Technical Note. FEEP Feasibility Report. Prepared by: Bernd Schuerenberg Date: 25/02/2003

Transcription

1 Technical Note HYPER 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 thereof, are forbidden without express authority. Offenders are liable to the payment of damages. All rights are reserved in the event of the grant of a patent or the registration of a utility model or design. Doc. No: HYP-5-02 Page 1

2 Change Record Issue Date Sheet Description of Change Release /09/02 30/01/03 25/02/03 All All - first draft issue for Progress Meeting # 3 chapters 1-5 edited and updated; completion of chapters 6-9 (as part B of this report) Re-edited version Doc. No: HYP-5-02 Page A-I

3 Table of Contents 1 INTRODUCTION AND SCOPE INTRODUCTION SCOPE 3 2 DOCUMENTS AND DEFINITIONS APPLICABLE DOCUMENTS REFERENCE DOCUMENTS HYPER SMART GOCE CS SLIT EMITTER DOCUMENTS IN NEEDLE EMITTER AND IN CAPILLARY EMITTER DOCUMENTS COLLOID THRUSTER DOCUMENTS NEUTRALISERS OTHER REFERENCE DOCUMENTS DEFINITIONS CENTRE OF MASS AND DRAG-FREE POINT CO-ORDINATE SYSTEMS ACRONYMS 10 3 FEEP PROPULSION SYSTEMS: TECHNOLOGIES INDIUM THRUSTERS INDIUM NEEDLE EMITTERS INDIUM MULTI-CAPILLARY EMITTERS CAESIUM THRUSTERS NEUTRALISERS 29 4 REQUIREMENTS AND CONSTRAINTS FOR THE FEEP SYSTEM FEEP APPLICATION FOR HYPER FEEP THRUSTER USAGE DISTURBANCE FORCES AND DISTURBANCE MOMENTS BASELINED THRUSTER ARRANGEMENT FOR HYPER HYPER REQUIREMENTS ON FEEP SYSTEM SMART-2 REQUIREMENTS SMART-2 THRUSTER ARRANGEMENT SMART-2 REQUIREMENTS SUMMARY GOCE REQUIREMENTS GOCE THRUSTER ARRANGEMENT GOCE REQUIREMENTS SUMMARY 45 Doc. No: HYP-5-02 Page B-I

4 4.4 COMPARISON OF REQUIREMENTS CONCLUSIONS 54 5 FEEP TRADE-OFF CRITERIA 55 Doc. No: HYP-5-02 Page B-II

5 1 INTRODUCTION AND SCOPE 1.1 Introduction HYPER is one of several new scientific missions, which depend on the availability of suitable Micro-Newton Propulsion 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-Newton Propulsion 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 micropropulsion 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 future missions: FEEP thrusters (Cs-Slit and In-Needle) Colloid FEEP Thrusters Pulsed Plasma Thrusters Doc. No: HYP-5-02 Page 2

6 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 HYPER until 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 into account potential heritage and the expected maturity of technology summary of FEEP requirements for HYPER, SMART-2, and GOCE. For the detailed requirement specification for HYPER see HYP-5-01: "FEEP Requirements Specification for HYPER" comparison of requirements and identification of significant differences that may result in technology drivers the definition of trade-off criteria to compare different Micro-Newton Propulsion Systems based on FEEP technology. In addition, a comprehensive list of reference documents is given that may serve as an overview of the available documentation. Doc. No: HYP-5-02 Page 3

7 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 of documents about micro-n propulsion systems / technology HYPER RD Doc. No. Title 1-01 HYP-ASU-DD-CST-1 Description of the Atomic Sagnac Unit, Issue 1, 31 January CDF-09 Hyper CDF Study Report (as amended by Errata Corrigum ref HYP-CDF-E/C-1); Sept HYP-ESA-RS-SPA-3 Hyper Payload Requirement Specification, Issue 2, 7 June SMART-2 RD Doc. No. Title 2-01 SMT2.RP.005.EU.AST SMART-2 System Definition Study Final Report; issue 1, 2002 Doc. No: HYP-5-02 Page 4

