Technology Forum on Small Body Scientific Exploration – 4th Meeting of the NASA Small Bodies Assessment Group
Michael Patterson NASA Glenn Research Center
John Brophy Jet Propulsion Laboratory California Institute of Technology
January 24, 2011
National Aeronautics and Space Administration www.nasa.gov 1. In-Space Propulsion Overview 2. Current In-Space Investments for Robotic Missions 3. Description of Electric Propulsion 4. Current Technology Investments and Development Status
2 3 Flight System Development Future Applications
Non-toxic propellant-based propulsion systems (LOX, LH2, LCH4, Ethanol) Main Propulsion Systems (MPS)
Propulsion systems for orbit transfer, orbit injection, spacecraft maneuvering, landing, and ascent
Development of component technologies (igniters, exciters, injectors , combustion chambers, nozzles) for non-toxic propellants Current Focus Areas
• LOX/LCH4 propulsion technologies for main propulsion systems (MPS), reaction control systems (RCS), and propellant storage and distribution • Development of Orion Service Module propulsion system Reaction Control Systems (RCS) • Preliminary trade studies, requirements development, and planning for Altair ascent propulsion subsystem
Research and Technology Development
Materials Evaluation Combustion Diagnostics Ignition Performance Propellant Properties 4 Future Applications • Next generation ion propulsion system for deep-space science missions (NEXT) • Propulsion systems for Earth orbital applications including satellite servicing, repositioning, re-boost, orbital debris removal, and de-orbit • High power propulsion systems for cargo transportation in support of future human crewed missions beyond LEO Technology and Flight System Development of Ion and Hall Thrusters Current Focus Areas • Ion and Hall thruster design, fabrication, evaluation 200 kWe Class SEP Stage • VASIMR or other high power EP • Innovative concept development and evaluation technology • Reusable orbit transfer • Power electronics design, bread-boarding, evaluation • Robotic interplanetary cargo transfer • System integration, system qualification and acceptance testing • Thermal, stress, vibration analysis and test • Capability to fabricate laboratory and flight hardware
SEP Stage Upgrade • 90 kWe class system Solar Electric Propulsion (SEP) Stage • Orbital servicing and debris removal • 30 kWe class system with advanced PV and AR&D technologies • ISS reboost • NASA science missions • Science missions, e.g., Mars Sample/Return • Small commercial servicing applications in Earth orbit • Use of NEXT ion or Hall thruster technology 5 Future Applications • High-thrust/high-Isp injection stage to Mars and other destinations within inner solar system • Cargo and crewed transportation for human missions to the Moon, NEOs and Mars • Strong synergy with existing chemical rocket and stage hardware • Evolvability to propulsion and power (“Bimodal operation”), variable thrust and Isp (“LOX-afterburner” nozzle), ISRU and hybrid operation (BNTR with EP) Integrated NTP Crewed Vehicle Design for Mars DRA 5.0 Current Focus Areas • Vehicle / engine concept design and analysis, and requirements definition for Human Mars missions (DRA 5.0) • Engine conceptual design, analysis and modeling includes neutronics, thermal, fluid, stress and mass estimation • Fuels assessment and development planning with DOE • Innovative NTP engine design concept development • Infrastructure and test facility needs assessment
LOX Augmented Nuclear Rocket (LANTR) Engine Demonstration
Advanced Technology: Bimodal Propulsion and Power 6 In-Space Propulsion Investments for Robotic Missions Objective: “Develop in-space propulsion technologies that enable or benefit near to mid-term NASA science missions by significantly reducing travel times required for transit to distant bodies, increasing scientific payload capability or reducing mission costs.” ISP will enable access to more challenging and interesting science destinations, including enabling sample return missions.
