ALBERT

Description: No abstract available

Description: This paper discusses the methodology used to model common cause failures of thrusters on the International Space Station (ISS) Visiting Vehicles. The ISS Visiting Vehicles each have as many as 32 thrusters, whose redundancy and similar design make them susceptible to common cause failures. The Global Alpha Model (as described in NUREG/CR-5485) can be used to represent the system common cause contribution, but NUREG/CR-5496 supplies global alpha parameters for groups only up to size six. Because of the large number of redundant thrusters on each vehicle, regression is used to determine parameter values for groups of size larger than six. An additional challenge is that Visiting Vehicle thruster failures must occur in specific combinations in order to fail the propulsion system; not all failure groups of a certain size are critical.

Description: Gaseous and liquid oxygen are extremely reactive materials used in bipropellant propulsion systems. Both flight and ground oxygen systems require a high level of cleanliness to support engine performance, testing, and prevent mishaps. Solvents used to clean and verify the cleanliness of oxygen systems and supporting test hardware must be compatible with the system's materials of construction and effective at removing or reducing expected contaminants to an acceptable level. This paper will define the philosophy and test approach used for evaluating replacement solvents for the current Marshall Space Flight Center/Stennis Space Center baseline HCFC225 material that will no longer be available for purchase after 2014. MSFC/SSC applications in cleaning / sampling oxygen propulsion components, support equipment, and test system were reviewed then candidate replacement cleaners and test methods selected. All of these factors as well as testing results will be discussed.

Description: Separated flow over a bluff body is analyzed via large eddy simulations. The turbulent flow around a square cylinder features a variety of complex flow phenomena such as highly unsteady vortical structures, reverse flow in the near wall region, and wake turbulence. The formation of spanwise vortices is often times artificially suppressed in computations by either insufficient depth or a coarse spanwise resolution. As the resolution is refined and the domain extended, the artificial turbulent energy exchange between spanwise and streamwise turbulence is eliminated within the wake region. A parametric study is performed highlighting the effects of spanwise vortices where the spanwise computational domain's resolution and depth are varied. For Re=22,000, the mean and turbulent statistics computed from the numerical large eddy simulations (NLES) are in good agreement with experimental data. Von-Karman shedding is observed in the wake of the cylinder. Mesh independence is illustrated by comparing a mesh resolution of 2 million to 16 million. Sensitivities to time stepping were minimized and sampling frequency sensitivities were nonpresent. While increasing the spanwise depth and resolution can be costly, this practice was found to be necessary to eliminating the artificial turbulent energy exchange.

Description: Fundamental research for sonic boom reduction is needed to quantify the interaction of shock waves generated from the aircraft wing or tail surfaces with the nozzle exhaust plume. Aft body shock waves that interact with the exhaust plume contribute to the near-field pressure signature of a vehicle. The plume and shock interaction was studied using computational fluid dynamics and compared with experimental data from a coaxial convergent-divergent nozzle flow in an open jet facility. A simple diamond-shaped wedge was used to generate the shock in the outer flow to study its impact on the inner jet flow. Results show that the compression from the wedge deflects the nozzle plume and shocks form on the opposite plume boundary. The sonic boom pressure signature of the nozzle exhaust plume was modified by the presence of the wedge. Both the experimental results and computational predictions show changes in plume deflection.

Description: No abstract available

Description: We discuss thrust measurements of Dielectric Barrier Discharge (DBD) plasma actuators devices used for aerodynamic active flow control. After a review of our experience with conventional thrust measurement and significant non-repeatability of the results, we devised a suspended actuator test setup, and now present a methodology of thrust measurements with decreased uncertainty. The methodology consists of frequency scans at constant voltages. The procedure consists of increasing the frequency in a step-wise fashion from several Hz to the maximum frequency of several kHz, followed by frequency decrease back down to the start frequency of several Hz. This sequence is performed first at the highest voltage of interest, then repeated at lower voltages. The data in the descending frequency direction is more consistent and selected for reporting. Sample results show strong dependence of thrust on humidity which also affects the consistency and fluctuations of the measurements. We also observed negative values of thrust or "anti-thrust", at low frequencies between 4 Hz and up to 64 Hz. The anti-thrust is proportional to the mean-squared voltage and is frequency independent. Departures from the parabolic anti-thrust curve are correlated with appearance of visible plasma discharges. We propose the anti-thrust hypothesis. It states that the measured thrust is a sum of plasma thrust and anti-thrust, and assumes that the anti-thrust exists at all frequencies and voltages. The anti-thrust depends on actuator geometry and materials and on the test installation. It enables the separation of the plasma thrust from the measured total thrust. This approach enables more meaningful comparisons between actuators at different installations and laboratories. The dependence on test installation was validated by surrounding the actuator with a large diameter, grounded, metal sleeve.

Description: NTR: High thrust high specific impulse (2 x LOXLH2chemical) engine uses high power density fission reactor with enriched uranium fuel as thermal power source. Reactor heat is removed using H2propellant which is then exhausted to produce thrust. Conventional chemical engine LH2tanks, turbopumps, regenerative nozzles and radiation-cooled shirt extensions used --NTR is next evolutionary step in high performance liquid rocket engines During the Rover program, a common fuel element tie tube design was developed and used in the design of the 50 klbf Kiwi-B4E (1964), 75 klbf Phoebus-1B (1967), 250 klbf Phoebus-2A (June 1968), then back down to the 25 klbf Pewee engine (Nov-Dec 1968) NASA and DOE are using this same approach: design, build, ground then flight test a small engine using a common fuel element that is scalable to a larger 25 klbf thrust engine needed for human missions

Description: In-situ produced propellants have been identified in many architecture studies as key to implementing feasible chemical propulsion missions to destinations beyond lunar orbit. Some of the more noteworthy ones include: launching from Mars to return to Earth (either direct from the surface, or via an orbital rendezvous); using the Earth-Moon Lagrange point as a place to refuel Mars transfer stages with Lunar surface produced propellants; and using Mars Moon Phobos as a place to produce propellants for descent and ascent stages bound for the Mars surface. However successful implementation of these strategies require an ability to successfully transfer propellants from the in-situ production equipment into the propellant tankage of the rocket stage used to move to the desired location. In many circumstances the most desirable location for this transfer to occur is in the low-gravity environment of space. In support of low earth orbit propellant depot concepts, extensive studies have been conducted on transferring propellants in-space. Most of these propellant transfer techniques will be applicable to low gravity operations in other locations. Even ground-based transfer operations on the Moon, Mars, and especially Phobos could benefit from the propellant conserving techniques used for depot refueling. This paper will review the literature of in-situ propellants and refueling to: assess the performance benefits of the use in-situ propellants for mission concepts; review the parallels with propellant depot efforts; assess the progress of the techniques required; and provide recommendations for future research.

