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The NASA In-Space Propulsion Technology Project’s Current Products and Future Directions

46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit 25 - 28 July 2010, Nashville, TN

AIAA 2010-6648

The NASA In-Space Propulsion Technology Project’s Current Products and Future Directions
David J. Anderson 1 , Eric J. Pencil 2 , Larry C. Liou 3 NASA Glenn Research Center, Cleveland, OH 44135 John W. Dankanich 4 Gray Research, Inc., Cleveland, OH 44135 Michelle M. Munk 5 , David Hahne 6 NASA Langley Research Center, Hampton, VA 23681

The In-Space Propulsion Technology (ISPT) project, since its inception in 2001, has been developing and delivering in-space propulsion technologies that will enable or enhance NASA robotic science missions. These in-space propulsion technologies are applicable, and potentially enabling, for future NASA flagship and sample return missions currently being considered, as well as having broad applicability to future competed Discovery and New Frontiers mission solicitations. This paper provides status of the technology development, applicability, and availability of in-space propulsion technologies that have recently completed their technology development and will be ready for infusion into missions. The recently completed ISPT technologies are: 1) the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance for lower cost; 2) NASA’s Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttleable gridded ion system; and 3) Aerocapture technology development with investments in a family of thermal protection system (TPS) materials and structures; guidance, navigation, and control (GN&C) models of blunt-body rigid aeroshells; aerothermal effect models: and atmospheric models for Earth, Titan, Mars and Venus. The paper will also describe the ISPT project’s future focus on propulsion for sample return missions. The future ISPT technology development areas will be: 1) Planetary Ascent Vehicles (PAV); 2) multi-mission technologies for Earth Entry Vehicles (MMEEV) needed for sample return missions from many different destinations; 3) propulsion for Earth Return Vehicles (ERV) and transfer stages, and electric propulsion for sample return and low cost missions; 4) advanced propulsion technologies for sample return; and 5) Systems/Mission Analysis focused on sample return propulsion.

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Advanced Materials Bipropellant Rocket Advanced Xenon Feed System Composite Overwrap Pressure Vessel Digital Control Interface Unit Entry, Descend, and Landing Earth Entry Vehicle Engineering Model Electro-Form process Electromagnetic Compatibility Electromagnetic Interference

Acting ISPT Project Manager, ISPT, 21000 Brookpark Road/MS 77-4, AIAA Member Propulsion Project Manager, ISPT, 21000 Brookpark Road/MS 77-4, AIAA Associate Fellow 3 Advanced Chemical Project Manager, ISPT, 21000 Brookpark Road/MS 77-4, AIAA Senior Member 4 Lead Systems Engineer, ISPT, 21000 Brookpark Road/MS 77-4, AIAA Senior Member 5 Aerocapture Project Manager, 1 North Dryden Street,/MS 489, AIAA Senior Member 6 MMEEV Project Manager, 1 North Dryden Street,/MS 489

1 American Institute of Aeronautics and Astronautics
This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.


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Earth Return Vehicle Flow Control Module Gross Life-off Mass Guidance, Navigation and Control Glenn Research Center Goddard Space Flight Center High Performance Apogee Thruster High Voltage Hall Accelerator High Pressure Assemblies Integrated Product Development Team Specific impulse, second(s) In-Space Propulsion Technology project or office Jet Propulsion Laboratory Johnson Space Center Langley Research Center Low-Pressure Assemblies Low-Thrust Trajectory Tool Mission Analysis Low Thrust Optimization Mars Ascent Vehicle Multi-Mission Earth Entry Vehicle Marshall Space Flight Center Mars Sample Return Mixture Ratio Mars Science Laboratory Hydrazine—the fuel of the bipropellant National Aeronautics and Space Administration NASA Evolutionary Xenon Thruster NASA Research Announcement NASA Solar Electric Propulsion Technology Readiness Nitrogen Tetroxide—the oxidizer of the bipropellant Optimal Trajectories by Implicit Simulation Planetary Ascent Vehicles Pressure Control Module Preliminary Design Review Prototype Model Program to Optimize Simulated Trajectories Power-Processing Unit Science Definition Team Solar Electric Propulsion Science Mission Directorate at NASA Headquarters State of the Art Thermal Protection Systems Technology Readiness Level Two Stage to Orbit Velocity increment for propulsion system or spacecraft

I. Introduction
This paper provides a brief overview of the ISPT project with development status, near-term mission benefits, applicability, and availability of in-space propulsion technologies in the areas of aerocapture, electric propulsion, advanced chemical propulsion, planetary ascent vehicles, Earth return vehicles, other advanced propulsion technologies, and mission/systems analysis tools. These in-space propulsion technologies are applicable, and potentially enabling for future NASA flagship and sample return missions currently under consideration, as well as having broad applicability to future Discovery and New Frontiers mission solicitations. NASA’s Science Mission Directorate (SMD) missions seek to answer important science questions about our planet, the Solar System and beyond. To meet NASA’s future mission needs, the goal of the ISPT project has been the development of new enabling propulsion technologies that cannot be reasonably achieved within the cost or schedule constraints of mission development timelines, specifically achieving technology readiness level (TRL) 6 prior to preliminary design review (PDR). Since the ISPT goal is to develop products that realize near-term and mid-term benefits, ISPT primarily focuses on 2 American Institute of Aeronautics and Astronautics

technologies in the mid TRL range (TRL 3–6+ range) that have a reasonable chance of reaching maturity in 4–6 years provided adequate development resources. The project strongly emphasizes developing propulsion products for NASA flight missions. Since 2001, the In-Space Propulsion Technology (ISPT) project has been developing and delivering in-space propulsion technologies that will enable and/or benefit near and mid-term NASA robotic science missions by significantly reducing cost, mass, and/or travel times. ISPT technologies will help deliver spacecraft to SMD’s destinations of interest. In late 2006, the ISPT project office was transferred to the GRC where it has managed the ISPT project for Science Mission Directorate. From 2001 to 2006, the ISPT project office was located at MSFC, where it was initiated and managed.