Planetary Earth Entry Vehicles Systems & Sample Return Propulsion Ascent Vehicles /Aerocapture Mission Studies Propulsion Tech High-Temp Engine TPS & Structures (AMBR) Mars Ascent Vehicle Tools Ultra Lt Wt Tank
NEXT GN&C
Mission Studies
System Studies Hall Thruster Multi-Mission Earth Entry Vehicle
Component Tech
• The ISPT project addresses the primary propulsion technology needs for the agency’s future robotic science missions • The current ISPT focus is on TRL 3-6+ product development 7 8 Chemical propulsion converts the energy stored in the molecular bonds of a propellant into kinetic energy • Typically high thrust to weight (required for launch) • Exhaust velocity is limited by the chemical energy available • Higher exhaust velocities can reduce required propellant mass:
M f exp-ΔV v M0 e
• For a given change in velocity (ΔV), the
delivered mass (Mf) depends on the propellant exhaust velocity (ve) • Once in space, engines that provide a higher exhaust velocity can significantly reduce propellant mass requirements 9 Electric propulsion (EP) uses electrical power to provide kinetic energy to a propellant • Decouples kinetic energy from limitations of chemical energy • Provides higher exhaust velocities than chemical engines - Reduces propellant mass needed for a given impulse - Allows reduction in launch mass a/o increase in payload; can provide substantial benefits in mission cost • Increases launch window as compared to all-chemical systems in certain mission scenarios • Electric propulsion primarily benefits large total impulse missions - Orbit raising, repositioning, long-term station keeping - Cis-lunar, planetary and deep space missions - Precise impulse bits for formation flying
10 Additional considerations… • Significantly lower thrust to weight than chemical engines - Small but steady acceleration vs. short-burn chemical engines - EP engines must be designed for long life (thousands of hours)
• Increased dry mass due to: - Solar arrays - Power processing units - and Other EP-specific hardware • Spacecraft integration considerations: - Electric power requirements - Plasma plume and EMI
• Propulsion system trades performed to evaluate whether a given mission will benefit from the use of electric propulsion 11 Electric thrusters are categorized by their primary acceleration mechanism: • Electrothermal – Resistojets & Arcjets (commercial flight units available) heat gas and expand gas thru a nozzle – Auxiliary Propulsion; typically ≤ 1 kWe
Resistojet thrusters use resistive heating elements to increase the thermal energy of a gas propellant
Arcjet thrusters use an electric arc to increase the thermal energy of a gas propellant
• Electrostatic – Hall effect and Gridded Ion thrusters (commercial flight units + development) generate high voltages for ion (plasma) acceleration – Auxiliary and Primary Propulsion; typically 1 to 10’s of kWe
Ion thrusters use closely spaced high voltage grids to create an electrostatic field Hall thrusters use magnetically trapped electrons to create an electrostatic field 12 Electric thrusters are categorized by their primary acceleration mechanism:
• Electromagnetic – Pulsed plasma (commercial flight units available), Magnetoplasmadynamic and Pulsed Inductive thrusters (laboratory model) apply a Lorentz (JxB) force for plasma acceleration – Auxiliary and Primary Propulsion
Pulsed Plasma thrusters use a pulsed, repetitive current to ablate solid propellant, induce magnetic field (JxB)
Magnetoplasmadynamic thrusters use a high power, steady- state current to ionize gas propellant, induce magnetic field (JxB)
13 14 Objective: Develop advanced ion propulsion system Thruster Attribute with improved performance and life characteristics Thruster power range, kW 0.5 - 6.9 to reduce user costs and enhance/enable a broad Max. Specific Impulse, s 4,190 range of NASA SMD missions Thrust range, mN 26 - 236 Propellant Throughput, kg 450* Mass (with harness), kg 13.5 Envelope dimensions, cm 43.5 x 58.0 Power Processing Unit Attribute Power Processing Unit mass, kg 33.9 Envelope dimensions, cm 42 x 53 x 14 Input voltage range, V 80 - 160 Feed System Attribute High Pressure Assembly mass, kg 1.9 NEXT gridded ion NEXT PM ion thruster operation thruster Low Pressure Assembly mass, kg 3.1 NEXT Thruster NEXT addresses the entire ion propulsion String system PPU Thruster - Gridded ion thruster DCIU - Power processing unit (PPU) - Propellant management system (PMS) HPA - System integration (including gimbal and LPA control functions) Gimbal Primary Partners * Rated Capability Goal 300Kg Design/Qualification Goal (1.5x Rated) 450Kg - NASA Glenn Research Center: Lead Projected 1st Failure >750Kg Potential Rated Capability 500Kg - JPL, Aerojet Corp., L3 Comm. 15 Critical tests have been PM1 PM1R PPU Feed Gimbal System completed, or are imminent, on high fidelity hardware
