Title: FEEP Feasibility Report Prepared by: Bernd...

Technical Note HYPER

Doc. No: HYP-5-02 Issue: 3.0Date: 02/25/2003 File: HYP-5-02_v30final.doc

Title: FEEP Feasibility Report

Prepared by: Bernd Schuerenberg Date: 25/02/2003

Project Management: Walter Fichter

Distribution: See Distribution List

Copying of this document, and giving it to others and the use or communication of the contents there-of, are forbidden without express authority. Offenders are liable to the payment of damages. All rightsare reserved in the event of the grant of a patent or the registration of a utility model or design.

astrium GmbH FEEP Feasibility Report HYPER

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Change Record

Issue Date Sheet Description of Change Release1.02.0

3.0

26/09/0230/01/03

25/02/03

AllAll

-

first draft issue for Progress Meeting # 3chapters 1-5 edited and updated;completion of chapters 6 - 9 (as part B of this report)Re-edited version


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Table of Contents

1 INTRODUCTION AND SCOPE 2

1.1 INTRODUCTION 21.2 SCOPE 3

2 DOCUMENTS AND DEFINITIONS 4

2.1 APPLICABLE DOCUMENTS 42.2 REFERENCE DOCUMENTS 42.2.1 HYPER 42.2.2 SMART-2 42.2.3 GOCE 52.2.4 CS SLIT EMITTER DOCUMENTS 62.2.5 IN NEEDLE EMITTER AND IN CAPILLARY EMITTER DOCUMENTS 62.2.6 COLLOID THRUSTER DOCUMENTS 72.2.7 NEUTRALISERS 72.2.8 OTHER REFERENCE DOCUMENTS 72.3 DEFINITIONS 82.3.1 CENTRE OF MASS AND DRAG-FREE POINT 82.3.2 CO-ORDINATE SYSTEMS 82.4 ACRONYMS 10

3 FEEP PROPULSION SYSTEMS: TECHNOLOGIES 13

3.1 INDIUM THRUSTERS 133.1.1 INDIUM NEEDLE EMITTERS 133.1.2 INDIUM MULTI-CAPILLARY EMITTERS 233.2 CAESIUM THRUSTERS 253.3 NEUTRALISERS 29

4 REQUIREMENTS AND CONSTRAINTS FOR THE FEEP SYSTEM 31

4.1 FEEP APPLICATION FOR HYPER 314.1.1 FEEP THRUSTER USAGE 314.1.2 DISTURBANCE FORCES AND DISTURBANCE MOMENTS 324.1.3 BASELINED THRUSTER ARRANGEMENT FOR HYPER 334.1.4 HYPER REQUIREMENTS ON FEEP SYSTEM 384.2 SMART-2 REQUIREMENTS 384.2.1 SMART-2 THRUSTER ARRANGEMENT 384.2.2 SMART-2 REQUIREMENTS SUMMARY 404.3 GOCE REQUIREMENTS 444.3.1 GOCE THRUSTER ARRANGEMENT 444.3.2 GOCE REQUIREMENTS SUMMARY 45


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4.4 COMPARISON OF REQUIREMENTS 494.5 CONCLUSIONS 54

5 FEEP TRADE-OFF CRITERIA 55


Doc. No: HYP-5-02 Page 2Issue: 3.0Date: 02/25/2003 File: HYP-5-02_v30final.doc

1 INTRODUCTION AND SCOPE

1.1 Introduction

HYPER is one of several new scientific missions, which depend on the availability of suitable Micro-NewtonPropulsion Systems.

Micro-Newton Propulsion Systems based on FEEP thruster technologies are characterised by

� commandable thrust levels of high repeatability, starting at levels of approx. 0.1 micro-N

� high resolution / quantisation in the order of 0.1 micro-N

� low thruster noise

� fast response.

The following scientific missions will use and / or depend on the availability of suitable Micro-NewtonPropulsion Systems:

� MICROSCOPE (CNES)

� GOCE (ESA)

� SMART-2 (ESA / NASA)

� HYPER (ESA)

� LISA (ESA / NASA)

� DARWIN (ESA)

Also a number of NASA missions are planned, for which FEEPs will be high-ranking candidates as micro-propulsion systems, such as:

� Earth Science Experimental Mission EX-5

� Laser Interplanetary Ranging Experiment LIRE

� Terrestrial Planet Finder TPF

� Micro-Arcsecond X-Ray Interferometer Mission MAXIM

� Submillimeter Probe of the Evolution of Cosmic Structure SPECS.

In RD [8-01], four alternative micro-propulsion systems were analysed and evaluated for five different futuremissions:

� FEEP thrusters (Cs-Slit and In-Needle)

� Colloid FEEP Thrusters

� Pulsed Plasma Thrusters



� Micro-N Cold Gas Thrusters.

1.2 Scope

HYPER will be equipped with two propulsion systems:

� a small Cold Gas Propulsion System (CGPS) for initial attitude acquisition, for attitude control of HYPERuntil the FEEP system becomes operational, and for attitude re-acquisition in emergency cases,

� a FEEP Micro-Newton Propulsion System for the Science Phase of the mission.

This report addresses the FEEP Micro-Newton Propulsion System.

The study task "FEEP Feasibility" covers the following aspects:

� description of different micro-N thruster technologies based on FEEP technology, taking also intoaccount potential heritage and the expected maturity of technology

� summary of FEEP requirements for HYPER, SMART-2, and GOCE. For the detailed requirementspecification for HYPER see HYP-5-01: "FEEP Requirements Specification for HYPER"

� comparison of requirements and identification of significant differences that may result in technologydrivers

� the definition of trade-off criteria to compare different Micro-Newton Propulsion Systems based on FEEPtechnology.

In addition, a comprehensive list of reference documents is given that may serve as an overview of theavailable documentation.



2 Documents and Definitions

2.1 Applicable Documents

The following documents are applicable to the extent outlined in chapter 3.

AD Doc. No. Title

AD-01 HYP-2-05 Performance Requirements Breakdown for the HYPER Satellite

AD-02 HYP-FEEP-RS-SPA-1 FEEP Performance Specification;issue 1, 31 Jan 2002

AD-03 HYP-ESA-SOW-SPA-1 Statement of Work, HYPER Industrial Initial Feasibility Study;issue Draft (B)

AD-04 EHB-003 ROCKOT User's Guide;iss. 3, rev. 1; April 2001

AD-05 HYP-1-01 Orbit Trade-off Report

AD-06 HYP-2-01 Secondary AOCS Design

2.2 Reference Documents

Not all of the following documents are referenced within this note. The list serves as a useful overview ofdocuments about micro-N propulsion systems / technology.

2.2.1 HYPER

RD Doc. No. Title

1-01 HYP-ASU-DD-CST-1 Description of the Atomic Sagnac Unit, Issue 1, 31 January 2002

1-02 CDF-09 Hyper CDF Study Report (as amended by Errata Corrigum refHYP-CDF-E/C-1); Sept 2000

1-03 HYP-ESA-RS-SPA-3 Hyper Payload Requirement Specification, Issue 2, 7 June 2002

2.2.2 SMART-2

RD Doc. No. Title

2-01 SMT2.RP.005.EU.AST SMART-2 System Definition Study Final Report;issue 1, 2002



2-02 SM2-CAS-4100-TNO-003 Micro-Propulsion Trade-off;11 April 2002

2-03 SM2-CAS-5120-TNO-003 Mechanical Design Description;28 June 2002

2-04 MPE/380/DN Development of FEEP Neutralisers (Statement of Work);issue 3, 05 Feb 2001

2-05 SP-00000499-00-01 FEEP PCU Specification (ALTA thrust and RAL neutraliser);issue: 1 draft; 16 Sept 2002

2-06 ALTA/CL/SP-06 Development of Integrated FEEP Cluster Systems:Cluster Assembly Requirements Document;issue: 2; 29 March 2002

2-07 ALTA/LF/EA-01 Low thrust Propulsion System Characterization and Life Testingfor LISA and DARWIN:FEEP Preliminary Dynamic Model;issue 2; 22 May 2002

2-08 Direct communication Outline Drawings of the SMART-2 FEEP Cluster(status Dec. 2002)

2.2.3 GOCE

RD Doc. No. Title

3-01 GO-RQ-ASG-0008 Micro Thruster Assembly (MTA) Requirement Specification;issue 4 Draft, Dec. 2002

3-02 GO-RQ-ASG-0019 GOCE FEEP PCU Requirements Specification

3-03 GO-TN-ASG-0001 FEEP System Design Description,included in GOCE Platform Design Description

3-04 GOCE FEEP Propellant Budget

3-05 GO-TN-ARC-MP-0003 Plume Model,issue Draft, 17 Jan 2002

3-06 GO-MTA-ARCS-TN-001 Preliminary Test Results; Direct Thrust Measurement and ClusterTesting,issue: Draft, 09 July 2001

3-07 GO-DD-SDP-001 MTA Design Definition Report,issue: Draft, 19 June 2001



2.2.4 Cs Slit Emitter Documents

RD Doc. No. Title

4-01 ALTA/LF/EA-01 Low Thrust Propulsion System Characterisation and Life Testingfor LISA and DARWIN -- FEEP Preliminary Dynamic Model;issue2, 22 May 2002

4-02 ALTA/CL/SP-06 Development of Integrated FEEP Cluster Systems (ClusterAssembly Requirements Document);issue 1, 14 Feb 2002

