BP), which can be shimmed with respect to the S/C structure. The FUs are directly mountedonto the satellite structure. The TMAs are located on the north and south panels of the service module (SM), withthe thrusters nominally canted at approximately 45o to the spacecraft Z axis, to ensure the thrust axes pass throughthe centre of mass of the spacecraft. The TOMs allow the thrust vector to be steered during operation; the TOMsteering functionality is handled at ADCS level, driving through the MDE as discussed earlier. Each TOM also has2 microswitches, which change state at the reference position (one for each axis of movement). Figure. 2 shows thePPS Satellite configuration for the TMAs.


Figure 3. SPT-100 Thruster

Thruster module (TMA) – MY unit shown

+Z

+Y

+X

Thruster module (TMA) – MY unit shown

+Z

+Y

+X

+Z

+Y

+X

Figure 2. Satellite Configuration

The pyro valve isolates the tank from the remainder of the PPS.Tank pressurisation operations are only performed using the FDVplaced between the tank and the PV, whereas all the otherpressurisation tests are performed using the two other FDVs. AfterPV firing, the XMRS then isolates the Xenon stored in the XST fromthe plasma thrusters; during the on-orbit phase, the XMRS suppliesXenon to the TMAs at the required regulated pressure.

The validation of the overall flight PPS and interfaces has beenachieved by means of rigorous and incremental testing at equipmentand spacecraft level, and finally by in-orbit tests prior to flightoperations.

B. Thruster Operation1. HET Operation Overview

The SPT-100 HETs are plasma thrusters using Xenon aspropellant with high thrust density. The thruster consists of thefollowing major elements (see Fig. 3 below):

• The anode gas distributor is a metallic annular assembly,which provides uniform Xenon distribution into thechamber through a series of small orifices. The anode gas distributor and the propellant feeding line areat a potential of +300 Volts during the discharge and therefore are isolated from the inlet gas supply linethrough two redundant electrical isolators

• The discharge chamber is an annular U shaped canal made of a ceramic Boron Nitride and SiliconDioxide mix, which insulates the thruster body from the plasma

• The magnetic system consists of a single internal and four external electromagnetic coils which areelectrically fed in series with the discharge, and of a magnetic permeable path unit to produce the radialmagnetic fields in the discharge chamber

Two redundant cathode assemblies are located outside the thruster body, each of which includes:• A getter which traps all the oxygen traces in the Xenon before feeding the high temperature core• Heating coils used during start-up phase to bring the device to the necessary temperature• A Lanthanum Hexaboride thermal emitter which, when heated to a high temperature, ensures electron

emission• Thermal screens positioned around the high temperature cathode core, to thermally protect the casing and

thermally control the cathode• An ignitor used to initiate the discharge


The plasma is created within the discharge chamber by means of electron bombardment of the neutral xenon gas.The electron source is the hollow cathode located outside the thruster exit plane. The two cathodes are provided inthe SPT-100 design for redundancy, but only a single cathode is used during operation. The free electrons emittedby the cathode are attracted to the 300V potential which is applied to the anode located at the bottom of thedischarge chamber. These electrons, once in the discharge chamber, are subjected to the radial magnetic field andbecome “trapped” in spiral paths. They then collide with the xenon gas particles, which are introduced to thedischarge chamber through the anode injector orifices. This results in the creation of positively charged xenon ionswithin the discharge chamber, which are then subjected to the potential difference of approximately 300V betweenthe anode and the exit plane of the thruster, and are thus strongly expelled from the thruster.

The neutralisation of this positive ion beam is performed by the hollow cathode. Electrons supplied by thecathode are naturally attracted to the positive ion beam; this is a self-regulating process with no need for activeexternal control. The same hollow cathode is used to provide both the electrons for the main discharge plasma andthe neutralizing electrons.2. Thruster Control

The overall operational sequence of thePPS is controlled by the spacecraft on-boardsoftware, via the PPUs. Each PPU controlsthe selected SPT and its associated XFC onthe basis of programmed procedures andcommands received from the on-boardcomputer.