Description: Determining the adverse aerodynamic effects due to ice accretion often relies on dry-air wind-tunnel testing of artificial, or simulated, ice shapes. Recent developments in ice accretion documentation methods have yielded a laser-scanning capability that can measure highly three-dimensional features of ice accreted in icing wind tunnels. The objective of this paper was to evaluate the aerodynamic accuracy of ice-accretion simulations generated from laser-scan data. Ice-accretion tests were conducted in the NASA Icing Research Tunnel using an 18-inch chord, 2-D straight wing with NACA 23012 airfoil section. For six ice accretion cases, a 3-D laser scan was performed to document the ice geometry prior to the molding process. Aerodynamic performance testing was conducted at the University of Illinois low-speed wind tunnel at a Reynolds number of 1.8 x 10(exp 6) and a Mach number of 0.18 with an 18-inch chord NACA 23012 airfoil model that was designed to accommodate the artificial ice shapes. The ice-accretion molds were used to fabricate one set of artificial ice shapes from polyurethane castings. The laser-scan data were used to fabricate another set of artificial ice shapes using rapid prototype manufacturing such as stereolithography. The iced-airfoil results with both sets of artificial ice shapes were compared to evaluate the aerodynamic simulation accuracy of the laser-scan data. For four of the six ice-accretion cases, there was excellent agreement in the iced-airfoil aerodynamic performance between the casting and laser-scan based simulations. For example, typical differences in iced-airfoil maximum lift coefficient were less than 3% with corresponding differences in stall angle of approximately one degree or less. The aerodynamic simulation accuracy reported in this paper has demonstrated the combined accuracy of the laser-scan and rapid-prototype manufacturing approach to simulating ice accretion for a NACA 23012 airfoil. For several of the ice-accretion cases tested, the aerodynamics is known to depend upon the small, three dimensional features of the ice. These data show that the laser-scan and rapid-prototype manufacturing approach is capable of replicating these ice features within the reported accuracies of the laser-scan measurement and rapid-prototyping method; thus providing a new capability for high-fidelity ice-accretion documentation and artificial ice-shape fabrication for icing research.

Description: Reliable prediction of contra-rotating open rotor (CROR) noise is an essential element of any strategy for the development of low-noise open rotor propulsion systems that can meet both the community noise regulations and the cabin noise limits. Since CROR noise spectra typically exhibits a preponderance of tones, significant efforts have been directed towards predicting their tone spectra. To that end, there has been an ongoing effort at NASA to assess various in-house open rotor tone noise prediction tools using a benchmark CROR blade set for which significant aerodynamic and acoustic data had been acquired in wind tunnel tests. In the work presented here, the focus is on the near-field noise of the benchmark open rotor blade set at the cruise condition. Using an analytical CROR tone noise model with input from high-fidelity aerodynamic simulations, detailed tone noise spectral predictions have been generated and compared with the experimental data. Comparisons indicate that the theoretical predictions are in good agreement with the data, especially for the dominant CROR tones and their overall sound pressure level. The results also indicate that, whereas individual rotor tones are well predicted by the linear sources (i.e., thickness and loading), for the interaction tones it is essential that the quadrupole sources be included in the analysis.

Description: NASA has sought to utilize high-power solar electric propulsion as means of improving the affordability of in-space transportation for almost 50 years. Early efforts focused on 25 to 50 kilowatt systems that could be used with the Space Shuttle, while later efforts focused on systems nearly an order of magnitude higher power that could be used with heavy lift launch vehicles. These efforts never left the concept development phase in part because the technology required was not sufficiently mature. Since 2012 the NASA Space Technology Mission Directorate has had a coordinated plan to mature the requisite solar array and electric propulsion technology needed to implement a 30 to 50 kilowatt solar electric propulsion technology demonstration mission. Multiple solar electric propulsion technology demonstration mission concepts have been developed based on these maturing technologies with recent efforts focusing on an Asteroid Redirect Robotic Mission. If implemented, the Asteroid Redirect Vehicle will form the basis for a capability that can be cost-effectively evolved over time to provide solar electric propulsion transportation for a range of follow-on mission applications at power levels in excess of 100 kilowatts.

Description: The National Aeronautics and Space Administration (NASA) Science Mission Directorate In-Space Propulsion Technology office is sponsoring NASA Glenn Research Center to develop a 4 kW-class Hall thruster propulsion system for implementation in NASA science missions. A study was conducted to assess the impact of varying the facility background pressure on the High Voltage Hall Accelerator (HiVHAc) thruster performance and voltage-current characteristics. This present study evaluated the HiVHAc thruster performance in the lowest attainable background pressure condition at NASA GRC Vacuum Facility 5 to best simulate space-like conditions. Additional tests were performed at selected thruster operating conditions to investigate and elucidate the underlying physics that change during thruster operation at elevated facility background pressure. Tests were performed at background pressure conditions that are three and ten times higher than the lowest realized background pressure. Results indicated that the thruster discharge specific impulse and efficiency increased with elevated facility background pressure. The voltage-current profiles indicated a narrower stable operating region with increased background pressure. Experimental observations of the thruster operation indicated that increasing the facility background pressure shifted the ionization and acceleration zones upstream towards the thrusters anode. Future tests of the HiVHAc thruster are planned at background pressure conditions that are expected to be two to three times lower than what was achieved during this test campaign. These tests will not only assess the impact of reduced facility background pressure on thruster performance, voltage-current characteristics, and plume properties; but will also attempt to quantify the magnitude of the ionization.

Description: The Global Precipitation Measurement (GPM) mission is an international partnership between NASA and JAXA whose Core spacecraft performs cutting-edge measurements of rainfall and snowfall worldwide and unifies data gathered by a network of precipitation measurement satellites. The Core spacecraft's propulsion system is a blowdown monopropellant system with an initial hydrazine load of 545 kg in a single composite overwrapped propellant tank. At launch, the propulsion system contained propellant in the tank and manifold tubes upstream of the latch valves, with low-pressure helium gas in the manifold tubes downstream of the latch valves. The system had a relatively high beginning-of- life pressure and long downstream manifold lines; these factors created conditions that were conducive to high surge pressures. This paper discusses the GPM project's approach to surge mitigation in the propulsion system design. The paper describes the surge testing program and results, with discussions of specific difficulties encountered. Based on the results of surge testing and pressure drop analyses, a unique configuration of cavitating venturis was chosen to mitigate surge while minimizing pressure losses during thruster maneuvers. This paper concludes with a discussion of overall lessons learned with surge pressure testing for NASA Goddard spacecraft programs.

Description: Several minimum-mass aeroelastic optimization problems are solved to evaluate the effectiveness of a variety of novel tailoring schemes for subsonic transport wings. Aeroelastic strength and panel buckling constraints are imposed across a variety of trimmed maneuver loads. Tailoring with metallic thickness variations, functionally graded materials, composite laminates, tow steering, and distributed trailing edge control effectors are all found to provide reductions in structural wing mass with varying degrees of success. The question as to whether this wing mass reduction will offset the increased manufacturing cost is left unresolved for each case.