Technology Development Overview

ISPT emphasizes technology development with mission pull. Initially the ISPT goal was to develop technologies for Flagship missions. This goal led to the priorities of aerocapture (the use of aerodynamic drag for orbit capture) and electric propulsion. In 2006, the Solar System Exploration (SSE) Roadmap 1 identified technology development needs for Solar System exploration, and described transportation technologies as highest priority (new developments are required for all or most roadmap missions). According to the SSE Roadmap, the highest priority propulsion technologies are electric propulsion and aerocapture. The SSE Road map specifically stated that “Aerocapture technologies could enable two proposed Flagship missions, and solar electric propulsion could be strongly enhancing for most missions. These technologies provide rapid access, or increased mass, to the outer Solar System.”1 Electric propulsion and aerocapture are suited for enabling significant science return for the outer planetary moons under investigation. The ISPT technologies were quantified to allow greater science return with reduced travel times. The ISPT priorities and products are tied closely to the science roadmaps, the SMD’s science plan, and the decadal surveys. Excerpts from the science community are discussed in more detail in Ref. 2. The ISPT project is currently completing the development efforts in four technology areas. These include Advanced Chemical Propulsion, Aerocapture, Electric Propulsion, and Systems/Mission Analysis. It is one of ISPT’s objectives that all ISPT products be ultimately manufactured by industry and made equally available to all potential users for missions and proposals. The primary technology development in advanced chemical propulsion was the development of the Advanced Material Bi-propellant Rocket (AMBR) engine, which completed its developmental activities in 2009. Advanced chemical propulsion investments included the demonstration of active-mixture-ratio-control and lightweight tank technology. The advanced chemical propulsion technologies have an opportunity for rapid-technology infusion with minimal risk and broad mission applicability. Aerocapture technology developments result in better models for: 1) guidance, navigation, and control (GN&C) of blunt body rigid aeroshells, 2) atmosphere models for Earth, Titan, Mars and Venus, and 3) models for aerothermal effects. In addition to enhancing the technology readiness level (TRL) of rigid aeroshells, improvements were made in understanding and applying inflatable aerocapture concepts. Aerocapture technology was a contender for flight validation on NASA’s New Millennium ST9 mission. Electric propulsion (EP) technology development activities are focusing on completing NASA’s Evolutionary Xenon Thruster (NEXT) ion propulsion system. The NEXT system was selected under a competitive solicitation for an EP system applicable to a Flagship mission. NEXT is a 0.6-7-kW throttle-able gridded ion system suitable for future Discovery, New Frontiers, and flagship missions. At a sub-component level, ISPT is pursuing the development of a lightweight reliable xenon flow control system as well as standardized EP subcomponent designs. The ISPT project continues the development of other electric propulsion products, such as the High-Voltage Hall Accelerator (HIVHAC) thruster. The HIVHAC thruster is designed as a low cost, highly reliable thruster suited for cost-capped NASA Discovery-class missions. The systems analysis technology area performed numerous mission and system studies to guide technology investments and quantify the return on investment. Recent focus of the systems analysis area has been on developing tools to assist technology infusion. Tool development has included the development of low-thrust trajectory tools (LTTT), a suite of computer programs optimized for developing mission trajectories using EP, and an aerocapture quicklook tool. In 2009, ISPT was tasked to start development of propulsion technologies that would enable future sample return missions. Sample return missions could be quite varied, from collecting and bringing back samples of comets or asteroids, to soil, rocks, or atmosphere from planets or moons. Given this new focus, the future technology development areas for ISPT are: 1) Sample Return Propulsion, which includes: a. Electric propulsion for sample return and low cost Discovery-class missions b. Propulsion systems for Earth Return Vehicles (ERV) including transfer stages to the destination c. Low TRL advanced propulsion technologies 2) Planetary Ascent Vehicles (PAV), with a Mars Ascent Vehicle (MAV) being the initial development 3) Multi-mission technologies for Earth Entry Vehicles (MMEEV) 4) Systems/Mission Analysis that focuses on sample return propulsion The work on HIVHAC continues thruster development in FY2010 and then transitions into developing a HIVHAC 3 American Institute of Aeronautics and Astronautics

system under future Electric Propulsion for sample return (ERV and transfer stages) and low-cost missions. Previous work on the lightweight propellant-tanks will continue under advanced propulsion technologies for sample return with direct applicability to a Mars Sample Return (MSR) mission and with general applicability to all future planetary spacecraft. The Aerocapture efforts will merge with previous work related to Earth Entry Vehicles and transitions into the future multimission technologies for Earth Entry Vehicles (MMEEV). The Planetary Ascent Vehicles (PAV)/ Mars Ascent Vehicle (MAV) is a new development area to ISPT but builds upon and leverages the past MAV analysis and technology developments from the Mars Technology Program (MTP) and previous MSR studies.


Technology Infusion

The ISPT project is developing several technologies that have reached TRL 6 and are potentially applicable for infusion into future, Flagship, New Frontiers, and Discovery mission opportunities. Three technologies in particular are the NASA’s Evolutionary Xenon Thruster (NEXT) ion propulsion system, the Advanced Material Bi-propellant Rocket (AMBR) engine, and Aerocapture. ISPT and NASA are exploring several different paths to get its technology investments infused into future NASA, DOD, or commercial missions. NASA recognizes that it is desirable to fly new technologies that enable new scientific investigations or to enhance an investigation's science return. The SSE Roadmap states that NASA will strive to maximize the payoff from its technology investments, either by enabling individual missions or by enhancing classes of missions with creative solutions. Discovery, New Frontiers, and Flagship missions potentially provide opportunities to infuse advanced technologies developed by NASA, and advance NASA’s technology base and enable a broader set of future missions. To benefit from its technology investments, NASA provided an incentive to encourage the infusion of NEXT ion propulsion system or the AMBR engine into mission proposals in response to the New Frontiers 3 Announcement of Opportunity (AO). NASA is also offering an incentive to encourage the infusion of NEXT ion propulsion system, the AMBR engine, or aerocapture into mission proposals in response to the upcoming Discovery 2010 Announcement of Opportunity (AO). The Discovery 2010 AO was released on June 7, 2010 with proposals due September 2, 2010. Under these AO’s, proposers are offered an option of adopting one of the specific technologies for insertion into their missions. NASA would then share in the flight development costs of the proposed advanced technology, up to certain amounts specified in the AO depending upon which technology is proposed. Beyond the New Frontiers and Discovery opportunities, ISPT continues to seek opportunities to infuse NEXT, AMBR, Aerocapture, and its other technologies into a wide range of possible future mission opportunities. The ISPT project office and NEXT team personnel are actively supporting various flagship science definition team (SDT) studies. See the ISPT Overview paper in the 2010 IEEE Aerospace Conference for more details regarding these studies. 3 ISPT personnel supported several white papers that were developed in response to the current planetary science decadal survey development activities in 2009/2010. ISPT contributed to identifying the technology development that is required to accomplish the future missions being contemplated. And finally, NEXT and Aerocapture are showing up several times in Exploration Systems Mission Directorate (ESMD) and Office of the Chief Technologist (OCT) planning activities in 2010.



Aerocapture is the process of entering the atmosphere of a target body to practically eliminate the chemical propulsion requirements of orbit capture. Aerocapture is the next step beyond aerobraking, which relies on multiple passes high in the atmosphere using the spacecraft’s drag to reduce orbital energy. Aerobraking has been used at Mars on multiple orbiter missions. Aerocapture, illustrated in Fig. 1, maximizes the benefit from the atmosphere by capturing in a single pass. Aerocapture represents a major advance over aerobraking techniques, by generating more drag by flying at a lower altitude where the atmosphere is more dense. Keys to successful aerocapture are accurate arrival state knowledge, validated atmospheric models, sufficient vehicle control authority (i.e. lift-to-drag ratio), and robust guidance during the maneuver. A lightweight thermal protection system and structure will maximize the aerocapture mass benefits. The execution of the aerocapture maneuver itself is what enables the great mass savings over other orbital insertion methods. If the hardware subsystems are not mass efficient, or if performance is so poor that additional propellant is needed to adjust the final orbit, the benefits are significantly reduced. ISPT efforts in aerocapture Figure 1. Illustration of the aerocapture subsystem technologies are focused on improving the efficiency and maneuver. number of suitable alternatives for aeroshell structures and ablative thermal protection systems (TPS). These include development of 4 American Institute of Aeronautics and Astronautics