The XFC (shown in Fig. 4) consists ofvalves, filters and a thermally constrictingcapillary tube (thermothrottle) in series withflow restrictors to control the propellantflow ratio between the operating cathodeand the anode unit. The XFC modulesprovide independent flow paths for theanode and for each cathode. On each flowpath, two valves in series are provided. ThePPU provides a closed loop Xenon flowregulation to the thruster by adjusting thethermothrottle current; the density andviscosity of the gas are varied in order tokeep a constant delivered mass flow rate,and so the discharge current is maintained atits nominal set value (which maintains afixed nominal thrust level).

Each PPU supplies a pair of SPT-100thrusters (and their associated XFCs) withelectrical power to enable their correctoperation. The Thruster Selection Unit(TSU), which is included in the PPU, is aswitching unit receiving all PPU outputs andtransmitting them to one of two thruster;each PPU can only operate one thruster at atime. The PPU is provided by ETCA.

The PPUs exchange data with the on-board computer. According to the receivedorders the PPU enters into one of its sixpossible modes: OFF, STAND-BY,CONFIGURATION, VENTING,AUTOMATIC and REMOTE. Thesetransitions are as follows (see Fig. 5):

After the DTC (direct telecommand)“PPU ON” to turn ON the PPU DC/DC, the

s

s

s

s

s

ss

XFC SPT-100

Solenoide Valve

Screwed Connection

Flow Restrictor

Inox Tubing Weld Beam

Thermothrottle

Filter

Titanium Tubing

Figure 4. XFC Schematic

VENTING

From any state

STAND-BY

OFF

CONFIGURATION

AUTOMATICREMOTE

PPU OFF

reset

PPU ON

Configuration disable

Thruster ONRemote OFF

Venting OFF

Venting ON

Remote ON

Thruster OFF

Configuration enable

Figure 5. PPU States


sequencer enters into an initialisation phase before being put in stand-by mode.In the “OFF” mode the only acceptable command by the PPU is the ON command which puts the internal micro-

controller into the “STAND-BY” mode.In the “STAND-BY” mode, only low level electronics are active. The PPU then accepts any transition to

another mode.In the “CONFIGURATION” mode the PPU accepts orders to modify the selection of SPT and XFC units by

actuating the TSU relays. It is also possible to modify the setting of firing parameters.The “VENTING” mode is only used at the mission beginning to allow venting of the Xenon lines before the first

SPT firing. In this mode the PPU opens all the XFCs valves on the selected thruster.In the “REMOTE” mode the PPU micro-controller can be overridden by dedicated commands. The on-board

computer allows “step-by-step” management of the thrusters in order to provide flexibility to the ground operators ifnecessary.

The “AUTOMATIC” mode is the nominal mode of the PPU when the PPU drives the firing SPT in closed-loop.The power supplies of the PPU are active. The firing is driven with the set of parameters defined in theconfiguration mode. The logic of the “AUTOMATIC” mode is as follows:

• During the starting phase, the PPU sequences as follows:- Heat the selected cathode and thermothrottle up to the corresponding stabilised thrust level for the SPT- Open the XFC valves- Supply the cathode with ignition pulses once the Xe flow is stabilised

• The SPT has to be started within a given delay period from the start of this sequence. In case of failure thePPU goes into “STAND-BY” mode and sends a failed status signal

• During the stabilised phase (after SPT starting), the initiator electrode is no longer supplied and the closedcontrol loop becomes operational. In stabilised mode the discharge current control is made by adjusting theXenon flow rate through the thermothrottle in the XFC

III. Initial Flight Operations

A. Operations SequenceIntelsat 10-02 was launched from Baikanour Cosmodrome on 16 June 2004 aboard Proton/Breeze M. Following

the successful launch of the satellite, the PPS was initialised and tested in orbit. This sequence included thefollowing steps:

• Xe storage and regulation system health checks (pressure and temperature monitoring)• Release of the pointing mechanisms, and check of successful release by motion checks• Venting the low pressure section of the PPS in order to expel any contaminants, and refilling with flight-

quality Xenon• Firing each thruster and cathode combination

Eight firing tests were performed in total (two for each thruster, with the prime and then redundant cathode);both the prime and redundant branch of the regulation system were also used during this test programme. Eachfiring was approximately 60 minutes in duration.

