Xenon Cold Gas Thruster (XCGT)

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This document consists of general capabilities information that is not defined as controlled technical data under ITAR Part 120.10 or EAR Part 772. All images used by permission of Moog/Maxar. Xenon Cold Gas Thruster (XCGT)

IEPC-2019-941

Presented at the 36th International Electric Propulsion Conference University of Vienna • Vienna, Austria September 15-20, 2019

Ian Johnson1 and Dustin Warner2 Maxar, Palo Alto, CA, 94303, United States

Kevin Neff 3 and Sean McCormick4 Moog Inc., East Aurora, NY, 14052, United States

Satellites must have the ability to correct high rotational tip-off rates after initial release from the launch vehicle. For large geostationary communication satellites this is typically done with a chemical propulsion system. In the case of an all-electric propulsion (EP) satellite, the torque available to counter high rotational inertias immediately after tip-off are not available from traditional low-thrust EP thrusters without the solar array’s deployed, nor reaction wheels. In place of a dedicated, and expensive, chemical system for this singular beginning of life application, a simple solution developed by Maxar is to use xenon cold gas thrusters (XCGT) directly connected to the primary propellant tanks. Such a system is low-cost, low- risk, low-power, and easily integrated with existing spacecraft infrastructure and manufacturing processes. The XCGTs for this application were designed and built by Moog Inc and are based on heritage single-seat solenoid valves. Developmental testing showed that they operate up to 2700psi inlet pressures with 2.8N of thrust with a specific impulse of 22sec at 21C. The thrust decreases with pressure, while both the thrust and specific impulse are highly dependent on temperature and have non-linear features at the Xenon critical point. An overview of the use of the XCGT to correct satellite tip-off rotation, and a performance summary of the thruster will be reviewed, including performance predictions as compared to test results. The first Maxar spacecraft to utilize the XCGT system launched in Q2 2019.

Nomenclature F = thruster force, N �̇ = mass flow rate, kg/sec ISP = specific impulse, sec d = moment arm, m Τ = torque, Nm � = moment of inertia, kg-m2 I = impulse, Ns L = momentum, Nms � = propellant mass, kg tf = firing time, sec � = angular velocity, rad/sec 2 g0 = gravitational constant, kg/m

I. Introduction LECTRIC propulsion (EP) spacecraft offer unique advantages over traditional chemical propulsion, primarily E reduced launch mass, as detailed in numerous papers over the past 70 years. Most electric propulsion spacecraft throughout history have been a hybrid design, incorporating both EP and chemical propulsion systems. The

1 Propulsion Engineer, GNC Systems Engineering, [email protected] 2 Propulsion Engineer, External Product Development, [email protected] 3 Technical Lead, Space and Defense, [email protected] 4 Staff Design Engineer, Space and Defense, [email protected]