8 2-02 SM2-CAS-4100-TNO-003 Micro-Propulsion Trade-off; 11 April SM2-CAS-5120-TNO-003 Mechanical Design Description; 28 June MPE/380/DN Development of FEEP Neutralisers (Statement of Work); issue 3, 05 Feb SP FEEP PCU Specification (ALTA thrust and RAL neutraliser); issue: 1 draft; 16 Sept ALTA/CL/SP-06 Development of Integrated FEEP Cluster Systems: Cluster Assembly Requirements Document; issue: 2; 29 March ALTA/LF/EA-01 Low thrust Propulsion System Characterization and Life Testing for LISA and DARWIN: FEEP Preliminary Dynamic Model; issue 2; 22 May Direct communication Outline Drawings of the SMART-2 FEEP Cluster (status Dec. 2002) GOCE RD Doc. No. Title 3-01 GO-RQ-ASG-0008 Micro Thruster Assembly (MTA) Requirement Specification; issue 4 Draft, Dec 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 GO-MTA-ARCS-TN-001 Preliminary Test Results; Direct Thrust Measurement and Cluster Testing, issue: Draft, 09 July GO-DD-SDP-001 MTA Design Definition Report, issue: Draft, 19 June 2001 Doc. No: HYP-5-02 Page 5

9 2.2.4 Cs Slit Emitter Documents RD Doc. No. Title 4-01 ALTA/LF/EA-01 Low Thrust Propulsion System Characterisation and Life Testing for LISA and DARWIN -- FEEP Preliminary Dynamic Model; issue2, 22 May ALTA/CL/SP-06 Development of Integrated FEEP Cluster Systems (Cluster Assembly Requirements Document); issue 1, 14 Feb ALTA/P-06.1/G Micro-Newton Thruster Assembly for the GOCE Platform (Technical Proposal) issue 1, 22 June JPP Vol. 14, No. 5, Sept-Oct 1998 M. Marcuccio, A. Genovese, M. Andrenucci: Experimental Performance of Field Emission Microthrusters In Needle Emitter and In Capillary Emitter Documents RD Doc. No. Title 5-01 AIAA A. Genovese, M. Tajmar, N. Buldrini, W. Steiger: Extended Endurance Test of the Indium FEEP Micro-Thruster M. Thajmar, A. Genovese, W. Steiger: Indium FEEP Microthruster Experimental Characterisation; Journal of Propulsion and Power, submitted MPE/404/DN ARCS Indium Liquid Metal Ion Source Characterization Test Report issue 1, 22 Oct MAGNA STEYR proposal # 2825 Development and Supply of a Micro-N Thruster Assembly for the GOCE Platform (Technical Proposal); 19 June ARCS FEEP Infos (1) First 2,000 h Endurance Test of an Indium FEEP CLUSTER 15 July 2002 (2) First Extended Endurance Test of an Indium FEEP CLUSTER 30 October 2002 (3) 3500 h operation with an Indium FEEP thruster 10 January AIAA M. Tajmar, A. Genovese, W. Steiger: Indium-FEEP Microthruster: Experimental Characterisation Doc. No: HYP-5-02 Page 6

10 Electric Propulsion Plasma Simulations and Influence on S/C Charging AIAA Journal of S/C and Rockets, Vol 39, no. 6 (2002), Direct Thrust Model Assembly (MAGNA Steyr drawing) rd International Conference of Spacecraft Propulsion, Cannes; 2000 W. Steiger, A. Genovese, M. Tajmar: Micronewton Indium FEEP Thrusters 5-10 AIAA 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 indiummetal ion source 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 for SMART-2 Bill Folkner / JPL: DRS Architecture; 28 May Neutralisers RD Doc. No. Title 7-01 AIAA M. Tajmar: Survey on FEEP Neutralizer Options; July MPE/380/DN Development of FEEP Neutralisers (Statement of Work); ESA, issue 3, 05 Feb 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 Flying Satellites; Jan 2001 Doc. No: HYP-5-02 Page 7

11 2.3 Definitions 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 ASU planes, as shown in Figure Ideally, the spacecraft CoM and the DFP should coincide. In case of HYPER, 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 Further, it is assumed that SC_Y_p points into local zenith direction, and SC_Z_p points into flight direction at Ascending Node. These preliminary assumptions are used in the discussion and calculation of the disturbance forces and torques in sect SC_Z_p SC_Y_p Figure 2.3-1: Spacecraft Physical Coordinates, Centre-of-Mass (CoM), and Drag-free Point (DFP) (CoM shown in black; DFP shown in red) Co-ordinate Systems [A] Spacecraft Physical Co-ordinate System (see Figure 2.3-1): SC_O_p SC_X_p origin is the centre of the Adaptor Ring plane (= separation plane Launcher / Spacecraft) perpendicular to the adaptor ring plane; pointing negatively from centre of separation plane through centre of Solar Array (i.e. parallel to PST_X_f) Doc. No: HYP-5-02 Page 8