The Eurostar 3000 platform allows a number of telemetry channels to be acquired at high frequency (up to10Hz); this enables a reasonably well detailed investigation of the thruster start-up transients in the real spaceenvironment, as discussed below.

B. Initial In-Orbit Operations Results1. Xe System Health Check

The Xe tank and regulation systems were checked immediately after launch. An initial partial vent wasenvisaged within 24 hours of launch, in case the pressure in the low pressure section of the PPS exceeded a presetthreshold (due to XMRS internal leakage or temperature excursions). However, this was not required, as themeasured pressure remained below this level. LV exercising was still performed immediately after launch (each LVopened and then closed), to ensure the sealing was good prior to LEOP.


The Xe feed system was thenchecked immediately prior to PPSventing and initialisation (see below),with the results as shown in Table 2.

Note that the LPTs error could notbe checked correctly at this stage asthey saturate above 4 bar (as can beseen in Table 2); similarly, thedifference between the GLPT and LPTscould not be properly checked at thisstage. These checks were finalisedduring the venting activities asdescribed below. Otherwise, all valuesare in the expected range, confirmingthe good health of the feed system priorto any PPS operations.2. TOM Release

TOM unlocking occurred on 24th June 2004 at 21:10 GMT. Each TOM was successfully released by firing theprime and redundant initiators simultaneously on each pyro release device in turn (there are 2 pyro release deviceson each TOM). This was then followed by a reference position search, looking for the microswitch state changes asthe TOM is moved.3. Venting and XMRS Initialisation

This activity was performed on 29th June 2004, starting at approximately 13:00 (Washington local time). At thestart of this activity, the LP section pressure was approximately 6.2 bar as measured on GLPT; LPT1 and LPT2 weresaturated. During the venting and priming, the LPT readings were all acquired in dwell. The vent proceeded in 4stages:

• PPUA, YM thruster: 1 hour vent• PPUA, YP thruster: 30 minute vent (originally planned for 1 hour, but full period not needed)• PPU B, YM thruster: 30 minute vent (originally planned for 1 hour, but full period not needed)• PPU B, YP thruster: 4 hour vent; effectively hard vacuum achieved at end of this activity

This activity was completed with the priming of the LP section, as follows:• Ensure all XFC shut• Open prime LV• Open redundant LV• Close redundant LV

During the LV opening sequence, the peak pressure in the LP section indicated that the regulator slam startresulted in a pressure overshoot of just 50 mbar.

The difference between the LPT and GLPT readings during the post-vent priming is due to the GLPT error (thiscan be up to 200 mbar). In Fig. 6 which shows the priming activity, it should be noted that a TM high value (around2) corresponds to the LV being closed, and a value of around zero corresponds to the LV being open.

The pressure transducer and thermistors were checked for consistency throughout the above operations; allreadings were within the range of measurement accuracy, as detailed in Table 3.

During the priming transients (when the pressure is changing most rapidly), there were some minor deviations tothe above.

The Xenon tank pressureand temperature conditionswere also recorded duringthis activity; no significanttrend in tank conditions wasseen during the venting.