The 36th International Electric Propulsion Conference, University of Vienna, Austria September 15-20, 2019 This document consists of general capabilities information that is not defined as controlled technical data under ITAR Part 120.10 or EAR Part 772. All images used by permission of Moog/Maxar. chemical portion being used for launch vehicle tipoff control, contingencies, the bulk or entirety of orbit raising, some stationkeeping, and deorbit1, 2, 3. While the EP portion for most stationkeeping, or interplanetary deltaV as on DAWN4 and SMART-15. Completely eliminating the chemical system allows for either smaller spacecraft, or a lower cost launch vehicle – either of which can be mission enabling. With the advent of all-EP spacecraft, tipoff rate stabilization is a concern. In Low Earth Orbit, the relatively high magnetic field strength allows for a combination of torque rods and reaction wheels to capture the spacecraft and reduces rates to zero. After which, the arrays can be deployed and electric propulsion can be used for the remainder of the mission. In a Geosynchronous Transfer Orbit, the magnetic field is far lower, making torque rods ineffective. For a large all-EP spacecraft, the ability to capture entirely on wheels is exceeded at ~1deg/sec. Prior to reducing rates to zero, large deployables, such as arrays, cannot be utilized. Without arrays, there is limited power available for electric thrusters. High launch vehicle rotation separation errors would therefore result in a mission failure, without the addition of a lower-power, higher-thrust propulsion system on the spacecraft. Outside of a dedicated chemical system for this singular purpose, one such option is to use small thrusters running off the primary Xenon tanks. Maxar has a long history with electric propulsion equipped spacecraft, accumulating over 100,000 hours of on- orbit operation without a systems level failure.14, 15, 16 Since 2004, Maxar has operated high-power SPT-100 Hall Thrusters on our GEO COM spacecraft. Starting in 2018 the higher thrust SPT-140 Hall Thrusters were first flown. These thrusters allow for the possibility of completing an entire 15-year GEO communications mission with only electric propulsion; one of which was launched in 2019 and as of the writing this paper, is in the middle of its Electric Orbit Raising. Contingency tipoff rate mitigation for this all-EP spacecraft was available through the use of Moog build Xenon cold gas thrusters (XCGTs). Moog has played a significant role supporting the space industry’s EP community in Japan, Europe, and the US, for over 20 years with the delivery of more than 500 EP components, and nearly 100 EP subassemblies, used for satellite and deep space missions. This includes more than 75 EP flow control valves and or EP bang-bang regulators that are the same or similar design as the XCGT presented herein. Moog’s overall cold gas thruster heritage also spans more than 20 years with the delivery of more than 1000 thrusters of various designs over that period in support of space applications6. The Moog Xenon Cold Gas Thruster concept was previously discussed at the 2017 IEPC7. Xenon cold gas thrusters have been of interest to the community for some time. Recent work includes the European Space Agency’s Small GEO platform8, the University of Tokyo’s micro Xenon cold gas thrusters as part of the I- COUPS propulsion system9, and the Texas Spacecraft Laboratories In-Space demonstration10. A Fakel Resistojet11 and the Phase 4 RF Thruster12 have both been tested with Xenon in cold gas mode. JAXAR is planning the use of Xenon cold gas thrusters on the ETS-9 spacecraft17.

II. Trade Study: XCGT and Monopropellant To perform this singular beginning of life (BOL) maneuver a trade study was performed between a xenon cold gas system and a monopropellant hydrazine system. Amongst chemical propulsion subsystems, monopropellant hydrazine architectures are inexpensive, robust, and have extensive flight heritage. Given the Maxar application has such a small impulse requirement, and therefore very little fuel is required, the dry mass of a monopropellant hydrazine system greatly exceeds the quantity of propellant required to fulfill its purpose. Alternatively, a xenon cold gas system can be tied into the primary propulsion subsystem, without the need for additional tanks, and provide a small benefit to the overall mission length by making its unused propellant available for the primary mission, as was realized on Maxar’s first all-electric spacecraft launched in Q2 2019. As can be seen in Figure 1, a Monopropellant system would require a standalone tank, three seals between the tanks and the exterior, and four monopropellant thrusters. The total dry mass of the monopropellant system would be >10 times that of the components dedicated for the xenon cold gas system. Additionally, the smallest TRL9 Hydrazine diaphragm tank located at the time of the trade was roughly 1L in volume, capable of holding 1kg of propellant. Aside from the added cost of the tank, the additional volume must be accommodated, whereas the cold gas system simply ties into the existing propellant tanks. The monopropellant system has a catalyst present, which poses a FOD risk if the thrusters are inverted. This levies orientation restrictions during spacecraft integration that can, at times, slow production. At launch base, in addition to its own suite of stand-alone testes, the monopropellant system requires hazardous propellant loading. The cost associated with coordinating the personnel and ground support equipment for such a small propellant load are unjustified. The xenon cold gas system does not require its own tank; with the only additions being an isolation valve and the four single seat thrusters, in addition to the required tubing. Due to the cold gas system’s simplicity, the cost is less than that of a monopropellant hydrazine system. Lastly, ground handling for the xenon cold gas system is simple and straightforward. As part of the primary subsystem manufacturing, Xenon is loaded into the spacecraft at Maxar prior to shipment to launch base, with minimal follow-on testing.