12 SC_Y_p SC_Z_p normal to SC_X_p; (assumption: SC_Y_p = normal of the side panel pointing to local zenith at Ascending Node) completes the right hand co-ordinate system (assumption: SC_Z_p = normal of the side panel pointing in flight direction at Ascending Node) SC_Y_f SC_Z_f 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 SC_X_f SC_Y_f SC_Z_f origin is the centre of the separation plane Launcher / Spacecraft perpendicular to the adaptor ring plane; pointing negatively from centre of separation plane through centre of Solar Array (parallel to PST_X_f) normal to SC_X_f; (assumption: SC_Y_f nadir pointing at Descending Node) completes the right hand co-ordinate system (assumption: SC_Z_f flight direction at Ascending Node) Doc. No: HYP-5-02 Page 9

13 [C] FEEP Propulsion Co-ordinate System (see Figure 2.3-3): FEEP_O_p FEEP_X_p FEEP_Y_p FEEP_Z_p origin is the Center-of-Mass (CoM) parallel to SC_X_p parallel to SC_Y_p parallel to SC_Z_p (completes the co-ordinate system). 2a COM 2b FEEP_Z FEEP_X FEEP_Y Solar Array origin in CoM Figure 2.3-3: Co-ordinate System for the FEEP Propulsion System 2.4 Acronyms AD ASU ATOX BOL BoM CA CGPS Applicable Document Atomic Sagnac Unit Atomic Oxygen Begin of Life Begin of the (scientific) Mission Cluster Assembly (of Cs-Slit FEEPs; term used by SMART-2) Cold Gas Propulsion System Doc. No: HYP-5-02 Page 10

14 CM CMNT CNT CoM / CoG Cs CX DFAC DFP DFS DOF EM EoL EoM FEA FEEP GG GN2 GOCE HCM HV HYPER In LISA LMNIS LTP MAIT MBW MEMS MOI MTA MPA Calibration Mode Colloid Micro-N Thruster Carbon Nanotube Centre of Mass / Centre of Gravity Caesium Charge Exchange (ions) Drag-Free and Attitude Control Drag-Free Point Drag-Free Sensor Degree-of-Freedom Engineering Model End-of-Life End of (scientific) Mission Field Emission Array (technology of Electron emitting cathodes) Field Effect Electric Propulsion Gravity Gradient gaseous nitrogen Gravity and Steady-State Ocean Circulation Earth Explorer Health Check Mode High-Voltage Hyper-Precision Cold Atom Interferometry in Space Indium Laser Interferometer Space Antenna Liquid Metal Needle Ion Source LISA Technology Package Manufacturing, Assembly, Integration and Test Measurement Bandwidth Micro Electro-Mechanical System Moment of Inertia Micro Thruster Assembly (term used by GOCE) (consists of one ion emitter) Micro Propulsion Assembly (term used by GOCE) (MPA consists of several ion emitters pointing into a common thrust direction; equivalent to a 'thruster') Doc. No: HYP-5-02 Page 11

15 MPE OB PM PS PST QCM QM Rb RCS RD SAA SAM S/C scc SCM SMART-2 SSO STBY STEP TBC TBD TBR Micro Propulsion Electronics (term used by GOCE) (equivalent to PCU) Optical Bench Proof Mass Propulsion System Precision Star Tracker Quartz Crystal Microbalance (for contamination measurements) Qualification Model Rubidium Reaction Control Subsystem (see PS) Reference Document Solar Aspect Angle Secondary AOCS Mode Spacecraft standard cubic centimetre (i.e. at 1 bar) Science Mode Small Mission for Advanced Research & Technology Sun Synchronous Orbit Standby Mode Satellite Test of the Equivalence Principle to be confirmed to be defined to be reviewed Doc. No: HYP-5-02 Page 12

16 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 at ARCS / MAGNA: In-needle emitters. A suitable number of independently supplied and controlled emitters must be clustered to achieve the required maximum thrust level. In-multi-capillary thruster. This thruster is operated from one supply only Indium Needle Emitters Indium-needle emitters (see schematic in Figure 3.1-1and Figure 3.1-2) consist of an indium-wetted stainless-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 kv Extractor Electrode ground Plume Shield (collector) ground Doc. No: HYP-5-02 Page 13