Description: The National Aeronautics and Space Administration (NASA) Science Mission Directorate In- Space Propulsion Technology office is sponsoring NASA Glenn Research Center (GRC) to develop a 4 kW-class Hall thruster propulsion system for implementation in NASA science missions. Tests were performed within NASA GRC Vacuum Facility 5 at background pressure levels that were six times lower than what has previously been attained in other vacuum facilities. A study was conducted to assess the impact of varying the cathode-to-anode flow fraction and cathode position on the performance and operational characteristics of the High Voltage Hall Accelerator (HiVHAc) thruster. In addition, the impact of injecting additional xenon propellant in the vicinity of the cathode was also assessed. Cathode-to-anode flow fraction sensitivity tests were performed for power levels between 1.0 and 3.9 kW. It was found that varying the cathode flow fraction from 5 to approximately 10% of the anode flow resulted in the cathode-to-ground voltage becoming more positive. For an operating condition of 3.8 kW and 500 V, varying the cathode position from a distance of closest approach to 600 mm away did not result in any substantial variation in thrust but resulted in the cathode-to-ground changing from -17 to -4 V. The change in the cathode-to-ground voltage along with visual observations indicated a change in how the cathode plume was coupling to the thruster discharge. Finally, the injection of secondary xenon flow in the vicinity of the cathode had an impact similar to increasing the cathode-to-anode flow fraction, where the cathode-to-ground voltage became more positive and discharge current and thrust increased slightly. Future tests of the HiVHAc thruster are planned with a centrally mounted cathode in order to further assess the impact of cathode position on thruster performance.

Description: The National Aeronautics and Space Administration (NASA) Science Mission Directorate In-Space Propulsion Technology office is sponsoring NASA Glenn Research Center to develop a 4 kW-class Hall thruster propulsion system for implementation in NASA science missions. A study was conducted to assess the impact of varying the facility background pressure on the High Voltage Hall Accelerator (HiVHAc) thruster performance and voltage-current characteristics. This present study evaluated the HiVHAc thruster performance in the lowest attainable background pressure condition at NASA GRC Vacuum Facility 5 to best simulate space-like conditions. Additional tests were performed at selected thruster operating conditions to investigate and elucidate the underlying physics that change during thruster operation at elevated facility background pressure. Tests were performed at background pressure conditions that are three and ten times higher than the lowest realized background pressure. Results indicated that the thruster discharge specific impulse and efficiency increased with elevated facility background pressure. The voltage-current profiles indicated a narrower stable operating region with increased background pressure. Experimental observations of the thruster operation indicated that increasing the facility background pressure shifted the ionization and acceleration zones upstream towards the thruster's anode. Future tests of the HiVHAc thruster are planned at background pressure conditions that are expected to be two to three times lower than what was achieved during this test campaign. These tests will not only assess the impact of reduced facility background pressure on thruster performance, voltage-current characteristics, and plume properties; but will also attempt to quantify the magnitude of the ionization and acceleration zones upstream shifting as a function of increased background pressure.

Description: No abstract available

Description: Nuclear thermal propulsion has long been considered an enabling technology for human missions to Mars and beyond. One concept of operations for these missions utilizes the nuclear reactor to generate electrical power during coast phases, known as bimodal operation. This presentation focuses on the systems modeling and analysis efforts for a NERVA derived concept. The NERVA bimodal operation derives the thermal energy from the core tie tube elements. Recent analysis has shown potential temperature distributions in the tie tube elements that may limit the thermodynamic efficiency of the closed Brayton cycle used to generate electricity with the current design. The results of this analysis are discussed as well as the potential implications to a bimodal NERVA type reactor.

Description: This paper presents results from computational fluid dynamic (CFD) simulations of a three-stream plug nozzle. Time-accurate, Euler, quasi-1D and 2D-axisymmetric simulations were performed as part of an effort to provide a CFD-based approach to modeling nozzle dynamics. The CFD code used for the simulations is based on the space-time Conservation Element and Solution Element (CESE) method. Steady-state results were validated using the Wind-US code and a code utilizing the MacCormack method while the unsteady results were partially validated via an aeroacoustic benchmark problem. The CESE steady-state flow field solutions showed excellent agreement with solutions derived from the other methods and codes while preliminary unsteady results for the three-stream plug nozzle are also shown. Additionally, a study was performed to explore the sensitivity of gross thrust computations to the control surface definition. The results showed that most of the sensitivity while computing the gross thrust is attributed to the control surface stencil resolution and choice of stencil end points and not to the control surface definition itself.Finally, comparisons between the quasi-1D and 2D-axisymetric solutions were performed in order to gain insight on whether a quasi-1D solution can capture the steady and unsteady nozzle phenomena without the cost of a 2D-axisymmetric simulation. Initial results show that while the quasi-1D solutions are similar to the 2D-axisymmetric solutions, the inability of the quasi-1D simulations to predict two dimensional phenomena limits its accuracy.

Description: No abstract available

Description: No abstract available

Description: The NASA's Evolutionary Xenon Thruster (NEXT) program is developing the next-generation solar electric ion propulsion system with significant enhancements beyond the state-of-the-art NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) ion propulsion system to provide future NASA science missions with enhanced capabilities. A Long-Duration Test (LDT) was initiated in June 2005 to validate the thruster service life modeling and to quantify the thruster propellant throughput capability. Testing was recently completed in February 2014, with the thruster accumulating 51,184 hours of operation, processing 918 kg of xenon propellant, and delivering 35.5 MN-s of total impulse.As part of the test termination procedure, a comprehensive performance characterization was performed across the entire NEXT throttle table. This was performed prior to planned repairs of numerous diagnostics that had become inoperable over the course of the test. After completion of these diagnostic repairs in November 2013, a comprehensive end-of-test performance and wear characterization was performed on the test article prior to exposure to atmosphere. These data have confirmed steady thruster performance with minimal degradation as well as mitigation of numerous life limiting mechanisms encountered in the NSTAR design. Component erosion rates compare favorably to pretest predictions based on semi-empirical models used for the thruster service life assessment. Additional data relating to ion beam density profiles, facility backsputter rates, facility backpressure effects on thruster telemetry, and modulation of the neutralizer keeper current are presented as part of the end-of-test characterization. Presently the test article for the NEXT LDT has been exposed to atmosphere and placed within a clean room environment, with post-test disassembly and inspection underway.

Description: A key requirement of a low-gravity screen-channel liquid acquisition device (LAD) is the need to retain 100% liquid in the channel in response to propellant outflow and spacecraft maneuvers. The point at which a screen-channel LAD ingests vapor is known as breakdown, and can be measured several different ways such as: visual observation of bubbles in the LAD channel outflow; a sudden change in pressure drop between the propellant tank and LAD sump outlet; or, an indication by wet-dry sensors placed in the LAD channel or outflow stream. Here we describe a new type of sensor for gauging a screen-channel LAD, the Radio Frequency Mass Gauge (RFMG). The RFMG measures the natural electromagnetic modes of the screen-channel LAD, which is very similar to an RF waveguide, to determine the amount of propellant in the channel. By monitoring several of the RF modes, we show that the RFMG acts as a global sensor of the LAD channel propellant fill level, and enables detection of LAD breakdown even in the absence of outflow. This paper presents the theory behind the RFMG-LAD sensor, measurements and simulations of the RF modes of a LAD channel, and RFMG detection of LAD breakdown in a channel using a simulant fluid during inverted outflow and long-term stability tests.

