The status of spacecraft bus and platform technology development under the NASA ISPT program

The In-Space Propulsion Technology (ISPT) program is developing spacecraft bus and platform technologies that will enable or enhance NASA robotic science missions. The ISPT program is currently developing technology in four areas that include Propulsion System Technologies (electric and chemical), Entry Vehicle Technologies (aerocapture and Earth entry vehicles), Spacecraft Bus and Sample Return Propulsion Technologies (components and ascent vehicles), and Systems/Mission Analysis. Three technologies are ready for near-term flight infusion: 1) the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance; 2) NASA's Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttle-able 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; and aerothermal effect models. Two component technologies being developed with flight infusion in mind are the Advanced Xenon Flow Control System and ultra-lightweight propellant tank technologies. Future directions for ISPT are technologies that relate to sample return missions and other spacecraft bus technology needs like: 1) Mars Ascent Vehicles (MAV); 2) multi-mission technologies for Earth Entry Vehicles (MMEEV); and 3) electric propulsion. These technologies are more vehicles and mission-focused, and present a different set of technology development and infusion steps beyond those previously implemented. The Systems/Mission Analysis area is focused on developing tools and assessing the application of propulsion and spacecraft bus technologies to a wide variety of mission concepts. These in-space propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, and sample return missions currently under consideration, as well as having broad applicability to potential Flagship missions. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness of in-space propulsion technologies in the areas of electric propulsion, Aerocapture, Earth entry vehicles, propulsion components, Mars ascent vehicle, and mission/systems analysis.


INTRODUCTION
Missions carried out for the Planetary Science Division (PSD) of NASA's Science Mission Directorate (SMD) seek to answer important science questions about our Solar System. To enable or significantly enhance PSD's future planetary science missions, the In-Space Propulsion Technology (ISPT) program is developing critical propulsion, entry vehicle, and other spacecraft and platform subsystem technologies. The ISPT program aims to develop technologies in the mid TRL range (TRL 3 to 6+ range) that have a reasonable chance of reaching maturity in 4-6 years. The objective is to achieve technology readiness level (TRL) 6 and reduce risk sufficiently for mission infusion. ISPT strongly emphasizes developing propulsion products for NASA flight missions that will be ultimately manufactured by industry and made equally available to all potential users for missions and proposals. ISPT focuses on the development of new enabling technologies that cannot be reasonably achieved within the cost or schedule constraints of mission development timelines.
ISPT's technology investment focus has evolved over time. Since 2001 when ISPT was started, ISPT has been developing in-space propulsion technologies that will enable and/or benefit near and mid-term NASA robotic science missions by significantly reducing cost, mass, risk, and/or travel times. ISPT technologies will help deliver spacecraft to PSD's future destinations of interest. In 2009, the ISPT program was tasked to start development of propulsionrelated technologies that would enable future sample return missions.
In March, of 2011, the Planetary Science Decadal Survey [1] was released and made many references to ISPT technologies that were initiated in the previous decade such as aerocapture, NEXT, an advanced chemical rocket called AMBR, and advancements made in the areas of astrodynamics, mission trajectory and planning tools. The Decadal Survey validated the technology investments ISPT has made over the last 10 years, and it provides guidance for ISPT's future technology investments.
The Decadal Survey Committee supported NASA developing a multi-mission technology investment program that will "preserve its focus on fundamental system capabilities rather than solely on individual technology tasks." They highlighted the NEXT system development as an example of this "integrated approach" of "advancement of solar electric propulsion systems to enable wide variety of new missions throughout the solar system." The Decadal Survey made a recommendation for "making similar equivalent systems investments" in advanced solar array technology and aerocapture. The Decadal Survey also discussed the importance of developing those system technologies to TRL 6.
The ISPT program is currently developing technology in four areas. These include Propulsion System Technologies (Electric and Chemical), Entry Vehicle Technologies (Aerocapture and Earth entry vehicles), Spacecraft Bus and Sample Return Propulsion Technologies (components and ascent vehicles), and Systems/Mission Analysis. These inspace propulsion technologies are applicable, and potentially enabling, for future NASA Discovery, New Frontiers, and sample return missions currently under consideration, as well as having broad applicability to potential Flagship missions. This paper describes the technology development in the areas of electric propulsion, Aerocapture, Earth entry vehicles, propulsion components, Mars ascent vehicle, and mission/systems analysis. For more background on ISPT, please see References [2, 3, 4, 5, 6, 7, 8, 9 and 10].
The ISPT program is currently funded through FY2014, so the focus this year will be concluding on-going efforts, documenting the accomplishments, and systematically closing-out the program. The ISPT program is making a concerted effort to adjust our remaining development activities to improve the infusion paths for ISPT developed technologies. We are being active in seeking out infusion opportunities for the ISPT developed technologies, and are exploring a number of paths to get our technologies out of NASA and into the commercial world. ISPT is also leading or co-leading several strategic planning activities that include a Technology Infusion Study, a TRL Assessment Study, and the formulation of development plans for Halleffect electric propulsion applicable to Discovery-class missions.

