Integrated Vehicle and Trajectory Design of Small Spacecraft with Electric Propulsion for Earth and Interplanetary Missions

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California Institute of Technology Ben Longmier, University of Michigan; Derek Dalle, NASA/Ames Small Satellite

1 Integrated Vehicle and Trajectory Design of Small Spacecraft with Electric Propulsion for Earth and Interplanetary Missions Small Satellite Conference 2015 Sara Spangelo, NASA Jet Propulsion Laboratory (JPL)/ California Institute of Technology Ben Longmier, University of Michigan; Derek Dalle, NASA/Ames Small Satellite Conference, August 2015, Logan, Utah Copyright 2015 California Institute of Technology. Government sponsorship acknowledged.

Past, Present & Future of CubeSat Propulsion Systems Past: Low Earth Orbit (LEO) CubeSats passive drifters Present: Current State of the Art Cold gas systems for small ΔV<100 m/s, de-sats Large

makeup, deorbit CubeSats or debris in LEO Ability to perform formation flight (large apertures) Large maneuvers to transfer to comets, asteroids, planets!

2 Past, Present & Future of CubeSat Propulsion Systems Past: Low Earth Orbit (LEO) CubeSats passive drifters Present: Current State of the Art Cold gas systems for small ΔV<100 m/s, de-sats Large electric propulsion (EP) systems ~10 kg Future: Emerging EP solutions for CubeSats Game-changing and enabling/enhancing across a broad class of missions: Significant ΔV primary propulsion Drag makeup, deorbit CubeSats or debris in LEO Ability to perform formation flight (large apertures) Large maneuvers to transfer to comets, asteroids, planets! Ability to create/maintain constellations Hover, proximity ops, land on small bodies, rings, etc. Attitude control: de-saturate or replace reaction wheels Goal of this talk: Demonstrate future mission opportunities of small spacecraft with electric propulsion using systems-level perspective. Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. 2

Experience, Heritage, and Enabling Technology Significant flight experience and heritage in LEO and heritage components Telecommunication and Navigation systems High-rate X/Ka-Band radios (10+ Mbps

5 W Sinclair Interplanetary Reaction Wheels XB1 Blue Canyon System VACCO Cold Gas Systems (ΔV<80 m/s in 3U CubeSat) Solar arrays that are deployed and gimbaled for Sun-tracking Deployable Solar

3 Experience, Heritage, and Enabling Technology Significant flight experience and heritage in LEO and heritage components Telecommunication and Navigation systems High-rate X/Ka-Band radios (10+ Mbps in LEO) Iris Transponder and high gain antennas for deep space High-accuracy attitude control technology Blue Canyon s XACT: <2.5 kg, ~1 U, <2.5 W Sinclair Interplanetary Reaction Wheels XB1 Blue Canyon System VACCO Cold Gas Systems (ΔV<80 m/s in 3U CubeSat) Solar arrays that are deployed and gimbaled for Sun-tracking Deployable Solar Arrays (ehawk arrays ~130 W/kg) Integrated Computers, GNC, and Bus Architectures BCT XB1 Bus (GNC, C&DH, Telecom, Power, ACS) Radiation-tolerant flight computers (LEON, etc.) Companies developing buses like Tyvak, Blue Canyon Aluminum 3U CubeSat Structure (radiation shielding) Iris Transponder Image Credit: JPL, Blue Canyon, MMA Suppliers shown only for proof-of-concept; no selection is represented. Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. ehawk MMA Solar Arrays (130 W/kg) 3

4 Overview of Emerging Small Spacecraft EP Systems Thruster* (Point Design) Technology I sp Thrust System Power Units sec mn W Busek's 0.1 mn MEP Electrospray CAT Plasma Magnetoplasma Busek's 0.1 mn MEP Electrospray MIT ieps Electrospray Busek's Ion (BIT-1) Ion MiXI Ion Ion Busek's Ion (BIT-3) Ion JPL's MEP Electrospray *Thruster specs based on single operating points of publically available information Large variation in thrust to power with I sp Large variation in technology readiness levels (TRLs), making system-level comparison challenging Mass, volume, PPU efficiency also important factors Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. 4

CAT: Large V Engine Capability CAT: CubeSat Ambipolar Thruster Uses high-density plasma source Achieves high V and high thrust/power over range (3-300W) Fits within small spacecraft form-factor (<0.

