The GOCE Ion Propulsion Assembly Lessons Learnt from the First 22 Months of Flight Operations

Transcription

1 The GOCE Ion Propulsion Assembly Lessons Learnt from the First 22 Months of Flight Operations IEPC Presented at the 32nd International Electric Propulsion Conference, Wiesbaden Germany Neil Wallace 1, Peter Jameson 2 and Christopher Saunders 3 QinetiQ Ltd, Farnborough, UK Michael Fehringer 4, Clive Edwards 5 ESA-ESTEC, Noordwijk, The Netherlands Rune Floberghagen 6 ESA-ESOC, Darmstadt, Germany Abstract: The Gravity Field and steady state Ocean Circulation Explorer (GOCE) is a mission implemented in the Earth Explorer line of research oriented ESA-Earth observation missions. The objective of the mission is to produce high-accuracy, high resolution, global measurements of the Earth s gravity, leading to improved gravity field and geoid models for use in a wide range of applications. The GOCE spacecraft was launched successfully in March 2009 with a nominal 2 year mission. The spacecraft is flying in a near-circular, sun-synchronous dawn-dusk orbit, with 96.5 inclination and an altitude of around 260 km. During scientific measurement the Ion Propulsion Assembly is used to provide drag free control of the satellite. Two QinetiQ T5 ion thrusters are used, one in operation and one in redundancy, each capable of throttling over a thrust range mN at a resolution of 12 µn. The telemetry data generated by the spacecraft presents a good opportunity to investigate long term in-orbit performance of this type of propulsion system and to provide lessons learnt to improve future missions. This paper presents an evaluation of the in-flight Ion Propulsion Assembly data from the GOCE spacecraft. The evaluation covers the variation of thruster performance and operating parameters with thrust level and through time, including comparisons with commissioning and ground test data. ESA GOCE GSE IPA IPCU ITA PXFA XST Nomenclature = European Space Agency = Gravity and Ocean Circulation Explorer = Ground Support Equipment = Ion Propulsion Assembly = Ion Propulsion Control Unit = Ion Thruster Assembly = Proportional Xenon Feed Assembly = Xenon Storage Tank 1 Principal Engineer, Department Name, NCWALLACE@qinetiq.com. 2 Project Manager, Department Name, PJAMESON@qinetiq.com. 3 Engineer, Department Name, CJSAUNDERS@qinetiq.com. 4 GOCE System Manager, EOP-PGM, Michael.Fehringer@esa.int. 5 Electric Propulsion Engineer, TEC-MPE, Clive.Edwards@esa.int. 6 GOCE Mission Manager, EOP-GM, Rune.Floberghagen@esa.int. 1

2 I. Introduction to the GOCE mission The primary aim of GOCE is to provide unique models of the Earth s gravity field and of its equipotential reference surface, as represented by the geoid, on a global scale with very high spatial-resolution and accuracy. Such an advance in the knowledge of the Earth's gravity field and its geoid will help to develop a more detailed understanding of how the Earth's interior system works. New and fundamental insights are therefore expected in a wide range of research disciplines and applications, including solid Earth physics, oceanography and geodesy. To measure the Earth's gravitational field accurately from space, the spacecraft must operate a very sensitive gradiometer at the lowest possible altitude. At such a low altitude, the Earth's outer atmosphere causes a significant drag on the spacecraft, the magnitude of which varies around each orbit due to thinning of the atmosphere above the poles. The change in drag is cyclic in nature, varying between 1 and 20 mn, depending on the altitude and solar activity. Superimposed on these orbital variations are higher frequency fluctuations of up to ± 4 mn in magnitude. The GOCE spacecraft has a mass of approximately 1000 kg and flies in a near-circular, sun-synchronous, dawndusk orbit at 96.5 inclination, at an altitude of about 260 Km. This profile leads to a baseline mission duration of approximately 2 years, and since the ion thruster must be continuously operational during the measurement phases, the lifetime requirement is as high as 21,000 hours. All of these requirements can only realistically be met using a highly controllable ion thruster system. The QinetiQ T5mkV thruster is ideally suited for this mission, having been designed for a nominal thrust of between mN, and including solenoid magnets control, which allows the operating parameters to be efficiently and accurately controlled over the required thrust range. As such the T5mkV represents an enabling technology for the mission. The objective of the project on which this paper reports, is to provide a systematic evaluation of the flight data from the GOCE spacecraft relevant to the Ion Propulsion Assembly (IPA) system. This system is then used to provide understanding of the performance of the system in the space environment, which can be fed into future developments of this type of propulsion system, such as the Solar Electric Propulsion System (SEPS) on Bepi- Colombo. In order to achieve this objective the following targets were set at the start the project: To develop a tool for the analysis of the flight data on the IPA (and other flight systems) provided by the GOCE project. To use the above tool to investigate performance trends during thruster operations, and to compare flight data with available ground test data. To assess the impact of the IPA system on other spacecraft systems such as the Drag Free Attitude Control (DFAC) and power systems. Commissioning of the IPA system took place between 30th March 2009 and 23 rd June 2009, Following completion of commissioning activities the altitude was lowered to a nominal 259.4km, although the instantaneous value changes around the orbit due to the non spherical nature of the Earth. The baseline 22 month science mission was initiated on 11th September 2009, with drag free control operations performed continuously except for a period in February/March 2010 when thruster operation was performed manually. In addition, following a serious communications malfunction on 8th July 2010, the spacecraft altitude was raised to 263km and the ITA operated at a constant 2.5mN through manual commanding until early September 2010, when the altitude was gradually lowered back to 259.4km over a period of three weeks. The selected orbit leads to a near constant solar illumination around the orbit. However, there are seasonal variations which lead to a change in the solar aspect angle of ±23.5, and eclipses up to 37 minutes. The orbital period is approximately 90 minutes. 2