Figure 1. Example four thruster monopropellant system (left) vs. four thruster XCGT system (right). To calculate the required propellant mass to detumble a spacecraft, angular momentum (L) can be described by Eq. (1), which can be reduced to solve for angular velocity (�) via Eq. (2). Thruster parameters are related through Eq. (3), and the on-time varied via Eq. (4) to solve for a required propellant mass for a given starting angular velocity.

� = � × � × � = � × � (1) � × � × � � = (2) � � = �̇ × � × � (3) � = �̇ × � (4) A simple scenario is listed in Table 1 where a 5000kg-m2 moment of inertia spacecraft is detumbled from a 1deg/sec rotation rate using chemical, electric, and xenon cold gas thrusters. The high specific impulse electric propulsion system would obviously be the most efficient, while the XCGT system is the least. However, when coupled with the dry mass savings listed above, the net mass savings to the spacecraft is lowest with the XCGT system. Table 1. Detumbling a generic spacecraft from a 1deg/sec rotation rate comparison between chemical propulsion, high power electric propulsion, and Xenon cold gas thrusters. Chemical Electric XCGT Thruster Force (N) 1 0.180 1 Flow Rate (g/sec) 0.51 0.011 5 Specific Impulse (sec) 200 1670 20 Moment Arm (m) 1 1 1 Torque (N-m) 1 0.18 1 Moment of Inertia (kg-m^2) 5000 5000 5000 Firing Time (sec) 87 485 87 Impulse (Ns) 87 87 87 Momentum (Nms) 87 87 87 Propellant (g) 45 5 436 Ang. Velocity (rad/sec) 0.0175 0.0175 0.0175 Ang. Velocity (deg/sec) 1 1 1 For these reasons, a xenon cold gas system provided the simplest and most cost-effective solution to meet the requirements of this singular BOL contingency operation.

III. XCGT Design The Moog Xenon Cold Gas Thruster (XCGT) is an all welded stainless steel, normally closed solenoid valve, except with a nozzle in place of the outlet tube that is typical of a solenoid valve. The XCGT provides thrust in proportion to inlet pressure when in the opened position and a leak tight seal in the closed position. The XCGT has a single coil for operation with a nominal operating voltage of 28Vdc initially, and a hold open voltage of 10Vdc for

Figure 2. XCGT envelop in millimeters (left) and the unit undergoing vibration testing (right).

IV. Development Testing Individual thrust testing on two separate flight equivalent XCGTs was performed in a near vacuum environment (<1 psia), and ambient thermal conditions. Pressure, mass flow rate, thrust, and temperature data were taken for each test sequence. Variations in the testing included the inlet pressure range (300 – 2700 psia), pulse duration (1 – 30 seconds), and the number of pulses applied over a 10% duty cycle (2-17 pulses). Figure 3 is a photograph of the development thrust test setup in the vacuum chamber, with a thruster pointing vertically up.

Figure 3. XCGT Flight-Like Thruster in the Test Setup Moog hypothesized that the XCGT flow regime would transition from sonic flow to cavitating flow, as the supply pressure is increased, due to the supply entropy dropping below the critical point for xenon. The predicted supply pressure range where this transition would occur was between 900 to 1000 psia for gas supply temperatures around 21ºC. Figures 4 and 5 below compare, respectively, the thrust and mass flow rate data collected with the theoretical

Thrust vs. Supply Pressure at Vacuum 0.900 4

0.800 3.5

0.700 3

0.600 2.5

0.500 2 Thrust Thrust (N)

Thrust Thrust (lbf) 0.400

1.5 0.300

1 0.200

0.100 0.5

0.000 0 0.0 500.0 1000.0 1500.0 2000.0 2500.0 3000.0 Supply Pressure (psia) Measured Values Xe1 Measured Values Xe Dev Test #2 SN 0002 Dev Test #2 SN 0003 Predicted Vaues Figure 4. XCGT Predicted and Measured Thrust Results with Xenon

Mass Flow Rate vs. Supply Pressure at Vacuum 0.0350 0.016

0.014 0.0300

0.012 0.0250

0.01 0.0200

0.008

0.0150 0.006 Mass Mass Flow Rate (kg/sec) Mass Mass Flow Rate (lbm/sec)