17 Heater 10 V vs. ground. The heaters of the Indium reservoirs are electrically isolated to the HV of the reservoirs, in order to simplify the 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, redundant emitter would be needed for each thrusters, and would produce: Nominal thrust range 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 a different 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 48 independent 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 by surface tension between an internal vane system, which is not visible in Figure The vane system feeds the 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 total impulse requirement). Figure 3.1-3: Indium Needle Emitter with different Reservoir Sizes A prototype Indium FEEP is shown in Figure The extractor electrode of Figure has been reduced to a tantalum Extractor Ring. The reason for this design solution is the following: Indium micro-droplets from Doc. No: HYP-5-02 Page 14

18 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 heated every h, in order to evaporate the accumulated indium. For this purpose, the HV supply of the respective 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 outline drawing of a cluster of two emitters is shown. Doc. No: HYP-5-02 Page 15

19 Figure 3.1-5: Indium-Needle Multi-Emitter Assembly Figure 3.1-6: Cluster of two Prototype In-Needle FEEP Emitters (range: 100 N, each) (ref. [RD 5-06]) Parameters Thrust Thrust Noise Minimum Impulse Bit Total Impulse Specific Impulse Values µn / Emitter < 0.1 µn over period of 1,000 s < 5 nns 600 Ns per Emitter (for 15 g reservoir) 8,000-12,000 s Singly Charged Fraction 98% Electrical Efficiency 95% Table 3-1: Performance Data obtained from the Protype Testing Doc. No: HYP-5-02 Page 16

20 In the following, a number of characterisation and performance test results are provided: Figure shows the current-voltage curve Emitter Voltage [kv] Emitter Current [µa] Figure 3.1-7: Indium FEEP Current-Voltage Characteristic Curve Figure shows the measured ion beam profiles, which widen up at increasing thrust level, until limited by the plume shield geometry (approx 60 half-cone). Figure summarises the observed beam divergence as function of thrust level (95 % of ion current). Figure shows the measured thrust-current characteristic curve, measured on the JPL Thrust Balance 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 µn is shown in Figure Up to approx 2 µn (20 µa), mass efficiency is almost 100 %. Above 2 µn (20 µa), there is a sharp decrease, which is due to the increased emission of micro-droplets. Test results and model predictions are compared in [RD 5-11]. The reported thrust level stability / resolution is illustrated in Figure 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 of improving the Extractor Ring mechanical stability. With this improvement implemented in the GOCE FEEP emitters, the expected directional stability is estimated to less than 1 half-cone. In Figure and Figure , measured thrusters noise with the emitter operating at 12 µn and at 50 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 Doc. No: HYP-5-02 Page 17

21 µn 15.6 µn Current [µa] Current [µa] Angle [deg] Angle [deg] µn 2.0 Current [µa] Angle [deg] Figure 3.1-8: Ion Current Profile at (a) 0.49 N, (b) 15.6 N, (c) 55.6 N 70 Beam Divergence [deg] Thrust [µn] Figure 3.1-9: Beam Divergence as function of thrust level (95 % of ion current) Doc. No: HYP-5-02 Page 18

22 80 Thrust [µn] Polynomial Fit - Beam Divergence Loss 23% Thrust [µn] Emitted Current [µa] Figure : Direct Thrust Measurement of the In-Needle FEEP Thruster at JPL I Measurements Mass Efficiency [%] Current [µa] Figure : Mass efficiency vs. Emitter Current I (ref [RD 5-11]) Doc. No: HYP-5-02 Page 19

23 Thrust [µn] Time[s] Thrust [µn] Time [h] Figure : Thrust Resolution in Thrust Stabilised Operation (a) at 3.0 N (b) at 18.0 N Doc. No: HYP-5-02 Page 20

24 1 0.1 LISA Thrust Noise Requirement Thrust noise [µn/hz 1/2 ] E-3 1E-4 Thrust stabilization Thrust = 12 µn Sampling Frequency = 12.5 Hz Digital Accuracy = 12 bits 1E-5 1E-4 1E Frequency [Hz] Figure : Thruster Noise Performance at 12 N (thrust stabilized, one emitter) Figure : Thruster Noise Performance at 50 N (thrust stabilized; one emitter) In Figure , a view of a multi-emitter cluster is shown. The size of an emitter cluster consisting of 4 emitters is 162 x 162 x 78mm (l x w x h). Doc. No: HYP-5-02 Page 21