4-03 ALTA/P-06.1/G Micro-Newton Thruster Assembly for the GOCE Platform(Technical Proposal)issue 1, 22 June 2001

4-04 JPP Vol. 14, No. 5,Sept-Oct 1998

M. Marcuccio, A. Genovese, M. Andrenucci:Experimental Performance of Field Emission Microthrusters

2.2.5 In Needle Emitter and In Capillary Emitter Documents

RD Doc. No. Title

5-01 AIAA 2002-3688 A. Genovese, M. Tajmar, N. Buldrini, W. Steiger:Extended Endurance Test of the Indium FEEP Micro-Thruster

5-02 -- M. Thajmar, A. Genovese, W. Steiger:Indium FEEP Microthruster Experimental Characterisation;Journal of Propulsion and Power, submitted 2002

5-03 MPE/404/DN ARCS Indium Liquid Metal Ion Source Characterization TestReportissue 1, 22 Oct 2001

5-04 MAGNA STEYRproposal # 2825

Development and Supply of a Micro-N Thruster Assembly for theGOCE Platform (Technical Proposal);19 June 2001

5-05 ARCS FEEP Infos (1) First 2,000 h Endurance Test of an Indium FEEP CLUSTER15 July 2002

(2) First Extended Endurance Test of an Indium FEEP CLUSTER30 October 2002

(3) 3500 h operation with an Indium FEEP thruster10 January 2003

5-06 AIAA-2002-5718 M. Tajmar, A. Genovese, W. Steiger:Indium-FEEP Microthruster: Experimental Characterisation



5-07 -- Electric Propulsion Plasma Simulations and Influence on S/CChargingAIAA Journal of S/C and Rockets, Vol 39, no. 6 (2002), 886 - 893

5-08 -- Direct Thrust Model Assembly (MAGNA Steyr drawing)

5-09 3rd International Confe-rence of Spacecraft Pro-pulsion, Cannes; 2000

W. Steiger, A. Genovese, M. Tajmar:Micronewton Indium FEEP Thrusters

5-10 AIAA-2002-5718 M. Tajmar, A. Genovese, N. Buldrini, W. Steiger:Miniaturized Indium FEEP Multiemitter Design and Performance

5-11 Applied Physics A 76, 1-4(2002)

M. Tajmar, A. Genovese:Experimental validation of a mass efficiency model for an indium-metal ion source

2.2.6 Colloid Thruster Documents

RD Doc. No. Title

6-01 IEPC-99_014 M. Martinez-Sanchez, J. Fernandez de la Mora; V. Hruby, M.Gamero-Castano, V. Khyms:Research on Colloid Thrusters

6-02 JPL Presentation forSMART-2

Bill Folkner / JPL:DRS Architecture;28 May 2002

2.2.7 Neutralisers

RD Doc. No. Title

7-01 AIAA-2002-4243 M. Tajmar: Survey on FEEP Neutralizer Options;July 2002

7-02 MPE/380/DN Development of FEEP Neutralisers (Statement of Work);ESA, issue 3, 05 Feb 2001

2.2.8 Other Reference Documents

RD Doc. No. Title

8-01 SERC #1-01 J.G. Reichbach, R.J. Sedwick, M. Martinez-Sanchez:Micropropulsion System Selection for Precision Formation FlyingSatellites;Jan 2001



2.3 Definitions

2.3.1 Centre of Mass and Drag-Free Point

According to [AD-01], the Drag-Free Point (DFP) shall lie in the middle of the intersection line of the two ASUplanes, as shown in Figure 2.3-1. Ideally, the spacecraft CoM and the DFP should coincide. In case ofHYPER, the large mass of the Payload Module will not permit such a perfect matching.

The assumptions made with respect to S/C axes and Payload Module orientation are shown in Figure 2.3-1.Further, it is assumed that SC_Y_p points into local zenith direction, and SC_Z_p points into flight directionat Ascending Node. These preliminary assumptions are used in the discussion and calculation of thedisturbance forces and torques in sect 4.1.2.

Figure 2.3-1: Spacecraft Physical Coordinates, Centre-of-Mass (CoM), and Drag-free Point (DFP)(CoM shown in black; DFP shown in red)

2.3.2 Co-ordinate Systems

[A] Spacecraft Physical Co-ordinate System (see Figure 2.3-1):

SC_O_p origin is the centre of the Adaptor Ring plane (= separation plane Launcher / Spacecraft)

SC_X_p perpendicular to the adaptor ring plane; pointing negatively from centre of separation planethrough centre of Solar Array (i.e. parallel to PST_X_f)

SC_Z_p

SC_Y_p



SC_Y_p normal to SC_X_p; (assumption: SC_Y_p = normal of the side panel pointing to local zenithat Ascending Node)

SC_Z_p completes the right hand co-ordinate system(assumption: SC_Z_p = normal of the side panel pointing in flight direction at AscendingNode)

Figure 2.3-2: Spacecraft Functional Coordinates, Centre-of-Mass (CoM), and Drag-free Point (DFP)(CoM shown in black; DFP shown in red)

[B] Spacecraft Co-ordinate System (see Figure 2.3-2):

SC_O_f origin is the centre of the separation plane Launcher / Spacecraft

SC_X_f perpendicular to the adaptor ring plane; pointing negatively from centre of separation plane through centre of Solar Array (parallel to PST_X_f)

SC_Y_f normal to SC_X_f; (assumption: SC_Y_f � nadir pointing at Descending Node)

SC_Z_f completes the right hand co-ordinate system(assumption: SC_Z_f � flight direction at Ascending Node)

SC_Z_fSC_Y_f



[C] FEEP Propulsion Co-ordinate System (see Figure 2.3-3):

FEEP_O_p origin is the Center-of-Mass (CoM)

FEEP_X_p parallel to SC_X_p

FEEP_Y_p parallel to SC_Y_p

FEEP_Z_p parallel to SC_Z_p (completes the co-ordinate system).

FEEP_Y

FEEP_ZFEEP_X

Solar Array

2b

2aCOM

origin in CoM

Figure 2.3-3: Co-ordinate System for the FEEP Propulsion System

2.4 Acronyms

AD Applicable Document

ASU Atomic Sagnac Unit

ATOX Atomic Oxygen

BOL Begin of Life

BoM Begin of the (scientific) Mission

CA Cluster Assembly (of Cs-Slit FEEPs; term used by SMART-2)

CGPS Cold Gas Propulsion System



CM Calibration Mode

CMNT Colloid Micro-N Thruster

CNT Carbon Nanotube

CoM / CoG Centre of Mass / Centre of Gravity

Cs Caesium

CX Charge Exchange (ions)

DFAC Drag-Free and Attitude Control

DFP Drag-Free Point

DFS Drag-Free Sensor

DOF Degree-of-Freedom

EM Engineering Model

EoL End-of-Life

EoM End of (scientific) Mission

FEA Field Emission Array (technology of Electron emitting cathodes)

FEEP Field Effect Electric Propulsion

GG Gravity Gradient

GN2 gaseous nitrogen

GOCE Gravity and Steady-State Ocean Circulation Earth Explorer

HCM Health Check Mode

HV High-Voltage

HYPER Hyper-Precision Cold Atom Interferometry in Space

In Indium

LISA Laser Interferometer Space Antenna

LMNIS Liquid Metal Needle Ion Source

LTP LISA Technology Package

MAIT Manufacturing, Assembly, Integration and Test

MBW Measurement Bandwidth

MEMS Micro Electro-Mechanical System

MOI Moment of Inertia

MTA Micro Thruster Assembly (term used by GOCE) (consists of one ion emitter)

MPA Micro Propulsion Assembly (term used by GOCE) (MPA consists of several ionemitters pointing into a common thrust direction; equivalent to a 'thruster')



MPE Micro Propulsion Electronics (term used by GOCE) (equivalent to PCU)

OB Optical Bench

PM Proof Mass

PS Propulsion System

PST Precision Star Tracker

QCM Quartz Crystal Microbalance (for contamination measurements)

QM Qualification Model

Rb Rubidium

RCS Reaction Control Subsystem (see PS)

RD Reference Document

SAA Solar Aspect Angle

SAM Secondary AOCS Mode

S/C Spacecraft

scc standard cubic centimetre (i.e. at 1 bar)

SCM Science Mode

SMART-2 Small Mission for Advanced Research & Technology

SSO Sun Synchronous Orbit

STBY Standby Mode

STEP Satellite Test of the Equivalence Principle

TBC to be confirmed

TBD to be defined

TBR to be reviewed



3 FEEP Propulsion Systems: Technologies

In this chapter, the existing FEEP technologies available in Europe are described.

3.1 Indium Thrusters

Two different types of Indium thrusters / emitters have been developed, respectively are in development atARCS / MAGNA:

� In-needle emitters. A suitable number of independently supplied and controlled emitters must beclustered to achieve the required maximum thrust level.

� In-multi-capillary thruster. This thruster is operated from one supply only.

3.1.1 Indium Needle Emitters

Indium-needle emitters (see schematic in Figure 3.1-1and Figure 3.1-2) consist of an indium-wettedstainless-steel needle protruding out of a liquid metal reservoir. The following voltages are applied:

Figure 3.1-1: Indium FEEP Thruster Principle Figure 3.1-2: Indium FEEP Thruster Schematic

� Needle + 5 - 9 kV

� Extractor Electrode ground

� Plume Shield (collector) ground



� Heater 10 V vs. ground.