Functional & Complete Complete Complete Complete Complete Performance Testing
Qual-Level Vibration Test Complete Complete FY11 Complete Complete
Qual-Level Thermal/ Not Complete Complete FY11 Complete Vacuum Test planned
FY11 Key activities : • PPU re-verification test, and environmental test • Single-String System Integration Test: Complete • Long-Duration Test (LDT) • Multi-String System Integration Test: Complete • Phase II Review • Thruster Life Test: Completed goal of 450Kg throughput •Technology Maturation Review • >34,000 hours and >570 kg of xenon processed to date FY11-13 Key activities: • Life Test will continue through 750Kg or first failure • PPU design iteration and/or DCIU maturation • LDT extension to demo 750 kg • Residual, high-priority risk reduction activities 16 NSTAR Improve- CHARACTERISTIC NEXT NEXT BENEFIT (SOA) ment
Max. Thruster Power (kW) 2.3 6.9 3x Enables high power missions with fewer Max. Thrust (mN) 91 236 2.6x thruster strings Throttling Range Allows use over broader range of distances 4.9 13.8 3x (Max./Min. Thrust) from Sun
Reduces propellant mass, enabling more Max. Specific Impulse (sec) 3120 4190 32% payload and/or lighter spacecraft
6 Total Impulse (10 N-sec) 4.6 >18 >3.9x Enables low power, high V Discovery-class Propellant Throughput (kg) 150 450 3x missions with a single thruster
Mission Performance Finding Discovery - Small Body Missions Higher net payload mass with fewer thrusters than NSTAR system New Frontiers - CSSR: Higher net payload mass than NSTAR, with, simpler EP System: 2+1 NEXT vs 4+1 NSTAR thrusters • Comet Surface Sample Return Titan: > 700 kg entry package with 1+1 NEXT system • Titan Direct Lander Flagship - Saturn System Missions > 2400 kg to Saturn Orbit Insertion with 1+1 NEXT system, EGA + Atlas V EELV - •Titan Doubles delivered mass of chemical/JGA approach • Enceladus > 4000 kg to Saturn Orbit Insertion with 3+1 NEXT system, EGA + Delta IV Heavy
17 Ongoing Technology Development: High-Isp Hall High-Isp Hall thrusters needed for expanded mission capture
Input Power 0.3 - 3.5 kW Specific Impulse 1600 - 2700 s Efficiency > 55% @ 3.5 kW Thrust 20 – 150 mN Propellant > 300 kg Throughput Specific Mass 2.4 kg/kW HIVHAC EM Operational Life > 10,000 hrs
FY11 key activities: Technical Status TRL • EM thruster performance test HIVHAC Components • EM thruster environmental test Thruster 3-4 • Long Duration Test Readiness Review HIVHAC-related • Hall-related component evaluation Components Power Processor 3 FY11-15 planned key activities: Unit • Component development for remaining Xenon Feed System 6 system: PPU, feed system, etc. Gimbal 4-5 • System integration • Life testing 18 Concepts – Performance Regimes
16000
14000
12000
10000
(s) NEP Robotic NEP Piloted sp 8000 I Ion Interplanetary Interplanetary
6000 MPD/LFA PIT 4000 SEP Hall Interplanetary
OM 2000 Orbit LEO to GEO&Escape & SK PPT ACS Insertion & Repo. Electro-Thermal Cold Gas/Chemical 0 0.001 0.01 0.1 1 10 100 1000 P (kW per thruster unit) 1. Flight applications to date 2. Future driving requirements and associated mission applications and capabilities provided by this technology 3. Specific mission examples of the application of near-term and longer-term propulsion options
20 21 Multiple rendezvous for small bodies Control of arrival conditions Enables many asteroid and comet missions Achieve lower speed arrival or control that are impractical without SEP arrival time for Mars or Venus entry Change direction and velocity of approach to reach more landing sites Reduced number of mission critical events e.g., orbit insertion, earth avoidance, response to More mass delivered to destination anomalies…… Could enable more mass on smaller (and cheaper) launch vehicles Provides performance margin and resilience to mass growth