25 Thrust [µn] Time [s] Figure : Response to commanded Thrust Steps (one emitter) Figure : Typical view of an Indium-Needle Emitter Assembly Doc. No: HYP-5-02 Page 22

26 Development status: Several engineering / lab models have been built and performance tested. A test campaign at the thrust balance of JPL confirmed the good correspondence between electrical measurements and actual thrust balance measurements. For GOCE, two prototype emitters were successfully subjected to a 2,000 h Endurance Test (completed in spring 2002) Currently, a new 3000 h Endurance Test is performed with one emitter from the first endurance test and with to new emitters. This test had to be halted at approx. 1,800 h. An unexpected erosion of the needle tip was observed. The failure analysis is not yet conclusive. It is planned to repeat the 3000h Endurance Test with 3 new emitters at end of March 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 Indium Multi-Capillary Emitters At ARCS / MAGNA, also an Indium multi-capillary emitter was developed (see [RD 5-10], sect. 3). The Prototype Model with 3 x 3 capillary emitters is shown without and with the extractor plate in Figure Figure : 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 It is much easier to have similar emission characteristics with capillary emitters, because the Indium flow towards the emission site is well-defined by the inner diameter of the capillary. It is therefore possible to operate all capillary emitters in parallel with one HV supply (i.e. only one PCU channel!), which greatly Doc. No: HYP-5-02 Page 23

27 reduces the complexity of the electronics. The capillary emitters also share a common Indium propellant reservoir. The number of capillaries can be adjusted to the thrust demand. In [RD 5-10], sect. 3, the following information 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 , the characteristic curve measured with the prototype model is shown from 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 in Figure Eventually, the Indium multi-capillary emitters could become an attractive alternative to the Indium-needle emitters, e.g. emitters with capillaries, in order to operate with high mass efficiency. The thrusters noise of the multi-capillary emitters is lower than that of the needle emitters Voltage Thrust Voltage [V] Thrust [µn] Current [µa] Figure : In-Multi-Capillary Emitter - Current-Voltage and Current-Thrust Characteristic Curves Doc. No: HYP-5-02 Page 24

28 40 Thrust [µn] Polynomial Fit - Beam Divergence Loss 20% Thrust [µn] Emitted Current [µa] Figure : 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 involvement and support of ESA. The performance characterisation and modelling of this emitter type is well-documented in reports and publications. The Cs-Slit Thrusters are single-emitter thrusters. The length of the slit can be adapted to the respective needs of different missions. Rule of thumb: approx µn per millimetre of slit length. The thrust range of the HYPER FEEPs is identical to that of MICROSCOPE and SMART-2: µn, nominal thrust range. Precision slits (approx. 1.4 micrometers in width over the slit length) of 8 mm length have 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 abort of the 2000 h Endurance Test for GOCE, a number of improvements were initiated at ALTA, in order to overcome the experienced problems. Within the FEEP consultancy for HYPER, ALTA have identified these areas: 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 Doc. No: HYP-5-02 Page 25

29 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 to start in mid-2003 / second half of This test program shall demonstrate the maturity of the design, of the critical procedures, performance characteristics, and plume contamination using QCMs. The Caesium propellant is very difficult to handle, because of its extreme reactivity. Outmost care must be taken 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 be provided with a vacuum tight Container with a tightly closing Cover Lid + Release Mechanism. After Acceptance Test of the thrusters, the Cover Lid must be closed again, before the test chamber can be vented and opened. It must be pointed out, that contaminations of the propellant can not be directly measured. It is only the quality of contamination control in each step, which provides assurance that the propellant is not contaminated. Some contamination will be removed, when the thrusters are started up. This start up capability has been observed, but not quantified. It can be expected that spectrum of difficulties associated with the Caesium propellant will be overcome, within the MICROSCOPE and SMART-2 programs. 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 Doc. No: HYP-5-02 Page 26

