Enabling High Performance Green Propulsion for SmallSats

Transcription

1 Space Propulsion Redmond, WA Enabling High Performance Green Propulsion for SmallSats Robert Masse, Aerojet Rocketdyne Ronald Spores, Aerojet Rocketdyne May Allen, Aerojet Rocketdyne Scott Kimbrel, Aerojet Rocketdyne Chris McLean, Ball Aerospace Small Satellite Conference Logan, UT, USA 8-13 Aug 2015

2 Introduction NASA s Green Propellant Infusion Mission (GPIM) Program has just shipped the first mission-ready microsat green propulsion system delivering 50% greater density-isp than a comparable hydrazine system Funded by NASA Space Technology Mission Directorate (STMD) Hosted on Ball Aerospace BCP-100 ESPA-class spacecraft bus Aerojet Rocketdyne GR-1 (1N-class) thrusters Blow-down system with heritage propellant tank Heritage hydrazine system components with little or no modification 100% 90% 80% 70% 60% 50% 40% 30% 20% 10% 0% Monopropellant Market Share by Thrust Class >100-N 22-N 2/4 -N 1-N AF-M315E Ionic Liquid Propellant 6% higher Isp than N 2 H 4 45% more dense Low toxicity ~zero vapor pressure Open container safe Green Safe to handle in open containers 2

3 GPIM System Functional Schematic and Performance Service Valve Service Valve Orifice P GN 2 AF-M315E Pressure Transducers P Filter Design Characteristics Propellant... AF-M315E Total Mass (Wet) kg Propellant Load kg Thrusters...5 GR-1 Thrusters MEOP bar (480 psia) Specific Impulse sec (lbf-sec/lbm) Blow-down Range bar ( C Power: Catalyst bed Heaters V System Heaters.42 28V Thruster Valves V L Latch Valve T T T T T Temperature Sensors GPIM system configuration and operation are identical to conventional N 2 H 4 system ACS Thrusters Divert Thruster 3

4 GPIM System Configuration ACS Thrusters Component Panel, (Latch valve & filter located on back side) Service Panel V Thruster Safe/Arm Plug Pressure transducers Propellant tank Propulsion Interface Unit (PIU) (prop/telemetry electronics, Ball) Payload Interface Platform (PIP) (provided by Ball Aerospace) 4

5 GR-1 and GR-22 Thrusters Equal or greater utility compared to conventional hydrazine thrusters stipulated at outset of design effort: Delivered Isp consistent with propellant maximum capability No operational duty cycle limitations (as is the case for standard hydrazine thrusters) Power requirements consistent with what is available on missions being performed by likely early adopters (e.g. ESPA-class) Thermal soak-back to spacecraft similar to standard hydrazine thrusters Structurally capable of withstanding same launch environments as standard hydrazine thrusters Envelope same or smaller than comparable hydrazine thruster Redundancy on all fracture-critical features GR-1 (1-N class) GR-22 (20-N class) All design criteria met or exceeded 5

6 Aerojet Rocketdyne Advanced Propellants Test Lab 1.8 m dia 2.6 m Stainless Steel High-Altitude Chamber 64-channel Dewetron 204 ks/s data acquisition system Dual Stokes 1739 combination vacuum pumps Type B and K thermocouples Jenoptik VarioCAM thermal imaging camera 6

7 Prototype Thruster Testing Thrust Stand Mount Ar Purge Ring Ar Purge Ring Pitot Probe Pitot Probe GR-1 Thruster Thrust Stand Mount GR-22 Thruster GR-1 GR-22 Nozzle Expansion Ratio: 100:1 100:1 Valve 28VDC, 10 ºC (W): Preheat 28VDC (steady-state maximum, W) Feed Pressure (bar): Thrust (N): Maximum Steady-State Isp (s): Total Pulses Demonstrated: 11, Estimated Max Propellant Throughput Capability* (kg): Prototype Test Propellant Throughput (kg): Prototype Test Longest Burn (min): 20 1 Prototype Test Shortest Commanded Pulse Width (ms): *Based on heavyweight thruster testing 7

8 (GR-1) Thruster in Operation Photo Thermal Imaging 8

9 GR-1 Prototype vs. Lab Thruster Performance Transition after thermal wave reaches end of catalyst bed Heavyweight Test Unit GR-1 Flight-weight Prototype GR-1 prototype performs similarly to predecessor lab (heavyweight) thruster Thermal transitions moderately faster in flight-weight thruster, but similarity of ignition profiles shows lab thruster thermal insulation scheme yields good approximation of flight thruster Catalyst expected to demonstrate similar life capability in GR-1 as lab thruster Heavyweight testing will continue to be useful tool in ongoing efforts to further advance GR-1 thruster capabilities 9