More flexible launch opportunities Shorter trip times More frequent launch opportunities Might expand feasible mission e.g., Dawn delay was possible to accommodate set beyond the asteroid belt Phoenix launch including return of samples to Decreased reliance on JGA availability Earth
22 Deep Space 1
Deep Space 1: Technology Demonstration Mission Retired the following risks: Thruster life Guidance, navigation and control of an SEP spacecraft Mission operations Costs Spacecraft contamination Communications impact Dawn Electromagnetic compatibility Dawn: The use of SEP on Dawn reduced the cost of a multiple main belt asteroid rendezvous mission from New Frontier-class to Discovery- class – a difference of over $200M
23 Will orbit both the main-belt asteroid Vesta and the dwarf planet Ceres Launched: September 2007 1218 kg launch mass (dry mass of 750 kg) 10-kW Solar Array (at 1 AU) ~20,000 hours of thrusting with the ion propulsion system and operating flawlessly Approximate V delivered to date: 5.8 km/s Xenon used to date: 215 kg (425 kg loaded) July 2011 arrival at Vesta
24 SMART-1: Small Mission for Advanced Research in SMART-1 Technology Launched: September 2003 Hayabusa: Near-earth asteroid sample return Launched: May 2003 Return: June 13, 2010 GOCE: Gravity field and steady-state Ocean Circulation Explorer Launched: March 2009 Hayabusa GOCE
25 53 commercial satellites now flying with xenon ion Loral FS1300 propulsion S/C bus with SPT-100 Hall Commercial satellites now flying with up to 24 kW of thrusters solar power at beginning of life Boeing HS702 S/C bus with the 25-cm XIPS ion propulsion system
AF Advanced EHF Satellite with the BPT-4000 Hall thruster system
High power (> 20 kW) is now routine on commercial satellites Electric propulsion now used by almost all major satellite providers because it provides a significant economic benefit to the end user 26 During the 1980’s and 90’s, a rapid infusion of NASA electric propulsion technology onto COMSATs occurred – from 1983 when there were no satellites with electric propulsion to the present where about 43% of all active COMSATs use electric propulsion (144 of 336 spacecraft) 27 28 Low-mass body rendezvous Multiple flybys of NEOs, e.g., flyby 39 NEOs in 10 years with SEP Sample return missions from primitive bodies
Target Mission FlightTime Power Benefit (years) Near Earth Objects • Sample Return < 4 Solar: 5 to 15 kW Increases number of reachable (0.98 to 1.3 AU) • Multiple Flybys 10 objects from a few dozen to thousands Jupiter Family Comets • Rendezvous 3 to 5 Solar: 10 to 15 kW Enables missions to large, hard- (1 to 6 AU) • Sample Return 9 (round trip) Solar: 15 to 30 kW to-reach comets (Tempel 1, 2, etc.) Main Belt Asteroids • Multiple ~4 (per object) Solar: 10 to 15 kW Enables multiple main-belt (1.7 to 5 AU) Rendezvous asteroid rendezvous missions Trojan Asteroids • Rendezvous 3 - 5 Solar: 30 kW Enables rendezvous missions (~5 AU) RPS*: ~1 kW Centaurs • Rendezvous 9 - 10 RPS: ~1 kW Enables rendezvous missions (5 to 30 AU) Kuiper Belt Objects • Multiple Flybys 10 - 20 RPS: ~1 kW Enables rendezvous missions (> 40 AU) • Rendezvous *Radioisotope Power System at ~8 W/kg Electric propulsion greatly expands the number of reachable target bodies for affordable primitive body missions. 29 1,000,000
ISS 100,000
Dawn 10,000 Skylab Power levels needed for 1,000 SERT II near-term missions Power Doubles every four years
100 Solar Array Power (W) Power Array Solar
10 The Discovery-sized Dawn S/C has more power than Skylab Vanguard 1 1950 1960 1970 1980 1990 2000 2010 2020 Launch Year
Maximum solar power per spacecraft has doubled approximately every 4 years for the last 50 years Power levels needed for deep-space missions are now routine, but light-weight, low-risk arrays are required to improve performance 30 Primary Drivers Mission BOL Specific Specific Power Power Cost Power Level (kW) (W/kg) ($/W) Cost DS1 (1998) 2.5 42 “free” Risk (real and perceived) Dawn (2007) 10.3 82 ~900 Specific Power (W/kg) Near-Term 10 to 30 > 200 < 500 Secondary Drivers Far-Term > 100 > 300 < 250 Low Intensity Low Temp. (LILT) Natural Frequency • ISS arrays are about 27 W/kg and $3,500/W Packaging High-temperature operation (for Venus gravity assist trajectories) Need low-risk, light-weight solar arrays in the 10- to 30-kW range At 30 kW, a 200W/kg array saves > 200kg relative to an 82 W/kg array 31 SEP combined with multiple gravity assist flybys Deep Space 1 had no gravity assists is a powerful combination for outer planet Dawn uses a single Mars gravity missions, but the vast number of possible assist combinations can make it a daunting task to find the good trajectories The Titan Saturn System Mission The number of possible trajectories grows study considered the following geometrically with the number of encounters trajectory options