3 II. The Ion Propulsion Assembly A. IPA Function and Performance The Ion Propulsion Assembly (IPA) on GOCE compensates in real-time for the drag force experienced by the satellite as it orbits. This requires a continuously variable thrust between 0.6mN and 20.6mN that is generated in real time on the spacecraft. The Drag-free Attitude Control (DFAC) sub-system on the spacecraft senses the acceleration imparted to the body of the vehicle, and commands the IPA to produce a thrust that counter balances this force. In this way, the only acceleration that the satellite experiences is due to gravity (it is drag free ). This enables a sensitive gradiometer on the vehicle to very accurately probe higher order harmonics of the Earth s gravitational field. In addition, the IPA supports satellite maintenance phases by providing sufficient thrust for orbit raising manoeuvres. Primary and redundant branches of the IPA are included in the spacecraft. There are two T5 engines on GOCE, known as the Flight Model 1 (FM1) and Flight Model 2 (FM2) engines. The FM1 system is the primary engine, with the FM2 being redundant, and apart from a brief period during the system commissioning when FM2 was operational, FM1 has been used throughout the mission so far. There are two IPA branches used for FM1 and FM2 and these are known as Branch A and Branch B respectively. Operation during the mission phase (post commissioning) has been entirely with Branch A. B. IPA Physical Architecture The IPA on the GOCE satellite is made up of the Ion Thruster Assembly (ITA), the Ion Propulsion Control Unit (IPCU), the Proportional Xenon Feed Assembly (PXFA) and the Xenon Storage Tank (XST). Figure 1. Principle IPA constituents 1. Ion Thruster Assembly (ITA) The ITA is known as a QinetiQ T5 ion engine, and is a 10 cm diameter gridded ion thruster with a direct current discharge between a hollow cathode and a cylindrical anode used to ionise propellant gas (Xenon). This configuration is often referred to as a Kaufman thruster after the originator of the discharge chamber concept. The efficiency of the plasma production process is enhanced by the application of a weak and variable magnetic field within the discharge chamber generated using 6 circumferential solenoids. A 10cm diameter spherically dished twin grid system, forming the exit to the discharge chamber, extracts and accelerates the ions (via an applied potential difference) to provide the required thrust. The inner grid (known as the screen grid ) is fabricated from pressed and heat treated molybdenum, whilst the Accel grid is machined from solid graphite. An external hollow cathode, referred to as the neutraliser, emits the electrons necessary to neutralize the space charge of the emerging ion beam. On GOCE the thruster was equipped with an integral alignment interface (referred to as the alignment bracket) to allow the geometric axis of the thruster to be canted by up to 2.5 degrees to the spacecraft mounting interface plane. 3