The heaters of the Indium reservoirs are electrically isolated to the HV of the reservoirs, in order to simplifythe heater power supply.

The In-FEEP emitters for GOCE produce the following thrust levels:

� Nominal thrust range per emitter 0 - 35 µN

� Peak thrust up to 60 µN

For HYPER, 3 + 1 redundant emitter would be needed for each thrusters, and would produce:

� Nominal thrust range 0 - 140 µN, per thruster

� Peak thrust up to 240 µN, per thrusters.

Due to variations and differences in needle wetting (indium film thickness), each needle emitter will have adifferent current-voltage characteristic. For this reason, every emitter requires its separate power supply.This results in a large Power Control Unit (PCU): GOCE has 96 independent PCU channels; for HYPER 48independent PCU channels would be needed.

In Figure 3.1-3, emitters with different reservoir sizes are shown. In the liquid state, indium will be held bysurface tension between an internal vane system, which is not visible in Figure 3.1-3. The vane system feedsthe Indium propellant towards the needle, and allows almost complete usage of the indium mass (approx.99 %). For GOCE, a 30 gram reservoir was developed, which would also suit HYPER needs (see totalimpulse requirement).

Figure 3.1-3: Indium Needle Emitter with different Reservoir Sizes

A prototype Indium FEEP is shown in Figure 3.1-4. The extractor electrode of Figure 3.1-1 has been reducedto a tantalum Extractor Ring. The reason for this design solution is the following: Indium micro-droplets from



the emitter needle impinge and accumulate on the extractor. This will change the needle-extractor geometry,if the material is not removed regularly. For GOCE, it is predicted that the Extractor Ring must be heatedevery 150 - 200 h, in order to evaporate the accumulated indium. For this purpose, the HV supply of therespective emitter is switched OFF, the ring is heated for 1 - 3 minutes, and HV supply is switched ON again.The need to maintain the extractor geometry will also exist for HYPER.

Figure 3.1-4: Indium-FEEP with heatable Extractor Ring

In Figure 3.1-5, an exploded view of a multi-emitter assembly is depicted. And in Figure 3.1-6, an outlinedrawing of a cluster of two emitters is shown.



Figure 3.1-5: Indium-Needle Multi-EmitterAssembly

Figure 3.1-6: Cluster of two Prototype In-Needle FEEPEmitters (range: 100 �N, each)(ref. [RD 5-06])

Parameters Values

Thrust 0.1 – 100 µN / Emitter

Thrust Noise < 0.1 µN over period of 1,000 s

Minimum Impulse Bit < 5 nNs

Total Impulse 600 Ns per Emitter (for 15 g reservoir)

Specific Impulse 8,000-12,000 s

Singly Charged Fraction 98%

Electrical Efficiency 95%

Table 3-1: Performance Data obtained from the Protype Testing



In the following, a number of characterisation and performance test results are provided:

� Figure 3.1-7 shows the current-voltage curve.

0 100 200 300 400 500

5

6

7

8

9

10Em

itter

Vol

tage

[kV]

Emitter Current [µA]

Figure 3.1-7: Indium FEEP Current-Voltage Characteristic Curve

� Figure 3.1-8 shows the measured ion beam profiles, which widen up at increasing thrust level, untillimited by the plume shield geometry (approx 60° half-cone).

� Figure 3.1-9 summarises the observed beam divergence as function of thrust level(95 % of ion current).

� Figure 3.1-10 shows the measured thrust-current characteristic curve, measured on the JPL ThrustBalance Facility. The rule of thumb is 10 µA produce 1 µN.

� The mass efficiency of the FEEP needle emitter as function of the emitter current over 0 - 100 µN isshown in Figure 3.1-11. Up to approx 2 µN (20 µA), mass efficiency is almost 100 %. Above 2 µN (20µA), there is a sharp decrease, which is due to the increased emission of micro-droplets. Test resultsand model predictions are compared in [RD 5-11].

� The reported thrust level stability / resolution is illustrated in Figure 3.1-12. The FEEPs were operated in'thrust stabilised mode' (i.e. closed loop control).

� First measurement of the thrust vector directional stability were also made, and led to the need ofimproving the Extractor Ring mechanical stability. With this improvement implemented in the GOCEFEEP emitters, the expected directional stability is estimated to less than 1° half-cone.

� In Figure 3.1-13 and Figure 3.1-14, measured thrusters noise with the emitter operating at 12 µN and at50 µN with thrust stabilisation (at 12.5 Hz) is shown.

� A plot of the FEEP emitter response to commanded thrusters steps is shown in Figure 3.1-15.



-30 -20 -10 0 10 20 300.000

0.005

0.010

0.015

0.020

0.025

0.0300.49 µN

Cur

rent

[µA]

Angle [deg] -60 -40 -20 0 20 40 600.00

0.15

0.30

0.45

0.60

0.75

15.6 µN

Cur

rent

[µA]

Angle [deg]

-60 -40 -20 0 20 40 600.0

0.4

0.8

1.2

1.6

2.0

2.4

55.6 µN

Cur

rent

[µA]

Angle [deg]

Figure 3.1-8: Ion Current Profile at (a) 0.49 �N,(b) 15.6 �N, (c) 55.6 �N

0 25 50 75 10020

30

40

50

60

70

Beam

Div

erge

nce

[deg

]

Thrust [µN]

Figure 3.1-9: Beam Divergence as function of thrust level(95 % of ion current)



0 100 200 300 400 500 600 7000

10

20

30

40

50

60

70

80

Thrust [µN] Polynomial Fit - Beam Divergence Loss 23%

Thru

st [µ

N]

Emitted Current [µA]

Figure 3.1-10: Direct Thrust Measurement of the In-Needle FEEP Thruster at JPL

0 100 200 300 400 500 600 700 800 900 100010

20

30

40

50

60

70

80

90

100

110

��I-0.42

Measurements

Mas

s Ef

ficie

ncy

[%]

Current [µA]

Figure 3.1-11: Mass efficiency � vs. Emitter Current I(ref [RD 5-11])



0 200 400 600 800 1000 1200 1400

3.16

3.17

3.18

3.19

3.20

Thru

st [µ

N]

Time[s]

24 27 30 33 36 39 42 45 480

5

10

15

20

25

Thru

st [µ

N]

Time [h]

Figure 3.1-12: Thrust Resolution in Thrust Stabilised Operation(a) at 3.0 �N

(b) at 18.0 �N



1E-4 1E-3 0.01 0.1 1 101E-5

1E-4

1E-3

0.01

0.1

1

Thrust stabilizationThrust = 12 µNSampling Frequency = 12.5 HzDigital Accuracy = 12 bits

LISA Thrust Noise Requirement

Thru

st n

oise

[µN

/Hz1/

2 ]

Frequency [Hz]

Figure 3.1-13: Thruster Noise Performance at 12 �N (thrust stabilized, one emitter)

Figure 3.1-14: Thruster Noise Performance at 50 �N (thrust stabilized; one emitter)

In Figure 3.1-16, a view of a multi-emitter cluster is shown. The size of an emitter cluster consisting of 4emitters is 162 x 162 x 78mm (l x w x h).



38200 38400 38600 38800 39000

0

20

40

60

80

100

Thru

st [µ

N]

Time [s]

Figure 3.1-15: Response to commanded Thrust Steps(one emitter)

Figure 3.1-16: Typical view of an Indium-Needle Emitter Assembly



Development status:

� Several engineering / lab models have been built and performance tested. A test campaign at the thrustbalance of JPL confirmed the good correspondence between electrical measurements and actual thrustbalance measurements.

� For GOCE, two prototype emitters were successfully subjected to a 2,000 h Endurance Test (completedin spring 2002)

� Currently, a new 3000 h Endurance Test is performed with one emitter from the first endurance test andwith to new emitters. This test had to be halted at approx. 1,800 h. An unexpected erosion of the needletip was observed. The failure analysis is not yet conclusive. It is planned to repeat the 3000h EnduranceTest with 3 new emitters at end of March 2003.

� The Qualification Model for GOCE, consisting of a cluster of 4 emitters is currently in production.Performance mapping is scheduled to start at end of March 2003.

3.1.2 Indium Multi-Capillary Emitters

At ARCS / MAGNA, also an Indium multi-capillary emitter was developed (see [RD 5-10], sect. 3). ThePrototype Model with 3 x 3 capillary emitters is shown without and with the extractor plate in Figure 3.1-17.

Figure 3.1-17: In-Multi-Capillary FEEP Thruster(distance between capillaries: 5 mm)

The development status can be described as follows:

� Initial performance tests performed

� New types of capillaries were performance tested in February 2003

� Detailed characterisation tests with new types of capillaries will start in summer 2003

� Endurance test will start lateron in 2003.

It is much easier to have similar emission characteristics with capillary emitters, because the Indium flowtowards the emission site is well-defined by the inner diameter of the capillary. It is therefore possible tooperate all capillary emitters in parallel with one HV supply (i.e. only one PCU channel!), which greatly



reduces the complexity of the electronics. The capillary emitters also share a common Indium propellantreservoir.

The number of capillaries can be adjusted to the thrust demand. In [RD 5-10], sect. 3, the followinginformation on thrust range of a 9-capillary emitter is given:

� Thrust range with near 100 % mass efficiency: up to approx 25 µN (i.e. 2 - 3 µN / capillary)

� Peak thrust capability: 1,000 µN with approx. 20% mass efficiency.