families of low and medium density (14-36 lbs/ft3) TPS, and the related sensors, development of a carbon-carbon ribstiffened rigid aeroshell, and high temperature honeycomb structures and adhesives. Development occurred on inflatable decelerators through concept definition and initial design and testing of several inflatable decelerator candidates. Finally, progress has been made through improvement of models for atmospheres, aerothermal effects, and algorithms and testing of a flight-like guidance, navigation and control (GN&C) system. Aerocapture enables rapid access to orbital missions at the outer planets and is enabling for two of the potential flagship missions in the last Roadmap—Titan Explorer and Neptune–Triton Explorer. For targets in the outer Solar System, aerocapture technology would reduce the trip time and deliver a larger payload mass, enabling these missions to be implemented with the current generation of heavy lift launch vehicles. The SSE Roadmap recommends "Aerocapture technologies and flight validation are a high priority to solar system exploration."1 The March 2008 OPAG meeting minutes recommends that "Aerocapture is a key enabling technology for the outer solar system, particularly at Titan, and some gas giant planets." 4 Titan Explorer could be the first to use this technology in a Flagship mission. Because of the deep atmosphere, large– scale height, and modest entry velocities, Titan is an attractive target for the use of aerocapture. For a potential Neptune– Triton Explorer (NTE) mission, aerocapture enables transit from Earth to Neptune in less than ten years. Because of the much higher entry velocity and exit velocity near escape, Neptune aerocapture requires a higher-lift vehicle and is a more challenging maneuver than at Titan. The majority of investment in aerocapture technology occurred in advancing the TRL of efficient rigid aeroshell systems. A family of low-density TPS materials carrying the identifier “SRAM” was developed under a competitively awarded contract with Applied Research Associates (ARA). These have a density range between 14 Figure 2. Solar Tower lb/ft3 and 24 lb/ft3 with testing of aeroshell. the variable performance achieved by adjusting the ratios of constituent elements. These are applicable for heating rates up to 150 W/cm2 and 500 W/cm2 respectively. They could be used on missions Figure 3. 1.0-meter aeroshell with destinations to small bodies such as Titan and Mars. The SRAM family of ablators was tested in both arcjet and solar tower facilities (Fig. 2) at the coupon level; one-ft and two-ft square flat panels, and on a one meter, 70 degree, blunt body aeroshell structure; shown in Fig. 3. Another ARA family of low-to-medium density TPS systems (PhenCarb) is phenolic-based, ranges in density between 20 and 36 lb/ft3, and is applicable for heating rates between 200 and 1,500 W/cm2. In support of the rigid TPS system, ISPT funded testing of higher temperature adhesives and development of higher temperature composite structures effectively increasing the allowable bond-line temperature from 250C to 325 or 400C depending on the adhesive and composite construction. This work was performed by ATK, in the division formerly known as Composite Optics. Sensors that measure aeroshell recession with sub-millimeter accuracy were developed at NASA’s Ames Research Center and are currently planned for use on the Mars Science Laboratory (MSL) mission. Instrumenting entry systems to gather flight data is of primary importance to understand the environments and resulting vehicle requirements for future missions. Another advancement, enabled by ISPT funding, is the development of a Carbon-Carbon aeroshell that is rib stiffened, reducing the need for an additional structure system. The reduced mass of the structure, coupled with low-density insulation on the side of the shell, results in a 30 percent mass density improvement over the same size Genesis-like aeroshell. When this system was mechanically tested to levels that are representative of expected aerocapture loading environments, the system response compared within 10 percent to the finite element model, validating that model for use in predicting system response to other environments. This effort was competitively awarded and completed in early 2007 by Lockheed Martin and their partner Carbon-Carbon Advanced Technologies (C-CAT), and resulted in a TRL-6 product applicable for use in multiple NASA science missions. Ames Research Center developed and enhanced models that predict the entry thermal environments for aerocapture at Titan, Mars, Venus, and Neptune. In some cases, previous heating estimates were overly conservative because of the lack of resources available to produce validation data or to develop more complicated analysis methods. Coupled models updated with the most current Cassini data reveal, that aerocapture at Titan will load the TPS system at less than 20 W/cm2 verses prior predictions of 150-300 W/cm2. Through multiple years of concentrated effort, researchers funded by ISPT made modeling improvements that benefit all future entry missions, and published over 50 papers documenting these results. ISPT funds also supported the generation or update of engineering level atmospheric models for all primary aerocapture destinations except Earth work led at the Marshall Space Flight Center. 5 American Institute of Aeronautics and Astronautics

ISPT developed a rigorous, peer-reviewed plan as part of the ST9 New Millennium Proposal to take the ablative aerocapture system to a TRL 6 by FY09. Though the ST9 flight opportunity was cancelled, ISPT has continued to implement the ground maturation plans preparing the technology for a flight demo or first mission infusion. A 2.65-m diameter high-temperature aeroshell, with ARA’s SRAM TPS and ARC’s instrumentation plugs, is being built as a manufacturing demonstration, to be completed by late 2010 (Fig 4). Another effort to raise the TRL for TPS materials includes Space Environmental Effects (SEE) testing. Conducted at the Marshall Space Flight Center, this testing includes radiation exposure, cold soak, and Figure 4. 2.65-meter high-temperature aeroshell micrometeoroid impact on the ISPT-matured TPS and structure in curing oven. hot structure materials, to levels representative of a deep space mission. Figure 5 shows the shroud manufactured to cold soak the samples prior to a 7-km/s micrometeoroid impact by the Micro Light Gas Gun. Following exposure to these environments, samples will be arcjet tested to aerocapture heat rates and loads, in the Interaction Heating Facility at NASA-Ames. The results will be compared to arcjet tests of unexposed samples. The testing is expected to be complete in late CY2010. The aerocapture guidance algorithm used in all ISPT systems analyses, and selected for flight on ST9, is a fully analytic solution of about 300 lines of code, called HYPAS. Ball Aerospace has converted HYPAS to flight software and has completed development of a real-time hardware-inthe-loop test bench of a representative GN&C system for robotic planetary missions (shown in Fig. 6). The test bench demonstrates that the guidance performs well even when realistic hardware response times are present. Figure 5. Cold shroud for This development 5 brings the TRL of the aerocapture guidance to TRL 6, micrometeroroid testing (facility ready for flight infusion. Additional information on aerocapture located at NASA-MSFC). 6 technology developments can be found in the Discovery program library and in Ref. 7, 8, 9, 10, 11, 12. The use of aerocapture has been studied extensively through detailed systems analysis, most notably for use at Titan, Neptune, Venus and Mars. Mass benefits for all solar system destinations were derived and are documented in Ref. 13. The largest mass benefit from aerocapture was observed for Neptune, low Jupiter orbits, followed by Titan, Uranus, Venus, and then only marginal gains for Mars (the mass benefit is directly correlated to the amount of velocity change required for each mission). Alternatively, cost benefits are realized for multiple missions. When the overall system mass is reduced, the mission can utilize a smaller launch vehicle, saving tens of millions of dollars. Detailed mission assessment results are in Ref. 14, 15, 16. The mission mass benefits to Mars are expected to be about 5-15 percent, depending on the scale of the spacecraft. These benefits can be enabling. A multi-center team from ARC, JPL, JSC, LaRC, and MSFC conducted detailed mission and cost analyses for various Mars opportunities. An opposition-class sample return mission can be enabled in less than two years using aerocapture. Aerocapture enhances conjunction-class sample-return missions and large Mars orbiters. No new technology gaps were identified that would delay aerocapture implementation on such a mission. Venus was studied extensively to identify any needs for TPS, guidance, atmospheric or heating models. Detailed analyses evaluated the potential for aerocapture for a Venus Discovery class mission. Aerocapture delivered more than 80 percent additional mass over aerobraking and more than 600 percent Figure 6. Test Bench for over a chemical insertion. Aerocapture reduces Deep Space Network (DSN) Aerocapture GN&C) time by 121 days. No critical technology gaps were identified for aerocapture at Venus, but investments in TPS are recommended for achieving maximum mass benefits. 6 American Institute of Aeronautics and Astronautics