nDates nDates Earth-Earth-Venus-Venus-Earth-Saturn nLaunch Dates 3 flyby bodies 5 flyby bodies nDates Earth-Venus-Venus-Earth-Saturn Earth-Earth-Earth-Saturn Earth-Mars-Venus-Earth-Saturn
Earth-Venus-Earth-Earth-Saturn
Earth
GABody
GA Body
Target
st
nd
1 2
For n = 12 there are 12 x (3 x 12) x (5 x 12) x 12 = 311,040 trajectories
Advanced trajectory analysis capabilities are required to find the best low- thrust trajectories, which are mission enhancing and in some cases mission enabling. 32 Flagship and New Frontier Missions Discovery Missions Solar Array Power Levels of 15- to 30 kW Solar Array Power Levels of 5- to 15 kW Primary Drivers (per thruster) Primary Drivers (per thruster) Life (Propellant Throughput): > 900 kg Life (Propellant Throughput): > 450 kg Power: 5 kW to > 10 kW Power: 3 kW to 5 kW Specific Impulse: > 4,000 s Specific Impulse: 2,500 to 3,500 s Throttle Range: 30-to-1 Throttle Range: 15-to-1 Affordable PPU Affordable PPU
• Thruster life limitations forced Dawn to carry 3 thrusters cross-strapped to 2 PPUs – an expensive arrangement • The simplest, lowest-cost electric propulsion system has just two thrusters – one primary and one spare. To achieve this, high-power, long-life thruster development is needed for missions with power levels > ~15 kW
Need high-power thrusters for directed missions Need low-cost, low-risk electric propulsion systems for competed missions
33 Electric TRL Demonstrated Gap Benefit of Filling the Gap Propulsion Performance Technology NSTAR 9 Max. Power: 2.5 kW 1. PPU too expensive due to 1. This technology has been overtaken by /Dawn Max. Isp: 3100 s manufacturablility problems more modern versions (i.e., NEXT) Throughput: 235 kg XIPS 9 Max. Power: 4.5 kW 1. Demonstrated propellant throughput of 1. NEXT throughput capability makes Max. Isp: 3500 s 170 kg is too low for most missions redoing the XIPS life test questionable Throughput: 170 kg 2. PPU modifications required for deep- 2. PPU could be modified to operate the space missions NEXT thruster resulting in significant cost savings BPT-4000 7 Max. Power: 4.5 kW 1. PPU redesign needed for deep-space 1. Lowest cost EP system for Discovery Max. Isp: 2000 s missions and NF missions Throughput: 272 kg 2. Demonstrated propellant throughput of 2. Reduces system cost and complexity 272 kg is too low for some missions, for many missions need > 600 kg 3. Isp of 2000s is too low for many 3. Increase mission capture to all near- missions, need ~2800 s term competed missions NEXT 4 Max. Power: 7.2 kW 1. Too expensive for competed missions 1. Repackaging required to fix technical Max. Isp: 4200 s • PPU difficult to manufacture and issues, improve manufacturability, and Throughput: 450 kg troubleshoot reach TRL6 in order to lower cost and • Significant PPU development issues risk to acceptable levels remain • Digital Interface and Control Unit requires significant development 2. Power too low for far-term missions 2. Reduces EP system complexity and cost for far-term missions HIVHAC 3 Max. Power: 3.5 kW 1. PPU needed for deep-space missions 1. Same PPU could also be used with the Max. Isp: 2700 s BPT-4000 Throughput: TBD kg 2. Significant life issues remain 2. High propellant throughput capability essential for mission capture 34 Meeting the near-term goals is necessary to 20 reduce cost and risk for 18
Discovery and New 16 Frontiers users 14
12 Long-term goal represents an order of 10 magnitude increase in 8 NEXT power and propellant Near-term goals for Discovery (BPT-4000) 6 and New Frontiers (NEXT) missions throughput relative to MaximumInput Power (kW) XIPS 4 the Dawn ion thruster Dawn 2
0 Meeting the long-term 0 500 1000 1500 2000 goals is necessary to Demonstrated Propellant Throughput (kg) enable exciting new mission possibilities