5 CAT: Large V Engine Capability CAT: CubeSat Ambipolar Thruster Uses high-density plasma source Achieves high V and high thrust/power over range (3-300W) Fits within small spacecraft form-factor (<0.1 U, <0.5 kg) Awarded a CSLI Launch on PATRIOT mission awarded in 2014 Successful Kickstarter Campaign resulting in seed funding ($100 K) Some commercial funding supporting tech development Design of a 3U CubeSat with CAT engine performing initial testing in Low Earth Orbit. CAT Thruster Performance (I 2 used in analysis) Photo Credit: PEPL, CAT Performance Refernce: Longmier, B. and Sheehan, J. P., Initial Experiments of a New Permanent Magnet Helicon Thruster, International Electric Propulsion Conference, Washington, DC, Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. 5

Multidisciplinary System Level Modeling Propulsion Number of thrusters to achieve required thrust Propellant mass I sp,, Thrust Trajectory Design I sp, Thrust ΔV for orbit maneuver Thrust time, trip

) Deployable solar arrays, batteries BCT XB1 Bus (C&DH, ADCS, EPS, etc.) Structure, reaction wheels scale with size Iris transponder (tracking & communication) Total 6U dry mass ~ 6 kg S.

6 Multidisciplinary System Level Modeling Propulsion Number of thrusters to achieve required thrust Propellant mass I sp,, Thrust Trajectory Design I sp, Thrust ΔV for orbit maneuver Thrust time, trip time Power, propellant Available Power Subsystem Sizing Energy Capacity & Balance Solar Power Collection Thermal, Radiation Vehicle design based on other JPL deep space CubeSats (INSPIRE, MarCO, etc.) Deployable solar arrays, batteries BCT XB1 Bus (C&DH, ADCS, EPS, etc.) Structure, reaction wheels scale with size Iris transponder (tracking & communication) Total 6U dry mass ~ 6 kg S. Spangelo and B. Longmier, Optimization of CubeSat System-level Design and Propulsion Systems for Earth-Escape Missions", Journal of Spacecraft and Rockets, accepted December Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. 6

7 Simulation Framework Orbit Trajectory (ΔV, time, thrust) Propulsion System Model (thrust, I sp, mass) Thermal and Power Model Assumptions: Thrusters fire perfectly in desired direction. Spacecraft mass includes propulsion system (propellant, etc.), bus, and solar panels. Mass margin includes payload and PPU mass. Thrusters are modular and can be fractional Solar panels sized for continual thrusting. Spacecraft Size Compute Power, # Thrusters, Propellant Mass Size Solar Arrays Check Feasibility of Thermal and Power Systems M sc =M prop +M bus +M sp Masses: M sc Spacecraft : M prop Propulsion System M bus Bus M sp Solar Panels M max Maximum mass for given size (~2 kg/u) Modeling approach from: S. Spangelo, D. Landau, N. Aurora, S. Johnson, T. Randolph, Defining the Optimal Requirements for the Micro Electric Propulsion Systems for Small Spacecraft Applications", Journal of Spacecraft and Rockets, Under Review. Thrust, I sp, propellant Compute Total System Mass (thrusters, propellant, power system) Mass Margin = (M max -M sc )/M max Check Total Mass Feasibility and Compute Margins Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. Pareto Trade-offs Mass vs. Time 7

8 Constant Thrusting in Velocity Direction Simplest and (usually) most time efficient approach to raise altitude Velocity Vector ΔV Direction Thrust Vector Orbit Resulting spiral out trajectory Red shows thrust/ Green shows cruise 8

Results: Altitude Orbital Transfers from LEO Orbital transfers starting from 500

Flight times improve with spacecraft size as more thrusters can be accommodated

Earth s SOI LEO (500 km) Science Instruments 10-15 kg Laser Altimeter 5-10 kg 1-5

Ground-penetrating radar Imaging and Infrared Spectrometers (Mini-M3) Science and