Titan continues to be of considerable scientific interest following the success of Cassini/Huygens. Because of its atmospheric structure, it is an ideal candidate for aerocapture. The Outer Planets Flagship (OPF) study considered aerocapture within the baseline mission concept since aerocapture has the capability to delivery more than double the scientific payload of the chemical alternative. Aerocapture may play a key role in accomplishing a reduced Titan mission for a less-than-Flagship budget or providing an alternate Flagship operational scenario. Aerocapture was proven repeatedly to be an enabling or strongly enhancing technology for several atmospheric targets. The ISPT project team continues to develop aerocapture technologies in preparation for a flight demonstration, and rapid aerocapture analysis tools are being developed and made available to a wider user community. The TPS materials developed through ISPT enhance a wide range of missions by reducing the mass of entry vehicles. The remaining gaps for technology infusion are efficient TPS for Venus and high-speed Earth return, and investments in aerothermal modeling. TPS structures and aerodynamics for Neptune. All of the other component subsystems for an aerocapture vehicle are currently at or funded to reach TRL 6 in the next year for the bodies of interest. This assessment of technology readiness is detailed in Ref. 3. The structures and TPS subsystems as well as the aerodynamic and aerothermodynamic tools and methods can be applied to small-scale entry missions even if the aerocapture maneuver is not utilized. Aerocapture cannot reach TRL 6 for the system without space flight validation, since it is impossible to match the flight environment in ground facilities. Missions must be willing to accept the small risk of this shortfall, to realize the tremendous benefits of the technology. If they are not willing, Aerocapture will need to be validated in space before its first mission infusion. A space flight validation is expensive, but the costs will be recouped very quickly. The validation immediately reduces the risk to the first user and validates the maneuver for application to multiple, potentially lower-cost, missions to Titan, Mars, Venus, and Earth. Moreover, once Aerocapture is proven a reliable tool, it is anticipated that entirely new mission possibilities will open up.


NASA’s Evolutionary Xenon Thruster (NEXT)

Solar Electric Propulsion (SEP) enables missions requiring large in-space velocity changes over time. SEP has applications to rendezvous and sample return missions to small bodies and fast trajectories towards the outer planets. This is particularly relevant to the Saturn-Titan-Enceladus and the Neptune-Triton missions. In particular, the Titan-Saturn System mission demonstrated that improvements to mass, trip-time, and launch flexibility provided by SEP resulted in significant benefits to the mission. Significant improvements in the efficiency and performance of SEP are underway. The resulting systems may provide substantial benefits to the SSE Roadmap’s planned missions to small bodies and the inner planets. Electric propulsion is both an enabling and enhancing technology for reaching a wide range of targets. The high specific impulse, or efficiency of electric propulsion system, allows direct trajectories to multiple targets that are chemically infeasible. The technology allows for rendezvous missions in place of fly-bys, and as planned in the Dawn mission can enable multiple destinations. This technology offers major performance gains, only moderate development risk, and has significant impact on the capabilities of new missions. Current plans include completion of the NASA’s Evolutionary Xenon Thruster (NEXT) Ion Propulsion System target at Flagship, New Frontiers and demanding Discovery missions. The GRC-led NEXT project was competitively selected to develop a nominal 40-cm gridded-ion electric propulsion system. 17,18 The objectives of this development were to improve upon the state-of-art NASA Solar Electric Propulsion Technology Readiness (NSTAR) system flown on Deep Space-1 to enable flagship class missions by achieving: Figure 7 NEXT thermal vacuum testing at JPL. lower specific mass higher Isp (4050 s) greater throughput (current estimates exceed 700 kg of xenon), greater power handling capability (6.9 kW), thrust (240 mN), and throttle range (12:1). The ion propulsion system components developed under the NEXT task include the ion thruster, the power-processing unit (PPU), the feed system, and a gimbal mechanism. The NEXT project is developing prototype-model (PM) fidelity , thrusters through Aerojet Corporation. In addition to the technical goals, the project has the goal of transitioning thruster7 American Institute of Aeronautics and Astronautics

manufacturing capability with predictable yields to an industrial source. A prototype model NEXT thruster passed qualification level environmental testing (Fig. 7). The prototype model thruster completed a short duration test in which overall ion engine performance was steady with no indication of performance degradation. As of June 30, 2010 the thruster achieved over 489-kg xenon throughput, 1.79 x 10^7 N-s of total impulse, and >30,000 hours at multiple throttle conditions. The NEXT wear test demonstrated the largest total impulse ever achieved by a gridded-ion thruster. ISPT funding for the thruster life test continues through FY12 with the aim of demonstrating up to 750 kg of xenon throughput. 19 In addition to the thruster, the system includes a powerprocessing unit (PPU). The PPU contains all the electronics to convert spacecraft power to the voltages and currents necessary to operate the thruster (Fig. 8). Six different power supplies are required to start and run the thruster with voltages reaching 1800 VDC and total power processing at 7 kW. L3 Communications designed and fabricated the NEXT Engineering Model (EM) PPU. After completing acceptance tests, the PPU was incorporated into the single-string integrated test. Environmental testing follows including electromagnetic interference/electromagnetic compatibility (EMI/EMC) testing to characterize the capability and emissions of the unit. A xenon feed system was developed, and is comprised of a single high-pressure assembly (HPA) with multiple lowpressure assemblies (LPA). The HPA regulates xenon flow Figure 8. NEXT Engineering Model PPU. from tank pressure to a controlled input pressure to the LPAs. Each LPA provides precise xenon flow control to the thruster main plenum, discharge cathode, or neutralizer cathode. The entire system constitutes the propellant management system (PMS). PMS development is complete and the system passed all performance and environmental objectives. The system is single fault tolerant, 50 percent lighter than the Dawn xenon feed system, and can regulate xenon flow to the various components to better than three percent accuracy. An engineering-model (EM) fidelity gimbal mechanism was developed that can articulate the thruster approximately 18 degrees in pitch and yaw (Fig. 9). The NEXT project team successfully demonstrated performance of the EM gimbal. The gimbal sub-system incorporates a design that improves specific mass over the Dawn gimbal. The gimbal was mated with the thruster, and successfully completed vibration testing first with a mass simulator and then with the NEXT PM thruster. The project team completed development of the digital control interface unit (DCIU) simulator. This allows communication and control of all system components during testing. A flight DCIU is the interface between the ion propulsion system and the spacecraft. Life models, system level tests, such as a multi-thruster plume interaction test, and various other supporting tests and activities are part of recent NEXT system developments. JPL, Aerojet and L3 Communications are providing major support for the project. The integrated NEXT system was tested in relevant Figure 9. NEXT Thruster and Gimbal Mechanism. space conditions as a complete string. With the exception of the PPU environmental tests, this brings the system to a TRL level of 6 and makes it a candidate for all upcoming mission opportunities. The life test demonstrated sufficient throughput for many science destinations of interest. The test plan is to continue into the coming years validating greater total impulse capability with the aim of demonstrating 750 kg of xenon throughput. Additional information on the NEXT system can be found in the NEXT Ion Propulsion System Information Summary in the New Frontiers and Discovery program libraries.19,20,21 In the original solicitation, NEXT was selected as an electric propulsion system for flagship missions. NEXT is the most capable electric propulsion system ever developed. A single NEXT thruster: 8 American Institute of Aeronautics and Astronautics

uses seven kilowatts of power, has an estimated propellant throughput capability of over 750 kg, has a lifetime of over 35,000 hours of full power operation, has a total impulse capability of approximately 30 million N-s, or about three times that of the SOA DAWN thrusters. This performance leads to benefits for a wide range of potential mission applications. The NEXT thruster has clear mission advantages for very challenging missions. For example, the Dawn Discovery Mission only operates one NSTAR thruster at a time, but requires a second thruster for throughput capability. For the same mission, the NEXT thruster could deliver mass, equivalent to doubling the science package, with only a single thruster. Reducing the number of thrusters reduces propulsion system complexity and spacecraft integration challenges. The missions that are improved through the use of the NEXT thruster are those requiring post-launch V, such as sample returns, highly inclined, or deep-space body rendezvous missions. The comet sample return mission was studied for several destinations because of its high priority within the New Frontiers mission category. In many cases, chemical propulsion was considered infeasible due to launch vehicle limitations. Specifically for Temple 1 in Ref. 22-23, the NSTAR thruster was able to complete the mission, but required large solar arrays and four or five thrusters to deliver the required payload. NEXT would be able to deliver 10 percent more total mass and require half the number of thrusters. NEXT can not only deliver larger payloads, but can reduce trip times and increase launch window flexibility. Chemical options exist for several missions of interest. However, the large payload requirements of flagship missions often require multiple gravity assists that both increase trip time and decrease the launch opportunities. In the recent Enceladus flagship mission study, the NEXT SEP option was able to deliver comparable payloads as the chemical alternative using a single Earth gravity assist. The chemical option for Enceladus required a Venus-Venus-Earth-Earth gravity-assist. This adds thermal requirements and increased the trip time by 57 months, from 7.5 to 12.25 years. The ISPT portfolio of the NEXT system, HIVHAC thruster, and subsystem improvements offer electric propulsion solutions for scientific missions previously unattainable. The systems are compatible with spacecraft designs that can inherently provide power for additional science instruments and faster data transfer rates. Scientists can open their options to highly inclined regions of space, sample return or multi-orbiter missions, or even deep-space rendezvous missions with more science and reduced trip times.