Description: To support future NASA deep space missions, a radioisotope power system utilizing Stirling power conversion technology was under development. This development effort was performed under the joint sponsorship of the Department of Energy and NASA, until its termination at the end of 2013 due to budget constraints. The higher conversion efficiency of the Stirling cycle compared with that of the Radioisotope Thermoelectric Generators (RTGs) used in previous missions (Viking, Pioneer, Voyager, Galileo, Ulysses, Cassini, Pluto New Horizons and Mars Science Laboratory) offers the advantage of a four-fold reduction in Pu-238 fuel, thereby extending its limited domestic supply. As part of closeout activities, system-level testing of flight-like Advanced Stirling Convertors (ASCs) with a flight-like ASC Controller Unit (ACU) was performed in February 2014. This hardware is the most representative of the flight design tested to date. The test fully demonstrates the following ACU and system functionality: system startup; ASC control and operation at nominal and worst-case operating conditions; power rectification; DC output power management throughout nominal and out-of-range host voltage levels; ACU fault management, and system command / telemetry via MIL-STD 1553 bus. This testing shows the viability of such a system for future deep space missions and bolsters confidence in the maturity of the flight design.

Description: Presentation on the European Service Module mission description, propulsion subsystem, and propulsion and guidance navigation and control interface requirements. The content focuses on the updates to these areas between Constellation and the Multi-Purpose Crew Vehicle.

Description: Supersonic inlets with external compression, having a good level performance at the critical operating point, exhibit a marked instability of the flow in some subcritical operation below a critical value of the capture mass flow ratio. This takes the form of severe oscillations of the shock system, commonly known as "buzz". The underlying purpose of this study is to indicate how Detached Eddy Simulation (DES) analysis of supersonic inlets will alter how we envision unsteady inlet aerodynamics, particularly inlet buzz. Presented in this paper is a discussion regarding the physical explanation underlying inlet buzz as indicated by DES analysis. It is the normal shock wave boundary layer separation along the spike surface which reduces the capture mass flow that is the controlling mechanism which determines the onset of inlet buzz, and it is the aerodynamic characteristics of a choked nozzle that provide the feedback mechanism that sustains the buzz cycle by imposing a fixed mean corrected inlet weight flow. Comparisons between the DES analysis of the Lockheed Martin Corporation (LMCO) N+2 inlet and schlieren photographs taken during the test of the Gulfstream Large Scale Low Boom (LSLB) inlet in the NASA 8x6 ft. Supersonic Wind Tunnel (SWT) show a strong similarity both in turbulent flow field structure and shock wave formation during the buzz cycle. This demonstrates the value of DES analysis for the design and understanding of supersonic inlets.

Description: NASA Ames Research Center and the George Washington University are developing an electric propulsion subsystem that will be integrated into the PhoneSat bus. Experimental tests have shown a reliable performance by firing three different thrusters at various frequencies in vacuum conditions. The interface consists of a microcontroller that sends a trigger pulse to the Pulsed Plasma Unit that is responsible for the thruster operation. A Smartphone is utilized as the main user interface for the selection of commands that control the entire system. The propellant, which is the cathode itself, is a solid cylinder made of Titanium. This simplicity in the design avoids miniaturization and manufacturing problems. The characteristics of this thruster allow an array of CATs to perform attitude control and orbital correction maneuvers that will open the door for the implementation of an extensive collection of new mission concepts and space applications for CubeSats. NASA Ames is currently working on the integration of the system to fit the thrusters and the PPU inside a 1.5U CubeSat together with the PhoneSat bus. This satellite is intended to be deployed from the ISS in 2015 and test the functionality of the thrusters by spinning the satellite around its long axis and measure the rotational speed with the phone gyros. This test flight will raise the TRL of the propulsion system from 5 to 7 and will be a first test for further CubeSats with propulsion systems, a key subsystem for long duration or interplanetary small satellite missions.

Description: A power processing unit for a 15 kW Hall thruster is under development at NASA Glenn Research Center. The unit produces up to 400 VDC with two parallel 7.5 kW discharge modules that operate from a 300 VDC nominal input voltage. Silicon carbide MOSFETs and diodes were used in this design because they were the best choice to handle the high voltage stress while delivering high efficiency and low specific mass. Efficiencies in excess of 97 percent were demonstrated during integration testing with the NASA-300M 20 kW Hall thruster. Electromagnet, cathode keeper, and heater supplies were also developed and will be integrated with the discharge supply into a vacuum-rated brassboard power processing unit with full flight functionality. This design could be evolved into a flight unit for future missions that requires high power electric propulsion.

Description: Numerical and Analytical methods developed to determine damage accumulation in specific engine components when speed variation included. Dither Life Ratio shown to be well over factor of 2 for specific example. SteadyState assumption shown to be accurate for most turbopump cases, allowing rapid calculation of DLR. If hotfire speed data unknown, Monte Carlo method developed that uses speed statistics for similar engines. Application of techniques allow analyst to reduce both uncertainty and excess conservatism. High values of DLR could allow previously unacceptable part to pass HCF criteria without redesign. Given benefit and ease of implementation, recommend that any finite life turbomachine component analysis adopt these techniques. Probability Values calculated, compared, and evaluated for several industryproposed methods for combining random and harmonic loads. Two new excel macros written to calculate combined load for any specific probability level. Closed form Curve fits generated for widely used 3(sigma) and 2(sigma) probability levels. For design of lightweight aerospace components, obtaining accurate, reproducible, statistically meaningful answer critical.

Description: In September 2013 the NASA Innovative Advanced Concept (NIAC) organization awarded a phase I contract to the PuFF team. Our phase 1 proposal researched a pulsed fission-fusion propulsion system that compressed a target of deuterium (D) and tritium (T) as a mixture in a column, surrounded concentrically by Uranium. The target is surrounded by liquid lithium. A high power current would flow down the liquid lithium and the resulting Lorentz force would compress the column by roughly a factor of 10. The compressed column would reach criticality and a combination of fission and fusion reactions would occur. Our Phase I results, summarized herein, review our estimates of engine and vehicle performance, our work to date to model the fission-fusion reaction, and our initial efforts in experimental analysis.

Description: The PROPEL flight mission concept will demonstrate the safe use of an electrodynamic tether for generating thrust. PROPEL is being designed to be a versatile electrodynamic-tether system for multiple end users and to be flexible with respect to platform. As such, several implementation options are being explored, including a comprehensive mission design for PROPEL with a mission duration of six months; a space demonstration mission concept design with configuration of a pair of tethered satellites, one of which is the Japanese H-II Transfer Vehicle; and an ESPA-based system. We report here on these possible implementation options for PROPEL. electrodynamic tether; PROPEL demonstration mission; propellantless propulsion

Description: No abstract available

Description: A mode transition study is conducted in magnetically shielded thrusters where the magnetic field magnitude is varied to induce mode transitions. Three different oscillatory modes are identified with the 20-kW NASA-300MS-2 and the 6-kW H6MS: Mode 1) global mode similar to unshielded thrusters at low magnetic fields, Mode 2) cathode oscillations at nominal magnetic fields, and Mode 3) combined spoke, cathode and breathing mode oscillations at high magnetic fields. Mode 1 exhibits large amplitude, low frequency (1-10 kHz), breathing mode type oscillations where discharge current mean value and oscillation amplitude peak. The mean discharge current is minimized while thrust-to-power and anode efficiency are maximized in Mode 2, where higher frequency (50-90 kHz), low amplitude, cathode oscillations dominate. Thrust is maximized in Mode 3 and decreases by 5-6% with decreasing magnetic field strength. The presence or absence of spokes and strong cathode oscillations do not affect each other or discharge current. Similar to unshielded thrusters, mode transitions and plasma oscillations affect magnetically shielded thruster performance and should be characterized during system development.