9 Results: Altitude Orbital Transfers from LEO Orbital transfers starting from 500 km circular orbits. Flight times improve with spacecraft size as more thrusters can be accommodated (power, thermal) Transfer Distances and Vs *SOI: Sphere of Influence GEO Moon Earth s SOI LEO (500 km) Science Instruments kg Laser Altimeter 5-10 kg 1-5 kg <1 kg Multi-channel Microwave Radiometer Imaging Spectrometer Ground-penetrating radar Imaging and Infrared Spectrometers (Mini-M3) Science and OpNav Camera Magnetometer Infrared Imager Gamma Ray Spectrometer Microwave Radiometer Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. 9

10 Interplanetary Targets for Future Missions JPL Interplanetary CubeSat Roadmap JPL Interplanetary Roadmap Venus 0.72 AU 193% power (1 AU) Mercury 0.39 AU 670% power (1 AU) Moon 382 K km 100% power (1 AU) Mars 0.52 AU 44% power (1 AU) Plot of showin 10

11 Approach: Interplanetary Transfers and Flybys Phases to achieve flyby: 1. Initalization: Start trajectory in circular GEO. 2. Earth-Escape: Thrust in velocity direction until reach Moon/ escape Earth s SOI*. 3. Orbit Boost: Thrust in velocity direction until aphelion is equal to the distance to the plant from the Sun. Example Mars Transfer Trajectory (1.52 AU) 4. Cruise Phase: No thrusting until performs flyby. Opposite approach to get to inner planets. *SOI: Sphere of Influence (for Earth, radius: 925,000 km) Example Mars Transfer (16 kg, 100 W) Orbit Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. 11

Interplanetary Trajectories for Flybys Trajectory file: Alex Davis (JPL intern)

13 Results for Interplanetary Flybys Transfer Time (years) Flybys from GEO to planets in ~ 1 year for < 20 kg! Moon Mercury Venus Mars Earth SOI to Target SOI Time in Earth Orbit Spacecraft Mass (kg) Mercury 20 kg wet mass 9 months Venus 11 kg wet mass 6.7 months Moon 7.5 kg wet mass <1 months Mars 16 kg wet mass 13 months Moon Mercury Venus Mars Plot of showin Dry Mass Propellant Mass All spacecraft generate 70W at 1 AU 13

Results for Interplanetary Captures From GEO to capture at planets in ~ 1 year for < 20 kg! 1.2 1 0.8 0.6 0.4 0.

14 Results for Interplanetary Captures From GEO to capture at planets in ~ 1 year for < 20 kg! Transfer Time (years) Moon Venus Mars Capture Time Earth SOI to Target SOI Time in Earth Orbit Spacecraft Mass (kg) 25 Venus 15 kg wet mass 8.8 months Moon 9 kg wet mass 1.8 months Mars 20 kg wet mass 11.5 months Moon Venus Mars Plot of showin All spacecraft generate 150W at 1 AU Dry Mass Propellant Mass 14

Summary & Future Work Summary Systems-level framework for evaluating diverse thrusters and trajectories Integrated trajectory

interplanetary transfers and flybys (~1 year, < 20 kg) Future Work Model radiation, thermal, and attitude control in

) Consider higher-fidelity orbit transfer models and lifetimes issues Comparison of EP thrusters, solar sail technologies,

15 Summary & Future Work Summary Systems-level framework for evaluating diverse thrusters and trajectories Integrated trajectory and design decisions, inputs, constraints, objectives Trade-offs of performance metrics for Earth orbit changes Feasible interplanetary transfers and flybys (~1 year, < 20 kg) Future Work Model radiation, thermal, and attitude control in optimization problem Improve realism of operations (thrust strategy, radiation, lifetime, etc.) Consider higher-fidelity orbit transfer models and lifetimes issues Comparison of EP thrusters, solar sail technologies, chemical systems, etc. Extend work to different orbital transfers and destinations Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. 15

16 Questions? 16

17 Multidisciplinary Systems Modeling Approach Propulsion Dynamics CAT Thruster Performance Orbit Dynamics Energy Spacecraft Mass, Power, Volume Sizing Power Thermal Bus Telecom Payload S. Spangelo and B. Longmier, Optimization of CubeSat System-level Design and Propulsion Systems for Earth-Escape Missions", Journal of Spacecraft and Rockets, accepted December Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. 17

Power Collection Thermal Radiation Propulsion Number of thrusters to achieve required thrust

18 Multidisciplinary System Level Constraints and Interactions Apogee Trajectory Design I sp, Thrust ΔV to achieve orbits Thrust time, trip time Perigee Mass, Power, Thermal Battery Capacity Solar Power Collection Thermal Radiation Propulsion Number of thrusters to achieve required thrust Propellant mass Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. 21