30 The Propellant Reservoir is sized for 0.12 kg of Cs, but can be adapted in length to the actual need, without problems. In order to transport the propellant to the emitter, a vane system is integrated into the reservoir cylinder (see Figure 3.2-2). Whether the liquid Cs inside the reservoir will move under the action of gravity gradient, or under the influence of period temperature changes on the HYPER orbit cannot be judged at this stage. Figure 3.2-2: Propellant Management Device inside Cs-Reservoir (emitter interface is on the left side) In Figure and Figure 3.2-4, the current mechanical design of the Cluster Assembly (CA) for SMART-2 is shown. Four (4) FEEPs are arranged 90 apart and with a tilt angle of 60. On top of the CA, the neutralisers are mounted. The accommodation of such CA's on the HYPER S/C (at mid-height) would require 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 Doc. No: HYP-5-02 Page 27

31 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 lead away from a thruster arrangement as used on SMART-2 (see sect. 4.2), and towards thrusters arrangement option 1, which is defined in sect ALTA are currently working on three contracts for ESA: MICROSCOPE (with CNES): max. thrust 150 µn; Status: close to PDR; launch planned 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. Doc. No: HYP-5-02 Page 28

32 3.3 Neutralisers For FEEP applications, neutralisers are needed - not for the proper operation of the FEEP emitters themselves, but for the compensation of S/C charging due to the emitted positive ion current. For this reason, neutralisers positions can be chosen on any suitable location on the S/C, assuming adequate conductive paths. The charged particle environment on the 1,000 km orbit will provide some charge compensation. But at this stage of definition, it is assumed, that the neutraliser(s) must compensate the full charging due to the FEEP thruster ion beams. Specifics of HYPER: maximum thrust level of all FEEPs combined < 500 N corresponding to a total emission current of ~ 5 ma residual atmosphere at 1,000 km is << 10-8 mbar. In this environment, cathode degradation due to residual atmosphere is negligible. neutralisers can be positioned outside the areas with charge exchange ions. Also this degrading environment 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 was performed 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) carbon nanotube cathode (NASA-JPL). Rank 2: Scandate thermionic cathode by HeatWave; other FEA cathodes (CEA-LETI and Stanford Research Institute), Carbon nanotube cathode by BUSEK. The survey performed in [RD 7-01] is exhaustive and fully adequate for the HYPER Feasibility Study. The ranking 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 Doc. No: HYP-5-02 Page 29

33 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 FEA cathode of RAL-CMT! It is therefore recommended to reduce the number of neutralisers onboard HYPER to 2 units (one main and one cold redundant neutraliser, total). This also reduces the number of neutraliser supply & control channels in the PCU. In a next step towards neutraliser selection, cost and development status / 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. This neutraliser must be protected against atmosphere (requires container with cover lid). Doc. No: HYP-5-02 Page 30

34 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. Brief conclusions are given in sect FEEP Application for HYPER FEEP Thruster Usage After S/C initial rate damping and attitude acquisition, and after the Micro-Newton Propulsion System has been checked-out and commissioned, the FEEP Micro-Newton Propulsion System will be activated and its thrusters 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 sole control 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 reactivate the CGPS upon need, e.g. in case of emergency. In sect. 5, the trade-off between vented CGPS vs. stand-by CGPS is performed. The trade-off concludes with the recommendation to maintain the CGPS in stand-by. Disturbances originating from the CGPS can be held below the limiting levels imposed by the Payload. In Table 4.1-1, the FEEP thruster usage during the different mission phases is outlined. As sole control authority during Science Phase, the FEEP Micro-N Propulsion System must support both, Primary AOCS and Secondary AOCS. Secondary AOCS has to provide disturbance compensation and attitude control at the performance levels required during the science measurements. Primary AOCS will operate either in a mode identical to that of Secondary AOCS, but with reduced performance requirements, or in a mode with a coarser sensor complement. In [AD-05], an orbit insertion error of δh = +/- 20 km and δi = +/ is quoted. Without any orbit insertion error correction, the drift of the LTAN is in the order of a few degrees over the complete mission life time. An assessment by Astrium UK indicate, that the small orbit drift is tolerable for the experiments. Orbit insertion error 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 allows to reduce mission risk during Science Phase. Space Debris Mitigation Standards / Guidelines demand that LEO satellites "should, whenever possible, be safely 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-Propulsion System with respect to: required total impulse capability (approx. 80,000 N s), additional to the total impulse required for the Science Phase Doc. No: HYP-5-02 Page 31