ELECTRIC PROPULSION TECHNOLOGIES
ISPT's propulsion system technology investments are currently being made in the area of Solar Electric Propulsion (SEP). SEP is both an enabling and enhancing technology for reaching a wide range of targets. Several key missions of interest: sample return, small body rendezvous, multirendezvous, Titan/Saturn System Mission (TSSM), Uranus Orbiter w/Probe, etc., require significant post-launch ΔV and therefore can benefit greatly from the use of electric propulsion. [11,12] High performance in-space propulsion can also enable launch vehicle step down; significantly reducing mission cost. [13] The performances of the electric propulsion systems allow direct trajectories to multiple targets that are otherwise infeasible using chemical propulsion. The technology allows for multiple rendezvous missions in place of fly-bys and, as planned in the Dawn mission, can enable multiple destinations. SEP offers major performance gains, moderate development risk, and significant impact on the capabilities of new missions. ISPT's approach to the development of chemical propulsion technologies is primarily the evolution of component technologies that still offer significant performance improvements relative to state-of-art technologies. The investments focus on items that would provide performance benefit with minimal risk with respect to the technology being incorporated into future fight systems.

NASA's Evolutionary Xenon Thruster (NEXT)
Current plans include completion of the NEXT Ion Propulsion System targeted at Flagship, New Frontiers and demanding Discovery missions. The GRC-led NEXT project was competitively selected to develop a nominal 40cm gridded-ion electric propulsion system. [2] The objectives of this development were 1) to improve upon the state-of-art (SOA) NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) system flown on Deep Space-1 and Dawn, 2) to enable flagship class missions by achieving the performance characteristics listed in Table 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 the Aerojet Corporation. In addition to the technical goals, the project has the goal of transitioning thruster-manufacturing capability with predictable yields to an industrial source. To demonstrate the performance and life of the NEXT thruster, a test program is underway. The NEXT PM thruster completed a short-duration test in which overall ion-engine performance was steady with no indication of performance degradation. A NEXT PM thruster has passed qualification level environmental testing. As of September 17, 2013 the Long Duration Test (LDT) of the NEXT engineering model (EM) thruster achieved over 902-kg xenon throughput, 34.9 x 10 6 N-s of total impulse, and over 50,197 hours at multiple throttle conditions ( Figure  1). The NEXT LDT wear test has set the world record by demonstrating the largest total impulse ever achieved by a gridded-ion thruster.
ISPT funding for the thruster life test continues into FY14.
The goal was to demonstrate thruster operation to 800 kg which, depending on the relative rates of the pit and groove erosion of the screen grid, may or may not represent the end-of-life condition for the NEXT thruster. The facility impact assessment review for the NEXT LDT assessed the impact of back-sputtered carbon on accelerator grid erosion rates. Analysis results showed that back-sputtered carbon reduced accelerator grid wear by less than four percent. Enhanced charge-exchange impact analysis determined approximately 8 to 10 percent increase for charge-exchange.
Repair of the NEXT LDT diagnostics equipment was initiated in November 2013. The equipment wore out over the eight-year test. The purpose was to collect end-of-test data with a fully functional diagnostic suite to compare with beginning-of-life data. The NEXT LDT diagnostics suite is shown in Figure 2. Equipment repaired is listed below: (1) Repair of the planar probe rake: to measure near-field current densities, provide an independently verified beam current, and assess potential changes in divergence (2) Repair Residual Gas Analyzer (RGA): to determine changes in background gas pressures that can affect thruster erosion and life (3) Replacement of the QCM: to determine changes in back-sputter rates (4) Replace Ion gage next to thruster: to confirm pressure changes near thruster (wall mounted gages still functional) (5) Replace coated facility windows: to photo-document thruster clearly, and image grid gap & groove erosion patterns (6) Improved imaging system lighting: to measure grid aperture diameters (downstream chamfers) & Discharge Cathode Assemble (DCA) orifices The final end-of-test characterization of the NEXT LDT began in December 2013 with the full diagnostics suite. The restart sequence went smoothly with no issues. The LDT is presently at 50,520 hours of high-voltage operation, corresponding to 909 kg of processed xenon and 35.1 MN-s of total impulse delivered. Data collection is expected to be completed by early to mid-January. This will be followed

Figure 2 ─ NEXT LDT Diagnostics
One of the challenges of developing the NEXT ion propulsion system was the development of the Engineering Model PPU. The demanding test program has flushed out a number of part problems that required extensive investigations to resolve and implement corrective actions. [15] It should be noted that such part problems are not unique in a technology development phase, and can still be experienced in the transition-to-flight hardware development phase. Technology development projects like NEXT are attempting to identify and mitigate these kinds of issues, before the PPU moves into a flight development phase.