Table 2. Xe System Initial Health CheckPass / Fail criteria

ParameterMin Max

Result

LV1 statusLV2 status

ClosedClosed

ClosedClosed

XMRS prime temperatureXMRS redundant temperature

1414

6363

40.10640.495

XMRS temperature difference -6 +6 -0.389HPT1HPT2

121.71121.5

83.05782.989

HPT difference -3.21 +3.21 0.068LPT1LPT2

-0.035-0.035

4.0354.035

4.0004.000

LPT difference -0.07 +0.07 0.000GLPT -0.2 +7.2 6.125

Table 3. Pressure and Temperature Consistency During Vent

Pass / Fail criteria ResultsParameter

Min Max Min MaxTemperature difference (°C) -6 +6 -0.597 -0.247HPT difference (bar) -3.21 +3.21 -0.108 0.244LPT difference (bar) -0.07 +0.07 -0.026 0.000GLPT error (bar) -0.2 0.2 -0.111 0.000


4. Initial Thrust FiringsThe test consisted of firing each of the 4 thrusters using prime and then redundant cathodes. The sequence of the

firings is detailed in Table 4.All these tests were completed successfully, with no unexpected behaviour. The results for all 8 firings were

similar; curves illustrating the main performance characteristics for firings 1 and 5 are shown below.

A number of points are of particular interest:• The cathode on each thruster is electrically floating, but during thruster operation is tied to the external

plasma conditions; measurement of the cathode reference potential (CRP) then gives some indication ofthe environment potential of the spacecraft. In general, the CRP was positive, but on a number of firingsthe TM was saturated

• For a few minutes after thruster ignition, there is significant noise on a number of the telemetries, inparticular the discharge current and thermothrottle current (as can be seen below); this is associated without-gassing of the thruster ceramic, and is a well-known phenomenon

0

1

2

3

6.4 6.41 6.42 6.43 6.44 6.45 6.46 6.47 6.48 6.49 6.5

Time (hours)

Pre

ssur

e(b

ar)

LPT1 LPT2 GLPT Prime LV status Redundant LV status

Figure 6. Post-Vent Priming

Table 4. Initial Thruster Firing Sequence

FiringDate and Time

(GMT)Thruster

Thrusterposition

PPU CathodeXMRS

branch in use1 30 June 2004 18:00 #59 HET2MY A A Prime2 1 July 2004 06:00 #60 HET1MY B A Prime3 1 July 2004 18:00 #74 HET2PY B A Prime4 2 July 2004 06:00 #54 HET1PY A A Prime5 3 July 2004 00:00 #59 HET2MY A B Redundant6 3 July 2004 12:00 #60 HET1MY B B Redundant7 4 July 2004 00:00 #74 HET2PY B B Redundant8 4 July 2004 12:00 #54 HET1PY A B Redundant


• The duration and amplitude of the initial transients and noise were seen to be decreasing from thebeginning to the end of the IOT, indicating a near completion of out-gassing activity

In addition to the above, all other PPS TM have been checked for all the above firings; no unexpected valueshave been seen.

Throughout each firing, the telemetry has enabled the calculation of the thrust, flow rate and specific impulse.The thrust and flow rate are shown in the above curves; it can be seen that the flow rate in particular has to betreated with some caution during the initial out-gassing.

0

50

100

150

200

250

300

350

-5 0 5 10 15 20 25 30 35

Time from Preheat Start (Minutes)

Ano

deV

olta

ge(V

)

-20

-15

-10

-5

0

5

10

15

CR

P(V

)

Anode voltage CRP

0

1

2

3

4

5

6

-5 0 5 10 15 20 25 30 35


Cur

rent

Anoode current (A) Thermothrottle current (A) Oscillation current (mA)

2.5

2.55

2.6

2.65

2.7

-5 0 5 10 15 20 25 30 35


Pre

ssur

e(b

ar)

LPT1 LPT2 LPT average

0

10

20

30

40

50

60

70

80

90

100

-5 0 5 10 15 20 25 30 35


Thr

ust

(mN

)

0

1

2

3

4

5

6

7

8

9

10

Flo

wR

ate

(mg/

s)

