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
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Future Applications Non-toxic propellant-based propulsion systems (LOX, LH2, LCH4, Ethanol) 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 Preliminary trade studies, requirements development, and planning for Altair ascent propulsion subsystem Flight System Development Main Propulsion Systems (MPS) Reaction Control Systems (RCS) 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 Current Focus Areas Ion and Hall thruster design, fabrication, evaluation Innovative concept development and evaluation Power electronics design, bread-boarding, evaluation System integration, system qualification and acceptance testing Thermal, stress, vibration analysis and test Capability to fabricate laboratory and flight hardware Technology and Flight System Development of Ion and Hall Thrusters 200 kwe Class SEP Stage VASIMR or other high power EP technology Reusable orbit transfer Robotic interplanetary cargo transfer Solar Electric Propulsion (SEP) Stage 30 kwe class system with advanced PV and AR&D technologies NASA science missions Small commercial servicing applications in Earth orbit Use of NEXT ion or Hall thruster technology SEP Stage Upgrade 90 kwe class system Orbital servicing and debris removal ISS reboost Science missions, e.g., Mars Sample/Return 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) 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 Integrated NTP Crewed Vehicle Design for Mars DRA 5.0 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. Propulsion Planetary Ascent Vehicles Earth Entry Vehicles /Aerocapture Systems & Mission Studies Sample Return Propulsion Tech High-Temp Engine (AMBR) Mars Ascent Vehicle TPS & Structures Tools Ultra Lt Wt Tank NEXT GN&C Mission Studies Hall Thruster Multi-Mission Earth Entry Vehicle System Studies 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
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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 M f 0 exp - ΔV v e For a given change in velocity (ΔV), the delivered mass (M f ) depends on the propellant exhaust velocity (v e ) 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, steadystate current to ionize gas propellant, induce magnetic field (JxB) 13
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Objective: Develop advanced ion propulsion system with improved performance and life characteristics to reduce user costs and enhance/enable a broad range of NASA SMD missions NEXT gridded ion thruster NEXT addresses the entire ion propulsion system - Gridded ion thruster - Power processing unit (PPU) - Propellant management system (PMS) - System integration (including gimbal and control functions) Primary Partners - NASA Glenn Research Center: Lead - JPL, Aerojet Corp., L3 Comm. NEXT PM ion thruster operation HPA DCIU Thruster Attribute Thruster power range, kw 0.5-6.9 Max. Specific Impulse, s 4,190 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 Low Pressure Assembly mass, kg 3.1 PPU NEXT Thruster String LPA Thruster Gimbal * Rated Capability Goal 300Kg Design/Qualification Goal (1.5x Rated) 450Kg Projected 1 st Failure >750Kg Potential Rated Capability 500Kg 15
Critical tests have been completed, or are imminent, on high fidelity hardware PM1 PM1R PPU Feed System Gimbal Functional & Performance Testing Complete Complete Complete Complete Complete Qual-Level Vibration Test Complete Complete FY11 Complete Complete Qual-Level Thermal/ Vacuum Test Complete Complete FY11 Complete Not planned Single-String System Integration Test: Complete Multi-String System Integration Test: Complete Thruster Life Test: Completed goal of 450Kg throughput >34,000 hours and >570 kg of xenon processed to date Life Test will continue through 750Kg or first failure FY11 Key activities : PPU re-verification test, and environmental test Long-Duration Test (LDT) Phase II Review Technology Maturation Review FY11-13 Key activities: PPU design iteration and/or DCIU maturation LDT extension to demo 750 kg Residual, high-priority risk reduction activities 16
CHARACTERISTIC NSTAR (SOA) NEXT Improvement NEXT BENEFIT Max. Thruster Power (kw) 2.3 6.9 3x Max. Thrust (mn) 91 236 2.6x Enables high power missions with fewer thruster strings Throttling Range (Max./Min. Thrust) 4.9 13.8 3x Allows use over broader range of distances from Sun Max. Specific Impulse (sec) 3120 4190 32% Total Impulse (10 6 N-sec) 4.6 >18 >3.9x Propellant Throughput (kg) 150 450 3x Reduces propellant mass, enabling more payload and/or lighter spacecraft Enables low power, high V Discovery-class missions with a single thruster Mission Discovery - Small Body Missions New Frontiers - Comet Surface Sample Return Titan Direct Lander Flagship - Saturn System Missions Titan Enceladus Performance Finding Higher net payload mass with fewer thrusters than NSTAR system CSSR: Higher net payload mass than NSTAR, with, simpler EP System: 2+1 NEXT vs 4+1 NSTAR thrusters Titan: > 700 kg entry package with 1+1 NEXT system > 2400 kg to Saturn Orbit Insertion with 1+1 NEXT system, EGA + Atlas V EELV - Doubles delivered mass of chemical/jga approach > 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 Specific Impulse Efficiency Thrust Propellant Throughput Specific Mass Operational Life 0.3-3.5 kw 1600-2700 s > 55% @ 3.5 kw 20 150 mn > 300 kg 2.4 kg/kw > 10,000 hrs HIVHAC EM Technical Status TRL HIVHAC Components Thruster 3-4 HIVHAC-related Components Power Processor 3 Unit Xenon Feed System 6 Gimbal 4-5 FY11 key activities: EM thruster performance test EM thruster environmental test Long Duration Test Readiness Review Hall-related component evaluation FY11-15 planned key activities: Component development for remaining system: PPU, feed system, etc. System integration Life testing 18