In Figure 3.1-18, the characteristic curve measured with the prototype model is shown from 0 - 450 µN.Please note the very flat voltage-current curve, which is favourable for thrust stabilisation.

The performance measured at the ONERA Thrust Balance Facility in the range from 0 - 36 µN is shown inFigure 3.1-19.

Eventually, the Indium multi-capillary emitters could become an attractive alternative to the Indium-needleemitters, e.g. emitters with 25 - 30 capillaries, in order to operate with high mass efficiency. The thrustersnoise of the multi-capillary emitters is lower than that of the needle emitters.

0 1000 2000 3000 4000 50000

2000

4000

6000

8000

10000

Thru

st [µ

N]

Voltage

Volta

ge [V

]

Current [µA]

0

100

200

300

400

500

Thrust

Figure 3.1-18: In-Multi-Capillary Emitter - Current-Voltage and Current-Thrust Characteristic Curves



0 50 100 150 200 250 3000

5

10

15

20

25

30

35

40

Thrust [µN] Polynomial Fit - Beam Divergence Loss 20%

Thru

st [µ

N]

Emitted Current [µA]

Figure 3.1-19: Direct Thrust Measurement of the In Capillary FEEP Thruster at ONERA

3.2 Caesium Thrusters

The development of the Caesium-Slit FEEP thrusters has a long history and had the continuing involvementand support of ESA. The performance characterisation and modelling of this emitter type is well-documentedin reports and publications.

The Cs-Slit Thrusters are single-emitter thrusters. The length of the slit can be adapted to the respectiveneeds of different missions. Rule of thumb: approx. 15 - 20 µN per millimetre of slit length.

The thrust range of the HYPER FEEPs is identical to that of MICROSCOPE and SMART-2: 0 - 150 µN,nominal thrust range. Precision slits (approx. 1.4 micrometers in width over the slit length) of 8 mm lengthhave already been produced and tested successfully. This design is considered sound.

Also the size of the propellant reservoir (0.12 kg of Cs for SMART-2) closely matches the needs of HYPER.

Difficulties were experienced in 2002 with a slit length designed for milli-N applications. Following the abortof the 2000 h Endurance Test for GOCE, a number of improvements were initiated at ALTA, in order toovercome the experienced problems. Within the FEEP consultancy for HYPER, ALTA have identified theseareas:

� need to redesign the milli-N slit (which is longer than that used for HYPER)

� excessive Cs back-scattering in the test facility

� modification of the Cs feed system, in order to operate also in 1-g environment



� consequent improvement and validation of production and filling procedures and of the related GSE.

Currently, ALTA prepare for the FEEP Lifetime Test (planned duration: 1.5 years), which is anticipated tostart in mid-2003 / second half of 2003. This test program shall demonstrate the maturity of the design, of thecritical procedures, performance characteristics, and plume contamination using QCMs.

The Caesium propellant is very difficult to handle, because of its extreme reactivity. Outmost care must betaken in order to ensure that that the Cs propellant inside the thrusters / reservoir is free from contamination(oxygen, residual water). Filling operations must be undertaken in vacuum, and the thrusters must beprovided with a vacuum tight Container with a tightly closing Cover Lid + Release Mechanism. AfterAcceptance Test of the thrusters, the Cover Lid must be closed again, before the test chamber can bevented and opened. It must be pointed out, that contaminations of the propellant can not be directlymeasured. It is only the quality of contamination control in each step, which provides assurance that thepropellant is not contaminated. Some contamination will be removed, when the thrusters are started up. Thisstart up capability has been observed, but not quantified. It can be expected that spectrum of difficultiesassociated with the Caesium propellant will be overcome, within the MICROSCOPE and SMART-2programs.

The Container + Cover Lid make the individual thrusters quite large (see Figure 3.2-1).

Figure 3.2-1: ICD of the individual CS-Slit Thruster (Cover Lid closed



The Propellant Reservoir is sized for 0.12 kg of Cs, but can be adapted in length to the actual need, withoutproblems. In order to transport the propellant to the emitter, a vane system is integrated into the reservoircylinder (see Figure 3.2-2). Whether the liquid Cs inside the reservoir will move under the action of gravitygradient, or under the influence of period temperature changes on the HYPER orbit cannot be judged at thisstage.

Figure 3.2-2: Propellant Management Device inside Cs-Reservoir(emitter interface is on the left side)

In Figure 3.2-3 and Figure 3.2-4, the current mechanical design of the Cluster Assembly (CA) for SMART-2is shown. Four (4) FEEPs are arranged 90° apart and with a tilt angle of 60°. On top of the CA, theneutralisers are mounted. The accommodation of such CA's on the HYPER S/C (at mid-height) wouldrequire an octagonal S/C body and would significantly reduce the available volume inside the S/C.

Figure 3.2-3: ICD of the Cs-FEEP Cluster Assembly for SMART-2



Figure 3.2-4: ICD of the Cs-FEEP Cluster Assembly for SMART-2

The 60° tilt angle is probably not compatible with the circular Solar Array. The above concerns may leadaway from a thruster arrangement as used on SMART-2 (see sect. 4.2), and towards thrusters arrangementoption 1, which is defined in sect. 4.1.4.

ALTA are currently working on three contracts for ESA:

� MICROSCOPE (with CNES): max. thrust 150 µN;Status: close to PDR; launch planned 2005 -2006

� Development of the FEEP Cluster for SMART-2: max thrust 150 µN;Status: CDR completed; manufacturing of EM / QM thrusters on-going; qualification of a cluster

� Lifetime Test, including performance characterisation, direct thrust measurements in the Alenia facility,plume contamination measurements with QCMs, etc.



3.3 Neutralisers

For FEEP applications, neutralisers are needed - not for the proper operation of the FEEP emittersthemselves, but for the compensation of S/C charging due to the emitted positive ion current. For thisreason, neutralisers positions can be chosen on any suitable location on the S/C, assuming adequateconductive paths.

The charged particle environment on the 1,000 km orbit will provide some charge compensation. But at thisstage of definition, it is assumed, that the neutraliser(s) must compensate the full charging due to the FEEPthruster ion beams.

Specifics of HYPER:

� maximum thrust level of all FEEPs combined < 500 �Ncorresponding to a total emission current of ~ 5 mA

� residual atmosphere at 1,000 km is << 10-8 mbar. In this environment, cathode degradation due toresidual atmosphere is negligible.

� neutralisers can be positioned outside the areas with charge exchange ions. Also this degradingenvironment would not be relevant for the neutralisers on HYPER.

In [RD 7-01], a thorough survey of the various neutraliser technologies and of their development status wasperformed and a preliminary trade-off performed. Candidate neutralisers for FEEP application are:

� Low-power thermionic cathodes

� Field Emission Array (FEA) cathodes produced in MEMS technology

� Carbon Nanotube (CNT) FEA cathodes.

For all 3 categories / technologies suitable cathodes exist or are promising and under development. In[RD 7-01], the preliminary trade-off led to the following ranking, based on performance reasons:

� Rank 1: FEA cathode by RAL-CMF, followed by FE Picture Element Technology (FEPET) carbonnanotube cathode (NASA-JPL).

� Rank 2: Scandate thermionic cathode by HeatWave; other FEA cathodes (CEA-LETI and StanfordResearch Institute), Carbon nanotube cathode by BUSEK.

The survey performed in [RD 7-01] is exhaustive and fully adequate for the HYPER Feasibility Study. Theranking was independently confirmed by SMART-2 Phase A, which also selected the FEA cathode of RAL-CMT. Some characteristics are given below:

� extractor electrode potential 60 V

� electron current emitted per micro-tip approx. 5 �A

� technology qualified for a ROSETTA instrument



� life-time test without significant degradation 1,000 h(cf. HYPER mission life-time 17,500 h).

However, it must be pointed out that neutralisers are no negligible cost element, in particular the FEAcathode of RAL-CMT! It is therefore recommended to reduce the number of neutralisers onboard HYPER to2 units (one main and one cold redundant neutraliser, total). This also reduces the number of neutralisersupply & control channels in the PCU. In a next step towards neutraliser selection, cost and developmentstatus / SMART-2 heritage for the different neutralisers must be taken into account.

At the time being, only the thermionic neutraliser from Thales-Thomson is space and lifetime qualified. Thisneutraliser must be protected against atmosphere (requires container with cover lid).



4 Requirements and Constraints for the FEEP System

In this chapter, the requirements relevant for the Micro-Propulsion System of HYPER, as specified in HYP-5-01, are compared with the requirements on the Micro-Propulsion Systems of SMART-2 and of GOCE. Briefconclusions are given in sect. 4.5.

4.1 FEEP Application for HYPER

4.1.1 FEEP Thruster Usage

After S/C initial rate damping and attitude acquisition, and after the Micro-Newton Propulsion System hasbeen checked-out and commissioned, the FEEP Micro-Newton Propulsion System will be activated and itsthrusters will become the primary actuators. At the beginning of the Science Phase, two alternatives exist:

(a) the Cold Gas Propulsion System (CGPS) is fully vented, and the FEEP System will become the solecontrol authority, without any backup to cope with unforeseen events

(b) the CGPS is put into a minimum-leakage stand-by configuration. In this case, it will be possible to re-activate the CGPS upon need, e.g. in case of emergency. In sect. 5, the trade-off between vented CGPSvs. stand-by CGPS is performed. The trade-off concludes with the recommendation to maintain theCGPS in stand-by. Disturbances originating from the CGPS can be held below the limiting levelsimposed by the Payload.