35 Thruster Maximum Projected Maximum Current Time to TRL6 Input Propellant Isp TRL Power Throughput (s) (kW) (kg) NEXT 7.2 750 4200 6 NearTerm: ≤ 3 years
NEXT STEP 13.7 750 4400 4 Mid Term: ≤ 6 years
High-Power Hall 20 2000 3000 4 Mid Term: ≤ 6 years
High-Power Ion 28 2000 8500 3 Far Term: ≤ 9 years NEXT/NEXT STEP High-Power Hall High-Power Ion
7.2 – 13.7 kW 20 kW
It is now possible to build EP thrusters with capabilities that were unheard of only a few years ago enabling new mission concepts. 28 kW 36 Thruster Maximum Demonstrated Demonstrated Current Time to TRL6 Input Power Propellant Maximum TRL (kW) Throughput Isp (kg) (s) NSTAR/Dawn 2.5 235 3100 9 N/A
XIPS 4.5 155 3500 9 N/A
BPT-4000 Hall 4.5 272 2200 7 NearTerm: ≤ 3 years
NEXT 7.2 450 4200 5 NearTerm: ≤ 3 years
High-Isp Hall 4.5 --- 2800 4 Mid Term: ≤ 6 years
NSTAR/Dawn XIPS NEXT BPT-4000 For near-term Discovery missions available thrusters have excellent
performance, but the power processor units (PPU) are problematical 37 Power Processor Unit (PPU) Converts the solar array power to the Dawn currents and voltages needed to start and operate the thruster PPUs are complicated and expensive Commercial PPUs are designed to operate from a regulated high- voltage bus NEXT PPUs for deep-space missions must accept a variable input voltage (typically 80 V to 140 V) A modular PPU design (like XIPS) is needed to improve manufacturability and enable XIPS tailoring for different power levels
Need a modular PPU that is manufacturable and affordable to improve mission capture for competed missions 38 39 What? Identify a very small Near Earth Asteroid (NEA) with a mass of ~10,000 kg (corresponding to a diameter of ~2 m) Use a high-power Solar Electric Propulsion (SEP) System to rendezvous with this object, capture it, and return it to the International Space Station (ISS) When? Launch before the end of the decade using a single evolved expendable launch vehicle (EELV) with a total flight time of ~5years Why? Assess resource potential of NEAs for exploration and commercial use Use the ISS as a geology lab Use the ISS as a test bed for learning how to handle/process asteroid material in space
40 Key Feasibility Issues: Is it possible to find a sufficiently small, scientifically interesting NEA in an accessible orbit? How would you capture, secure, and transport a 10,000-kg asteroid? How would you safely approach and dock with the ISS while transporting a 10,000-kg asteroid? International Space Station (ISS) What would you do with The asteroid would be curated at ISS where numerous possible scientific a 10,000-kg asteroid at and resource utilization experiments would be conducted. This would provide valuable experience with tools and techniques, prior to a human the ISS? mission to a NEA.
41 Reasonable projections suggest that several dozen candidate NEAs of the right size and orbit The Panoramic Survey Pan-STARRS Telescope & Rapid could be found through the use of new Response System telescopes by the end of the decade from which a suitable target could be selected.
Low-thrust trajectory analyses suggest that the mission could be performed from a single EELV with a total flight time of approximately 5 years and a SEP power level of 40 kW.
A concept for capturing, securing, and de- spinning the asteroid has been identified.
Multiple approaches for docking the SEP vehicle and its asteroid cargo with the ISS have been Large Synoptic Survey Telescope identified.
Safely handling the asteroid at the ISS would require care and planning, but is definitely feasible.
Planetary protection issues were not addressed High-Power Electric Propulsion in this study. 42 Complete NEXT to TRL6 for near-term New Frontiers and Flagship missions
Complete High-Isp Hall development for Discovery missions
Develop advanced solar array technologies for near-term and far-term missions with electric propulsion
Develop high-power, very-long-life, electric propulsion technologies for far-term New Frontiers and flagship missions
Develop advanced trajectory techniques needed to find the best performing low-thrust trajectories from among millions of possible combinations
Develop thruster life validation modeling and simulations necessary for risk management of EP system implementation
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