Description: MSFC has embarked on use of green propellant replacement of hydrazine for a variety of applications. This paper focused on activities for auxiliary power unit but MSFC is actively investigating use of green propellants for thruster applications. MSFC is interested in partnership with the international community to address the infusion of green propellant into greater use.

Description: In 2012, the National Aeronautics & Space Administration (NASA) Space Technology Mission Directorate (STMD) began the process of building an integrated technology roadmap, including both technology pull and technology push strategies. Technology Area 1 (TA-01) for Launch Propulsion Systems and TA-02 In-Space Propulsion are two of the fourteen TAs that provide recommendations for the overall technology investment strategy and prioritization of NASA's space technology activities. Identified within these documents are future needs of green propellant use. Green ionic liquid monopropellants and propulsion systems are beginning to be demonstrated in space flight environments. Starting in 2010 with the flight of Prisma, a 1-N thruster system began on-orbit demonstrations operating on ammonium dinitramide based propellant. The NASA Green Propellant Infusion Mission (GPIM) plans to demonstrate both 1-N, and 22-N hydroxyl ammonium nitrate (HAN)-based thrusters in a 2015 flight demonstration. In addition, engineers at MSFC have been evaluating green propellant alternatives for both thrusters and auxiliary power units (APUs). This paper summarizes the status of these development/demonstration activities and investigates the potential for evolution of green propellants from small spacecraft and satellites to larger spacecraft systems, human exploration, and launch system auxiliary propulsion applications.

Description: Affordable and convenient access to electrical power is critical to consumers, spacecraft, military and other applications alike. In the aerospace industry, an increased emphasis on small satellite flights and a move toward CubeSat and NanoSat technologies, the need for systems that could package into a small stowage volume while still being able to power robust space missions has become more critical. As a result, the Marshall Space Flight Center's Advanced Concepts Office identified a need for more efficient, affordable, and smaller space power systems to trade in performing design and feasibility studies. The Lightweight Inflatable Solar Array (LISA), a concept designed, prototyped, and tested at the NASA Marshall Space Flight Center (MSFC) in Huntsville, Alabama provides an affordable, lightweight, scalable, and easily manufactured approach for power generation in space or on Earth. This flexible technology has many wide-ranging applications from serving small satellites to soldiers in the field. By using very thin, ultraflexible solar arrays adhered to an inflatable structure, a large area (and thus large amount of power) can be folded and packaged into a relatively small volume (shown in artist rendering in Figure 1 below). The proposed presentation will provide an overview of the progress to date on the LISA project as well as a look at its potential, with continued development, to revolutionize small spacecraft and portable terrestrial power systems.

Description: A second wind tunnel test of the FAST-MAC circulation control semi-span model was recently completed in the National Transonic Facility at the NASA Langley Research Center. The model allowed independent control of four circulation control plenums producing a high momentum jet from a blowing slot near the wing trailing edge that was directed over a 15% chord simple-hinged flap. The model was configured for transonic testing of the cruise configuration with 0deg flap deflection to determine the potential for drag reduction with the circulation control blowing. Encouraging results from analysis of wing surface pressures suggested that the circulation control blowing was effective in reducing the transonic drag on the configuration, however this could not be quantified until the thrust generated by the blowing slot was correctly removed from the force and moment balance data. This paper will present the thrust removal methodology used for the FAST-MAC circulation control model and describe the experimental measurements and techniques used to develop the methodology. A discussion on the impact to the force and moment data as a result of removing the thrust from the blowing slot will also be presented for the cruise configuration, where at some Mach and Reynolds number conditions, the thrust-removed corrected data showed that a drag reduction was realized as a consequence of the blowing.

Description: This paper describes the test campaigns designed to investigate and demonstrate viability of using classical magnetoplasmadynamics to obtain a propulsive momentum transfer via the quantum vacuum virtual plasma. This paper will not address the physics of the quantum vacuum plasma thruster (QVPT), but instead will describe the recent test campaign. In addition, it contains a brief description of the supporting radio frequency (RF) field analysis, lessons learned, and potential applications of the technology to space exploration missions. During the first (Cannae) portion of the campaign, approximately 40 micronewtons of thrust were observed in an RF resonant cavity test article excited at approximately 935 megahertz and 28 watts. During the subsequent (tapered cavity) portion of the campaign, approximately 91 micronewtons of thrust were observed in an RF resonant cavity test article excited at approximately 1933 megahertz and 17 watts. Testing was performed on a low-thrust torsion pendulum that is capable of detecting force at a single-digit micronewton level. Test campaign results indicate that the RF resonant cavity thruster design, which is unique as an electric propulsion device, is producing a force that is not attributable to any classical electromagnetic phenomenon and therefore is potentially demonstrating an interaction with the quantum vacuum virtual plasma.

Description: The Orion Crew Module Propulsion Reaction Control System is currently complete and ready for flight as part of the Orion program's first flight test, Exploration Flight Test One (EFT-1). As part of the first article design, build, test, and integration effort, several key lessons learned have been noted and are planned for incorporation into the next build of the system. This paper provides an overview of those lessons learned and a status on the Orion propulsion system progress to date.

Description: The spray characteristics of a liquid-liquid double swirl coaxial injector were studied using non-invasive optical, laser, and X-ray diagnostics. A parametric study of injector exit geometry demonstrated that spray breakup time, breakup type and sheet stability could be controlled with exit geometry. Phase Doppler interferometry was used to characterize droplet statistics and non-dimensional droplet parameters over a range of inlet conditions and for various fluids allowing for a study on the role of specific fluid properties in atomization. Further, X-ray radiography allowed for investigation of sheet thickness and breakup length to be quantified for different recess exit diameters and inlet pressures. Finally, computed tomography scans revealed that the spray cone was distinctively non-uniform and comprised of several pockets of increased mass flux.

Description: Reduced-order modeling (ROM) methods are applied to the CFD-based aeroelastic analysis of the AGARD 445.6 wing in order to gain insight regarding well-known discrepancies between the aeroelastic analyses and the experimental results. The results presented include aeroelastic solutions using the inviscid CAP-TSD code and the FUN3D code (Euler and Navier-Stokes). Full CFD aeroelastic solutions and ROM aeroelastic solutions, computed at several Mach numbers, are presented in the form of root locus plots in order to better reveal the aeroelastic root migrations with increasing dynamic pressure. Important conclusions are drawn from these results including the ability of the linear CAP-TSD code to accurately predict the entire experimental flutter boundary (repeat of analyses performed in the 1980's), that the Euler solutions at supersonic conditions indicate that the third mode is always unstable, and that the FUN3D Navier-Stokes solutions stabilize the unstable third mode seen in the Euler solutions.