10 GR-22 Prototype vs. Lab Thruster Performance Measured Thrust (N) FW PrototypeThrust LM Chamber Pressure Measured Thrust (N) FW Prototype Pitot Pressure P/Pmax Time (sec) Compared cases started from similar catalyst bed preheat temperature General character of lab (heavyweight) and GR-22 prototype curves similar 22-N LM thruster heat-up time-constant observed to be 2 that of EM thruster (1-N lab and flight thrusters bore greater similarity) Heavyweight test articles will continue in use to prove out basic thruster operation, but are less thermally representative for larger thrusters 10

11 GR-1 Pulse Mode Performance High pulse-to-pulse repeatability Duty cycle-independent stability of thruster design verified Ppitot (psia) Ppitot (psia) q P f = 425 psia q On Time (sec) Run Pulses 0.02/0.65 sec on/off 0.09/2.91 sec on/off Run Pulses 0.20/6.47 sec on/off Duty Cycle 0.5% 1% 3% 5% 10% 50% 75% 90% X 0.02 * * X * X Tested at 425 psia (29.3 bar) 0.03 X * Tested at multiple feed pressures 0.09 X X X 0.2 * * * * X 0.5 X X q * X X 1 X 1.5 X X Run Pulses 0.50/4.50 sec on/off Run Pulses Pulses proceed sequentially from blue to red Time from Command On (sec) Time from Command On (sec) 11

12 GR-1 Impulse Bit Repeatability 20 msec pulses showed Ibit RMSSD <0.6% at BOL Substantially greater pulse-to-pulse repeatability than comparable hydrazine thrusters Likely due largely to reduced hysteresis associated with Ti-cladding of magnetic valve components 300 Impulse Bit (mn-sec) msecon/178 on/1.78 msec off 20 msec on/180 msec off 20 msec on/180 msec off Calculated from postprocessed fast-response thrust stand data Pulse Number 0.20/1.78 sec on/off pulses did not reach steady-state within 20 pulses executed Smooth trend indicates at least similar pulse-to-pulse repeatability Long pulses demonstrate inherently greater repeatability Aerojet Rocketdyne Proprietary 12

13 GR-1 Impulse Bit Repeatability (cont d) Comparison of Ibit at different times during test shows expected trends Small drop during early thruster operation due to catalyst burn-in Thereafter Ibit remains nearly constant Small increase in Ibit variability is observed as thruster continues to accumulate life Test switched to longer duration life-accumulation firings for remainder of test performance tracking well with prior heavyweight thruster tests Transient where Ibit affected by difference in start temperatures Comparable portion of pulse trains Impulse Bit (mn-sec) Cumulative propellant throughput at start of sequence (Kg) Pulse Number 13

14 GR-1 and GR-22 Performance Summary GR-1 (psia) Isp = Isp o + [ F P] P where [ F ] = N bar Thrust = P ( Isp Isp ) f o 1- e Pf P Γ ref Ispo = sec Isp = sec where Pref = bar Γ = bar Thrust (N) Low flow rates challenge flow meter accuracy Isp (sec) GR-22 Thrust = Isp = Isp o [ F P] + P f Γ 1+ Po + Pf ( Isp Isp ) o 1- e where Pf P Γ ref [ F P] = N bar Γ = 1594 bar P = bar o Ispo = 222.1sec Isp = sec where Pref = bar Γ = bar Thrust (N) Isp (sec) Feed Pressure (bar) 14

15 Prototype Thruster Vibration Testing GR-1 GR-22 Power Spectral Density (G 2 /Hz) GPIM Vehicle Y,Z Axis GPIM Vehicle X Axis Solid lines indicate design levels. Dotted lines indicate levels tested to date. GPIM Vehicle Y,Z Axis GPIM Vehicle X Axis Frequency (Hz) Frequency (Hz) Both thrusters designed to maximum mission projected envelope (MPE) +6dB Vib testing performed along three GPIM spacecraft axes (skew to the GR-1 thrust axis) Tested levels MPE+3dB, except only at MPE along the Y and Z axes for the GR-1 due to interfering test fixture resonance GPIM flight units tested to MPE+3dB on all axes Both thrusters completed testing with no observable shift in fundamental frequencies or observable damage 15

16 GPIM Program Status and Conclusion 1-N and 22-N prototype thruster fabrication and testing complete Performance consistent with predictions Comprehensive stability testing verifies operability at any duty cycle High pulse-to-pulse repeatability Performance on track to repeat life demonstrated by lab thrusters GR-22 design/analysis updates (to mitigate chamber thermal fatigue issue discovered in test) complete to be verified by test Several producibility and performance-related upgrades have also been identified for the GR-1 First system delivered as GPIM demonstration mission payload Ready for integration with ESPA-class microsat bus architectures 16