0.0100 0.004

0.0050 0.002

0.0000 0 0.0 500.0 1000.0 1500.0 2000.0 2500.0 3000.0 Supply Pressure (psia)

Measured Values Xe Calculated Based on Thrust Measure Dev Test #2 SN 0002 Dev Test #2 SN 0003 Predicted Values Figure 5. XCGT Predicted and Measured Mass Flow Rates with Xenon

Isp vs. Supply Pressure at Vacuum 40

(sec) 20 sp I

0 0 500 1000 1500 2000 2500 3000 Supply Pressure (psia)

Prediction Xe SN 0002 Xe SN 0003 Calculated Flow Rates Figure 6. XCGT Predicted and Calculated Specific Impulse with Xenon

Figure 7. XCGT Xenon Plumes at Various Inlet Pressures

Moog enveloped the cold temperature thruster system performance with a test set-up that simulated worst case flight conditions. Two thrusters were operated simultaneously in parallel, with a single upstream isolation valve, and the xenon gas supply was cooled to minimum potential operating temperatures with no heaters powered. The results of this case as shown in Figure 8 indicate a significant decrease in both thrust and Isp as the temperature is reduced from ambient. The decrease in performance at reduced supply temperatures is mostly attributed to a significant pressure drop of the upstream isolation valve and feed lines due to higher viscosity of the xenon at lower temperatures.

Thrust vs. Line Temp. 0.300

0.250

0.200

0.150 Thrust Thrust (lbf)

0.100

0.050

0.000 -80.00 -60.00 -40.00 -20.00 0.00 20.00 40.00 60.00 80.00 100.00 Line Temperature (ºF) SN0002 Cold SN0002 Ambient SN0003 Ambient

Isp vs. Line Temp. 30.0

25.0

20.0

15.0 Isp Isp (sec)

10.0

5.0

0.0 -80.00 -60.00 -40.00 -20.00 0.00 20.00 40.00 60.00 80.00 100.00 Line Temperature (ºF) SN0002 Cold SN0002 Ambient SN0003 Ambient Figure 8. XCGT thrust (top) and specific impulse (lower) as function of temperature for 1300psi inlet pressures.

Figure 9. Generic four thruster layout on spacecraft for complete 3-axis control.

Figure 10. XCGT schematics for single (no redundancy) and dual face (single-fault tolerance) configurations.

As a practical example, with an 80Nms reaction wheel on large spacecraft (moment of inertia 5000kg-m^2), the ability to capture entirely on wheels is exceeded at rates above ~1deg/sec. To capture a 1deg/sec tip-off rate on the thrusters, two 1N XCGTs with 1m moment arms would require 44sec of dual thruster firing (88sec of total firing) as shown in Figure 11. With a 20sec specific impulse, 440g of Xenon would be required. This is compared to only 45g

Figure 11. Theoretical capture profile from high spin rate with the Xenon cold gas thrusters.

Excluding pressure drops through the pneumatic system, the expected thruster inlet pressure is a function of tank volume, loaded propellant mass, and temperature. As shown below for a generic 500L tank, the pressure will decrease dramatically as temperature drops. Assuming 45C/2700psi are the maximum allowable temperatures and pressures, a fully loaded tank (950kg) would likely launch at temperatures of 30-35C in order to hold both temperature and pressure margin. This would result in XCGT inlet pressures of ~2000psi. Running the system at 17C (the Xenon critical point) would further reduce the pressure to 1350psi. For the purpose of this paper, ambient temperature (21C) is assumed, resulting in BOL pressures of 1500psi. Per Figures 4-6, 1500psi and 21C (70F) would result in a thruster performance of 1.7N and 22sec specific impulse. Higher temperatures would increase both the thrust and specific impulse. As the mission progresses and the bulk of the Xenon is used for station keeping, the thrust of the XCGT will decrease (due to pressure decrease), although the specific impulse improves slightly due to the gas quality as discussed in Section IV. At end of life Hall Thruster pressures (~40psi), the XCGT thrust would be ~30mN (extrapolated from Figure 4), lower than that of the Hall Thrusters. As such, the XCGTs are only designed for a BOL application where the Electric Thrusters are unavailable.