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Description: Mars Sample Return (MSR) missions could benefit from the high specific impulse of Solar Electric Propulsion (SEP) to achieve lower launch masses than with chemical propulsion. SEP presents formulation challenges due to the coupled nature of launch vehicle performance, propulsion system, power system, and mission timeline. This paper describes a SEP orbiter-sizing tool, which models spacecraft mass & timeline in conjunction with low thrust round-trip Earth-Mars trajectories, and presents selected concept designs. A variety of system designs are possible for SEP MSR orbiters, with large dry mass allocations, similar round-trip durations to chemical orbiters, and reduced design variability between opportunities.

Description: Kennedy Space Center (KSC) has led the efforts for lunar and Martian landing site preparation, including excavation, soil stabilization, and plume damage prediction. There has been much discussion of sintering but until our team recently demonstrated it for the lunar case there was little understanding of the serious challenges. Simplistic sintering creates a crumbly, brittle, weak surface unsuitable for a rocket exhaust plume. The goal of this project is to solve those problems and make it possible to land a human class lander on Mars, making terminal landing of humans on Mars possible for the first time.

Description: The purpose of this work is to develop a set of theoretical and experimental techniques to characterize the aeroelasticity of the thermal protection system (TPS) on the NASA Hypersonic Inflatable Aerodynamic Decelerator (HIAD). A square TPS coupon experiences trailing edge oscillatory behavior during experimental testing in the 8' High Temperature Tunnel (HTT), which may indicate the presence of aeroelastic flutter. Several theoretical aeroelastic models have been developed, each corresponding to a different experimental test configuration. Von Karman large deflection theory is used for the plate-like components of the TPS, along with piston theory for the aerodynamics. The constraints between the individual TPS layers and the presence of a unidirectional foundation at the back of the coupon are included by developing the necessary energy expressions and using the Rayleigh Ritz method to derive the nonlinear equations of motion. Free vibrations and limit cycle oscillations are computed and the frequencies and amplitudes are compared with accelerometer and photogrammetry data from the experiments.

Description: This report presents the results of several thermodynamic vent system (TVS) tests with liquid oxygen plus a test with liquid nitrogen. In all tests, the liquid was heated above its normal boiling point to 111 K for oxygen and 100 K for nitrogen. The elevated temperature was representative of tank conditions for a candidate lunar lander ascent stage. An initial test series was conducted with saturated oxygen liquid and vapor at 0.6 MPa. The initial series was followed by tests where the test tank was pressurized with gaseous helium to 1.4 to 1.6 MPa. For these tests, the helium mole fraction in the ullage was quite high, about 0.57 to 0.62. TVS behavior is different when helium is present than when helium is absent. The tank pressure becomes the sum of the vapor pressure and the partial pressure of helium. Therefore, tank pressure depends not only on temperature, as is the case for a pure liquid-vapor system, but also on helium density (i.e., the mass of helium divided by the ullage volume). Thus, properly controlling TVS operation is more challenging with helium pressurization than without helium pressurization. When helium was present, the liquid temperature would rise with each successive TVS cycle if tank pressure was kept within a constant control band. Alternatively, if the liquid temperature was maintained within a constant TVS control band, the tank pressure would drop with each TVS cycle. The final test series, which was conducted with liquid nitrogen pressurized with helium, demonstrated simultaneous pressure and temperature control during TVS operation. The simultaneous control was achieved by systematic injection of additional helium during each TVS cycle. Adding helium maintained the helium partial pressure as the liquid volume decreased because of TVS operation. The TVS demonstrations with liquid oxygen pressurized with helium were conducted with three different fluid-mixer configurations-a submerged axial jet mixer, a pair of spray hoops in the tank ullage, and combined use of the axial jet and spray hoops. A submerged liquid pump and compact heat exchanger located inside the test tank were used with all the mixer configurations. The initial series without helium and the final series with liquid nitrogen both used the axial jet mixer. The axial jet configuration successfully demonstrated the ability to control tank pressure; but in the normal-gravity environment, the temperature in the upper tank region (ullage and unwetted wall) was not controlled. The spray hoops and axial jet combination also successfully demonstrated pressure control as well as temperature control of the entire tank and contents. The spray-hoops-only configuration was not expected to be a reliable means of tank mixing because there was no direct means to produce liquid circulation. However, surprisingly good results also were obtained with the sprayhoops- only configuration (i.e., performance metrics such as cycle-averaged vent flowrate were similar to those obtained with the other configurations). A simple thermodynamic model was developed that correctly predicted the TVS behavior (temperature rise or pressure drop per TVS cycle) when helium was present in the ullage. The model predictions were correlated over a range of input parameters. The correlations show that temperature rise or pressure drop per cycle was proportional to both helium mole fraction and tank heat input. The response also depended on the tank fill fraction: the temperature rise or pressure drop (per TVS cycle) increased as the ullage volume decreased.

Description: Dr. Harold "Sonny" White has developed the physics theory basis for utilizing the quantum vacuum to produce thrust. The engineering implementation of the theory is known as Q-thrusters. During FY13, three test campaigns were conducted that conclusively demonstrated tangible evidence of Q-thruster physics with measurable thrust bringing the TRL up from TRL 2 to early TRL 3. This project will continue with the development of the technology to a breadboard level by leveraging the most recent NASA/industry test hardware. This project will replace the manual tuning process used in the 2013 test campaign with an automated Radio Frequency (RF) Phase Lock Loop system (precursor to flight-like implementation), and will redesign the signal ports to minimize RF leakage (improves efficiency). This project will build on the 2013 test campaign using the above improvements on the test implementation to get ready for subsequent Independent Verification and Validation testing at Glenn Research Center (GRC) and Jet Propulsion Laboratory (JPL) in FY 2015. Q-thruster technology has a much higher thrust to power than current forms of electric propulsion (~7x Hall thrusters), and can significantly reduce the total power required for either Solar Electric Propulsion (SEP) or Nuclear Electric Propulsion (NEP). Also, due to the high thrust and high specific impulse, Q-thruster technology will greatly relax the specific mass requirements for in-space nuclear reactor systems. Q-thrusters can reduce transit times for a power-constrained architecture.

Description: This document serves as the final report for the SMAAART AFC-Enabled Vertical Tail System Integration Study. Included are the ground rule assumptions which have gone into the study, layouts of the baseline and AFC-enabled configurations, critical sizing information, system requirements and architectures, and assumed system properties that result in an NPV assessment of the two candidate AFC technologies.

Description: Reynolds-Averaged Navier-Stokes (RANS) computational fluid dynamics (CFD) analysis was conducted to study the low-speed stall aerodynamics of a GIII aircraft's swept wing modified with a laminar-flow wing glove. The stall aerodynamics of the gloved wing were analyzed and compared with the unmodified wing for the flight speed of 120 knots and altitude of 2300 ft above mean sea level (MSL). The Star-CCM+ polyhedral unstructured CFD code was first validated for wing stall predictions using the wing-body geometry from the First American Institute of Aeronautics and Astronautics (AIAA) CFD High-Lift Prediction Workshop. It was found that the Star-CCM+ CFD code can produce results that are within the scattering of other CFD codes considered at the workshop. In particular, the Star-CCM+ CFD code was able to predict wing stall for the AIAA wing-body geometry to within 1 degree of angle of attack as compared to benchmark wind-tunnel test data. Current results show that the addition of the laminar-flow wing glove causes the gloved wing to stall much earlier than the unmodified wing. Furthermore, the gloved wing has a different stall characteristic than the clean wing, with no sharp lift drop-off at stall for the gloved wing.