Thrust Flowrate

Figure 8. Manoeuvre 5 Results and Performance

0

50

100

150

200

250

300

350

-5 0 5 10 15 20 25 30 35


Ano

deV

olta

ge(V

)

-20

-15

-10

-5

0

5

10

15

CR

P(V

)

Anode voltage CRP

0

1

2

3

4

5

6

-5 0 5 10 15 20 25 30 35


Cur

rent

Anode current (A) Thermothrottle current (A) Oscillation current (mA)

2.50

2.55

2.60

2.65

2.70

-5 0 5 10 15 20 25 30 35


Pre

ssur

e(b

ar)

LPT1 LPT2 LPT average

0

10

20

30

40

50

60

70

80

90

100

-5 0 5 10 15 20 25 30 35


Thr

ust

(mN

)

0

1

2

3

4

5

6

7

8

9

10

Flo

wR

ate

(mg/

s)

Thrust Flowrate

Figure 7. Manoeuvre 1 Results and Performance


0.05

0.07

0.09

0.11

0.13

0.15

0.17

0.19

0.21

0.23

0.25

08/01/2004 09/20/2004 11/09/2004 12/29/2004 02/17/2005 04/08/2005

Del

taV

-RS

S(m

/s)

0

20

40

60

80

100

120

140

160

180

Cu

mu

lati

veM

aneu

ver

Du

rati

on

(ho

urs

)

H2PA-DV

H2MA-DV

H2PB-DV

H2MB-DV

H2PA-Duration

H2MA-Duration

H2PB-Duration

H2MB-Duration

Figure 9. Delta-V and Cumulative Thrusting Durations of the 10-02 PPSsince 27 August 2004

The performance figures and firing timing data for all 8 firings are shown below, along with the correspondingthruster acceptance test data for comparison. The in-orbit data can be seen to correlate well with the equivalentground test data; the maximum discrepancies are 0.21mN thrust and 78s specific impulse (which should becompared with allowable tolerances of 3.2mN and 96s on thrust and specific impulse respectively due to themodelling and TM errors); hence all the in-orbit test data can be considered to be well within the allowabletolerances compared to acceptance tests.

5. Summary of In-Orbit Test ResultsAll in-orbit tests were successfully completed, with all the data analysed and shown to be as expected within the

limits of the TM accuracy.

C. Operational Use of PPSRegular North-South Station Keeping manoeuvres on Intelsat 10-02 began on 27th August 2004, with eight (four

North and four South) manoeuvres per weekly cycle. Both North-South and East-West manoeuvres are performedautomatically by onboard software, based on the manoeuvre plans uploaded every week.

Intelsat 10-02 is equipped to impart corrections for (1) variation in inclination due to lunar-solar attraction; (2)variation in the major axis and in the mean longitude due to the geopotential; and (3) variation in eccentricity due tosolar radiation pressure.Taking advantage of thelarge cant angle of the PPSthrusters, the eccentricitycontrol is achieved duringNorth-South station-keepingmanoeuvres. The delta-vdemands on these thrusterstherefore vary with the yearof operation (for inclinationcontrol) and the seasonalvariation in the eccentricitygrowth, as shown in Figure9. The cumulative hours offiring recorded by the primeYM (H2M) and YP (H2P)thrusters with both A and Bcathodes are also shown inFigure 9. It can be seen thatthe hours accumulated bythe thrusters on two sidesdiffer significantly due tothe eccentricity corrections.