19 Modeling Assumptions Systems-level integrated models (trajectories, spacecraft, propulsion) Approach generally applicable to all MEP technologies The thrusters generate thrust perfectly in the desired direction. Attitude control is accomplished by on-board reaction wheels in the case of primary propulsion and by the thrusters when they perform attitude control. The PPU, heater, and neutralizer are sized to accommodate the thruster. There are no solar eclipses or occultations in the trajectories. The solar panels are sized to support continuous thrusting and nominal bus. The spacecraft volume and mass are constrained by conventional CubeSat form-factors for small spacecraft and extrapolated for larger ones. We investigate study 6-12 U (12-24 kg) spacecraft sizes The payload system power is ignored, although this is expected to be significantly less than the thruster power. When multiple thrusters are operated simultaneously they each have the same performance as an individual thruster. Deployed panels When multiple thrusters are used successively, the performance of each thruster is identical. Body fixed panels Maximum thermal power that can be dissipated Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. 22

20 Model- MEP Propulsion System Micro Electro Spray (MEP) technology Liquid metal propellant micro-fabricated with Indium propellant Capillary-force driven propellant management system with no pressurization, valves, or moving parts Small, compact, scalable technology pushing limits of microfabrication techniques Total Propulsion System Power Propulsion System Efficiency Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. 23

21 Modeling: Power System Mass Power system scales with required power to support propulsion system Parameters: P s : Average power consumption of the system P max : Maximum Power Generated by Fixed (fix) or Deployed (dep) panels M sp : Solar Panel Mass for Fixed (fix) or Deployed (dep) panels Clyde Space Double Deployed 2-Sided 30 W Solar Panels ehawk MMA Solar Arrays (130 W/kg) Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. 24

22 Modeling: Spacecraft Components Scaling Scaling for Solar Panels Scaling for Thermal System Deployed panels Body fixed panels Maximum thermal power that can be dissipated Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. 25

23 Where do Emerging EP Systems Fit? Small Electric Propulsion (EP) technologies: Aim to provide high thrust, maximize thrust to power ratio V Provide long duration thrusting Satisfy small spacecraft mass, volume, power, thermal constraints Power is input power to PPU Thrust to Power, mn/kw CAT Plasma Busek s BHT-200 Busek s MEP (HARPS) MIT ieps Mini Helicon Plasma MiXI Ion Busek s Ion (BIT-3) 20 Busek s Ion MIT MEMs (BIT-1) Ion MEP JPL s MEP Specific Impulse (Isp), sec CAT Plasma: low I sp maximizes Thrust to Power; MEP mid-range I sp Most propulsion systems actually span a range but point design plotted Reference: Review of MEP Technology, Marrese-Reading, John Zimer, et a.., MEP A-Team Study, Sept. 17, 2014 Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. 26

24 Approach- Goals, Decisions, Constraints Objective: Maximize payload mass fraction M p M i M p : Payload mass of all components except propulsion system M i : Mass of initial wet spacecraft Why? To maximize capability to carry science instruments/spacecraft components Decision Variables: Number of MEP thrusters (or size) Specific Impulse (I sp ) which is directly related to thrust level Driving system interface parameter that defines trades between mission design (trip time), power system, and propulsion system (mass, power) Substantial driver in the propulsion system design and defines required operating voltages and total propellant throughput Constraints: Due to voltage limitations, I sp < 7000 sec; lower voltages preferred to reduce complexity of high-voltage systems Short lifetimes preferred; missions with lifetimes less than 3 years Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. 27

CAT: Optimal Solutions for Different Objectives Missions start in 500 km circular orbit until they escape Earth s SOI (925 K km) Case 1: Constant Thrust in Velocity Direction Case 2: Optimized

Time Case 1-108 days Case 2-175 days Minimize Propellant Minimize Propellant & Battery Mass/ Volume Minimize Radiation (~6mm Al assumed) Case 2-1.34 kg Case 1-2.5 kg/ 0.5 U Case 2-1.