Description: Viewgraphs showing the combustion chamber pressure measured in a heavy weight chamber during testing at the NASA Ames Hybrid Combustion Facility.

Description: J-2X engine testing on the A-2 test stand at the NASA John C. Stennis Space Center (SSC) has recently concluded. As part of that test campaign, the engine was operated at lower power levels in support of expanding the use of J-2X to other missions. However, the A-2 diffuser was not designed for engine testing at the proposed low power levels. To evaluate the risk of damage to the diffuser, computer simulations were created of the rocket engine exhaust plume inside the 50ft long, water-cooled, altitude-simulating diffuser. The simulations predicted that low power level testing would cause the plume to oscillate in the lower sections of the diffuser. This can possibly cause excessive vibrations, stress, and heat transfer from the plume to the diffuser walls. To understand and assess the performance of the diffuser during low power level engine testing, nine accelerometers and four strain gages were installed around the outer surface of the diffuser. The added instrumentation also allowed for the verification of the rocket exhaust plume computational model. Prior to engine hot-fire testing, a diffuser water-flow test was conducted to verify the proper operation of the newly installed instrumentation. Subsequently, two J-2X engine hot-fire tests were completed. Hot-Fire Test 1 was 11.5 seconds in duration, and accelerometer and strain data verified that the rocket engine plume oscillated in the lower sections of the diffuser. The accelerometers showed very different results dependent upon location. The diffuser consists of four sections, with Section 1 being closest to the engine nozzle and Section 4 being farthest from the engine nozzle. Section 1 accelerometers showed increased amplitudes at startup and shutdown, but low amplitudes while the diffuser was started. Section 3 accelerometers showed the opposite results with near zero G amplitudes prior to and after diffuser start and peak amplitudes to +/- 100G while the diffuser was started. Hot-Fire Test 1 strain gages showed different data dependent on section. Section 1 strains were small, and were in the range of 50 to 150 microstrain, which would result in stresses from 1.45 to 4.35 ksi. The yield stress of the material, A-285 Grade C Steel, is 29.7 ksi. Section 4 strain gages showed much higher values with strains peaking at 1600 microstrain. This strain corresponds to a stress of 46.41 ksi, which is in excess of the yield stress, but below the ultimate stress of 55 to 75 ksi. The decreased accelerations and strain in Section 1, and the increased accelerations and strain in Sections 3 and 4 verified the computer simulation prediction of increased plume oscillations in the lower sections of the diffuser. Hot-Fire Test 2 ran for a duration of 125 seconds. The engine operated at a slightly higher power level than Hot-Fire Test 1 for the initial 35 seconds of the test. After 35 seconds the power level was lowered to Hot-Fire Test 1 levels. The acceleration and strain data for Hot-Fire Test 2 was similar during the initial part of the test. However, just prior to the engine being lowered to the Hot-Fire Test 1 power level, the strain gage data in Section 4 showed a large decrease to strains near zero microstrain from their peak at 1500 microstrain. Future work includes further strain and acceleration data analysis and evaluation.

Description: There has been significant work recently in the development of iodine-fed Hall thrusters for in-space propulsion applications.1 The use of iodine as a propellant provides many advantages over present xenon-gas-fed Hall thruster systems. Iodine is a solid at ambient temperature (no pressurization required) and has no special handling requirements, making it safe for secondary flight opportunities. It has exceptionally high I sp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing system level advantages over mid-term high power electric propulsion options. Iodine provides thrust and efficiency that are comparable to xenonfed Hall thrusters while operating in the same discharge current and voltage regime, making it possible to leverage the development of flight-qualified xenon Hall thruster power processing units for the iodine application. Work at MSFC is presently aimed at designing, integrating, and demonstrating a flight-like iodine feed system suitable for the Hall thruster application. This effort represents a significant advancement in state-of-the-art. Though Iodine thrusters have demonstrated high performance with mission enabling potential, a flight-like feed system has never been demonstrated and iodine compatible components do not yet exist. Presented in this paper is the end-to-end integrated feed system demonstration. The system includes a propellant tank with active feedback-control heating, fill and drain interfaces, latching and proportional flow control valves (PFCV), flow resistors, and flight-like CubeSat power and control electronics. Hardware is integrated into a CubeSat-sized structure, calibrated and tested under vacuum conditions, and operated under under hot-fire conditions using a Busek BHT-200 thruster designed for iodine. Performance of the system is evaluated thorugh accurate measurement of thrust and a calibrated of mass flow rate measurement, which is a function of reservoir temperature/pressure, the flow resistors, and the setting of the PFCV. The calibration is performed using independent flow control monitoring techniques, providing an in situ measure of the flowrate as a function of controllable parameters. The reservoir temperature controls the iodine sublimation rate, providing propellant to ths thruster by pressurizing the propellant feed system to approx.1-2 psi. Control of the temperature and the PFCV are used to maintain reservoir pressure and keep the thruster discharge current constant.

Description: Solar and drag sail technology is entering the mainstream for space propulsion applications within NASA and around the world. Solar sails derive propulsion by reflecting sunlight from a large, mirror- like sail made of a lightweight, reflective material. The continuous sunlight pressure provides efficient primary propulsion without the expenditure of propellant or any other consumable, allowing for very high V maneuvers and long-duration deep space exploration. Drag sails increase the aerodynamic drag on Low Earth Orbit (LEO) spacecraft, providing a lightweight and relatively inexpensive approach for end-of-life deorbit and reentry. Since NASA began investing in the technology in the late 1990's, significant progress has been made toward their demonstration and implementation in space. NASA's Marshall Space Flight Center (MSFC) managed the development and testing of two different 20-m solar sail systems and rigorously tested them under simulated space conditions in the Glenn Research Center's Space Power Facility at Plum Brook Station, Ohio. One of these systems, developed by L'Garde, Inc., is planned for flight in 2015. Called Sunjammer, the 38m sailcraft will unfurl in deep space and demonstrate solar sail propulsion and navigation as it flies to Earth-Sun L1. In the interim, NASA MSFC funded the NanoSail-D, a subscale drag sail system designed for small spacecraft applications. The NanoSail-D flew aboard the Fast Affordable Science and Technology SATellite (FASTSAT) in 2010, also developed by MSFC, and began its mission after it was was ejected from the FASTSAT into Earth orbit, where it remained for several weeks before deorbiting as planned. NASA recently selected two small satellite missions as part of the Advanced Exploration Systems (AES) Program, both of which will use solar sails to enable their scientific objectives. Lunar Flashlight, managed by JPL, will search for and map volatiles in permanently shadowed Lunar craters using a solar sail as a gigantic mirror to steer sunlight into the shaded craters. The Near Earth Asteroid (NEA) Scout mission will use the sail as primary propulsion allowing it to survey and image one or more NEA's of interest for possible future human exploration. Both are planned for launch in 2017. As the technology matures, solar sails will increasingly be used to enable science and exploration missions that are currently impossible or prohibitively expensive using traditional chemical and electric rockets. For example, the NASA Heliophysics Decadal Survey identifies no less than three such missions for possible flight before the mid-2020's. Solar sail propulsion technology is no longer an intesting theoretical possibility; it has been demonstrated in space and is now a critical technology for science and solar system exploration.