One of the recent PPU part problems was the catastrophic failure of the multi-layer ceramic (MLC) capacitor in multiple beam power supplies. The investigation process utilized an extensive and knowledgeable team that investigated all branches of the fault tree. The corrective actions identified that a custom-built MLC had piezoelectric properties that made it susceptible to an oscillating current in the beam supply circuit. The corrective actions in this case were to replace the custom-build MLC capacitor as well as to eliminate the oscillating current. Recently, another part problem was uncovered, which manifested itself as a shorted diode. The diagnosis was that a void in the printed circuit board contributed to an overvoltage condition on the diode that caused it to short. This conclusion was confirmed with x-ray inspection of the printed circuit board. The corrective actions for the MLC capacitor issues were implemented in the EM PPU, and resolved the problems. The PPU has been refurbished to complete the planned test matrix, which includes PPUthruster integration testing, electromagnetic interference testing, and breadboard digital control interface unit (DCIU) integration tests. The NEXT PPU is shown in Figure 3.
An area in which further NEXT work has been needed is that of precise plume, particle, and field characterization. A non-reimbursable Space Act Agreement (SAA) was drafted by NASA and The Aerospace Corporation (TAC) to establish a collaborative measurement program intended to examine the plume, particle, and field environments of the latest generation NASA ion propulsion technology. A series of measurements has been completed to verify basic characteristics of NEXT operation and expand on the available public-domain and internal databases regarding NASA technology and its potential use on non-NASA spacecraft systems. [16] Figure 4 shows the NEXT thruster installed in the vacuum facility at TAC. Among the work elements planned are in-depth EMI/EMC, plume particle and plasma probe, optical emission and laser diagnostic measurements. This work is of considerable relevance to future spacecraft integration of the subject thrusters.
The NEXT evaluation at Aerospace also includes measurement of ion beam flux and divergence, charge state ratios, charge exchange ion flux, plume optical emission spectrum and absolute flux, radio frequency and microwave absolute emission spectrum plus timedomain emissions, carrier wave attenuation and phase effects, plume erosion and molybdenum contamination effects, absolute thrust and thrust correction factors. Plume characterization tests with the NEXT ion thruster were performed using the EM and PM thrusters. Examinations of the beam current density and xenon charge-state distribution as functions of position on  the accelerator grid have been completed. [17] The angular dependence of beam current was measured at intermediate and far-field distances to assist with plume modeling and to evaluate the thrust loss due to beam divergence. Thrust correction factors were derived from the data. [17] Transmission and phase noise measurements were made through the plume of an EM NEXT ion thruster. [18] Attenuation measurements were taken at multiple operating points at frequencies between 1 and 18 GHz. Attenuation was observed between 1 and 3 GHz and scaled with plasma density. [18] Phase noise spectra were also taken. Direct thrust measurements have been made on the NEXT PM ion thruster using a standard pendulum style thrust stand constructed specifically for this application. [19,20] Values have been obtained for the full 40-level throttle table as well as for a few off-nominal operating conditions. [19,20] A particle-based model with a Monte Carlo collision model has been developed by Wright State University (WSU) to study the plasma inside the discharge model of the generic ion thruster. This model tracks five major particle types inside the discharge chamber in detail: xenon neutrals, singly and doubly charged xenon ions, secondary electrons and primary electrons. [21] Both electric and magnetic field effects are included in the calculation of the charged particle's motion. Validation of this computational model has been made with comparisons to the NSTAR discharge chamber. Comparison of numerical simulation results with experimental measurements was found to have good agreement. [21] The model has been applied to the NEXT discharge chamber design at multiple thruster operating conditions. [22,23,24,25,26] A Close-out Review of NEXT NRA Phase 2 was conducted in 2012. Feedback from the customer community was provided. The primary remaining concern is completion of the development of the NEXT power processing unit. 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. [27,28,14,29,30]

Electric Propulsion for Sample Return and Discovery-class Missions
ISPT is investing propulsion technologies for applications to low-cost Discovery-class missions and Earth-Return Vehicles for large and small bodies. The first example leverages the development of a High-Voltage Hall Accelerator (HIVHAC) thruster into a lower-cost electric propulsion system. [3,30] 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 and sample return applications.
Advancements in the HIVHAC thruster include a large throttle range from 0.3-3.9kW 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. The benefits include cost savings with a reduced part count and less-complex lowercost propulsion system.
Wear tests of the NASA-103M.XL thruster validated and demonstrated a means to mitigate discharge channel erosion as a life-limiting mechanism in Hall thrusters. The thruster, operated in excess of 5500 hours (115 kg of xenon throughput) at a higher specific impulse (thruster operating voltage) as compared to SOA Hall thrusters.