I sp (s) Concepts Performance Regimes 16000 14000 12000 10000 8000 Ion NEP Robotic Interplanetary NEP Piloted Interplanetary 6000 4000 Hall SEP Interplanetary MPD/LFA PIT 2000 0 Cold Gas/Chemical PPT ACS 0.001 0.01 0.1 1 10 100 1000 OM & SK P (kw per thruster unit) Orbit LEO to GEO&Escape Insertion & Repo. Electro-Thermal
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
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Multiple rendezvous for small bodies Enables many asteroid and comet missions that are impractical without SEP Reduced number of mission critical events e.g., orbit insertion, earth avoidance, response to anomalies Control of arrival conditions Achieve lower speed arrival or control arrival time for Mars or Venus entry Change direction and velocity of approach to reach more landing sites More mass delivered to destination Could enable more mass on smaller (and cheaper) launch vehicles Provides performance margin and resilience to mass growth More flexible launch opportunities More frequent launch opportunities e.g., Dawn delay was possible to accommodate Phoenix launch Decreased reliance on JGA availability Shorter trip times Might expand feasible mission set beyond the asteroid belt including return of samples to Earth 22
Deep Space 1 Deep Space 1: Technology Demonstration Mission Retired the following risks: Dawn: Thruster life Guidance, navigation and control of an SEP spacecraft Mission operations Costs Spacecraft contamination Communications impact 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 Discoveryclass 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 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 SMART-1 Hayabusa GOCE 25
53 commercial satellites now flying with xenon ion propulsion Commercial satellites now flying with up to 24 kw of solar power at beginning of life Loral FS1300 S/C bus with SPT-100 Hall thrusters 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
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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 Flight Time (years) Near Earth Objects (0.98 to 1.3 AU) Jupiter Family Comets (1 to 6 AU) Main Belt Asteroids (1.7 to 5 AU) Trojan Asteroids (~5 AU) Centaurs (5 to 30 AU) Kuiper Belt Objects (> 40 AU) Sample Return Multiple Flybys Rendezvous Sample Return Multiple Rendezvous *Radioisotope Power System at ~8 W/kg < 4 10 3 to 5 9 (round trip) Power Solar: 5 to 15 kw Solar: 10 to 15 kw Solar: 15 to 30 kw Benefit Increases number of reachable objects from a few dozen to thousands Enables missions to large, hardto-reach comets (Tempel 1, 2, etc.) ~4 (per object) Solar: 10 to 15 kw Enables multiple main-belt asteroid rendezvous missions Rendezvous 3-5 Solar: 30 kw RPS*: ~1 kw Enables rendezvous missions Rendezvous 9-10 RPS: ~1 kw Enables rendezvous missions Multiple Flybys Rendezvous 10-20 RPS: ~1 kw Enables rendezvous missions Electric propulsion greatly expands the number of reachable target bodies for affordable primitive body missions. 29
Solar Array Power (W) 1,000,000 100,000 10,000 1,000 100 SERT II Skylab ISS Dawn Power levels needed for near-term missions Power Doubles every four years 10 1 Vanguard The Discovery-sized Dawn S/C has more power than Skylab 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 Power Level Cost Risk (real and perceived) Specific Power (W/kg) Secondary Drivers Low Intensity Low Temp. (LILT) Natural Frequency Packaging High-temperature operation (for Venus gravity assist trajectories) Mission BOL Power (kw) Specific Power (W/kg) Specific Cost ($/W) DS1 (1998) 2.5 42 free Dawn (2007) 10.3 82 ~900 Near-Term 10 to 30 > 200 < 500 Far-Term > 100 > 300 < 250 ISS arrays are about 27 W/kg and $3,500/W 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
Earth 1 st GA Body 2 nd GA Body Target SEP combined with multiple gravity assist flybys is a powerful combination for outer planet missions, but the vast number of possible combinations can make it a daunting task to find the good trajectories The number of possible trajectories grows geometrically with the number of encounters n Launch Dates n Dates 3 flyby bodies n Dates 5 flyby bodies n Dates Deep Space 1 had no gravity assists Dawn uses a single Mars gravity assist The Titan Saturn System Mission study considered the following trajectory options Earth-Earth-Venus-Venus-Earth-Saturn Earth-Venus-Venus-Earth-Saturn Earth-Earth-Earth-Saturn Earth-Mars-Venus-Earth-Saturn Earth-Venus-Earth-Earth-Saturn 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 lowthrust trajectories, which are mission enhancing and in some cases mission enabling. 32