4 2. Proportional Xenon Feed System (PXFA) The PXFA was developed, qualified and supplied by Bradford Engineering (The Netherlands) and is split into pressure control and flow control sections. The pressure control section receives xenon from the tank at high pressure (between 5 and 125 bar) and regulates down to a pressure of 2.50 ± 0.05bar. This section also includes particle filters, pressure transducers for monitoring the tank and supply pressures, and isolation valves to allow switching between primary and redundant units. The flow control section provides constant flowrates to the T5 Cathode (0.11mg/s) and Neutraliser (0.041mg/s) using advanced viscosity/thermal controlled passive flow restrictors, and varies the Main flow according to the demanded thrust level ( mg/s). This variable flow function is provided by a combined proportional flow control valve and flow sensor system, which allows the flow sensor output to be used for closed loop control of the flow control valve. Elimination of micro disturbances has been achieved by designing isolation features at component and flow control section level, as well as performing highly accurate modal stiffness modelling of the complete system. 3. Ion Propulsion Control Unit (IPCU) & IPA control architecture The IPCU was developed, qualified and supplied by EADS Astrium Crisa (Spain) and was designed to provide all of the power supplies (eg beam, anode, magnets, cathodes, anode, heaters) and control functions for the ITA and the PXFA as follows: Control electronics provide TC/TM communication with the spacecraft via the two MIL-1553 interfaces, timing synchronisation with the spacecraft using a PPS signal, and implements the PXFA interface. AC inverter converts the DC spacecraft power into two AC power outputs for the low voltage (LV) and high voltage (HV) power supplies. Ion Beam Converter converts the DC spacecraft power into the HV DC source required for the ion beam. LV control provides auxiliary DC/DC conversion for internal IPCU functions and provides TM/TC links between the Control Electronics and the LV supplies and HV control. LV supplies implements the LV power supplies, interfacing directly with the ITA. HV control provides auxiliary DC/DC conversion for internal IPCU functions and provides the TM/TC link with the LV control. HV supplies implements the HV power supplies, interfacing directly with the ITA. The IPCU also contains an integral ERC 32 CPU microprocessor that allowed the implementation of the IPA Thrust Control Algorithm, developed by QinetiQ. C. IPA Control Architecture The control algorithm allows the IPA to be commanded by a single thrust demand from the spacecraft computer, with an architecture as shown at Figure 2. The flow rate and anode current are adjusted at 10Hz, in an open loop mode to provide coarse control of the thrust, while the magnet current is adjusted at 100Hz in a closed loop mode to provide fine control for higher accuracy and quick response. A thrust demand is input to the IPCU from the spacecraft DFACS at 10Hz. Figure 2. IPA Thrust Control Algorithm Schematic 4

5 Within the IPCU software the raw demand signal is filtered to provide smoothed demand inputs for the flow rate and anode current controls. The IPCU then uses the control algorithm schedules to calculate the anode current and flow rate demands for inputs to the ITA and PXFA respectively. The PXFA uses a separate control loop to ensure that the flow rate is provided to the ITA with high accuracy. Fine control of the output thrust is provided by closed loop control of the magnet current. The IPCU software compares the raw demand with the measured thrust, and the resulting error signal used to drive the magnet current. The measured thrust is calculated directly from the ion beam current and voltage telemetry, with a correction factor applied to take into account physical effects on the actual thrust such as beam divergence, doubly charged ions, and neutral propellant flows. Because of the interactive nature of the three control parameters, the magnet current gain is a complex function of thrust demand, thrust error, flow rate and anode current. III. IPA Thrust Data A. Thrust Profile and Achieved Thrust Errors The variation in thrust over the mission is presented in Figure 3. Thrust demand can be seen to be lower than expected, but with a trend towards an increase in the mean. Some short term increases in thrust demand are also evident as a consequence of solar activity and altitude. Figure 3. Thrust history of the IPA since the start of GOCE science operations The difference between thrust demand and actual thrust is presented in Figure 4, which shows that the majority of the thrust error occurs within ±5% of the thrust demand. Operation of the engine has occurred predominantly around the 2mN level, and this corresponds to a noisier part of the schedules as observed during ground testing. This noise corresponds to coupling between the IPCU power supplies and the ITA discharge plasma. For comparison, the thrust error presented in Figure 5 shows significant improvement, typically within ±1%, for thrust greater than 3mN. 5