Components for two Engineering Development Units (EDU-1) thrusters were designed and fabricated. Preliminary performance mapping of the EDU-1 thruster at various operating conditions was performed at NASA Glenn Research Center (GRC) as shown in Figures 5 and 6. [3,31] The EDU-1 thruster hardware was operated in vacuum test environments for operations and performance assessments. The results indicated that several design changes were needed to resolve problems with thermal design, boronnitride advancement mechanisms, magnetic topology, and high-voltage isolation. A list of rework items was compiled and design corrections were identified and evaluated by either analysis and/or test.
The design improvements were implemented in a reworked engineering model design, which is designated as EDU-2. Vacuum Facility 12 (VF-12) was used to conduct the official performance acceptance test (PAT), given the pumping speed and resulting vacuum chamber background pressure. The results indicate that performance and operational requirements met expectations, with significant improvement to the thermal margins of key components. Vibration testing was completed with performance tests conducted both before and after vibration tests. The HIVHAC EMR thruster was successfully vibrated to approximately 11.5 g in three axes, which were consistent with the specifications used to qualify the NASA Evolutionary Xenon Thruster ion thruster. Preliminary visual inspection of the thruster indicates that the thruster passed the vibration testing with no visual damage evident and no change in thruster performance was measured.
Single String Integration Test of the NASA HIVHAC Hall thruster was conducted in NASA GRC VF5. During the test a number of plasma diagnostics were implemented to study the effect of varying the facility background pressure on thruster operation. These diagnostics include thrust stand, Faraday probe, ExB probe, and retarding potential analyzer. The test results indicated a rise in thrust and discharge current with background pressure. There was a decrease in ion energy per charge, an increase in multiply-charged species production, a decrease in plume divergence, and a decrease in ion beam current with increasing background pressure. A simplified ingestion model was applied to determine the maximum acceptable background pressure for thrust measurement. The maximum acceptable ingestion percentage was found to be around one percent. Examination of the diagnostics results suggest the ionization and acceleration zones of the thruster were shifting upstream with increasing background pressure. [32] The HIVHAC EDU-2 thruster advancement mechanism on inner and outer boron nitrate channels was successfully demonstrated immediately after thruster hot-fire operation in VF-12. The advancement mechanism showed smooth advancement of both channels as a full qualification vibration test post-test validation of the mechanism. The actuation test was conducted immediately following thruster shutdown, assuring high-temperature conditions within the thruster. In the future, the test sequence will include performance acceptance tests, the remaining thermal vacuum environmental tests, and a long duration wear test. Current plans include the design, fabrication and assembly of a full Hall propulsion system that can meet a variety of Discovery and Earth Return Vehicle needs.
In addition to the thruster development, the HIVHAC project is evaluating power processing unit (PPU) and xenon feed system (XFS) development options. These were developed under other efforts, but can apply directly to a Hall Propulsion system. The goal is to advance the TRL level of key components of a Hall propulsion system (thruster, PPU/DCIU, feed system) to level 6 in preparation for a first flight.
The functional requirements of a HIVHAC PPU (Figure 4) 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 [3,30]   substantially lower cost than SOA, or decreasing the solar array provided equivalent performance at even greater mission cost savings. [3,31] The

Multi-Mission Earth Entry Vehicle (MMEEV)
The Entry Vehicle Technologies (EVT) project has been developing the Entry, Descent, and Landing (EDL) trade space and system analysis tools for the Multi-Mission Earth Entry Vehicle (MMEEV). The Multi-Mission Earth Entry Vehicle (MMEEV) is a flexible design concept. It 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. The Mars Sample Return (MSR) Earth Entry Vehicle (EEV) design, due to planetary protection requirements, is designed to be the most reliable space vehicle ever flown. Other EEVs have different and less restrictive requirements. The MMEEV approach provides an effective foundation for many sample return missions. By leveraging common design elements, this approach can significantly reduce the risk and associated cost in development across all sample return missions. [4] Detailed studies show that to meet the stringent containment requirements for a Mars sample return mission, the MMEEV should possess three 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 "self-righting." It needs to quickly stabilize itself in a heatshield-forward orientation in the event of perturbations. Second, the heat shield TPS needs to be robust and have a high level of reliability derived from developmental testing and flight test heritage for both nominal and off-nominal (such as MMOD impacts) environments. The reliability requirement has traditionally implied the use of heritage carbon phenolic TPS, which is limited in supply and manufacturability. NASA has held two workshops, in 2010 and 2012, to assess the availability of carbon phenolic and possible replacement materials. The forward path will depend on funding availability, and is not yet defined. Third, the MMEEV has no limited-reliability system, such as a parachute or other deployable drag device that could fail upon entry.