35 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 Science Phase - eclipse seasons, outages, emergency operations, recovery Slew to and acquisition of new Guide Stars Scientific Measurement periods 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-Propulsion System. Sun-synchronicity would be lost, complicating S/C operations, unless also inclination change manoeuvres 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 with respect to maximum thrust level and total propellant demand, and would not be commensurate for a small and optimised scientific satellite like HYPER. EOL de-orbiting is therefore not baselined Disturbance Forces and Disturbance Moments The different disturbances (forces and torques) that must be compensated with the FEEP System are summarised in Table The major disturbances are shown in bold. All major disturbances occur as cyclic effects (approximately periodic with orbit or with 2 orbit ), because the satellite 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 "FEEP Requirements Specification". Doc. No: HYP-5-02 Page 32

36 Sources of disturbance forces Sources of disturbance torques Gravity Gradient radial effect (due to distance between CoM and Drag-Free Point) Gravity Gradient torque (due to unsymmetric MOI and offset angles between SC_X_p and orbit normal) Gravity Gradient perpendicular effect (due to distance between CoM and Drag-Free Point) Residual S/C magnetic moment interacting with Earth Magnetic Field Air drag Air drag induced torques (due to geometric asymmetry, in particular at offset angles between SC_X_p and orbit normal) Solar pressure Solar pressure induced torques (due to geometric asymmetry, in particular at offset angles between SC_X_p and orbit normal) Earth Albedo and IR Earth Albedo and IR induced torques Parasitic accelerations due to thruster alignment errors Parasitic moments due to thrust vector alignment and position errors Parasitic accelerations due to thruster mismatch errors Parasitic moments due to thruster mismatch errors Table 4.1-2: List of Disturbance Forces and Moments Baselined Thruster Arrangement for HYPER For HYPER, several thruster arrangements were investigated. Two alternatives exist, depending on the FEEP 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. Doc. No: HYP-5-02 Page 33

37 Figure 4.1-1: HYPER FEEP Thruster Arrangement (Option 1) with 4 x 3 thrusters Table outlines which control forces and moments can be produced by different combinations of thrusters: 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 the translational 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 All of these four thrusters point through S/C CoM, and provide only linear forces. Doc. No: HYP-5-02 Page 34

38 Function Primary Set Secondary Set Force in + X ( ) Force in - X ( ) Force in + Y ( ) Force in - Y ( ) Force in + Z ( ) Force in - Z ( ) torque about X (+) torque about X (-) torque about Y (+) torque about Y (-) torque about Z (+) torque about Z (-) 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 Two groups are defined: group {1, 2, 3, 4, 5, 6, 7, 8} firing tangentially in the Y-Z plane, producing forces in +/- Y and +/- Z, and moments about +/- X group {11, 12, 13, 14, 15, 16, 17, 18} firing tangentially in the X-Z plane, producing forces in +/- X, and moments 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 requires an octagonal S/C body, in order to fit into the ROCKOT fairing. Doc. No: HYP-5-02 Page 35

39 Figure 4.1-2: HYPER FEEP Thruster Arrangement (Option 1 a) with 4 x 3 plus 4 x 1 thrusters Solar Array 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} Doc. No: HYP-5-02 Page 36

40 Function Primary Set Secondary Set Third Set force in + X Torque x (-) pair, or Torque x (+) pair force in - X Torque x (-) pair, or Torque x (+) pair force in + Y force in - Y force in + Z force in - Z torque about X (+) torque about X (-) torque about Y (+) torque about Y (-) torque about Z (+) Force z (+) pair, or Force z (-) pair torque about Z (-) Force z (+) pair, or 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 other side, the thrusters can be operated in quadruples! Some thrusters pairs produce parasitic torques or parasitic forces, which must be compensated by an additional pair. Doc. No: HYP-5-02 Page 37

41 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 thrusters arrangements from the Phase A studies, and a summary of the requirements specified in [RD 2-05] and [RD 2-06] is given below SMART-2 Thruster Arrangement Two SMART-2 Phase A Studies were performed, one by CASA (E) and one by Astrium (UK). These two studies 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 cluster assemblies on the middle of the side panels (see Figure 4.2-1). Doc. No: HYP-5-02 Page 38

42 In Figure and Figure 4.2-3, the two arrangements are shown. Astrium have proposed to rotate the Clusters by 45 compared to the CASA version. Figure also illustrates the 90% ion beam width of the 16 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) Doc. No: HYP-5-02 Page 39