6 Number of Occurances %-90% -80% -70% -60% -50% -40% -30% -25% -20% -15% -10% -5% 0% 5% 10% 15% 20% 25% 30% 40% 50% 60% 70% 80% 90% 100% Thrust Error Bin 15th Sep th Feb th July nd Dec th Apr 2011 Figure 4. Histogram for thrust errors for 5 epochs during science mission phase Figure 5. Histogram for thrust errors for 5 epochs during science mission phase, limited to a thrust demand >3mN B. Geophysical and Solar-Terrestrial Influences on Thrust In general, thrust demand varies continuously due to fluctuations in atmospheric density which occur as a function of altitude, latitude, longitude, time of day and solar activity levels. Figure 6 shows how the thrust varies around a typical sequence of orbits, where the altitude varies due to the Earth s oblateness (ie the altitude is higher at the poles). At higher altitude the atmospheric density is lower, leading to a reduced thrust demand. In addition, the effects of atmospheric heating can be observed in differences between ascending and descending nodes. Figures 7 6

7 and 8 show how the atmospheric density and hence thrust demand is higher for the ascending node after solar heating during the day, and conversely for the descending node Thrust (µn) Altitude (km) :00 01:12 02:24 03:36 04:48 06:00 Date 07:12 08:24 09:36 10:48 Thrust Altitude Figure 6. Thrust and orbital altitude for several orbits during March :00 Figure 7. GOCE at its ascending node 7

8 Figure 8. GOCE at its descending node IV. Commissioning Data A. Discharge Parameters During the commissioning phase, the IPA was operated first in open loop control, with initial Branch A operations on 2nd April 2009 (ramped up to 8.3mN) and Branch B operated up to 20mN on 3rd April The magnet current flight data is broadly consistent with test data, displaying noise at low thrust as previously identified. Initially it appears that FM1 can be seen to approximately follow the ITA data, ie with GSE power supplies in place of an IPCU. However, the IPA test measurements were performed at only 3 thrust levels, so the magnet current profile is not apparent between 1mN and 8.3mN. In addition, the commissioning phase only set each thrust level for a short period, hence thermal equilibrium was not achieved. The exceptions to this occurred at 1mN and 8.3mN, where the dwell time was longer. At these conditions the magnet current demand can be seen to be reducing to a level more consistent with the IPA test data. The high magnet current between 1mN and 2mN is a feature of the gain in the schedules, which is not evident in the test data without an intermediate data point. 8

9 Figure 9. Magnet current as a function of thrust for FM1 during commissioning Similar effects are also displayed in Figure 10 for the Discharge Cathode Keeper voltage. Once again the flight data is consistent with test data, with low thrust noise and dwell time effects evident. Figure 10. DC Keeper voltage as a function of thrust for FM1 during commissioning The neutraliser cathode keeper voltage for FM1 is presented in Figure 11. This shows the neutraliser keeper at much lower voltages than found in ground testing, and a trend towards decreasing voltages as the thrust is increased. This may be explained by the neutraliser plasma coupling with the beam in flight, causing the plasma impedance to decrease. This effect would increase as a function of thrust. 9

10 Figure 11. NC Keeper voltage as a function of thrust for FM1 during commissioning The anode voltage values recorded during the commissioning phase can also be seen to match the pre-launch test data. The general trend however, is for lower voltages than during ground test. The actual delivered thrust is a complex function of flow rate, anode voltage, magnet current and grid current, and these are balanced in the IPCU by the schedules to ensure that enough headroom exists in the system to rapidly throttle up or down from the current operating point. Figure 12. Anode voltage as a function of thrust for FM1 during commissioning 10

11 B. Ion Beam Parameters Figure 13 shows the grid current as a function of thrust for FM1 during commissioning. Similar to the discharge parameters the flight data is consistent with ground test behaviour. The effect of the dwell time at a given thrust level is evident at 4.8 and 8.3mN, where the temperature modifies the grid separation and hence grid current. During ground testing the measurements were taken after temperature stabilization. Figure 13. Grid current as a function of thrust for FM1 during commissioning C. Performance Parameters The power demand from the ITA during commissioning is shown at Figure 14 for FM1. The data is very closely matched to the pre-launch test values. The power demand is lower than the requirement for both engines. Figure 14. Power demand as a function of thrust for FM1 during commissioning 11