Advanced Materials Bipropellant Rocket (AMBR)

ISPT’s approach to the development of chemical propulsion technologies was the evolution of subcomponent technologies that still offered significant performance improvements. The main area of investment focused on items that would provide performance benefit with minimal risk with respect to the technology being incorporated into future fight systems. The primary investment within the advanced chemical propulsion technology area was the Advanced Materials Bipropellant Rocket (AMBR) engine (Fig. 10), which Figure 10. AMBR engine test article. was awarded, through a competitive process, to Aerojet Corporation in FY2006. The AMBR engine is a high temperature thruster addressing the cost and manufacturability challenges by using iridium coated rhenium chambers. It expands the operating environment to higher temperatures with the goal of achieving a seven-second increase in Specific Impulse (Isp) for NTO/N2H4. The project included the manufacture and hot-fire tests of a prototype engine demonstrating increased performance and validating new manufacturing techniques. The project also completed vibration (Fig. 11), shock, and long duration testing to raise the TRL to 6. Additional information is found in the AMBR information summary in the New Frontiers and Discovery program libraries.19,20,24 Figure 11. AMBR shock test in To initiate the effort, NASA Marshall Space Flight Center (NASA-MSFC) and X-direction. NASA Jet Propulsion Laboratory (NASA-JPL) first conducted mission-level and system-level studies to translate the target engine performance into spacecraft performance. Four conceptual missions were selected and used for the analyses based on the current scientific interest, launch vehicle capability, and trends in spacecraft size. 9 American Institute of Aeronautics and Astronautics

GTO to GEO 4800 kg ΔV for GEO insertion only ~1830 m/s Enceladus Orbiter (Titan aerocapture) 6620 kg. ΔV ~2400 m/s Europa Orbiter 2170 kg ΔV ~2600 m/s Mars Orbiter 2250 kg ΔV ~1860 m/s Applying the target AMBR engine specific impulse of 335 seconds (approximately seven seconds higher than the stateof-the art), the study shows a 23 percent payload gain for the Mars Orbiter mission and similar payload gains are also evident for the other missions as shown in Table 1. Table 1. Total Propulsion System Mass Reduction Performance testing was conducted on the AMBR engine in October 2008 and February 2009 and long duration testing in June 2009. The results show an Isp of 333 seconds―the highest ever achieved for hydrazine/NTO propellant combination (Fig. 12). This result represents a five second Isp gain over the HiPAT engine, a thrust increase to 140 lbf, mixture ratio of 1.1, chamber pressure of 195 psia, and oxidizer inlet pressure of 250 psia. 25 While these numbers differ from the original goal of 335 seconds Isp, 200 lbf thrust, mixture ratio of 1.2, and an inlet pressure of 400 psia, the single-iteration results are very encouraging. The test results show that the engine, as currently operating, can benefit many space applications. Typically, planetary and commercial spacecraft operate at pressures more comparable to the lower 250 psia propellant inlet pressure obtained in the test. While there are no planned follow-on activities, a possible next step would be to iterate the injector design, improve the cooling properties for a wider range of thermal stability, or using an Engineered El-form process for improved mechanical properties. The AMBR engine development 26 benefits missions with large propulsion maneuvers through the reduction of wet mass. The expectation for the AMBR engine is to have a 30 percent cost reduction in the combustion chamber manufacturing with an increase in performance. The mission mass benefits are dependent on the mission-required V, but are approximately the size of scientific instrument packages flown on previous missions. Ref. 3 shows the potential reduction in system mass, which could result in increased payload or margins, due to the increased specific power for multiple missions. These early study results were based on the initial AMBR target performance targets of 335 seconds Isp and 200 lbf thrust. If one uses target performance data and Figure 12. Notional operating box for AMBR engine. corresponding benefit analysis, the approximate mass benefit can be determined from Ref. 3. The system would deliver additional mass, over 50 kg; which equates to a potential increase in scientific payload by 100 percent.


Propulsion Component Technologies

ISPT’s approach to the development of chemical propulsion technologies was the evolution of subcomponent technologies that still offered significant performance improvements. The main area of investment focused on items that would provide performance benefit with minimal risk with respect to the technology being incorporated into future fight systems. Current technology investments include tasks to improve mixture ratio control, and reliable lightweight propellant tanks.

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Mixture Ratio (MR) control either reduces the residuals propellants carried or allows for additional extended mission operation otherwise lost due to an imbalance in the oxidizer-to-fuel ratio experienced during operation. Small investments were made to characterize balance flow meters, validate MR control to maximize precision, and determine the potential benefits of MR control. Two hot-fire tests of the system hardware (the Balanced Flow Meters, Fig. 13) were held during the AMBR testing and results are being compiled. The need for mixture ratio control (MRC) stems from the propulsion system margin that must be carried due to MR uncertainty. It is common for spacecraft with bi-propellant propulsion systems to reach end-of-life with residual oxidizer or fuel. Controlling the Figure 13. Example Balanced mixture ratio allows for either reduced Flow Meter Orifice residuals at launch, decreased mission risk by increasing propellant margin, or increased mission duration. Because the savings are directly proportional to the amount of propellant consumed, benefits are more significant on missions requiring large V Figure 14. Light-weight maneuvers. Typically, those missions already use bi-propellant systems. thin-liner COPV. Investments were made to evaluate manufacturing techniques for thin liner composite overwrap pressure vessels (COPV) (Fig. 14). The task evaluated liner bonding and welding techniques. The improved fabrication processes meet manufacturing recommendations and standards to minimize risk and increase yields for COPVs. This activity worked directly with members of NASA’s COPV working group, who implement the standard processes in future COPV efforts. The use of lightweight tanks produces a direct savings by reducing the propulsion system dry mass. Mass benefits approximate 2.5 percent of the propellant mass, or net tank mass savings of 50 percent over state-of-the-art titanium tanks. The mission benefits in advanced chemical propulsion are synergistic, and the cumulative effects have tremendous potential. The infusion of the individual subsystems separately provides reduced risk, or combined provides considerable payload mass benefits. Ref. 27 has a thorough description of the complete Advanced Chemical Propulsion effort that was concluded in 2009. 28 NASA’s In-Space Propulsion Technology project is investing in the Advanced Xenon Feed System (AXFS) for electric propulsion systems. The feed system is designed for an increased reliability with decrease in system mass, volume, and cost of SOA flight systems and comparable TRL 6 technology. The final development module, the pressure control module (PCM), was completed in 2007. The NRL completed functional and environmental testing of the VACCO PCM in September of 2008. Following the environmental testing, the PCM was integrated with the FCMs and an integrated AXFS with controller was delivered to the project. NASA GRC completed hot-fire testing of the AXFS (Fig. 15) with the HIVHAC Hall thruster successfully demonstrating hot-fire operation using Figure 15. AXFS mounted in hot-fire configuration. closed-loop control with downstream pressure feedback and with the Hall thruster discharge current. Follow-on testing determines the viability of the AXFS to perform single-stage, single module, control from high-pressure xenon directly to a thruster. The AXFS technology is ready for transition into a qualification program. It achieved its objective 29 by demonstrating accurate xenon control with significant system reduction in mass and volume through the use of integrated modules for lowcost control options and/or reliability beyond practical SOA technology implementation. The resultant feed system represents a dramatic improvement over the NSTAR flight feed system and represents an additional 70 percent reduction in mass, 50 percent reduction in footprint, and 50 percent reduction in cost over the baseline NEXT feed system at TRL 6. The project successfully completed the integrated system testing and advanced the modules to TRL 6.28