While these design attributes were developed from MSR applications, they also work well for many other EEV missions providing an optimum balance of cost, risk, and payload requirements. The current MMEEV parametric configuration is presented in Figure 8 (basic vehicle architecture), and Table 2 (parametric variables). Because each individual sample return mission may have a unique set of performance metrics of highest interest, the goal is to provide a qualitative performance comparison across a specified trade space. Each sample return mission can then select the most desirable design point to begin a more optimized design.
MMEEV performance studies were integrated into the "Multi-Mission Systems Analysis for Planetary Entry" (M-SAPE) Tool. This is a low-fidelity EDL analysis tool that provides mission designers the ability to perform extensive trade analyses very early in the design process. The M-SAPE tool does contain low-, mid-, and high-fidelity models. However the M-SAPE tool contains at least lowfidelity models for vehicle geometry, flight mechanics, A parametric preliminary thermal soak model was developed at NASA-Ames to define the thermal environment of the returned sample canister after the vehicle experiences the heat pulse and subsequent impact and waits to be recovered. [40] Samples from various comets, asteroids, and planets may have differing thermal requirements and impact g-load requirements to preserve the science return. Active thermal control is considered for applications with extreme thermal requirements, though not currently modeled in M-SAPE. Feeding into the thermal soak model is actual test data on impact foams. Several closed-cell foam candidates have been impact tested [41] at NASA-Langley. Thermal conductivity testing of these impact foams have been completed at Southern Research Institute (SRI) in FY-13 in both the virgin and post-impact condition [42]. Results indicate that the effect of impact is to increase the foam's density by ~100 percent while not significantly affecting thermal conductivity. This result is believed to be due to the venting of the manufacturing gases and replacement with air during impact.. These parameters for various materials will be part of the closed-loop M-SAPE analysis capability. Finally, usable subsonic center of gravity limits and an aerodynamic model for an array of MMEEV designs will be established via spin tunnel testing at the NASA-Langley 20-FT Vertical Spin Tunnel ( Figure  9). This type of subsonic test, combined with aerodynamic parameter identification (PID) techniques provides unique dynamic aerodynamic results without the interference of a sting. Dynamically scaled vehicle models with various aftbody configurations (i.e., payload sizes) were tested in 2013. A video-based motion tracking system installed in the 20-FT VST provided location and attitude of the freeflying vehicle at 150 Hz. Post-processing PID techniques were applied to provide dynamic and static aerodynamic coefficients for the various vehicle geometries and mass characteristics.
The goal of this work is to provide tools for evaluating MMEEV designs and trade space in support of mission proposal development and preliminary design. Development and use of the capabilities will enable New Frontiers and Discovery missions to cost-effectively perform proposal development and preliminary design analyses. Although Science Mission Directorate management and the ISPT project team favor this approach, there are currently no manifested missions that use the MMEEV design.

Aerocapture
Aerocapture is the process of using the destination planet's atmosphere to decelerate the vehicle from aerodynamic forces to capture the spacecraft in a desirable orbit from a hyperbolic trajectory. Aerobraking is a gradual series of passes through the upper atmosphere (once a spacecraft is propulsively captured into a high ellipse) to reduce orbital energy. Aerocapture is capable of much larger decelerations and maximizes the benefit from the atmosphere by capturing a useful science orbit in a single pass. During Aerocapture, as a spacecraft flies at a lower altitude where the atmosphere is denser, the resultant drag and heating is much higher than for aerobraking. Aerocapture heat loads are generally much higher than for nominal entry, descent, and landing, atmospheric entries. An aeroshell is required to both protect the spacecraft from the environment, and provide an aerodynamic surface for control during the pass. 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. Aerocapture significantly reduces the chemical propulsion requirements of an orbit capture.
Aerocapture has shown repeatedly in detailed analyses to be an enabling or strongly enhancing technology for several targets with atmospheres. ISPT has been investing to mature Aerocapture subsystems since 2001. [4] The aerocapture project team continues to mature aerocapture components. A rapid aerocapture analysis tool has been developed and made available to the user community. The TPS materials developed through ISPT enhance a wide range of missions by reducing the mass of entry vehicles. All of the component technologies for an aerocapture vehicle are currently at TRL 5-6. This assessment of technology readiness is detailed in Reference [43]. The structures and TPS subsystems as well as the aerodynamic and aerothermodynamic tools and methods can be applied to planetary entry, descent, and landing or aerocapture applications.