12 V. Science mode Data A. Discharge Parameters During science phase operations, the Branch A IPA has been operated in closed loop DFAC control, with the thrust control algorithm setting the thrust demand. As can be seen in Figure 15 since the science phase has been operational, the majority of the data is clustered around 2mN, with a tail extending up to just less than 7mN. Figure 15. Histogram of thrust levels during science mission The time history of the GOCE thrust levels can be used to calculate the total impulse generated by the IPA in the mission so far. The total cumulative impulse at 20th May 2011 was 0.116MNs. The mission requirement was for 0.795MNs total impulse capability, so this would suggest significant lifetime is still left in the engines in terms of impulse generation. At current thrust levels, there should be no issue with reaching an extended mission end-of-life, in December This is illustrated at Figure 16, which shows the predicted cumulative total impulse assuming the ITA was to operate continuously at a constant thrust from the last available actual data point. This shows that even if the ITA was operated continuously from now on at an average of 10mN thrust, the mission would reach the end of 2012 (the nominal end of an extended science mission), with a total impulse on the thruster of only ~0.63MNs. At lower thrust levels (likely given the experience so far), the mission could extend beyond this point. It is also worth noting that the total impulse capability for the T5 is rated for 1.5MNs, so there is significant margin in the lifetime over the full thrust range. 12

13 Figure 16. Predicted cumulative total impulse from the last available actual data point, for a range of possible thrust levels (assumed constant over the rest of the mission) Any changes in thruster performance through life are automatically compensated for by the magnet current control loop, effectively leading to increases in the magnet current as the thruster performance degrades due to grid erosion (leading to reduced neutral density in the discharge chamber and consequently lower ionisation efficiency), and temperature changes (significant for ground testing due to sputter deposition, but should be lower in orbit). Figure 17 shows the magnet current values at a fixed thrust of 2mN during the mission science operations phase. The current has remained approximately constant throughout the mission. As the thrust levels have generally remained low, any grid erosion has been minimal and as such the magnet current has not had to compensate for other parameter changes. Figure 1. FM1 thruster magnet current at a fixed thrust of 2mN during the mission science phase 13

14 Figure 18 shows the magnet current at five epochs during the mission science operations phase. The current ramps up to peak at 1.5mN, before falling off again. The data matches the discrete test data points from the testing reasonably well apart from the obvious spike at 1.5mN. There is also significant scatter around 2mN which is to be expected due to the noisy engine operation at low thrust and the rapid magnet current modulation (100Hz). The spike in magnet current can be linked to the magnet gain f2 parameter in the IPCU look up tables. A fixed value of 0.09A/mN is applicable up to 1.5mN, subsequently dropping to 0.06A/mN at 2.0mN, A/mN at 2.5mN, and then a steady decline to about 0.01A/mN at 9.0mN. Overall the flight data appears consistent with the IPA test data at 1mN, and a slight increase through life is evident consistent with the PVM engine 5000hour test. Figure 2. Magnet current over two orbits for five epochs during the mission science operations phase The discharge cathode keeper voltage behaviour at a fixed thrust of 2mN over the course of the mission science phase is shown at Figure 19. For a fixed thrust, the discharge cathode voltage should fall initially through the early phase of the mission. This can be explained by conditioning which occurs as the oxide layer depletes, exposing more barium. Later in life, excessive barium depletion can lead to a subsequent rise in voltage, although depletion levels can be expected to be low due to the lower than anticipated thrust demand in orbit. Figure 3. FM1 thruster DC keeper voltage at a fixed thrust of 2mN during the mission science phase As expected, at a fixed thrust, the voltage initially drops through the first two months. Around May 2010 (~225 days into the science operations phase), the voltage starts to increase, and has been rising steadily ever since. The 14

