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

Transcription

1 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 Satellite Conference, June 2013

2 How Far Can CubeSats Go (Alone)? Can CubeSats go beyond Low Earth Orbit (LEO)? Is there a fundamental size, mass, power, cost limitation? No! BravoSat: CubeSat with CAT engine designed to escape Earth orbit Enabling factors: 1. Miniaturized thruster technology with high V capabilities 2. Heritage and experience operating CubeSats in LEO 3. Intelligent use of mass/ volume/ energy! Photo Credit: NASA Website 2

3 Is the Sky the Limit? Dream versus Reality Dream (Goal): Can we escape Earth orbit? How far can we go? Maximize altitude increase on a CubeSat.. [Not equal to maximizing V] Reality (Constraints) Maximum engine power (thrust) Initial orbital configuration (LEO) CubeSat or small satellite: Physical (mass/ volume) Limited potential to collect/ store energy CubeSat Missions From Small Satellite Survey Lifetime- degradation of batteries, solar panels, radiation, etc. Image Credit: CubeSat Team Websites, Survey: 3

4 How Much V Is Needed? Transferring from circular orbit to circular orbit Maximum LEO GEO Moon Orbit Earth SOI Closest Mars Approach 8 7 Velocity, km/sec Distance from Earth (Altitude), km Starting in LEO, escaping Earth s Sphere of Influence (SOI): V~ 7km/sec Image Credit: Google Images 4

5 CAT: Large V Engine Capability CAT: CubeSat Ambipolar Thruster Uses high-density plasma source Achieves high V and high thrust/power Fits within small spacecraft form-factor Design of a 5 kg 3U CubeSat with CAT engine performing initial testing in Low Earth Orbit. Photo Credit: PEPL 5

6 CAT: Large V Engine Specifications Used to select best trade-off between thrust, η thruster, and I sp Linear relationship: Mass flow power Fixed peak electron temperature T e.max 30 ev 6

7 Is this Really Feasible for a CubeSat? Let s review the major constraints: 1. Mass/ Volume is <.5U, <.5 kg Fuel ~2.5 kg required (~.5 U) Total: ~3kg, 1U, Remaining: 2kg, 2U 1 2. Instantaneous Power CAT engine with CubeSat subsystems Thruster can operate up to 100W, supportable by custom CubeSat bus 3. Energy and Time Trade between power * time (thruster is on) and available energy Constraint on mission duration! Summary: None of these constraints are show-stoppers for a CubeSat! 1 Assuming 5kg 3U CubeSat Image Credit: PEPL 7

8 How Much Mass Is Needed? Work through Rocket Equation for I 2 fuel and CAT Parameter Symbol Input/ Equation Value Units Specific Impulse Input 1010 sec Exhaust Velocity 9908 km/sec Dry Spacecraft Mass Input 2.5 kg Propellant Mass Input 2.5 kg Initial Mass 5.0 kg Final Mass 2.5 kg Delta V Capability 7.0 km/sec Ideal Rocket Equation We can escape Earth s Sphere of Influence ( V~ 7km/sec) with ~2.5 kg of fuel! g: gravity constant = 9.81 m/sec 2 8

9 How Much Mass Is Needed? Maximum LEO Let s review different fuels and destinations GEO Moon Orbit Earth SOI Closest Mars Approach 8 7 Destination Maximum Mean Earth's Closest Mars 6 Parameter Units LEO GEO Moon SOI Approach Distance from 5 Earth km 2,000 35, , ,000 56,000,000 Velocity, km/sec 4 V km/ sec I 2 Fuel kg Ga Fuel kg H 2 0 Fuel kg Distance from Earth (Altitude), km We may be able to escape with even <2.5 kg fuel! LEO: Low Earth Orbit, GEO: Geostationary Earth Orbit, SOI: Sphere of Influence 9

10 How Much Time Is Needed? Let s review different fuels and destinations GEO Constant thrust at 10 Watts Destination Maximum Mean Earth's Closest Mars 6 Parameter Units LEO GEO Moon SOI Approach Distance from 5 Earth km 2,000 35, , ,000 56,000,000 Velocity, km/sec V km/ sec I 2 Fuel days Maximum LEO Moon Orbit Earth SOI Closest Mars Approach Ga Fuel days H 2 0 Fuel days ,369 1,933 2,026 2, Distance from Earth (Altitude), km We can escape Earth s Sphere of Influence in < 1 year! LEO: Low Earth Orbit, GEO: Geostationary Earth Orbit, SOI: Sphere of Influence 10

11 Three for Trajectory Design 1-3: Consider additional realistic constraints and exploit environmental factors Case # Description Consider Energy? Consider Magnetics? Exploit Perigee? 1 Constant thrust along velocity vector 2 Phase 1: Thrust when α<30 Phase 2: Pulse thrust every 200 minutes 100W: No 10/20W: Yes No No Yes Somewhat No 3 Thrust at perigee along velocity vector 2 Yes Yes Yes 1 α: Angle between the velocity vector and thrust vector (aligned with Earth s magnetic field for passively stabilized craft) 2 Assume perigee is roughly when α is small 11