Figure 10. Pressures and corresponding XCGT thrust levels for a 500L Xenon tank at varying temperatures.

VI. Conclusion The problem of high launch vehicle tipoff rates with large all electric propulsion GEO spacecraft requires a relatively high thrust, low power solution. Maxar solved this issue with Moog built Xenon cold gas thrusters, directly connected to the Xenon tanks. Developmental testing at varying pressures and temperatures showed good correlation between theory and experiment. This solution for this singular BOL operation was simple, light weight, low risk, and low cost. The first Maxar all-EP spacecraft utilizing the Moog Xenon Cold Gas Thrusters was launched in 2019.

The 36th International Electric Propulsion Conference, University of Vienna, Austria September 15-20, 2019 This document consists of general capabilities information that is not defined as controlled technical data under ITAR Part 120.10 or EAR Part 772. All images used by permission of Moog/Maxar. References 1Meyers, R., “Overview of Major U.S. Industrial Electric Propulsion Programs”, AIAA Joint Propulsion Conference, AIAA- 2004-3331, 2004. 2Corey, R. and Pidgeon, D., “Electric Propulsion at Space Systems/Loral”, International Electric Propulsion Conference, IEPC- 2009-270, 2009. 3Casaregola, C., “Electric Propulsion for Commercial Applications: In-Flight Experience and Perspective at Eutelsat”, IEPC- 2013-332, 2013. 4Brophy, J., et. al. “The Ion Propulsion System on DAWN”, AIAA Joint Propulsion Conference, AIAA-2003-4542, 2003. 5Koppel, C., and Estublier, D. “The SMART-1 Hall Effect Thruster around the Moon: In Flight Experience”, International Electric Propulsion Conference, IEPC-2005-119, 2005. 6Bzibziak, R. “Update of Cold Gas Propulsion at Moog”, AIAA-2000-3718, 2000. 7Loghry, C., et. al. “LEO to GEO (and Beyond) Transfers using High Power Solar Electric Propulsion (HP-SEP)”, International Electric Propulsion Conference, IEPC-2017-396, 2017. 8De Tata, M., et. al. “SGEO Development Status and Opportunities for the EP based Small European Telecommunication Platform”, International Electric Propulsion Conference, IEPC-2011-203, 2011. 9Koizumi, H., et. al. “Unified Propulsion System to Explore Near-Earth Asteroids by a 50kg Spacecraft”, 28th Conference on Small Satellites, SSC14-VI-6, 2014. 10Imken, T., et. al. “Design and testing of a cold gas thruster for interplanetary cubesat mission”, Journal of Small Satellites, 2015. 11Jankovsky, R., and Sankovic, J. “Performance of a Fakel K10K Resistojet”, Joint Propulsion Conference, 1997. 12Siddiqui, M., et. al. “First Performance Measurements of the Phase Four RF Thruster”, International Electric Propulsion Conference, 2017. 13Delgado, J., et. al. “Qualification of the SPT-140 for use on Western Spacecraft”. 50th AIAA Joint Propulsion Conference, 2014. 14Lord, P. W., Tilley, S., Goebel, D. M., and Snyder, J. S., “Third Generation Commercial Solar Electric Propulsion: A Foundation for Space Exploration Missions,” 2018 IEEE Aerospace Conference, Big Sky, MT, March 2018. 15Johnson, I. K., Kay, E., Lee, T., Bae, R., and Feher, N., “New Avenues for Research and Development of Electric Propulsion Thrusters at SSL,” IEPC-2017-400, 35th International Electric Propulsion Conference, Atlanta, GA. October 2017. 16Johnson, I., G. Santiago, J. Li, and J. Baldwin, "100,000 hrs of on-orbit electric propulsion and Maxar’s first electric orbit raising," to be presented at the 2020 AIAA SciTech Forum, Orlando, FL, Jan. 6-10, 2020. 17Funaki, I. “Development Status of 6kW-class Hall Thrusters at JAXA”, 36th International Electric Propulsion Conference, Vienna, Austria, September 2019.

The 36th International Electric Propulsion Conference, University of Vienna, Austria September 15-20, 2019