43 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 and continuous availability of the FEEP Micro-Propulsion System. These requirements have not yet been formulated. In Figure 4.2-4, the thruster noise requirements of SMART-2 (shown green) applies for thrust levels above 25 N (from [AD-02]. In the mean time, these requirements have apparently been tightened to < N/Hz below 25 N, and < N/Hz above 25 N (see [RD 2-06]) Spectral Density N/sqrt(Hz) GOCE Smart Frequency [Hz] Figure 4.2-4: SMART-2 Thruster Noise Requirements (green line) Technology: Cesium-Slit Emitter; single-emitter thrusters Nomenclature: CA PCU Cluster Assembly consisting of 4 single-emitter thrusters and 2 neutralisers Power Control Unit for one Cluster Assembly Doc. No: HYP-5-02 Page 40

44 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 Thrusters at 60 cant angle; separated by 90 Cs reservoir with 0.12 kg of Cs propellant 4 PCUs 1. Ion Beam Ion beam divergence (> 90 %) 2. Thrust Range min.thrust per emitter +/- 20 along slit; +/- 40 cross slit 1 N (required) 0.1 N (goal) [RD 2-06] [RD 2-05] [RD 2-06] max. thrust per emitter 150 N [RD 2-05], [RD 2-06] 3. Characteristic Curve Bias No specification F achieved = F commanded + B Max. deviation from linearity Hysteresis Scale Factor Error No specification No specification 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= 1N 5. Quantisation step Quantisation step < 100 ms for Fc= 1 N [< 500 ms for Fc= 150 N] 0.1 N below 50 N 0.3 N from 50 N N due to PCU; emitters respond within a few ms; [RD 2-05] [RD 2-05] [RD 2-06] Table 4.2-1: Preliminary SMART-2 Requirements (continued) Doc. No: HYP-5-02 Page 41

45 6. Thrust Vector Noise thrust noise ( Hz) 7. Thrust Vector < N/Hz up to 25 N < N/Hz above 25 N no split between thruster and PCU; [RD 2-06] Thruster directions cant angle = w.r.t. common mounting plane of CA; [RD 2-05] thrust stability < 0.5 % over entire life [RD 2-06] deviation actual from nominal 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 at 150 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 mm 4 boxes (l x w x h); [RD 2-05] 11. FEEP System Power basic power for PCU operating; 0 N thrust delta-power for PCU operating; 150 N thrust W per PCU < 30 W for 4 x 150 N thrust Table 4.2-1: Preliminary SMART-2 Requirements (continued) Doc. No: HYP-5-02 Page 42

46 HV for thrusters emitter: between and 9 kv; accelerator: from - 1 to - 3 kv (fixed setting) [RD 2-06] Gate voltage for neutralisers + 75 V to V RAL FEA Neutraliser [RD 2-05] Neutraliser levels low: 2.5 ma (at < 50 N); high: 7.5 ma (at N); at each CA; I emitter 1.5 ma, each thruster [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 W high level: 3.56 W during nominal operation; [RD 2-05] 13. Redundancy / Failure Tolerance / Reliability Thruster redundancy / failure tolerance PCU redundancy Neutraliser redundancy FDIR failure tolerance: any 1 thruster out of 16 thrusters Main + Redundant Logic Section; independent thruster supply; independent neutraliser supply 1 main + 1 cold redundant unit at each CA only overload / overcurrent are managed autonomously by the PCU telemetry available to OBC adequacy to be verified against reliability! [RD 2-05] [RD 2-05] needed to avoid failure propagation [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 verification by test of each thruster no specification no specification Table 4.2-2: Preliminary SMART-2 Requirements Doc. No: HYP-5-02 Page 43

47 4.3 GOCE Requirements GOCE will be the first mission that will use FEEPs within a Drag Free and Attitude Control System. Micro-N FEEP 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 GOCE Thruster Arrangement Each thrust direction is equipped with 12 In-needle emitters. The thruster arrangement on GOCE is shown in Figure (see [RD 3-01] and [RD 3-03]). Each of the 8 Micro-Thruster Assemblies (MTA) shown in Figure consists of 12 Indium-Needle Emitters, in order to produce the required maximum operational thrust level of 400 N per MTA, respectively 650 N during accelerometer calibration periods. The reservoir size per Indium Emitter is 30 grams, which exceeds the required total impulse capability by a factor of 2. The number of Indium Emitters per MTA is solely driven by the high thrust levels needed for adequate 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 backflow ions can occur outside this cone. Due to the criticality for the GOCE Solar Array, the 3000 h Extended Endurance 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 Doc. No: HYP-5-02 Page 44