Fusion-Enabled Pluto Orbiter and Lander Presented by: Stephanie Thomas DIRECT FUSION DRIVE
Team Members Stephanie Thomas Michael Paluszek Princeton Satellite Systems 6 Market St. Suite 926 Plainsboro, NJ 08536 http://www.psatellite.com Dr. Samuel Cohen Princeton Plasma Physics Lab Plainsboro, NJ This material is based upon work supported by NASA under award No. NNX16AK28G 2
Why Fusion Propulsion? 1. Get there FASTER Faster can mean CHEAPER if it s also the same size or smaller 2. Do more SCIENCE Payload capabilities and data rates depend on POWER 3
DFD: Propulsion and Power in One Device The fusing plasma acts as the heating source for propellant flowing outside the confinement region. This process of thrust augmentation gives us substantial thrust with high exhaust velocity! 1-2 m 4
Pluto Explorer Mission with DFD Single launch from Earth, fly directly to Pluto with constant thrust Put a 1000 kg spacecraft in orbit around Pluto, beam 30 kw to a lander using optical transmission, return high-definition video and get there in only 4 years! 5
Pluto Explorer Vehicle Design DFD Engines 6 D, 3 He Tanks under Sun Shield Radiator Lander Solar Array Optical Comm Engines Fuel (D 2 ) Helium-3 Liquid D 2 tank Gas 3 He tank Fuel Tanks 2 @ 1350 kg 6000 kg 0.5 kg 2.1 m radius 0.95 m radius 300 kg Radiator Area 50 m 2 Radiators Lander Structure Orbiter payload 134 kg 230 kg 200 kg 436 kg Launch Mass 10000
STOP RIGHT THERE. Why is a fusion drive suddenly achievable?
DFD is DIFFERENT from other fusion reactor concepts 1. Unique heating method 2. Simple configuration 3. Small size 4. Clean operation low radiation à FRC has 10x better confinement than tokamak These features make it uniquely suited for use in space!
DFD Technology is Simple, Small, Clean Based on a PPPL experiment Simple array of magnetic coils Easily direct ion flow for thrust Small size promotes Clean operation (very few neutrons) D- 3 He fuel Tritium exits before it can fuse 1-10 MW per minivan-sized engine Enabled by new RF magnetic plasma heating method odd-parity heating Closed field lines FRC region Steady-state operation Plasma pulse in PPPL experiment 9 9/26/17
NIAC Study PHASE I RESULTS Engine design Mass and energy breakdowns Trajectory analysis PHASE II PLANS
What did we learn from Phase I? Area Progress Result Thrust Model Trajectories Magnet Model RF Subsystem 11 Use PPPL s UEDGE * fluid code results to generate thrust and Isp model as a function of power and flow rate Develop new trajectories using constant thrust, Isp, and efficiency results from UEDGE model Modeled the magnet mass and discovered tradeoff between HTS and LTS, materials and operating temperature, component mass Idea for switching amplifiers (class E) to lower RF subsystem mass and approach 100% efficiency Feasible envelope for operation, ~5 N/MW Pluto mission is feasible! Departing from LEO takes ~60 days and expands launch options New NASA STTR New NASA STTR * T. D. Roglien, LLNL, Users Manual for the UEDGE Edge-Plasma Transport Code UCRL-ID-137121
Analyzing Key DFD Subsystems Assess TRL and feasibility of all supporting subsystems, and increase fidelity of specific power estimates 12
Parametric Engine Model in MATLAB Assumptions Gas box power RMF power and efficiency Synchrotron reflection Superconductor properties Fusion Parameters Temperatures and densities Length & radius Beta Fuel ratio Fusion crosssections Subsystems Shielding Thermal conversion system Radiator mass Coil cooling Fusion Power Model Neutron load Synchrotron Bremsstrahlung Thrust power Thermal power 13