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In order to reduce costs for NASA science mission and leverage recent feed system flight experiences, JPL has developed a standard architecture for electric propulsion systems. One task under development is the maturation of the Digital Control Interface Unit (DCIU). The brassboard DCIU was designed and fabricated as shown in Fig. 16. The unit is undergoing functionality tests with flight software routines and operated with resistive loads. The feed system design approach is valid for either Hall or ion thruster systems and can utilize either commercial or NASA-specific components. Critical components of the simplified feed system were obtained for a demonstration test performed with an NSTAR-like laboratory-ion thruster. A single-string feed system was assembled using flight-like components consisting of a mechanical regulator and the proportional flow control valves, pressure transducers, and flow control devices necessary for a low-pressure assembly. The tests demonstrated operation over a representative throttle table and characterized system operation including flow stability and throttling performance. 30

Figure 16.

Standard Architecture Brassboard DCIU.


Electric Propulsion for Sample Return and Discovery-class Missions

ISPT is investing in Sample Return Propulsion technologies for applications such as Earth-Return Vehicles for large and small bodies. The first example leverages the development of a High-Voltage Hall Accelerator (HIVHAC) Hall thruster into a lower-cost electric propulsion systems. 31 HIVHAC is the first NASA electric propulsion thruster specifically designed as a low-cost electric propulsion option. It targets Discovery and New Frontiers missions and smaller mission classes. The HIVHAC thruster does not provide as high a maximum specific impulse as NEXT, but the higher thrust-to-power and lower power requirements are suited for the demands of some Discovery-class missions. Advancements in the HIVHAC thruster include a large throttle range from 0.3 – 3.5 kW allowing for a low power operation. It results in the potential for smaller solar arrays at cost savings, and a long-life capability to allow for greater total impulse with fewer thrusters. It allows for cost benefits with less complex systems. Wear tests of the NASA-103M.XL thruster validated and demonstrated the patented life extending innovation as a means to mitigate discharge channel erosion as a life limiting mechanism in Hall thrusters. Test priorities focused on the wear test of the laboratory thruster to validate the lifetime extending innovation to demonstrate throughput capabilities of the design. The thruster, shown in Fig. 17, operated in excess of 4700 hours (100 kg of xenon throughput). Components for two EM thrusters were fabricated. To date only one thruster is assembled for tests. Preliminary performance mapping of the EM thruster at various operating conditions was performed at NASA GRC. 32 The test sequence will include performance acceptance tests, environmental tests and a long duration test in FY10. Current plans include the design, fabrication and assembly of a full Hall propulsion system, but are pending final Figure 17. HIVHAC Thruster Engineering Model. approval to proceed. In addition to the thruster development, the HIVHAC project is evaluating PPU and xenon feed system XFS developments options that were sponsored by other projects but can apply directly to a HIVHAC system. The goal is to advance the TRL level of a HIVHAC Hall thruster propulsion system to level 6 in preparation for a first flight. The functional requirements of a HIVHAC PPU are operation over a power throttling range of 300 to 3,800 W, over a range of output voltages between 200 and 700 V, and output currents between 1.4 and 5 A as the input varies over a range of 80 to 160 V. A Performance map across these demanding conditions was generated for one candidate option.31 Beyond conventional feed system options, one option for feed systems that was demonstrated with the Hall thruster is the VACCO advanced xenon feed system. The ISPT project addresses the need for low-cost electric propulsion options. Studies23 indicate that a low-power Hall thruster is cost enabling, and enhances performance. Initial studies compared the HIVHAC thruster to SOA systems for Near12 American Institute of Aeronautics and Astronautics

Earth Object (NEO) sample returns, comet rendezvous, and the Dawn science mission. The HIVHAC thruster is expected to have both a greater throughput capability and a lower recurring cost than the SOA NSTAR thruster. For the NEO mission evaluated, the HIVHAC thruster system delivered over 30 percent more mass than the NSTAR system. The performance increase accompanied a cost savings of approximately 25 percent over the SOA NSTAR system. The Dawn mission was evaluated, and the expected HIVHAC Hall thruster delivered approximately 14 percent more mass at substantially lower cost than SOA, or decreasing the solar array provided equivalent performance at even greater mission cost savings. 33


Planetary Ascent Vehicle

For many years, NASA and the science community asked for a Mars Sample Return (MSR) mission. There were numerous studies to evaluate MSR mission architectures, technology needs and development plans, and top-level requirements. Because of the challenges, technologically and financially, of the MSR mission, NASA initiated a study to look at MSR propulsion technologies through the In-Space Propulsion Technology (ISPT) project office. 34 The objective of the ISPT project is to develop propulsion technologies that enhance or enable NASA science missions for the planetary science division by increasing performance while reducing cost, risk, and/or trip length. The largest propulsion risk element of the MSR mission is the Mars Ascent Vehicle (MAV). The development of a major subsystem of the Mars Sample Return mission requires a direct and in-depth analysis on technology sensitivities to the overall MSR architecture and the mission’s concept of operations (CONOPS). The MSR architecture dictates the physical and thermal environments, power requirements, and system interface of the MAV system. The current architecture (Fig. 18) for the MSR lander is to use the Mars Science Laboratory (MSL) entry, descent, and landing (EDL) system. The MSL EDL requires minor modifications such as. a larger parachute, additional propellant, and COPV propellant tanks to accommodate for a lander that will slightly exceed the lander mass of the MSL rover. Using the MSL sky crane concept places restrictions on the MAV system options.

Figure 18.

MSR baseline architecture.

Beyond the limitations of the EDL system, the MAV (Fig. 19) has specific requirements to deliver the orbiting sample (OS) in an orbit suitable for the Earth Return Vehicle (ERV). The basic requirements include: 500km +/- 100km circular orbit +/- 0.2o inclination Ability to launch from +/- 30o latitudes Accommodate ~5kg, 16cm diameter payload Continuous telemetry Storage for 90 Sols, potentially up to one year 13 American Institute of Aeronautics and Astronautics