As flight aeroshells become larger (over 3 meters in diameter), it is more difficult to hand-pack them, as was done with the Apollo capsules and every successful Mars heatshield before the Mars Science Laboratory (MSL). ARA developed a modular TPS approach, in which large modules of TPS are pre-packed into honeycomb, cured, and precisely milled to fit the aeroshell structure. Because SRAM and PhenCarb are somewhat elastic, a small number of modules (less than ten) are needed to cover the aeroshell (compared to tens of PICA segments used on MSL). Gaps between modules were packed with the same ablator and cured. The result is a seamless heatshield. To mature this approach, ISPT has manufactured a 2.65-meter (Discovery-class size) low-density Aeroshell ( Figure 10). The TPS is applied to the ATK 400 ˚C bondline structure. Lawrence Livermore National Laboratory (LLNL) scientists performed nondestructive Computed Tomography (CT) scan of the completed aeroshell to mature the diagnostics methodology and verify the manufacturing methods. Figure 11 shows the 2.65-m Aeroshell with LLNL LINAC CT scan radiation source in the background. Analysis of the results will be completed in FY-14. Manufacturing at this scale will mature the high-temperature aeroshell system to TRL 5.  Figure 12 shows an impacted SRAM backshell material before, during, and after arcjet testing. Micrometeoroid cavity volumes pre-and post-test can be compared using laser and CT scanning techniques. The testing was completed in August 2012 with the results provided in Reference [44]. Additional information on aerocapture technology developments can be found in the Discovery Program library [27], and in References [45, 46, 47, 48, 49 and 50].
In order to develop and assess the ability of new technologies to enable science missions as well as to define requirements for new technologies, a series of mission studies were performed by the EVT project. A mission study for a direct ballistic entry probe missions to Saturn and Venus were performed in FY-12 using a 45-deg spherecone rigid aeroshell design over a range of entry mass and velocities, aeroshell diameter, and entry flight path angles. Analysis of the results shows the existence of a range of "critical" ballistic coefficients beyond which the steepest possible entries are determined by the pressure limit of the TPS material rather than the deceleration load limit.
In FY-13, a Uranus mission study was performed. The objectives of the Uranus mission study were to: 1) Establish a range of probe atmospheric entry environments based on the Uranus Flagship mission outlined in the Planetary Science Decadal Survey for two launch windows: Years  Beyond the limitations of the EDL system, the MAV has specific requirements to deliver the orbiting sample (OS) into an orbit suitable for the Earth Return Vehicle (ERV) to rendezvous with and capture the sample. Many of the subsystem requirements of the MAV are still to be determined, with many to be defined by the prime integrator during development. However, the driving top-level requirements of the MAV are described in References [4,54].
Another challenge for the MAV is to meet the environmental requirements for the mission. The environmental requirements include the Earth launch, transit within the cruise stage, the Mars EDL, and finally a long surface stay on Mars. The environments anticipated to influence the system design are the vacuum environment during cruise, the 15g quasi-static lateral load during EDL, and the diurnal temperature cycling, as low as -99°C during the surface stay. The thermal requirements necessitate a thermal enclosure or "igloo" in order to maintain practical lander power requirements. A detailed set of requirements and system design standards and guidelines has been established for all study participants to ensure comparable system capability and margins. [55] Through the NASA Research Announcement (NRA) process, the ISPT program solicited MAV system designs and plans to initiate propulsion system development. Multiple contractors were selected to proceed in October of 2010 and efforts were initiated in February 2011. Awards were made to ATK, Lockheed Martin, and Northrop Grumman to develop MAV concepts using solid-solid, solid-liquid, and liquid-liquid 1st and 2nd stage propulsion systems respectively. During the NRA efforts, the contractors completed Principal Investigator led collaborative engineering designs of the MAV and will begin contract options to develop the required technologies in early FY12. Additionally, Firestar Technologies is working, under an SBIR, to develop a Nitrous Oxide Fuel Blend propulsion system applicable to the MAV. [56] The results of the industry efforts indicate that while technology development remains, there are multiple paths to meet performance and requirements of the Mars Ascent Vehicle. The industry efforts and designs are documented in four 2012 IEEE Aerospace Conference papers. [54,57,58,59] The baseline MAV concept design is shown in Figure 13. The Government baseline design is pre-decisional and for understanding design trades and sensitivities, and does not represent any concept selection.
NASA performed system design studies with the Jet Propulsion Laboratory's (JPL) Team-X and GRC's COMPASS teams. [50] The collaborative designs included a system level optimization using the industry designs and an internal "leveled" design to allow comparison of system mass, complexity, and maturity. The trades included the MAV support systems and lander impacts to minimize the total landed mass. The preliminary results of the studies indicate that the baseline solid-solid system appears to offer the lowest mass solution, but it may have challenges achieving the required orbit dispersion accuracies The solidliquid option has a slightly higher mass, imposing more thermal requirements on the lander, but can reduce dispersion errors. The liquid-liquid option has the highest mass growth potential due to its mass fraction relative to a solid motor, but requires the least lander resources and has very tight dispersions. The preliminary NOFBx system evaluation indicates it may be a competitive option, but is unlikely to offer a single stage to orbit solution with a lower mass than the two-stage solid.