In Table 4.1-1, the FEEP thruster usage during the different mission phases is outlined. As sole controlauthority during Science Phase, the FEEP Micro-N Propulsion System must support both, Primary AOCSand Secondary AOCS. Secondary AOCS has to provide disturbance compensation and attitude control atthe performance levels required during the science measurements. Primary AOCS will operate either in amode identical to that of Secondary AOCS, but with reduced performance requirements, or in a mode with acoarser sensor complement.

In [AD-05], an orbit insertion error of δh = +/- 20 km and δi = +/- 0.05° is quoted. Without any orbit insertionerror correction, the drift of the LTAN is in the order of a few degrees over the complete mission life time. Anassessment by Astrium UK indicate, that the small orbit drift is tolerable for the experiments. Orbit insertionerror correction is therefore not required, and a small GN2 propellant tank will suffice for HYPER.

It is recommended to maintain a well-defined cold gas reserve of up to 0.5 kg of GN2 at BoM, which allowsto reduce mission risk during Science Phase.

Space Debris Mitigation Standards / Guidelines demand that LEO satellites "should, whenever possible, besafely de-orbited within 25 years".

In [AD-05], de-orbiting with 500 �N and 1,000 �N was analysed. The conclusions drawn from [AD-05] are,that de-orbiting from the 1,000 km reference orbit would be design driving for the FEEP Micro-PropulsionSystem with respect to:

� required total impulse capability (approx. 80,000 N s), additional to the total impulse required for theScience Phase



Phase FEEP Thruster Usage and Comments

Acquisition Phase � transition from Cold Gas RCS control to FEEP control� sun pointing attitude acquisition / hold

Commissioning Phase � In-flight FEEP calibration� Commissioning of Secondary AOCS

Science Phase -Measurement Seasons

� Slew to and acquisition of new Guide Stars� Scientific Measurement periods

Science Phase -eclipse seasons, outages,emergency operations,recovery

� ground station outage over several days� on-board anomalies leading to autonomous Sun pointing attitude

acquisition� regular thruster maintenance, if needed

Table 4.1-1: FEEP Thruster Usage during HYPER Mission Phases

� additional 10 years of required life-time and operational support for S/C Bus and FEEP Micro-PropulsionSystem.

� Sun-synchronicity would be lost, complicating S/C operations, unless also inclination changemanoeuvres are performed (requiring an extra of approx. 90,000 Ns).

EOL de-orbiting is not compatible with the HYPER satellite concept. It would be overly design-driving withrespect to maximum thrust level and total propellant demand, and would not be commensurate for a smalland optimised scientific satellite like HYPER. EOL de-orbiting is therefore not baselined.

4.1.2 Disturbance Forces and Disturbance Moments

The different disturbances (forces and torques) that must be compensated with the FEEP System aresummarised in Table 4.1-2. The major disturbances are shown in bold.

All major disturbances occur as cyclic effects (approximately periodic with �orbit or with 2 �orbit), because thesatellite attitude is held inertially fixed.

Based on disturbance force and torque simulations reported in HYP-2-01 "Secondary AOCS Design Report",the nominal thrust range and the peak thrust capability of each thrusters are specified in HYP-5-01 "FEEPRequirements Specification".



Sources of disturbance forces Sources of disturbance torques

� Gravity Gradient radial effect (due todistance between CoM and Drag-Free Point)

� Gravity Gradient torque (due tounsymmetric MOI and offset angles betweenSC_X_p and orbit normal)

� Gravity Gradient perpendicular effect (dueto distance between CoM and Drag-FreePoint)

� Residual S/C magnetic moment interactingwith Earth Magnetic Field

� Air drag � Air drag induced torques (due to geometricasymmetry, in particular at offset anglesbetween SC_X_p and orbit normal)

� Solar pressure � Solar pressure induced torques(due to geometric asymmetry, in particular atoffset angles between SC_X_p and orbitnormal)

� Earth Albedo and IR � Earth Albedo and IR induced torques

� Parasitic accelerations due to thrusteralignment errors

� Parasitic moments due to thrust vectoralignment and position errors

� Parasitic accelerations due to thrustermismatch errors

� Parasitic moments due to thruster mismatcherrors

Table 4.1-2: List of Disturbance Forces and Moments

4.1.3 Baselined Thruster Arrangement for HYPER

For HYPER, several thruster arrangements were investigated. Two alternatives exist, depending on theFEEP thrusters technology ultimately selected.

Option 1:

Four clusters, each with 3 thrusters, are located at the corners of the S/C on opposite diagonals:

� Group {1, 2, 3, 4} is pointing into +/- X

� group {11, 12, 13, 14} into +/- Y

� group {21, 22, 23, 24} into +/- Z.



Figure 4.1-1: HYPER FEEP Thruster Arrangement (Option 1)with 4 x 3 thrusters

Table 4.1-3 outlines which control forces and moments can be produced by different combinations ofthrusters:

� All three rotational DOFs can be controlled with with full redundancy.

� No redundancy exists for the three translational DOFs.

This configuration is well-suited for the Indium-Needle Thrusters. The missing redundancy for thetranslational DOFs is achieved by adding one additional emitter to each of the 12 thrusters.

In case of In-Multi-Capillary thrusters or Cs-Slit thrusters, the missing redundancy can be achieved by either:

� doubling of the 12 thrusters, or by

� introducing of a fourth group of tilted thrusters {31, 32, 33, 34} - see Figure 4.1-2. All of these fourthrusters point through S/C CoM, and provide only linear forces.



Function Primary Set Secondary Set

Force in + X 3 + 4 (33 + 34)

Force in - X 1 + 2 (31 + 32)

Force in + Y 11 + 14 (31 + 34)

Force in - Y 12 +13 (32 + 33)

Force in + Z 22 + 24 (31 + 33)

Force in - Z 21 + 23 (32 + 34)

torque about X (+) 13 + 14 21 + 22

torque about X (-) 11 + 12 23 + 24

torque about Y (+) 2 + 3 21 + 24

torque about Y (-) 1 + 4 22 + 23

torque about Z (+) 2 + 4 11 + 13

torque about Z (-) 1 + 3 12 +14

Table 4.1-3: Thruster combinations to produce control forces and moments

Option 2:

Four clusters are arranged at the four corners of the S/C, at mid-height - see Figure 4.1-3. Two groups aredefined:

� group {1, 2, 3, 4, 5, 6, 7, 8} firing tangentially in the Y-Z plane, producing forces in +/- Y and +/- Z, andmoments about +/- X

� group {11, 12, 13, 14, 15, 16, 17, 18} firing tangentially in the X-Z plane, producing forces in +/- X, andmoments about +/- X, +/- Y and +/- Z.

This configuration is well-suited for the Cs-Slit thrusters, which have an ion beam width (90%) of +/- 20° x+/- 40°. It is very similar to the thrusters arrangement proposed for SMART-2. If used on HYPER, it requiresan octagonal S/C body, in order to fit into the ROCKOT fairing.



Figure 4.1-2: HYPER FEEP Thruster Arrangement (Option 1 a)with 4 x 3 plus 4 x 1 thrusters

1

2

3

4

6

8

5

7

� � 45°

Solar Array

11

12

13

14

18

16

17

15

Figure 4.1-3: HYPER FEEP Thruster Arrangement (Option 2)(a) showing thrusters {1, 2, 3, 4, 5, 6, 7, 8} and

(b) showing thrusters {11, 12, 13, 14, 15, 16, 17, 18}

33

31

34

32



Function Primary Set Secondary Set Third Set

force in + X 11+ 13 + 15 + 17 11 + 15+ Torque x (-) pair,

or13 + 17

+ Torque x (+) pair

force in - X 12 + 14 +16 + 18 12 + 16+ Torque x (-) pair,

or14 + 18

+ Torque x (+) pair

force in + Y 2 + 3 5 + 8 2 + 3 + 5 + 8

force in - Y 6 + 7 1 + 4 1 + 4 + 6 + 7

force in + Z 4 + 5 2 + 7 4 + 5 + 2 + 7

force in - Z 1 + 8 3 + 6 1 + 8 + 3 + 6

torque about X (+) 1 + 5 3 + 7

torque about X (-) 2 + 6 4 + 8

torque about Y (+) 11 +14 16 + 17

torque about Y (-) 12 + 13 15 + 18

torque about Z (+) 11 + 13 + 16 + 18 11 + 18+ Force z (+) pair,

or13 + 16

+ Force z (-) pair

torque about Z (-) 12 + 14 + 15 + 17 12 + 17+ Force z (+) pair,

or14 + 15

+ Force z (-) pair

Table 4.1-4: Thruster combinations to produce control forces and moments (Option 2)

Due to the tilt angles (approx. 45°), lower thrust levels are produced with one thrusters pair. On the otherside, the thrusters can be operated in quadruples ! Some thrusters pairs produce parasitic torques orparasitic forces, which must be compensated by an additional pair.



4.1.4 HYPER Requirements on FEEP System

See HYP-5-01 "FEEP Requirements Specification".

4.2 SMART-2 Requirements

In the SMART-2 Phase A studies, Cs-Slit FEEP thrusters were baselined. The recommended thrustersarrangements from the Phase A studies, and a summary of the requirements specified in [RD 2-05] and[RD 2-06] is given below.