The following technology development strategy is pre-decisional and is an approach under consideration. The strategy for technology development is the employment of an Integrated Product Development Team (IPDT) with updates as necessary to a technology steering community and host workshops as appropriate. The IPDT consists of members from ISPT project office, the Mars Exploration Program at JPL for intimate knowledge of the system interfaces, requirements, and sensitivities to the overall MSR mission, and NASA launch vehicle system design and test support. Management of the subsystem and system development is based in NASA’s ISPT project with lead systems engineering support to maintain interface controls and guide system integration activities. It is recommended that the initial tasks clearly define the requirements of component technology and calculate the potential return on investment. The definition of component level requirements and interfaces, and potential payoff are conducted through detailed collaborative engineering design, such as. JPL Team X, studies. It is anticipated that the component level developments would be competed; through the NASA Research Announcement (NRA) process. Figure 19. MAV Launch Information was solicited in December 2008 – January 2009 for enhancing and Platform enabling technologies for the Mars Ascent Vehicle. A task with The Aerospace Corporation was initiated in January of 2009 to investigate military technology applicability to the MAV. The responses and results were evaluated, ranked, and used to develop text for an NRA solicitation. The NRA was released in February 2010 with proposals due in May, 2010. Proposals are currently under review with an expected to award in early FY11. Some of the proposed technologies have potential to converge within the existing MSL EDL capabilities while other options may be enabled by enhancements to the EDL system. MAV propulsion concepts are ranging from solid, liquid, and hybrid systems with state-of-the-art or advanced propellants, and numerous subsystem architectures. These various concepts all have gross life-off mass (GLOM) implications to the MAV system. Vehicle sensitivities were established for the MAV through the minimization of the GLOM for a wide range of launch conditions, vehicle performance, and final payload orbit. Table 2. TSTO MAV Baseline Masses Mass(kg) 27.7 158.6 38.4 Description Motor casing, nozzle and interstage. Solid propellant stretched Star 17. Motor casing, avionics, payload attach structure. Solid 13B. propellant Star MAV Element Stage 1 Dry Stage 1 Propellant Stage 2 Dry Table 3. TSTO MAV Baseline Thrust and Isp Engine Parameter Stage 1 Thrust (N) Stage 1 Isp (s) Stage 1 Exit Area (m ) Stage 2 Thrust (N) Stage 2 Isp (s) Stage 2 Exit Area (m2)

Value 21576.8 285.7 0.032 6318.9 285.5 0.0093

Stage 2 Propellant Payload Fairing Payload

34.7 3.1 5.0

PLF jettisoned with first stage @ 200km. Sample and Container

Table 4. TSTO MAV Baseline Launch Site and Mission Assumptions Trajectory Assumption Latitude (degrees) Longitude (degrees) Launch Elevation Angle (degrees) Launch Azimuth (degrees) Target Circular Orbit Altitude (km) Target Orbit Inclination (degrees) Value 45.0 0.0 90.0, Vertical Optimal 500 45

Using the payload, dry masses, thrust and Isp assumptions listed in Table 2 and Table 3, an OTIS model was developed at NASA’s Glenn Research Center (GRC) along with a POST model by the team at NASA's Marshall Space Flight Center (MSFC). The existing TSTO baseline mission and launch site assumptions were also used as defined in Table 4.

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POST (Program to Optimize Simulated Trajectories) is a Table 5. Comparison of OTIS and POST Baseline TSTO Baseline Trajectories generalized point mass, discrete Comparison Parameter POST OTIS POST - OTIS parameter targeting and optimization program. POST Gross Lift-off Mass (kg) 271.75 272.63 -0.88 (0.32%) provides the capability to target and optimize point mass Launch Azimuth (deg) 88.77 87.56 1.21 (1.37%) trajectories for a powered or unpowered vehicle near an First Stage Propellant Load (kg) 149.31 153.85 -4.54 (-3.04%) arbitrary rotating, oblate planet. Second Stage Propellant Load (kg) 48.24 44.67 3.58 (7.41%) Optimal Trajectories by Implicit Simulation program (OTIS) is a Mission time (s) 656.67 659.09 -2.42 (-0.37%) general-purpose program, which is Maximum Dynamic Pressure (Pa) 11956.65 12290.78 -334.13 (-2.79 %) used to perform trajectory performance studies. A user can Maximum Dynamic Pressure (psf) 249.72 256.70 -6.98 (-2.79%) simulate a wide variety of vehicles such as aircraft, missiles, re-entry vehicles, ascent vehicles, satellites, and interplanetary vehicles. A detailed comparison of the POST and OTIS TSTO MAV baseline cases is provided in Table 5 to illustrate the agreement between both analysis programs.

Figure 20.

MAV GLOM vs. launch inclination Figure 21. MAV GLOM bs Launch inclination

The agreement of the GLOM between POST and OTIS is better than one percent. A large number of trades have been completed for various launch inclinations, elevations, and azimuths in addition to propulsion system trades of thrust and Isp, various system mass assumptions, and various final orbits. A notional launch configuration of the MAV prior to launch is illustrated in Fig. 20. An example trade of sensitivity of the MAV GLOM to launch site inclination is shown in Fig. 21. The analyses will be all completed in July, 2010 and will be available as a NASA Technical Memorandum.


Multi-Mission Earth Entry Vehicle

The Multi-Mission Earth Entry Vehicle (MMEEV) is a flexible design concept which can be optimized or tailored by any sample return mission, including lunar, asteroid, comet, and planetary (e.g. Mars), to meet that mission’s specific requirements. Based on the Mars Sample Return (MSR) EEV design, which due to planetary protection requirements, is designed to be the most reliable space vehicle ever flown, the MMEEV concept provides a logical foundation by which any sample return mission can build upon in optimizing an EEV design which meets their specific needs By leveraging common design elements, this approach could significantly reduce the risk and associated cost in development across all sample return missions, while also providing significant feed-forward risk reduction in the form of technology development, testing, and even flight experience, for an eventual MSR implementation.

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The current MMEEV parametric configuration and mass model, developed for a range of vehicle parameters, including vehicle diameter and payload mass/size, is presented in Fig. 22 (basic vehicle architecture) and Table 6 (range of parametric variables). Engineering estimates of MMEEV vehicle and trajectory performance are generated using the NASA Langley Research Center’s 6-DOF simulation software, Program to Optimize Simulated Trajectories (POST2). Fully integrated with POST2 are MMEEV specific models, including vehicle mass properties, aerodynamics, aerothermodynamics, TPS thickness/sizing, and a simplified 1-D impact analysis. Preliminary estimates for heat rates, heat loads, impact environments, and other vehicle and trajectory performance characteristics are provided across the MMEEV design trade space, which includes vehicle configuration and payload considerations, as well as a broad range of likely sample return mission entry conditions (e.g. entry velocity and entry flight path angle). Since each individual sample return mission may have a unique set of performance metrics Figure 22. Basic MMEEV architecture of highest interest, the goal is to provide a qualitative performance comparison across the specified trade space. From this, each sample return mission can select the most desirable design point from which to begin a more optimized design. Table 6. MMEEV parametric variables Parametric Variable Payload Vehicle Diameter Inertial Entry Velocity Inertial Entry Flight Path Angle Range 5 to 30 kg 0.5 to 2.5 m 10 to 16 km/s -5° to -25°

Continued development of the MMEEV models is planned to include: more sophisticated parametric configuration, including payload accommodation, models; higher fidelity impact dynamics model (e.g. Finite-Element Model); updated aerodynamics models based on ground (e.g. wind tunnel and ballistic range) testing as well as CFD analysis; and high fidelity TPS mass/thickness sizing models for additional candidate TPS materials. MMEEV performance studies will also continue, with the eventual integration of the MMEEV models into an EDL “Quicklook” Tool, a prototype EDL analysis tool, originally developed in support of ISPT aerocapture studies, and currently being developed to support mission studies to any celestial body with an atmosphere. Detailed studies show that to meet the stringent containment requirements for a Mars sample return mission, the MMEEV should possess particular design attributes. First, the vehicle aerodynamics must be very well understood. This means utilizing a shape with extensive analysis, testing, and flight experience. The vehicle aerodynamics must also be “selfrighting,” so it will quickly stabilize itself in a heatshield-forward orientation if the release from the ERV, a micrometeoroid impact, or some other anomaly, cause it to enter the atmosphere in any other orientation. Second, the heat shield TPS needs to be robust enough to ensure a high level of reliability for both nominal and off-nominal (such as MMOD impacts) environments. Third, the MMEEV has no parachute or other deployable drag device, since the reliability of such a device is several orders of magnitude less than the level likely required (i.e. the capsule would still need to be designed to survive and safely contain the sample after an Earth impact in the event of a failure of the drag device). The biggest challenge for any space vehicle, including the MMEEV, is to adequately prove the reliability of the components, subsystems, and the flight system as a whole. The current estimate to develop the EEV technology for MSR to TRL6 is approximately $41 million. This does not include a dedicated flight test, which many experts agree is needed to achieve the one-in-a-million system reliability, since the entry flight environment cannot be replicated in ground-based facilities. This is a fairly expensive flight test due to the high entry velocities that are required. One way to achieve a flight validation for little extra cost to NASA is to use the MMEEV design concept, or at least the major components of the design, in sample return missions likely to fly prior to MSR, such as New Frontiers or Discovery. NASA Headquarters managers and the In-Space Propulsion Technology (ISPT) team are pursuing this approach, but currently there are no manifested missions that are planning to use an MSR EEV design.