Each of the MAV concepts was evaluated for risk and technology maturation and was recommended, primarily in the propulsion elements. The MAV NRA work initially focused on the key risks of the individual propulsion systems at the component level. The MAV project team expects to achieve a milestone in late FY12 to address the key risks of each option and determine the final viability of various concepts. If the most promising MAV concept(s) is viable with respect to mass, volume, and risks, an integrated propulsion stage demonstration would be the next step. If sufficient risk can be reduced through the technology development activities, the final step would be an engineering model MAV development with an objective of a vehicle terrestrial flight demonstration. However, the MAV technology development for the most part is on hold pending the completion of the Mars Program Planning Group (MPPG) activities. Some on-going MAV related studies are being completed, and a long-lead activity to assess the aging of solid rocket motor propellants under Mars environmental conditions (landing shocks and thermal cycling) will proceed until future decisions determine the future MSR architecture and MAV requirements ( Figure  14). New in 2013, NASA initiated the development of a new propellant formulation activity with ATK. NASA and ATK traded a wide range of solid motor propellant formulation options to increase the mechanical properties at low temperatures for hydroxyl terminated polybutadiene (HTPB) and to increase performance for carboxyl terminated polybutadiene (CTPB) formulations.
Both HTPB and CTPB propellant options have been found to meet the requirements of the MAV. The higher risk HTPB formulation began a long duration aging test in November of 2013. The propellant aging facility is shown in Figure  14. The propellant will undergo 18 months of testing including an initial simulation of the Mars transit at high vacuum followed by Mars surface environment of surface pressure and temperature conditions. The sample will be removed at 6-month intervals for performance and mechanical property testing.

Ultra-lightweight Tank Technology (ULTT)
ISPT invests in the evolution of component technologies that offer significant performance improvements without increasing system level risk. The ISPT Program invested in ultra-lightweight tank technology (ULTT) led by JPL. The ULTT efforts in the past focused on manufacturability and non-destructive evaluation of the lightweight tanks. The tank effort continues to validate defect-detection techniques to maintain NASA standard compliance for ultra-thin wall tanks. The follow-on potential is to develop and qualify positive expulsive ultra-lightweight tanks specifically for the MSL Sky Crane. The Mars Science Laboratory (MSL) Sky Crane, with large propellant tanks, is shown in Figure  15.
The Sky Crane tanks offer mass savings on the order of 24 kg. This is dependent on the final tank wall thickness. The mass reduction would increase the landed mass capability of Sky Crane for a relatively low cost per kg. The Sky Crane Entry Descent Lander (EDL) system could be used again in a future Mars Sample Return (MSR) mission. Both are highly mass constrained. While this particular tank design will be qualified for the Sky Crane application (Figure 16), the ultra-lightweight technology will be applicable for a wide range of future science missions. Propulsion tanks remain the highest dry-mass reduction potential within chemical propulsion systems. This technology would significantly push the state-of-the-art with the promise of a 2X improvement over conventional tank designs.
The development effort is divided into two main tasks: a Non-Destructive Inspection (NDI) task and the ultralightweight tank design/manufacturing/testing task. The NDI task completed an initial assessment of several NDI techniques, such as eddy-current and surface wave ultrasonic techniques. The results from the tests indicate that these techniques are adequate to find cracks as small as 0.003 inches in the titanium lining. The objective for the NDI task is to establish the crack size that can be detected consistently using these new methods. The ultra-lightweight tank development task would incorporate the NDI technique in the manufacturing and qualification of the new tank.
In order for the tank design to be a success, the approach must demonstrate "safe life." Safe life for non-toxic materials requires proving a design will leak-before-burst. Safe life for toxic liquids, like hydrazine, is more stringent. The NDI technique must be able to detect small cracks in the thin liners, then the NDI results need to be verified, by test, that worst-case crack growth will not grow to failure. An automated eddy current inspection technique has been developed and tested for the detection of small fatigue cracks in thin titanium panels. In this work, a commercially available eddy current probe was deployed on a motion control system in order to obtain high-resolution eddy current C-Scan images of 48 individual samples.
A data processing technique was developed and deployed to enhance the flaw response and automate detection of cracklike indications in the samples. The noise floor of the inspection technique was calculated as three times the standard deviation of the eddy current response in the two unflawed control samples. The remaining 46 samples had fatigue cracks with estimated depths varying between 0.0021 and 0.0067 inches. All the fatigue crack panels registered crack-like indications at a level greater than three times the calculated noise floor. The improved detection capability promises to find 0.003 inch cracks reliably, which represents a 2x improvement over SOA detection techniques.