4.2.1 SMART-2 Thruster Arrangement

Two SMART-2 Phase A Studies were performed, one by CASA (E) and one by Astrium (UK). These twostudies arrived at almost the same thruster arrangements.

Figure 4.2-1: Arrangement of the 4 Clusters of 4 Thrusters, each(SMART-2, CASA-Study)

Smart-2 has an octagonal S/C body. Both Phase A studies recommend to mount the FEEP clusterassemblies on the middle of the side panels (see Figure 4.2-1).



In Figure 4.2-2 and Figure 4.2-3, the two arrangements are shown. Astrium have proposed to rotate theClusters by 45° compared to the CASA version. Figure 4.2-3 also illustrates the 90% ion beam width of the16 FEEP thrusters (+/- 20° x +/- 40°, each).

Figure 4.2-2: FEEP Thruster Arrangement (CASA Phase A)(FEEP thrusters shown red)

Figure 4.2-3: FEEP Thruster Arrangement (Astrium Phase A)(FEEP thrusters shown red; ion beam width shown yellow)



4.2.2 SMART-2 Requirements Summary

A summary of the FEEP requirements specified in [RD 2-01] and [RD 2-06] are given below in tabular form.

The SMART-2 mission lifetime is 5 years in orbit. This long mission lifetime, requires high reliability andcontinuous availability of the FEEP Micro-Propulsion System. These requirements have not yet beenformulated.

In Figure 4.2-4, the thruster noise requirements of SMART-2 (shown green) applies for thrust levels above25 �N (from [AD-02]. In the mean time, these requirements have apparently been tightened to < 1 10-7

N/�Hz below 25 �N, and < 0.5 10-6 N/�Hz above 25 �N (see [RD 2-06]).

10-4

10-3

10-2

10-1

10010

-7

10-6

10-5

10-4

10-3

Frequency [Hz]

SpectralDensityN/sqrt(Hz)

GOCESmart-2

Figure 4.2-4: SMART-2 Thruster Noise Requirements(green line)

Technology: Cesium-Slit Emitter; single-emitter thrusters

Nomenclature:

CA Cluster Assembly consisting of 4 single-emitter thrusters and 2 neutralisers

PCU Power Control Unit for one Cluster Assembly



Table 4.2-1: Preliminary SMART-2 Requirements:

Requirements Value Remarks

0. Elements of FEEP System

4 CAs, each with 4 thrusters, each with 1 emitter, 1 reservoir, 1 container with Cover Lid 2 neutralisers

4 PCUs

Thrusters at 60° cantangle; separated by 90°

Cs reservoir with 0.12 kgof Cs propellant

1. Ion Beam

Ion beam divergence(> 90 %)

+/- 20° along slit;+/- 40° cross slit

[RD 2-06]

2. Thrust Range

min.thrust per emitter 1 �N (required)0.1 �N (goal)

[RD 2-05][RD 2-06]

max. thrust per emitter 150 �N [RD 2-05], [RD 2-06]

3. Characteristic Curve

Bias No specification Fachieved = Fcommanded + B

Max. deviation from linearity No specification

Hysteresis No specification

Scale Factor Error No specification

thrust accuracy +/- 0.5 % repeatability over lifetime [RD 2-06]

4. Thrust Response

thrust command time 500 ms 2 Hz update rate[RD 2-05]

rise and fall time to 90% for��Fc�= 1�N

< 100 ms for ��Fc�= 1 �N[< 500 ms for ��Fc�= 150 �N]

due to PCU; emittersrespond within a few ms;[RD 2-05]

5. Quantisation step

Quantisation step 0.1 �N below 50 �N0.3 �N from 50 �N - 150 �N

[RD 2-05][RD 2-06]

Table 4.2-1: Preliminary SMART-2 Requirements (continued)



6. Thrust Vector Noise

thrust noise (1 - 10-4 Hz) < 1 10-7 N/�Hz up to 25 �N< 0.5 10-6 N/�Hz above 25 �N

no split between thrusterand PCU; [RD 2-06]

7. Thrust Vector

Thruster directions cant angle � = 30° - 60° w.r.t. common mountingplane of CA;[RD 2-05]

thrust stability < 0.5 % over entire life [RD 2-06]

deviation actual fromnominal thrust direction

3° (half-cone, 3 sigma) over lifetime [RD 2-06]

thrust direction stability < 0.5 ° over entire life [RD 2-06]

thrust vector position no specification

8. MTA Total Impulse

total impulse per thruster 10,000 N s [RD 2-05]

minimum specific impulse > 7,000 s [RD 2-05][RD 2-06]

mass efficiency at150 �N

no specification

size of Cs reservoir 0.12 kg of Cs per thruster [RD 2-05]

9. MPA Total Mass

mass of each CA < 3.75 kg ALTA input

mass of each PCU 4.65 kg +/- 5 % 1 PCU for each CA;[RD 2-05]

10. Dimensions

dimensions of each CA -- Not relevant for HYPER

dimensions of each PCU 119 x 338 x 187.5 mm 4 boxes (l x w x h);[RD 2-05]

11. FEEP System Power

basic power for PCUoperating; 0 �N thrust

23.89 W per PCU

delta-power for PCUoperating; 150 �N thrust

< 30 W for 4 x 150 �N thrust

Table 4.2-1: Preliminary SMART-2 Requirements (continued)



HV for thrusters emitter: between + 2.5 and 9 kV;accelerator: from - 1 to - 3 kV(fixed setting)

[RD 2-06]

Gate voltage forneutralisers

+ 75 V to + 200 V RAL FEA Neutraliser[RD 2-05]

Neutraliser levels low: 2.5 mA (at < 50 �N);high: 7.5 mA (at 50 - 150 �N);

at each CA;Iemitter 1.5 mA, eachthruster[RD 2-05]

14. Heating

FEEP thruster < 5 W per emitter;< 5 W per reservoir

during nominal operation;[RD 2-05]

neutraliser: low level: 2.16 Whigh level: 3.56 W

during nominal operation;[RD 2-05]

13. Redundancy / Failure Tolerance / Reliability

Thruster redundancy /failure tolerance

failure tolerance: any 1 thruster outof 16 thrusters

adequacy to be verifiedagainst reliability !

PCU redundancy Main + Redundant Logic Section;independent thruster supply;independent neutraliser supply

[RD 2-05]

Neutraliser redundancy 1 main + 1 cold redundant unit ateach CA

[RD 2-05]

FDIR only overload / overcurrent aremanaged autonomously by the PCU

telemetry available to OBC

needed to avoid failurepropagation[RD 2-05]

Mission Lifetime 5 years of operational flight [RD 2-05], [RD 2-06]

Reliability no specification found

13. Verification

Life-time 1.5 years Lifetime Test

Plume / Contamination no specification

verification by test of eachthruster

no specification

Table 4.2-2: Preliminary SMART-2 Requirements



4.3 GOCE Requirements

GOCE will be the first mission that will use FEEPs within a Drag Free and Attitude Control System. Micro-NFEEP thrusters are used for:

� side force compensation

� fine attitude control

� instrument calibration by producing precise accelerations.

GOCE has selected Indium-Needle Emitters as the currently best-proven and available FEEP technology,and as a technology which is less sensitive to the ATOX environment.

Note, that due to its low altitude, air drag compensation in flight direction is performed with Ion Thrusters,which operate in the milli-N range.

4.3.1 GOCE Thruster Arrangement

Each thrust direction is equipped with 12 In-needle emitters.

The thruster arrangement on GOCE is shown in Figure 4.3-1 (see [RD 3-01] and [RD 3-03]). Each of the 8Micro-Thruster Assemblies (MTA) shown in Figure 4.3-1 consists of 12 Indium-Needle Emitters, in order toproduce the required maximum operational thrust level of 400 �N per MTA, respectively 650 �N duringaccelerometer calibration periods.

The reservoir size per Indium Emitter is 30 grams, which exceeds the required total impulse capability by afactor of 2. The number of Indium Emitters per MTA is solely driven by the high thrust levels needed foradequate control authority. One of the 12 emitters is implemented to provide redundancy (1 out of 12).

A major area of concern for GOCE is surface contamination. The beam of energetic Indium ions is 100 %limited to its beam divergence of 60°, half-cone. However, low-energetic liquid metal micro-droplet and back-flow ions can occur outside this cone. Due to the criticality for the GOCE Solar Array, the 3000 h ExtendedEndurance Test includes contamination measurements on witness surfaces at 90° from the beam direction.This test started in October 2002, and will provide useful information at end of March 2003.



Solar Array

MTA at 30°

Figure 4.3-1: Thruster Arrangement on GOCE

4.3.2 GOCE Requirements Summary

In Table 4.3-1, the requirements on the Micro-Thruster Assemblies (MTA) and on the Micro PropulsionElectronics (MPE) are summarised (from [RD 3-01]). It must be emphasised that these requirements are stillin the process of improved definition and consolidation. GOCE has already entered Phase C/D (whileSMART-2 is in Phase A-Extension). FEEP CDR is planned in summer 2003.

In Figure 4.3-2, the thruster noise requirements of GOCE are shown.

Due to its low altitude, and due to the neutralisers of the Ion Thrusters, no extra neutralisers for the FEEPthrusters are needed. S/C charging by the FEEP thrusters is no issue for GOCE.