Flagship and New Frontier Missions Solar Array Power Levels of 15- to 30 kw Primary Drivers (per thruster) Life (Propellant Throughput): > 900 kg Power: 5 kw to > 10 kw Specific Impulse: > 4,000 s Throttle Range: 30-to-1 Affordable PPU Discovery Missions Solar Array Power Levels of 5- to 15 kw Primary Drivers (per thruster) Life (Propellant Throughput): > 450 kg Power: 3 kw to 5 kw Specific Impulse: 2,500 to 3,500 s Throttle Range: 15-to-1 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 Propulsion Technology TRL Demonstrated Performance Gap Benefit of Filling the Gap NSTAR /Dawn 9 Max. Power: 2.5 kw Max. Isp: 3100 s Throughput: 235 kg XIPS 9 Max. Power: 4.5 kw Max. Isp: 3500 s Throughput: 170 kg BPT-4000 7 Max. Power: 4.5 kw Max. Isp: 2000 s Throughput: 272 kg NEXT 4 Max. Power: 7.2 kw Max. Isp: 4200 s Throughput: 450 kg HIVHAC 3 Max. Power: 3.5 kw Max. Isp: 2700 s Throughput: TBD kg 1. PPU too expensive due to manufacturablility problems 1. Demonstrated propellant throughput of 170 kg is too low for most missions 2. PPU modifications required for deepspace missions 1. PPU redesign needed for deep-space missions 2. Demonstrated propellant throughput of 272 kg is too low for some missions, need > 600 kg 3. Isp of 2000s is too low for many missions, need ~2800 s 1. Too expensive for competed missions PPU difficult to manufacture and troubleshoot Significant PPU development issues remain Digital Interface and Control Unit requires significant development 2. Power too low for far-term missions 1. PPU needed for deep-space missions 2. Significant life issues remain 1. This technology has been overtaken by more modern versions (i.e., NEXT) 1. NEXT throughput capability makes redoing the XIPS life test questionable 2. PPU could be modified to operate the NEXT thruster resulting in significant cost savings 1. Lowest cost EP system for Discovery and NF missions 2. Reduces system cost and complexity for many missions 3. Increase mission capture to all nearterm competed missions 1. Repackaging required to fix technical issues, improve manufacturability, and reach TRL6 in order to lower cost and risk to acceptable levels 2. Reduces EP system complexity and cost for far-term missions 1. Same PPU could also be used with the BPT-4000 2. High propellant throughput capability essential for mission capture 34
Maximum Input Power (kw) Meeting the near-term goals is necessary to reduce cost and risk for Discovery and New Frontiers users 20 18 16 14 Long-term goal represents an order of magnitude increase in power and propellant throughput relative to the Dawn ion thruster 12 10 8 6 4 2 XIPS Dawn NEXT Near-term goals for Discovery (BPT-4000) and New Frontiers (NEXT) missions Meeting the long-term goals is necessary to enable exciting new mission possibilities 0 0 500 1000 1500 2000 Demonstrated Propellant Throughput (kg) 35
Thruster Maximum Input Power (kw) Projected Propellant Throughput (kg) Maximum Isp (s) Current TRL Time to TRL6 NEXT 7.2 750 4200 6 Near Term: 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 Input Power (kw) Demonstrated Propellant Throughput (kg) Demonstrated Maximum Isp (s) Current TRL NSTAR/Dawn 2.5 235 3100 9 N/A XIPS 4.5 155 3500 9 N/A Time to TRL6 BPT-4000 Hall 4.5 272 2200 7 Near Term: 3 years NEXT 7.2 450 4200 5 Near Term: 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 currents and voltages needed to start and operate the thruster PPUs are complicated and expensive Commercial PPUs are designed to operate from a regulated highvoltage bus 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 tailoring for different power levels XIPS NEXT Dawn Need a modular PPU that is manufacturable and affordable to improve mission capture for competed missions 38
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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? What would you do with a 10,000-kg asteroid at the ISS? International Space Station (ISS) The asteroid would be curated at ISS where numerous possible scientific and resource utilization experiments would be conducted. This would provide valuable experience with tools and techniques, prior to a human mission to a NEA. 41
Reasonable projections suggest that several dozen candidate NEAs of the right size and orbit could be found through the use of new telescopes by the end of the decade from which a suitable target could be selected. The Panoramic Survey Telescope & Rapid Response System Pan-STARRS 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 despinning the asteroid has been identified. Multiple approaches for docking the SEP vehicle and its asteroid cargo with the ISS have been identified. Large Synoptic Survey Telescope Safely handling the asteroid at the ISS would require care and planning, but is definitely feasible. Planetary protection issues were not addressed in this study. High-Power Electric Propulsion 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 43
Michael.J.Patterson@nasa.gov (216) 977-7481 John.R.Brophy@jpl.nasa.gov (818) 354-0446 44