The new technique enables the manufacturing of compositeoverwrapped titanium tanks with an anticipated 48 percent mass savings as compared to the heritage Sky Crane tank design. In parallel the ultra-lightweight development work will be completed through a contracted effort with ATK, the suppliers of the MSL tanks. The work will be divided into several phases: design, manufacturing and acceptance/qualification tests. The test phase will include cyclic testing of the flawed liner tank design to demonstrate leak-before-burst and safe life requirements. The design phase led to the PDR, which was held in February 2012, and activities are progressing towards CRD in December 2013, which is the likely stopping point of this development effort unless there is interest in a mission user to co-fund the manufacturing and acceptance/qualification test phases.  Systems analysis is critical prior to investing in technology development. In today's environment, advanced technology must maintain its relevance through mission pull. Systems analysis is used to identify the future mission needs for decadal missions and Discovery design reference mission (DRMs). The mission studies identify technology gaps, and are used to quantify mission benefits at the system level. This allows studies to guide the investments and define metrics for the technology advancements. Recent systems analysis efforts include quantitative assessment of higher specific impulse Hall thrusters [60], higher thrust-to-power gridded-ion engines, and evaluation of monopropellant system anomalies to assess failure modes and potential mitigation options. In addition to informing project decisions, the mission design studies provide an opportunity to work with the science and user community.
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 allow the potential mission users to quantify the benefits and understand implementation of new technologies. A common set of tools increases confidence in the benefit of ISPT products both for mission planners as well as for potential proposal reviewers. The ISPT program awarded three Astrodynamics research grants in 2013. The three awards are research and tool development for outer planet moon tours, low-energy trajectories, and a guess tool to initial Mystic trajectory optimization. The awards were provided to University of Texas-Austin, Purdue University and University of California at Irvine respectively. The efforts were solicited through the SMD ROSES call, started in the spring of 2013, and will conclude in the spring of 2014. The resulting products will be made available to the entire community when complete. Figure 17 is a screenshot from Purdue's low-energy trajectory tool that will interface with GMAT led by NASA GSFC.

Figure 17 ─ Low-energy trajectory tool screenshot
ISPT aerocapture project released its Aerocapture Quicklook Tool, formally the multidisciplinary tool for Systems Analysis of Planetary EDL (SAPE). [39] SAPE is a Python based multidisciplinary analysis tool for entry, decent, and landing (EDL) at Venus, Earth, Mars, Jupiter, Saturn, Uranus, Neptune, and Titan. The purpose of the SAPE tool is to provide a method of rapid assessment of aerocapture or EDL system performance, characteristics, and requirements. SAPE includes integrated analysis modules for geometry, trajectory, aerodynamics, aerothermal, thermal protection system, and structural sizing. For aerocapture and EDL system designs, systems analysis teams include systems engineers and disciplinary specific experts in flight mechanics, aerodynamics, aerothermodynamics, structural analysis, and thermal protection systems. The systems analysis process may take from several weeks to years to complete. While the role of discipline experts cannot be replaced by any tool, the integrated capabilities of SAPE can automate and streamline several parts of the analysis process significantly reducing the time and cost for preliminary assessment. SAPE continues to receive investment for assessment of Earth Entry Vehicles. [4]

CONCLUSION
The ISPT program is currently funded through FY2014. As a result, the focus this year will be concluding on-going efforts, documenting accomplishments, and systematically closing-out the program.
In 2013 and 2014 the NEXT team wraps-up long-duration testing and power processing unit development, and completes closeout documentation. In 2013 HIVHAC completed a test in GRC's VF-5 facility with the same diagnostics suite used for a test of the BPT-4000. This test will help to understand facility effects on Hall thruster testing. The VF-5 facility is undergoing improvements in 2014 to boost its already world-class capabilities, and the program hopes to get the HIVHAC thruster back into the improved facility for another test sequence. HIVHAC will conclude its FY14 activities with verification test of its life extension mechanism, magnetically shielded design iteration, and continued support of the CPE PPU SBIR development. The Ultra-light weight tank (ULTT) will conclude its development at a CDR in January of 2014. The The ISPT program is making a concerted effort to adjust the remaining development activities to improve the infusion paths for ISPT developed technologies. The program team is actively seeking out infusion opportunities for the ISPT developed technologies, and is exploring a number of paths to get the technologies out of NASA and into the commercial world. ISPT will continue to support mission infusion. ISPT is leading or co-leading several strategic planning activities that include a Technology Infusion Study, a TRL Assessment Study, and the formulation of development plans for Hall-effect electric propulsion applicable to Discovery-class missions.
The planetary decadal survey identified the need for future work in propulsion, entry vehicles, and spacecraft bus and other platform technologies. ISPT will continue to work with the PSD to identify the propulsion technologies that will be pursued in the future. ISPT will continue to look for ways to reduce system level costs and enhance the infusion process. If the ISPT program concludes in FY2014, the Space Science Projects Office at NASA Glenn will be available to users who have interest in the ISPT-developed technologies. 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.