Technology: Indium-Needle Emitters

Nomenclature:

MTA Micro-Thruster Assembly, consisting of 12 Indium-Needle Emitters

MPA Micro-Propulsion Assembly, consisting of 1 MTA + 1 Micro-Propulsion Electronics (MPE)supply channel

MPE Micro-Propulsion Electronics (synonym: Power Control Unit (synonym: PCU))

MTA

MTA

MTA

Z

Y



10-4

10-3

10-2

10-1

10010

-7

10-6

10-5

10-4

10-3

Frequency [Hz]

SpectralDensityN/sqrt(Hz)

GOCESmart-2

Figure 4.3-2: GOCE Thruster Noise Requirements(dashed blue line)

Table 4.3-1: Summary of Requirements applicable to GOCE's Micro-Thruster Assemblies (MTA) and to Micro-Propulsion Electronics (MPE)

Requirement Value Remarks

0. Elements of the FEEP System

8 MTAs

2 MPEs

12 emitters, each

48 independent channels,each

1. Ion Beam

beam divergence 60° half-cone;containing at least 100 % of primaryions

reason: Indium sputteringonto Solar Array

Effectiveness of plumeshields

under investigated on-going test within the3000 h extendedendurance test

2. Thrust Demand Profile

thrust demand profile duringmission

preliminary thrust profile given highly dynamic comparedto SMART-2



Table 4.3-2: (continued)

3. Characteristic Curve

min. thrust per MTA 2 �N with 12 needle emitters

max. thrust per MTA- normal operations- during P/L calibration

400 �N650 �N

Bias 10 �N (MTA + MPE) 6 �N (MTA alone)

Fachieved = Fcommanded + B

Max. deviation from linearity 10 �N (MPA)

Hysteresis 20 �N (MPA)

Scale Factor Error 5 % (MPA) 1 % (MTA alone)

4. Thrust Response

overshoot 5 % of commanded step (MPA)none (MTA)

rise time to 95% for��Fc� 100 �N

< 70 ms (MPA)< 5 ms (MTA alone)

��Fc� 100 �N total by 12emitters

delay time not specified (irrelevant)

5. Quantisation step

Quantisation step < 0.64 �N 12 emitters in parallel !

6. Thrust Vector Noise

thrust vector noise10-4 - 3 10-3 Hz3 10-3 - 10-2 Hz10-2 - 104 Hz

10 % of the following:5 10-4 N/�Hzslope6 10-7 N/�Hz

90 % allocated to MPE;10 % allocated to MTA

over full thrust range2 �N - 650 �N

7. Thrust Vector

nominal MTA thrust vectoralignment accuracy

0.1° (half-cone) normal to mounting plane

deviation MTA actual fromnominal thrust direction

1° (3 sigma) over thrust range;over life-time

MTA thrust vector position 5 mm (3 sigma) internal errors;1 mm MTA mounting error



Table 4.3-2: (continued)

8. MTA Total Impulse

total impulse per MTA 6,045 Ns assumptions not defined

9. MPA Total Mass

mass of one MTA ~ 5.0 kg status

mass of MPE < 66 kg for 96 emitters

10. Dimensions

dimensions MTA -- See ICD

dimensions MPE 2 boxes of 383 x 411 x 270 mm, or4 boxes of 383 x 238 x 270 mm

2 boxes are baseline forGOCE

11. MPA Power

basic power MPA;operating, 0 �N

175 W, total @ 28 VDC; for 96 emitters

power for 2 MTAs, each at400 �N

0.091 W per 1 �N @ 28 VDC; newinformation

12. Redundancy

MTA internal 11 + 1 redundant = 12 emittersfor each MTA

all emitter channels areindependent: furtherfailures lead to gracefuldegradation

MPE internal Main + Redundant Logic Section;independent emitter supplies

13. Verification

Life-time by technology life time test(s)

Plume / Contamination within technology demonstration (bytest)

= part of ExtendedEndurance Test

verification by test on eachMTA

performance test

thrust characterisation(ramp, steady-state, steps)

thrust vector noise

thrust vector stability (direction)

preliminary list



4.4 Comparison of Requirements

HYPER Requirements SMART-2 GOCE Remarks

No. Requirement

3.3.1-2 Neutralisersneeded

2 neutralisers perCluster of 4 thrusters(1 main + 1 coldredundant);2 commandable levels:2,5 mA (low) and 7.5mA (high)

Not needed forGOCE OK

3.3.1-3 Ion BeamDivergence< 60° half-cone(100%)

Ion Beam width is20° x 40° (90%)

Ion BeamDivergence <60° half-cone(100%)

OK

3.3.1-4 Provisions toavoid plumecontamination at90° from ionbeam direction

No equivalentrequirement yetdefined

Provisions toavoid plumecontamination at90° from ionbeam directionalso required forGOCE

in definition /verification

3.3.1-5 Health / IntegrityCheck in flight

compliant compliant OK

3.3.2-1 Total deliverableimpulse perthruster

compliant 2 x HYPERrequirement

OK

3.3.2-2 Maximum thrustlevel perthrusters:- 100 µNnominal max- 150 µN peak

150 µN max With 12 In-needle emitters:- 400 µNnominal max- 600 µN peak

OK with currentthrust demand

3.3.2-3 Min. Thrust level 0.5 µN

1 µN (required); 0.1µN (goal)

< 0.64 µN with 4emitters

OK, becausemin. thrust levelcan be produ-ced by oppositethrusters




No. Requirement

3.3.2-6 Required close-loop response:10 µN / s

no equivalentrequirement

no equivalentrequirement

OK(reportedperformances)

3.3.2-7 Thrust commandupdate rate: 10Hz

2 Hz 2 Hz (TBC) HYPER reqtcan probablybe relaxed to2 Hz withoutimpact

3.3.2-8 Thrust bias errorper thruster: < +/-4 µN

Contained in thrustrepeatabilityrequirementR3.3.2-14

Equivalentrequirement

OK

3.3.2-9 Scale Factor(Gain) error:< 5 %



OK

3.3.2-10 Deviation fromlinearity:< 4 µN perthruster



OK

3.3.2-11 Hysteresis < 8 µNper thruster



OK

3.3.2-12 Thrust responseto +/- 10 µN step:< 5 % overshoot< 100 ms risetime (95 %)

< 100 ms for +/- 1µN step


Probablybothcompliant

3.3.2-13 Thrustquantisation step:< 0.1 µN

0.1 µN for < 50 µN

0.3 µN for 50 - 150µN

< 0.2 µN with 4emitters

OK(no hardrequirement)



HYPER Requirements SMART-2 GOCE

No. Requirement

3.3.2-14 Thrustrepeatabilitywithin therequirements ofR3.3.2-8 toR3.3.2-12

Better than HYPERrequirement

Equivalentrequirement OK

3.3.2-15 Thrust noise: useGOCE reqt

Better than GOCErequirements

Identicalrequirement

OK

3.3.2-16 Thrust vectorangular stability:(a) < 0.5° (3)over thrust range0 - 150 µN(b) < 0.5° (3)over lifetime

Yet to be verified bytest

Some testevidence; requiredstability yet to beverified by test

Thisperformancemust yet beverified bytest

3.3.2-17 Thrust vectorangular noise(random variationof thrustdirection):< 0.1° half-cone(rms)

Yet to be measuredby test

Some testevidence; yet to bemeasured by test

Thisperformancemust yet beverified bytest

3.3.2-18 Only short andisolated sparkingevents

5 mm Slit probablycompliant

Confirmed by test OK

3.3.2-20 Thruster shalloperate in fullsun-light and inshade


3.4.1-1 Size of propellantreservoir to meetR3.3.2-1

Compliant;0.12 kg Cs perthruster

Compliant;0.03 kg Cs peremitter

OK

3.4.1-2 Suitablepropellantmanagementdevice

Vane systemcurrently underoptimisation

Compliant;99 % propellantusage

OK;Cs reservoirin redesign




No. Requirement

3.4.1-3 Protection ofpropellant againstcontamination

Major issue with Cs;under optimisation

(a) vacuum tightemitter container withCover LidMechanism for eachthrusters

(b) critical propellanthandling and loadingprocedure

Uncritical withIndium

Adequacy forCs propellantunderdevelopment /validation(preparationfor LifetimeTest program)

3.4.1-4 Heaters - reservoir

- emitter slit

- reservoir /needle

- extractor ring

both compliant

3.4.1-5 Plume Shield &collector for low-energy ions

compliant Identicalrequirement

OK

3.4.1-6 Analytical and testevidence of plumecontaminationoutside primarybeam to beprovided

See test program oflifetime test;compliant

See test programof 3000 hExtendedEndurance Test;compliant

Test evidenceto be revieweduponavailability

3.4.1-7 Management ofback-flow material/ avoid sputteringon S/C surfaces

Similar requirement Similarrequirement

See commentto R3.4.1-6

3.4.1-7 Stability ofEmitter-Accelerometergeometry

Not needed heating ExtractorRing (1 - 3 min)required in regularintervals ofapprox. 200 h

OK;In-needleemittersrequire heating

3.4.2-1 Mission Lifetime:2.25 years in orbit

5 years in orbit 2.5 years in orbit OK

3.4.2-2 Adequate failuretolerance

1 out of 16 thrusters 1 out of 12emitters perthruster

Adequacy tobe reviewedafter lifetimetests




No. Requirement

3.4.2-3 RedundancyConcept

16 independentthrusters

Redundant PCULogic

8 x 12independentemitter channels(i.e. PCU +emitterchannels)