1 MW Engine Energy Balance Total power Gas Box 0.11 MW Bremsstrahlung 0.23 MW Fusion 1.13 MW Synchrotron 0.27 MW Power to sustain heating 0.06 MW RMF0 0.12 MW 0.07 MW Neutron Thrust 3.48 kw 0.52 MW Thrust Power 0.24 MW Heat Engine 0.24 MW 0.61 MW Electric specific power = 0.16 kw/kg Power available to spacecraft Radiator Thrust specific power = 0.34 kw/kg Fusion specific power = 0.74 kw/kg 14
Engine Mass Analysis 1 MW Engine 1350 kg Shielding thickness 5 cm Magnets 558 kg Shielding 339 kg Power Generation 178 kg Structure 122 kg RMF system 108 kg Radiators 27 kg Coil cooling 18 kg Specific Power 0.74 kw/kg 15
UEDGE Plasma Transport Software Power Input to e - Gas Flow (D 2 ) Calculate thrust 16
Numerical Simulation of Energy to SOL New fluid code results confirm 5-10 N/MW thrust input power (power carried in SOL) 70 UEDGE Thrust with Thrust Fits 10 1 MW Thrust Performance 9400 60 9.5 9200 50 Thrust (N) 40 30 Thrust (N) 9 8.5 9000 8800 Isp (s) 20 10 8 8600 0 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 Flow Rate (g/s) Princeton Satellite Systems 7.5 8400 0.08 0.085 0.09 0.095 0.1 0.105 0.11 0.115 Princeton Satellite Systems Mass Flow (g/s) 17
Thrust and Velocity Model Input Power (MW) 6 5.5 5 4.5 4 3.5 3 2.5 2 1.5 1 Results indicate a feasible range of flow rates for the SOL to absorb the energy from the fusion products. UEDGE Thrust Model 60 50 40 30 20 10 Thrust (N) Input Power (MW) 6 5.5 5 4.5 4 3.5 3 2.5 2 1.5 1 0.5 UEDGE Velocity Model 0.1 0.2 0.3 0.4 0.5 0.6 Mass Flow (g/s) This gives us a numerical model to trade thrust and specific impulse. 120 115 110 105 100 95 90 85 80 Exhaust Velocity (km/s) 0.5 0.1 0.2 0.3 0.4 0.5 0.6 Mass Flow (g/s) 0 18
Earth Departure from LEO 30 days with 20 N 60 days with 10 N Short amount of time in radiation belts 7 km/s DV is small portion of mission total Expands launch options 19
Constant-Thrust Trajectories Straight-line, planar and 3D trajectories all confirm feasibility of Pluto mission concept Distance (AU) 30 20 10 100 Payload: 1000 kg Power: 0.42 MW Thrust: 4.19 N Mass: 6998 kg Trajectory 0 0 0.5 1 1.5 2 2.5 3 3.5 4 0-5 Planar Orbit v (km/s) 50 0 0 0.5 1 1.5 2 2.5 3 3.5 4 y -10-15 m f (kg) 6000 4000 2000 0 0 0.5 1 1.5 2 2.5 3 3.5 4 Time (years) Princeton Satellite Systems -20-25 -30-25 -20-15 -10-5 0 x Princeton Satellite Systems 20
Powered Injection into Pluto Orbit Feasible Uses twice the velocity change of an impulsive burn at closest approach 2-3 days 2.5 2 1.5 1 0.5 0-0.5-1 -1.5 10 4 Thrust = 10 N Mass = 2500 kg DV = 977 m/s Fixed Mass, Minimum Fuel Optimization 1.1 (day) 2.8 (day) 2.3 (day) 1.7 (day) -2-2.5 0.6 (day) 0.0 (day) -6-5 -4-3 -2-1 0 10 4 21
Phase II Plans Thrust augmentation experiment 22 Explore dynamics of adding plasma to the scrape-off-layer Synchrotron loss modeling Existing models are based on tokamaks Theoretical effort to create model specific to FRC Additional subsystem models Superconducting coils Subcontract with MIT s Dr. Joseph Minervini MIT has a new HTS magnet in-house RF heating system Closed-loop no-thrust power generation equipment
I want one now! Caveat: We have NOT yet demonstrated fusion PFRC-3 requires $50M and superconducting magnets Demonstrate fusion in 5-7 years Engineering of the space subsystems can occur in parallel with the fusion science experiments Once fusion near breakeven has been demonstrated, commercial development can take place Best-case timeline: 15 years to a space reactor 23
Thank You Stephanie Thomas sjthomas@psatellite.com Princeton Satellite Systems 6 Market St. Suite 926 Plainsboro, NJ 08536 http://www.psatellite.com (609) 275-9606 Dr. Samuel Cohen scohen@pppl.gov Princeton Plasma Physics Lab Plainsboro, NJ 24
Backup ADDDITIONAL TECHNICAL INFORMATION