15 data analysis ends on 20th May 2011, so the voltage has only risen by ~1V over the course of the last year. Extrapolating the second half of the science operations phase data from 350 days onwards to the end of 2012 when an extended GOCE science operations phase might be expected to finish, the predicted average voltage is 17.9V. This would still be within specification for a thrust level of 2mN. Figure 4. DC Keeper voltage over two orbits for five epochs during the mission science operations phase Discharge cathode keeper voltage as a function of thrust for five epochs during the mission science phase is shown at Figure 20. Flight voltages appear lower than test data, possibly due to higher outgassing of contaminants in the harder vacuum experienced in orbit. The cathode lifetime has been extensively investigated during ground testing. In these tests, a very high lifetime (of the order 15,000hr) was observed with only a 5% depletion of the barium in the emitter, whilst an additional investigation showed that the cathode would operate satisfactorily with 50% barium depletion levels (equivalent to 150,000 hours of operation). Further testing in support of GOCE simulated a 70,000hr duration mission at 200km altitude (the poor vacuum at low altitudes, can contaminate the keeper plate and tip). This test showed - despite contamination of the keeper - nominal cathode performance. Overall, it is apparent that significant lifetime still exists in the cathodes on GOCE. The neutraliser cathode keeper voltage behaviour at a fixed thrust of 2mN over the course of the mission science phase is shown at Figure 21. Neutraliser cathode keeper voltage as a function of thrust for five epochs during the mission science phase is shown at Figure 22. In both plots a reduction in the voltage over the course of the mission can be seen. At a fixed thrust, the voltage can be seen to drop by ~0.5V following the start of science operations. Since September 2010, the voltage has been relatively constant. Considering performance as a function of thrust, the data clearly shows that a) the voltage as been dropping as a function of time (this is consistent with the PVM 5000hr test), and b) at all times the performance has been considerably better (i.e. lower voltage) than predicted from the pre-launch ground tests. 15

16 Figure 5. FM1 thruster NC Keeper voltage at a fixed thrust of 2mN during the mission science phase Figure 6. NC Keeper voltage over two orbits for five epochs during the mission science operations phase Figure 23 shows the anode voltage at a fixed thrust of 2mN during the mission science phase. Overall the voltage has remained fairly constant, although there is some evidence of a sinusoidal type oscillation in the voltage, which may be due to seasonal temperature variations. 16

17 Figure 7. FM1 thruster Anode voltage at a fixed thrust of 2mN during the mission science phase Figure 24 shows the anode voltage as a function of thrust over the five analysis epochs. Also shown is the anode gain parameter from the IPCU lookup tables (parameter f3 ). Compared with the magnet current the change in gain is less pronounced between 1.5mN and 2.5mN, but the anode voltage is affected probably by the magnet current. Long term however, the anode does not appear to be degrading at any appreciable rate. Figure 8. Anode voltage over two orbits for five epochs during the mission science operations phase B. Ion Beam Parameters Figure 25 shows the grid current during normal mission operations at 2mN. The current has remained relatively constant through life, but there is evidence for an increase in the spread of grid current values over time, which 17

18 could be due to a variation in thermal conditions. In addition, small step changes in value can be observed following a re-start event when the engine has been off. This is consistent with experience during ground testing when temperature changes following re-starts would lead to different grid separations and hence small changes in grid currents. Figure 9. FM1 thruster grid current at a fixed thrust of 2mN during the mission science phase In flight data appears lower than observed during IPA testing, possibly due to the vacuum conditions. There is no significant evidence of degradation in the IPCU electronics or grid erosion leading to higher grid currents through increased thermal expansion. However, the total impulse generated over the first 22 months (0.11MNs) still represents only a small percentage of the lifetime capability. The low average thrust levels demanded in orbit compared to the qualification program, combined with the low grid currents suggest that substantial margin in the grid lifetime should still be available. Figure 10. FM1 grid current over two orbits for five epochs during the mission science phase. 18

19 Beam events occur due to a number of different processes, but all lead to a collapse of the plasma field in the thruster and an extinguishing of the ion beam. Beam-outs can occur for the following reasons: Variation in the screen and accel grid spacing due to thermal variations. Flakes from sputtered thruster material landing on the grids. Plasma perturbations if operating outside of a stable operating envelope. Beginning of Life (BOL) grid imperfections due to manufacturing processes (these are eroded away in the very early stages of operation, and hence are not a major beam-out contributor in the main stages of the mission). Outgassing of the thruster (again primarily at BOL). The cumulative number of beam trips and accel grid trips experienced on the ITA during the mission science phase is shown in Figure 27. The system is designed such that there should be an immediate restart of the engine if a beam-out is detected. Telemetry data is provided on both Beam Trips (a collapse of the ion beam in the discharge chamber) and Acc Grid Trips (a collapse of the beam due to material on the accel grids). The system requirement is to recover to thrust control from beam outs within 2 seconds. If not, an Event should be generated and indicated in the IPCU telemetry (subject to appropriate parameter monitoring). In all cases the beam can be seen to have recovered by the next housekeeping packet (ie within 8 seconds). Fortunately, diagnostic logs are generated on board at a higher data logging, in the event of a beam-out occurring. These can be requested for ground transmission, although they are not stored onboard for ever, and are subsequently overwritten on the next event. From a beam out event that occurred on the 21st June 2011, the ITA transitions to STANDBY mode but is back in THRUST mode exactly 0.5s later, and the thrust level recovers back to the pre-beam out levels exactly 1 second after the beam-out. This is the same behaviour as seen in pre-flight testing. Figure 11. Beam trip count, and accel grid trip count for GOCE during mission science phase C. Performance Parameters Figure 28 shows the power demand as a function of thrust over the five analysis epochs. The power demand appears to be quite constant over the mission lifetime, and is slightly lower than the pre-launch test data. This is likely due to the lower than expected voltages and currents in the cathodes and other components. Also shown is the ITA total power demand requirement. The data generally shows lower power demand that the requirements (by ~10W in the best case), although around 2mN thrust there is more evidence for the plasma noise. Although the total accumulated hours of operation are close to the design limit after 2 years in orbit, the low thrust implies lower degradation (hence effect on power demand) than expected. 19