Redundant PCULogic

both compliant

3.4.2-4 Commensuratereliability

yet TBD yet TBD TBD

3.4.3-1 Environmentalconditions ofROCKOT or DNEPRlaunch


3.4.4-1 Failure Detection andIsolation:(a) health info foronboard failuredetection andisolation

(b) adequate HK andhealth data for on-ground diagnostics


3.4.5-1 Mass :(a) 35 kg, if Cs-Slit

(b) 57 kg, if In-Needle

compliant Compliant(2 x number ofHYPERthrusters)

Mass of In-multicapillaryoption wouldapprox.correspond toCs-Slit option

3.4.5-2 Dimensionsaccording existingh/w

Cluster Assembly For HYPER,number ofthrusters and ofPCU are 50 %of GOCE

OK;

Cluster Assyfrom SMART-2may have tobe tailored forHYPER config.constraints




No. Requirement

3.5.1-2 Propellant Loading:proven procedures,GSE and relatedinterfaces

In validationprior to LifetimeTest

Automatic wettingprocedures;rest uncritical

Indium: OK

Cesium: to bereviewed uponavailability

3.5.2-1 Electrical interfaces compliant compliant OK

3.5.2-2 Maximum Power:(a) operating @ 0 µN

(b) operating @ 300 µNcompliant compliant OK

4.5 Conclusions

In the summary table of sect. 4.4, no essential new performance requirements for the FEEP System ofHYPER were identified, compared to those of SMART-2 and GOCE. Heritage from these two programs willbe usable for HYPER.

Both Indium thruster types and the Cs-Slit thruster must be considered feasible for HYPER, if the respectiveon-going / planned endurance and lifetime test programs are successfully completed.

The largest impact have probably the configurational constraints of the launcher fairing. Most likely, they willdemand modifications to the thrusters accommodation.

There are a number of minor modifications, which will not have a driving impact on the existing FEEPtechnologies. These are:

� Sampling rate

� Configuration of the cluster assemblies.

From a system point-of-view, thrusters arrangement, and mass and power values of the FEEP System mayneed further optimisation.



5 FEEP Trade-off Criteria

In [RD 8-01], trade-offs of micro-propulsion systems were performed. For almost all comparable missions,two alternatives ranked highest:

� colloid thrusters,

� FEEP thrusters.

Colloid thrusters are developed by the US-supplier BUSEK, and have been proposed as US-contribution forSMART-2. In SMART-2 Phase B, the technical assessment of this micro-propulsion alternative will beperformed. In this report, only the trade criteria for the Caesium and Indium FEEP alternatives wereelaborated.

In Table 5-1, the trade criteria are identified for two options: In-needle FEEPs and Cs-slit FEEPs. A weighingof the trade criteria is recommended at a later stage, when the Cs-Slit emitter for SMART-2 has completedthe various tests associated with the Lifetime Test program, i.e. when the mission lifetime requirement hasbeen demonstrated for the Cs-slit emitter technology. The same holds for the Indium FEEPs (needle emittersand multi-capillary emitters). At that stage, also more test evidence will be available for the Indium multi-capillary emitters, which are considered a promising third alternative.

Table 5-1: Trade Criteria In-Needle vs. Cs-Slit Emitters

Indium-Needle EmitterTechnology

Caesium-Slit Emitter Technology

(1) Development Status of Emitters

- similar emitters have been flight- proven

- successful 2000 h Endurance Test in 2002

- on-going Extended 3000 h Endu- rance Test until end of March 2003; includes some contamination measurements at 90° from beam direction

- manufacturing processes established; some design optimisations introduced for GOCE (Phase C/D)

- severe problems have led to abort the 2000 h Endurance Test

- on-going recovery actions, include test chamber improvements, more robust Cs filling procedure, Cs feed system improvement for 1-g environment, review of slit geometry / production

- various performance tests

- start of FEEP Lifetime Test in summer 2003, lasting 1.5 years

(2) Complexity of FEEP Thrusters

- total of 12 thrusters

- assembly of 4 emitters per thruster

- total of 16 single-emitter thrusters

- vacuum tight container and Cover Lid



Table 5-1: Trade Criteria (continued)



(3) Complexity of PCU

- total of 12 x 4 = 48 independent emitter supplies are needed for HYPER

- total of 16 x 1 = 16 independent emitter supplies are needed for HYPER

(4) Performance Characteristics

- performance verified by various test campaigns, including thrust balance test at JPL

- "thrust stabilisation" (closed loop) provides significant improvement, compared open loop operation

- performance as verified is well within HYPER requirements

- performance models are well-established and mostly validated by test

- performance within HYPER requirements

(5) Mass

- 48 emitters & PCU channels

- predicted mass of the 12 "thrusters" + 2 PCUs is 57 kg, total

- 16 emitters & PCU channels

- predicted mass for 16 thrusters (8x 2 thrusters clustered) + 4 PCUs is 35 kg, total

(6) Power Consumption

- stand-by operating power is ~ 100 W, total

- power to produced 300 µN, total is ~ 27 W (on top of stand-by power)

- stand-by operating power is ~ 100 W, total

- power to produced 300 µN, total is ~ 20 W (on top of stand-by power)

(7) Heritage

- selected for GOCE (Phase C/D) - baseline in SMART-2 Phase A Study

- supported by several ESA funded programs






(8) Contamination Protection for Propellant

- minimal - contamination of the very reactive Cs propellant is a critical issue with this technology

- propellant loading only in vacuum or in an inert atmosphere; strong handling constraints. Vacuum-tight emitter container & cover lid mechanisms are needed

- the loaded emitters should not participate in the S/C environmental test program � risk

- very late insertion and / or removal of each emitter assembly from the cluster housing must be possible. Cs loading as late as during launch preparation must be possible. � mobile Cs handling and loading GSE!

- monitoring of propellant contamination is not possible � mission risk (?)

(9) Emitter Assembly Self-contamination

- no problems reported from the test programs

- 1.5 years Lifetime Test will determine, whether there is a risk of self-contamination due to the high volatility of Cs exists or not.

(10) Emitter Start-up / Emitter Geometry Recovery

- Indium micro-droplets accumu- lates on the Extractor Ring, and must be removed approx. every 200 h by heating the Extractor Ring for a few minutes � short non-availability of individual emitters � emitter performance remains within specified tolerances of the characteristic curve

- Due to the extreme reactivity of Cs with oxygen and water, an initial start up of contaminated emitters must be possible.

- Contamination of the slit emitter must be removable by the start-up process (in orbit)

- Self-contamination or failure to start up is a potential mission risk.

- more test evidence is needed






(11) S/C Contamination

- Indium has a very low volatility and a higher melting point compared to Cesium.

- Indium sputtering is a feared surface contamination on Solar Arrays and on optical surfaces.

- plume shield limits high-energy beam to 60° half-cone (100%) or less

- plume shield at ground potential; collector ring at - 1 kV to collect charge transfer ions and charged micro-droplets

- test evidence of the contamination at 90° from the beam centre line will be available at end of March 2003 � witness surface from the Exten- ded Endurance Test

- Cs is very volatile

- 90 % of ion beam within 20° x 40°

- test evidence of the contamination outside the beam width also needed for the Cs-Slit emitters � witness surfaces participating in the 1.5 years Lifetime Test

(12) Thruster availability

- approx. every 200 h, the Extractor Rings must be heated for a few minutes, in order to remove accumulated Indium. The HV of the respective emitter must be switched off. Performed sequentially on all 48 emitters

- risk of sustained discharges at higher thrust levels not yet proven to be eliminated.






(13) Reliability

- 48 PCU channels

- higher failure tolerance (1 and even 2 failed emitter channels per thruster are tolerable !)

- effect of any first failure is low (soft failure; only loss of one emitters; which can be recovered by operating the remaining emitters at higher thrust level)

- proven reliability of the In-Needle emitters

- 16 PCU channels

- low failure tolerance (1 out of 16 thrusters)

- effect of any first failure is significant (loss of one complete thruster, and thus also of the respective pairs)

- reliable long-term operation not yet verified

(14) Thruster accommodation

- Either 4 clusters of 3 thrusters, each, or 4 x 2 thrusters, plus 4 x 1 thruster, dependent on accommodation constraints

- 2 cut-outs needed at the edge of the Solar Array

- Accommodation constraints require the HYPER thrusters to be arranged in 8 clusters of 2 thrusters, each. SMART-2 cluster design (housing) cannot be reused.

- 4 cut-outs needed at the edge of the Solar Array

- integration of the FEEP emitters into their housing (cluster) shall be possible on the integrated satellite at a late stage � cluster must provide suitable alignment accuracy.

(15) Status of Production Processes

- automatic filling and wetting procedures established

- established quality control procedures for all production processes for the GOCE FEEPs

- difficult production of the slit geometry � critical production process under optimisation

- "Robust" filling procedure, including GSE, under development



Name Dep./Comp.

Giorgio Bagnasco ESA/ESTEC

Phil Airey ESA/ESTEC

Ruedeger Reinhard ESA/ESTEC

Ernst Maria Rasel IQO

Philippe Bouyer IOTA

Arnaud Landragin SYRTE

Ulrich Johann Astrium

Walter Fichter Astrium

Title: FEEP Feasibility Report Prepared by: Bernd...

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