25 CAT: Optimal Solutions for Different Objectives Missions start in 500 km circular orbit until they escape Earth s SOI (925 K km) Case 1: Constant Thrust in Velocity Direction Case 2: Optimized Variable Thrust/Time at Perigee Comparison of solutions for various goals Optimization Goal Initial Orbit Sun Synchronous (P av : 25 W) Initial Orbit Not Sun Synchronous (P av : 11 W) Minimize Time Case days Case days Minimize Propellant Minimize Propellant & Battery Mass/ Volume Minimize Radiation (~6mm Al assumed) Case kg Case kg/ 0.5 U Case krad (Case 1: krad) S. Spangelo and B. Longmier, Optimization of CubeSat System-level Design and Propulsion Systems for Earth-Escape Missions", Journal of Spacecraft and Rockets, Accepted

Active Interplanetary CubeSat Projects INSPIRE (JPL) 1

Moon NEA Scout (MSFC/JPL) 2,3 Asteroid characterization

relay Lunar Flashlight (JPL/MSFC) 2,3 Lunar orbiter to

2,3 Biosensor to study impact of radiation on living

JPL, 3 NASA's Advanced Exploration Systems (AES) Copyright

26 Active Interplanetary CubeSat Projects INSPIRE (JPL) 1 Navigation demonstration with the IRIS radio beyond the Moon NEA Scout (MSFC/JPL) 2,3 Asteroid characterization mission [EM-1] MarCO (JPL) 2 InSight insertion real-time relay Lunar Flashlight (JPL/MSFC) 2,3 Lunar orbiter to search for ice in lunar craters [EM-1] BioSentinel (Ames) 2,3 Biosensor to study impact of radiation on living organisms [EM-1] 1 JPL/NASA Planetary Science Division, 2 JPL, 3 NASA's Advanced Exploration Systems (AES) Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. 29

Future Impact of Science-Driven Small Spacecraft Performing significant ΔV and high-precision attitude control enables: Escaping Earth-orbit, transferring to Moon, Mars, asteroids, coments, beyond

27 Future Impact of Science-Driven Small Spacecraft Performing significant ΔV and high-precision attitude control enables: Escaping Earth-orbit, transferring to Moon, Mars, asteroids, coments, beyond Creating and maintaining formation flight/constellations (e.g. large apertures) Autonomous Operations enabling: Autonomous navigation by imaging asteroids (e.g. DS1) Agile Science for on-board autonomy to locate Earth, detect objects (e.g. plumes) Dynamic observation planning, disruption-tolerant networking (DTN) Science Mission Applications to perform SMEX/Discovery-class science: Multi-spacecraft architectures: constellations, mother-daughtership, swarms, formation flying to perform distributed temporal/spatial measurements Pre-cursor missions to explore dirty/dangerous/unknown environments (comets, asteroids, moons, Earth-Sun Lagrange points) Agile Science Reference: D. R. Thompson, S. A. Chien, J. C. Castillo-Rogez Copyright 2015 California Institute of Technology. Government sponsorship acknowledged. 30

Perigee Apogee Velocity Vector Orbit Thrust Vector Thrust Location Perigee This

28 Case 2: Problem Description The scheme where we thrust only at perigee exploits the fact that increasing the ΔV at perigee (gravity well) results in greater apogee raises. Perigee Apogee Velocity Vector Orbit Thrust Vector Thrust Location Perigee This approach may be more (time and fuel) efficient relative to the constant thrust approach (Case 1). 31

29 Case 1: What s the Impact of Attitude Control Errors? Results given for orbit starting in 500 W circular orbit until Earth-escape (925,000 km) Even with γ=20 o, only requires an additional 13.1 days (10W)/ 5.2 days (25 W) Orbit shape and precession will also change with cross-track ΔV component Angular Error (γ) Actual/ Ideal Thrust Ratio Increase in Time Constant Thrust (10 W) 1 o % 0.02% 5 o % 0.4% 10 o % 1.5% 20 o % 6.4% Increase in Time Constant Thrust (25 W) Velocity Vector γ Actual Thrust Vector Ideal Thrust Vector 32

BravoSat: Optimizing the Delta-V Capability of a CubeSat Mission. with Novel Plasma Propulsion Technology ISSC 2013

BravoSat: Optimizing the Delta-V Capability of a CubeSat Mission with Novel Plasma Propulsion Technology Sara Spangelo, NASA JPL, Caltech Benjamin Longmier, University of Michigan Interplanetary Small