Table 5. Thruster Performance Summary

Thruster Firing time (min) Average overrun

Averagestabilised

Acceptancetest

Location Serial noManoeuvre Cathode

Total Outgas Stabilised Thrust Isp Thrust Isp Thrust Isp

1 A 31.0 4.9 26.1 83.84 1570 83.78 1544 83.9 1623HET2MY #59

5 B 31.4 3.1 28.2 82.91 1589 82.88 1582 83.0 1590

2 A 33.6 9.1 24.5 84.10 1614 84.07 1578 84.2 1613HET1MY #60

6 B 33.5 0.6 33.0 84.09 1607 84.17 1610 84.3 1612

3 A 30.8 5.2 25.5 84.38 1608 84.37 1587 84.5 1610HET2PY #74

7 B 30.5 4.6 25.9 84.33 1611 84.07 1584 84.2 1601

4 A 34.0 7.5 26.4 83.23 1564 83.17 1525 83.3 1575HET1PY #54

8 B 34.0 1.3 32.6 82.94 1602 82.97 1601 83.1 1575


Figure 10. Peak Oscillation Current values

The standard practice is to periodically alternate usage of prime and redundant propulsion equipment, to preventlong term deposition of sputtered plume contaminants on the unused units, perform periodic health checks, and toensure uniform performance levels. Redundant cathodes on prime thrusters were put in use on 30 December 2004,after the first four months of PPS operation. All the performance parameters were within specification limits duringmore than 200 hours of manoeuvre time accumulated on each of the prime plasma thrusters.

When a plasma thruster is in operation, it produces quasi-periodic oscillations on anode current at a frequency of10-70 kHz. These oscillations are filtered by a capacitor in the Filter Unit, which also houses discharge circuits tosuppress electrostatic discharge. It can be seen from Figure 10 that the oscillation current, which is the RMS valueof the discharge alternative current and common mode currents resulting from thruster plume and plasma presence,stabilizes gradually after about 150-200 hours of thruster operation.

The thrust and xenon flow rate values measured during the thruster acceptance tests were used for modeling thein-orbit performance parameters. As a function of the discharge current (Id), the stabilized thrust is found to havealmost a constant value, since the variation in discharge current, after the initial transients, is negligible within thetelemetry resolution. Xenon flow rate is modeled as a function of the thermothrottle current (Itt) and the variationsin the inlet pressure (Pxe). The mass flow rate and the specific impulse (Isp) derived from the telemetry records ofId, Itt and Pxe are shown in Figure 11. Small perturbations in the thermothrottle current result in correspondingvariations in the mass flow rate, which appear amplified, when used in the Isp computations. It can be seen that thethermothrottle in the Xenon Flow Controller (XFC) used for supplying xenon to Cathode-B of the YP (H2PB)thruster draws about 13% more current while the flow rate remains unchanged before and after the cathode switchover (30 December 2004). This characteristic of the individual thermothrottle capillaries operating at slightlydifferent current values to regulate the mass flow rate and hence the discharge current at almost constant levels ismainly attributed to the manufacturing tolerances in the capillary dimensions.


Figure 11. Thermothrottle Current, Xenon Flow Rate and the Specific Impulse

IV. ConclusionsThe Intelsat 10 PPS has performed as expected throughout all the initial in-orbit testing and subsequent

stationkeeping operations. The regulator, thruster and cathode combinations used during the IOT phase haveensured that all aspects of the PPS have been thoroughly checked out. The thruster behaviour has been as expected,with some outgassing at beginning of each firing showing up as current oscillations, which reduce as the firing timeon each thruster is accumulated. All the thruster performances deduced from the available telemetry are consistentwith the corresponding ground acceptance test data..

V. References1Gray, H., Demaire, A, and Didey, A., “Plasma Propulsion Systems For Astrium Telecommunications Satellites”, AIAA-

2003-6202, September 20032Stephan, J-M., “Electric propulsion activities for Eurostar 3000”, 3rd International Conference on Spacecraft Propulsion,

Cannes, October 2000, ESA SP-465, December 20003Cadiou, A., Gelas, C., Darnon, F., Jolivet, L. and Pillet, N, “An Overview of the CNES Electric Propulsion Program”, 27th

International Electric Propulsion Conference, Pasadena, California, 20014Stephan, J-M., Drube, M., Gray, H., and Demaire, A., “Plasma Propulsion System functional chain validation on Eurostar

3000”, AIAA-2003-2408, April 2003

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