20 Figure 12. Power demand over two orbits for five epochs during the mission science phase. Nominally, specific impulse may be expected to drop through life as the grids erode. However, on GOCE, the ITA is operated in a way that ensures the specific impulse is kept as constant as possible (at a given thrust). This is done by operating at a constant flow rate for a given thrust, and compensating for any changes in the thrust performance, but increasing the beam power (i.e. by increasing the magnet currents). Figure 29 shows the specific impulse to be consistent with test data and better than the requirement. Figure 13. Directly calculated specific impulse for 12th March 2011 to 20th May 2011 plus a 151µg/s allowance for the fixed flows to the discharge and neutraliser cathodes 20

21 VI. Conclusions and Lessons Learnt This document has reported on the results of a study into the in-flight performance of the GOCE Ion Propulsion Assembly (IPA), using QinetiQ T5 ion engines. The FM1 ITA has been the only operational engine since September The engine has exhibited very stable performance, with all major engine parameters remaining within expected limits, and in some cases performing better than during pre-launch testing. In terms of thrust, the required magnitude of thrust from the ITA has varied between ~1mN and 7mN during the main science operations phase of the mission, with the majority of time spent at thrust levels of ~2-3mN. Thrust error (the difference between the commanded thrust and the delivered thrust) has been low, with the majority of data points having a thrust error of <0.1mN. Of the main engine parameters, the grid current shows stable values over the course of the mission so far. Grid current dropped slightly following launch (possibly due to burn-in of the grid set), but now appears stable. The spread in grid current at a fixed thrust is getting larger with time, but only slowly; this suggests the grid erosion rate is low, and significant margin should still be available in the grid lifetime. The magnet current has also shown stable performance. The in-orbit data seems to match the ground test data well, and the influence of the IPCU gain parameters can be seen in the telemetry, indicating that this aspect is also working as expected. For both cathodes on the ITA (discharge and neutraliser), the in-orbit data shows lower operating voltages than recorded during pre-launch testing. This is possibly due to better outgassing of contaminants in the harder vacuum of space. The cathodes show some evidence of conditioning, whereby initially the voltage drops as oxide layers deplete on the cathode tip exposing barium. Later in life, as the barium is depleted, the required voltage rises, but there is significant margin on lifetime. The beam-out rate for the ITA has been assessed, and has been approximately once every 13days up to July 2010, and once every 7 days since October These are relatively low rates and in all cases, the ion beam automatically recovered back to full thrust mode within 8 seconds. The anode voltage has also been analysed. In the presence of grid wear and cathode performance degradation, the anode voltage could be expected to increase over time. However, the GOCE data shows that it has been stable over the course of the mission, and has been lower than anticipated from ground testing (related to the lower than expected cathode voltages). The power demand and specific impulse from the ITA have been slightly better than anticipated from pre-launch testing, probably related to the lower than expected power demand from the cathodes and anode. Consequently a mission extension to December 2012 appears readily achievable. Overall, the GOCE flight experience has shown that Kaufman ion thrusters can successfully operate for long periods of time in the space environment, and exhibit predictable performance. The low level of anomalies, and steady performance, should build confidence, enhance the reputation of the technology, and place it in a good position for use on future missions. 21