PFRC Status and Program Plan Confirmed: Field-reversal, high-beta, longterm stability (300 ms pulses), electron heating (>500 ev with 20 kw input power). In progress: Ion heating Next up: Particle exhaust, high power and field operation, demonstrate fusion Machine Objectives PFRC-1 PFRC-2 PFRC-3A PFRC-3B Electron Heating Ion Heating Heating above 5 kev D-He3 Fusion Fuel H H H D-3He Goals/ Achievements* 3 ms pulse* 0.15 kg field* e-temp = 0.3 kev* 0.1 s pulse* 1.2 kg field i-temp = 1 kev 10 s pulse 10 kg field i-temp = 5 kev 10 s pulse 80 kg field i-temp = 50 kev Plasma Radius 4 cm 8 cm 16 cm 16 cm Time Frame 2008-2011 2011-2015 2015-2019 2019-2020 Total Cost $2M $6M $20M $20M 26 9/25/17
Thrust Augmentation H or D is used as a propellant- it flows along the magnetic field lines outside of the separatrix; scrape-off layer (SOL) e- are heated by the fusion products that are ejected into the SOL; e- energy transferred to ions in plume expansion This reduces the exhaust velocity of the fusion products from 25,000 km/s to ~50 km/s and increases thrust to >20 N Thrust/I sp is adjustable based on rate that gas is injected into the gas box The exhaust plume is directed by a magnetic nozzle, consisting of a throat coil and nozzle coils to accelerate the flow. 27
Rotating Magnetic Fields (RMF) Parity refers to the symmetry of the magnetic field mirrored across the z=0 midplane Frequency is a fraction of the ion cyclotron frequency for the helium-3 Would be 0.3 to 2 MHz Provides all the startup power and a fraction of the heating power during operation RF antennas shown to the right PFRC-2 Design RMF 0 Antennas 28 9/25/17
Specific Power Plot on right based on work done assuming direct insertion Other electric propulsion systems Solar not applicable to outer planet missions Fission Electric 300 250 DFD Isp: 10198 s Theoretical Maximum Delta-V Structural Fraction (F): 0.02 V max = u e log( (1+F)/F ) Mass (kg) 10 5 10 4 10 3 Mass for 4 Year Mission/ 65 km/s Delta-V/ 2 MW Power Plant SLS Block 4 SLS Block 1 Falcon 9 Heavy Delta IV Heavy Total Fuel Engine Lambert V (km/s) 200 150 100 Nuclear Electric VASIMR Isp: 4997 s Nuclear Electric Hall Isp: 2040 s Fission DFD 10 2 0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2 Specific Power (kw/kg) Princeton Satellite Systems 50 Nuclear Thermal Isp: 900 s Chemical Isp: 462 s 29 0 0 5 10 15 20 25 30 Pluto Transfer Time (yrs)
Fusion Cross-Sections Mean Sigma V (m 3 /sec) 10 20 10 22 10 24 10 26 10 28 10 30 10 32 10 34 10 36 10 38 10 40 D-T D-D D- 3 He p- 11 B Reaction Rates x x x Reactor operating temperatures D D n D D p D T D He3 p B11 10 0 10 1 10 2 10 3 Temperature (KeV) 30
Neutron Reduction Mechanisms 1. Fuel D- 3 He 2. Size small radius FRC 3. Fast removal of tritium in scrape-off layer 4. Low D fuel ratio: D/ 3 He of 1:3 5. Non-equilibrium: D heating only ½ as much as 3 He by w RMF Beam-like distributions 31
DFD is REALLY small Typical tokamak reactor: 1000-4000 MW 60 m tall Development in 30-50 years PFRC reactor: 1-10 MW 2 m diameter Development in 5-10 years Small enough to fit on a single launch vehicle ITER research reactor Person for scale 32
Computational Tools LSP: particle in cell RMF: heating model 33 UEDGE: fluid model
Heat Recycling Brayton cycle heat engine Heat high pressure gas to do work Used in High Temperature Gas Cooled Fission Reactors Helium working fluid Paired compressor/turbine sets with counter-rotating turbines and compressors Common shaft for compressor, turbine and generator Need for multiple compressor and turbine stages and recuperator to be determined Large radiator wings 1 LP Compressor Intercooler Radiator HP Compressor Recuperator Cooling Radiator 2 Heat Source 3 4 HP Turbine Heat Source LP Turbine HTGCR 34