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Systems/Mission Analysis

Systems analysis is used during all phases of any propulsion hardware development. The systems analysis area serves two primary functions: 1) to help define the requirements for new technology development and the figures of merit to prioritize the return on investment, 2) to develop new tools to easily and accurately determine the mission benefits of new propulsion technologies allowing a more rapid infusion of the propulsion products. Systems analysis is critical prior to investing in technology development. In today’s environment, advanced technology must maintain its relevance through mission pull. Recent and ongoing systems analysis studies include Mars Earth Return Vehicle (MERV), main-belt asteroid sample return (Fig. 23), Discovery mission class electric propulsion study, and the MAV sensitivity trades. The second focus of the systems analysis project area is the development and maintenance of tools for the mission and systems analyses. Improved and updated tools are critical to clearly understand and quantify mission and system level impacts of advanced propulsion technologies. Having a common set of tools Figure 23. Notional Main-belt Asteroid increases confidence in the benefit of ISPT products both for Sample-Return spacecraft (shown without solar mission planners as well as for potential proposal reviewers. Tool array) development efforts were completed on the Low-Thrust Trajectory Tool (LTTT) and the Advanced Chemical Propulsion System (ACPS) tool. Low-thrust trajectory analyses are critical to the infusion of new electric propulsion technology. Low-thrust trajectory analysis is typically more complex than chemical propulsion solutions. It requires expertise to evaluate mission performance. Some of the heritage tools proved to be extremely valuable, but cannot perform direct optimization and require good initial guesses by the users. This leads to solutions difficult to verify quickly and independently. The ability to calculate the performance benefit of complex electric propulsion missions are intrinsic to the determination of propulsion system requirements. The ISPT office invested in multiple low-thrust trajectory tools that independently verify low thrust trajectories at various degrees of fidelity. The ability for the user community to rapidly and accurately access the mission level impacts of ISPT products can ease technology infusion. In addition to the tools currently available, there are on-going activities to develop an Aerocapture Quicklook tool to allow users an opportunity to quantify mission benefits of an aerocapture system including mass properties and geometry. Every effort will be made to have these tools validated, verified, and made publicly available. Instructions to obtain the tools currently available are provided on the ISPT project website: http://spaceflightsystems.grc.nasa.gov/Advanced/ScienceProject/ISPT/ The ISPT low-thrust trajectory tools suite includes Mystic, the Mission Analysis Low Thrust Optimization (MALTO) program, Copernicus, and Simulated N-body Analysis Program (SNAP). SNAP is a high fidelity propagator. MALTO is a medium fidelity tool for trajectory analysis and mission design. Copernicus is suitable for both low and high fidelity analyses as a generalized spacecraft trajectory design and optimization program. Mystic is a high fidelity tool capable of N-body analysis and is the primary tool used for trajectory design, analysis, and operations of the Dawn mission. While some of the tools are export controlled, the ISPT website does offer publicly available tools and includes instructions to request tools with limited distribution. The ISPT project team is planning a series of courses for training on the ISPT project tools. Copernicus training was held in November of 2009. Significant updates have also been made to Copernicus including a new release, version 2.3, in July 2010 with improvements of higher fidelity thruster models, power system modeling, output data generation, bug fixes etc. The Copernicus team also completed an update to the user’s guide to fully replace the user guide from version 1.0 released in 2006.


Future Plans and Conclusion

The future focus areas for ISPT are propulsion systems for sample return missions. Activity in these technology development areas increases in 2010. The direction focuses on: 1) Planetary Ascent Vehicles (PAV); 2) multi-mission technologies for Earth Entry Vehicles (MMEEV) needed for sample return missions; and 3) electric and chemical propulsion for Earth Return Vehicles (ERV), transfer stages, and low cost Discovery-class missions. These sample return missions are inherently propulsion intensive. Several of the ISPT technology areas may be involved in a single sample return mission. The mission may use EP for transfer to, and possibly back from, the destination. Chemical propulsion utilized for the ascent and descent to the surface. Aeroshells used for Earth re-entry and an aerocapture maneuver used to capture at the destination. 17 American Institute of Aeronautics and Astronautics

Studies in the three focus areas continue, and technology development activities are progressing. While the budget is tight for the next few years, the future is bright for ISPT. Future propulsion needs also include an electric propulsion system that is powered by a radioisotope-powered generator. Current EP systems are designed for widely varying input power levels to account for the spacecraft's motion around the solar system. If the vehicle does not need to rely on solar power, then the propulsion system is simpler and lighter. The system can be optimized around a known constant input power. Known future missions of interest for NASA and the science community, and those which are yet to be conceived, continue to demand propulsion systems with increasing performance and lower cost. This paper addressed how the ISPT project is developing propulsion technologies for NASA missions to address this demand. ISPT completes current developments to TRL 6 in the next year, and continues to support possible infusion. Among these is the NEXT electric propulsion system, which wraps-up PPU development and testing in 2010, but continues long-duration life testing for several more years. The NEXT system is available for all future mission opportunities. The AMBR engine reached TRL 6 with the completed development of the high temperature bi-propellant chemical thruster in 2009, and wraps-up the final reporting and documentation in early 2010. Finally, an aerocapture system comprised of a blunt body TPS system, the GN&C, sensors and the supporting models is to achieve its technology readiness by mid 2010. Regardless, if the mission requires electric propulsion, aerocapture, or a conventional chemical system, ISPT technology has the potential to provide significant mission benefits including reduced cost, risk, and trip times, while increasing the overall science capability and mission performance. Aerocapture and electric propulsion are frequently identified as enabling or enhancing technologies. ISPT continues to look for ways to reduce system level costs and enhance the infusion process. The cost of life testing of electric propulsion thrusters is one area where the savings are expected to be significant. Standardizing on common components or sub systems and utilizing modular stages for multiple missions is a way to reduce propulsion system costs. Performance enhancements tasks are anticipated in the area of electric propulsion through design and material improvements to achieve longer thruster life. Costs are addressed in the design process of the Hall thruster, and through modular design and shared hardware for NEXT and other electric propulsion systems. In the aerocapture area, the development plan for the rigid body aeroshell technologies follows a development plan proposed to the ST9 mission. In the chemical and component area, development in materials and engine designs continues to improve performance and reduce costs through advanced manufacturing techniques.

The results and findings presented here are based on work funded by the National Aeronautics and Space Administration (NASA), Science Mission Directorate (SMD). ISPT implements the project through task agreements with NASA centers, contracts with industry, and via grants with academic institutions. Implementing NASA centers include Ames Research Center (ARC), Dryden Flight Research Center (DRFC), Glenn Research Center (GRC), Goddard Space Flight Center (GSFC), Jet Propulsion Laboratory (JPL), Johnson Space Center (JSC), Langley Research Center (LaRC), and the Marshall Space Flight Center (MSFC). There are also numerous industry partners in the development of the ISPT products. The authors acknowledge the technical achievements by the respective NASA and contractor teams and the contributions of the respective technology area project managers.

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