12 Case 1: Constant Thrust Along Velocity Vector Work through Rocket Equation for I 2 fuel and CAT Thrust Vector Exhaust Vector Velocity Vector BravoSat Orbit *Not to scale 12

13 Case 1: Constant Thrust Along Velocity Vector 2.5 kg fuel, 2.5 kg dry mass Initial 700 km circular orbit 13

14 Case 1: Constant Thrust Along Velocity Vector How much mass do we need? Depends how far you want to go!!! Results are shown when all fuel is used Escape Earth s Sphere of Influence (SOI) requires ~2.5 kg Starting in 700km circular polar orbit with 2.5 kg dry mass 14

15 Exploiting CAT s Magnetic Dipole Body Axis Definition Thrust Vector Body Z axis 15

16 Exploiting CAT s Magnetic Dipole Passive Magnetic Stabilization Scheme BravoSat Orbit Magnetic Field lines 16

17 Exploiting CAT s Magnetic Dipole Desirable Configuration for Thrusting to Exploiting Passive Magnetic Stabilization Thrust Vector Body Z axis Velocity Vector BravoSat Orbit Body Z axis α Velocity Vector 17

18 Exploiting CAT s Magnetic Dipole Short, high-powered (100W) burns Example: Thrust only when α 1 < 40 o Body Z axis α Velocity Vector 18

19 Exploiting CAT s Magnetic Dipole Short, high-powered (100W) burns Challenge: Once orbit grows, must be smarter about thrust duration (energy, thermal) 19

20 Case 2: Multi-Staged Approach Strategy: 1. Phase 1: Thrust for 10 minutes when velocity and thrust vectors aligned 2. Phase 2: Thrust for 10 minutes every 200 minutes (pulsing) Rationale: Exploit passive magnetic stabilization in Phase 1 Thrust at perigee, where thrust vector ~parallel to velocity vector Energetically possible in both phases (thrust <10 minutes at a time) 20

21 Case 2: Multi-Staged Approach Phase 1: Thrust <10 mins when aligned Phase 2: Thrust every 200 mins for 10 mins 21

22 Case 2: Multi-Staged Approach Results sampled once a day Initial Fuel Mass Time to run out of fuel Radius when out of fuel 2 kg 320 days km 2.5 kg 389 days km Starting in 700km circular polar orbit with 2.5 kg dry mass 22

23 Case 3: Exploiting Perigee to Escape Earth! Strategy: 1. Start in elliptical orbit (assume 500km x 1500km) 2. Thrust at perigee (100W, 10mins), when α small 3. Continue until escape Earth orbit, or run out of fuel/time Rationale: Exploit passive magnetic stabilization Energetically feasible (thrust once per orbit for 10 mins) 23

24 Case 3: Exploiting Perigee to Escape Earth! Escaping with 1 kg fuel! Possible if apogee is between L1 and L5 or between L3 and L4 24

25 Case 3: Exploiting Perigee to Escape Earth! Interesting variations due to Earth-Moon interactions until escapes! Starting in 500km x 1500km orbit with 2.5 kg dry mass Note results are sampled once a day 25

26 Case 3: Exploiting Perigee to Escape Earth! Interesting interactions with the Moon! Note results are sampled once a day Magnetic field still high at perigee So can use passive stabilization Starting in 500km x 1500km orbit with 2.5 kg dry mass 26

27 Consider 3 Different Approaches: Summary Best-case results assuming firing with 2.5 kg dry mass # Description Power Fuel to Escape Time to Escape Challenges 1 Constant thrust along velocity vector 2 Phase 1: Thrust when α 1 <30 Phase 2: Pulse thrust every 200 minutes 3 Thrust at perigee along velocity vector 2 10 W Constantly 2.5 kg 269 days Require attitude control 20 W Constantly 135 days 100 W/ >1,000 times 100 W/ ~700 times 2 kg 320 days Require attitude control High number of battery cycles ~1 kg < 3 years Require initial elliptical orbit Require (fine) active attitude control Disclaimer: These are not optimized, but rather representative results! 1 α: Angle between the velocity vector and thrust vector (aligned with Earth s magnetic field for passively stabilized craft) 2 Assume perigee is roughly when α is small 27

28 / Future Work Demonstrated feasibility of CubeSat (CAT engine) to escape Earth orbit! Explored 3 firing strategies to: Minimize fuel mass, firing time, time to escape Earth orbit Consider realistic constraints/ environmental factors Future Design Considerations Battery degradation, radiation (South Atlantic Anomaly, Van Allen Belts) 1 Future Work Refining strategies, optimizing to minimize fuel mass/ power/ time Identifying design factors for CAT development (e.g. thermal, power, magnetic) 1 May be able to overcome these challenges with adequate radiation shielding (e.g. <1MeV Al shielding) 28

29 Acknowledgements JP Sheehan, University of Michigan PEPL Members, University of Michigan Derek Dalle, University of Michigan James Smith, Jet Propulsion Lab 29

30 Questions? BravoSat Orbit Magnetic Field lines expanded on side opposite to the Sun Magnetic Field lines 30