Description: The use of Trigas in rocket propulsion systems is somewhat of a new technology. This paper defines Trigas as a mixture of gases composed largely of helium with a small percentage of a stoichiometric mixture of hydrogen and oxygen. When exposed to a catalyst the hydrogen and oxygen in the mixture combusts, significantly raising the temperature of the mixture. The increase in enthalpy resulting from the combustion process significantly decreases the required quantity of gas needed to pressurize the ullage of the vehicle propellant tanks. The objective of this effort was to better understand the operating characteristics of Trigas in a pressurization system with low temperature applications. In conjunction with ongoing programs at NASA Marshall Space Flight Center, an effort has been undertaken to evaluate the operating characteristics of Trigas through modeling and bench testing. Through improved understanding of the operating characteristics, the risk of using this new technology in a launch vehicle propulsion system was reduced. Bench testing of Trigas was a multistep process that targeted gas characteristics and performance aspects that pose a risk to application in a pressurization system. Pressurization systems are vital to propulsion system performance. Keeping a target ullage pressure in propulsions tanks is necessary to supply propellant at the conditions and flow rates required to maintain desired engine functionality. The first component of testing consisted of sampling Trigas sources that had been stagnant for various lengths of time in order to determine the rate at which stratification takes place. Second, a bench test was set up in which Trigas was sent through a catalyst bed. This test was designed to evaluate the performance characteristics of Trigas, under low temperature inlet temperatures, in a flightlike catalyst bed reactor. The third, most complex, test examined the performance characteristics of Trigas at low temperature temperatures in a test configuration built to more closely resemble a vehicle pressurization system. The results of these bench tests address various risks that all relate to underperformance of Trigas pressurization systems.

Description: Outline: Introduction; Launch Ascent and Vehicle Aerodynamics (LAVA) Framework; 1st AIAA Sonic Boom Prediction Workshop Oblique Shock/Plume Interaction; Summary.

Description: An annular linear induction pump (ALIP) that could be used for circulating liquid-metal coolant in a fission surface power reactor system is modeled in the present work using the computational COMSOL Multiphysics package. The pump is modeled using a two-dimensional, axisymmetric geometry and solved under conditions similar to those used during experimental pump testing. Real, nonlinear, temperature-dependent material properties can be incorporated into the model for both the electrically-conducting working fluid in the pump (NaK-78) and structural components of the pump. The intricate three-phase coil configuration of the pump is implemented in the model to produce an axially-traveling magnetic wave that is qualitatively similar to the measured magnetic wave. The model qualitatively captures the expected feature of a peak in efficiency as a function of flow rate.

Description: High efficiency of rocket propul-sion systems is essential for humanity to venture be-yond the moon. Nuclear Thermal Propulsion (NTP) is a promising alternative to conventional chemical rock-ets with relatively high thrust and twice the efficiency of the Space Shuttle Main Engine. NASA is in the pro-cess of developing a new NTP engine, and is evaluat-ing ground test facility concepts that allow for the thor-ough testing of NTP devices. NTP engine exhaust, hot gaseous hydrogen, is nominally expected to be free of radioactive byproducts from the nuclear reactor; how-ever, it has the potential to be contaminated due to off-nominal engine reactor performance. Several options are being investigated to mitigate this hazard potential with one option in particular that completely contains the engine exhaust during engine test operations. The exhaust products are subsequently disposed of between engine tests. For this concept (see Figure 1), oxygen is injected into the high-temperature hydrogen exhaust that reacts to produce steam, excess oxygen and any trace amounts of radioactive noble gases released by off-nominal NTP engine reactor performance. Water is injected to condense the potentially contaminated steam into water. This water and the gaseous oxygen (GO2) are subsequently passed to a containment area where the water and GO2 are separated into separate containment tanks.

Description: Eagleworks Laboratory is an advanced propulsions physics laboratory with two primary investigations currently underway. The first is a Quantum Vacuum Plasma Thruster (QVPT or Q-thrusters), an advanced electric propulsion technology in the development and demonstration phase. The second investigation is in Warp Field Interferometry (WFI). This is an investigation of Dr. Harold "Sonny" White's theoretical physics models for warp field equations using optical experiments in the Electro Optical laboratory (EOL) at Johnson Space Center. These investigations are pursuing technology necessary to enable human exploration of the solar system and beyond.

Description: Several experiments which have reported a delay of transition are analyzed in terms of the frequencies of the induced disturbances generated by different flow control elements. Two of the experiments employed passive stabilizers in the boundary layer, one leading-edge bluntness, and one employed an active spark discharge in the boundary layer. It is found that the frequencies generated by the various elements lie in the damping region of the associated stability curve. It is concluded that the creation of strong disturbances in the damping region stabilizes the boundary-layer and delays the transition from laminar to turbulent flow.

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Description: A modular aerospike engine concept has been developed with the objective of demonstrating the viability of the aerospike design using additive manufacturing techniques. The aerospike system is a self-compensating design that allows for optimal performance over the entire flight regime and allows for the lowest possible mass vehicle designs. At low altitudes, improvements in Isp can be traded against chamber pressure, staging, and payload. In upper stage applications, expansion ratio and engine envelope can be traded against nozzle efficiency. These features provide flexibility to the System Designer optimizing a complete vehicle stage. The aerospike concept is a good example of a component that has demonstrated improved performance capability, but traditionally has manufacturing requirements that are too expensive and complex to use in a production vehicle. In recent years, additive manufacturing has emerged as a potential method for improving the speed and cost of building geometrically complex components in rocket engines. It offers a reduction in tooling overhead and significant improvements in the integration of the designer and manufacturing method. In addition, the modularity of the engine design provides the ability to perform full scale testing on the combustion devices outside of the full engine configuration. The proposed design uses a hydrocarbon based gas-generator cycle, with plans to take advantage of existing powerhead hardware while focusing DDT&E resources on manufacturing and sub-system testing of the combustion devices. The major risks for the modular aerospike concept lie in the performance of the propellant feed system, the structural integrity of the additive manufactured components, and the aerodynamic efficiency of the exhaust flow.

Description: For the past year, NASA Marshall Space Flight Center and Johnson Space Center have been working on a government version of a lunar lander design for the Resource Prospector Mission. A propulsion cold flow test system, representing an early flight design of the propulsion system, has been fabricated. The primary objective of the cold flow test is to simulate the Resource Prospector propulsion system operation through water flow testing and obtain data for anchoring analytical models. This effort will also provide an opportunity to develop a propulsion system mockup to examine hardware integration to a flight structure. This paper will report the work progress of the propulsion cold flow test system development and test preparation. At the time this paper is written, the initial waterhammer testing is underway. The initial assessment of the test data suggests that the results are as expected and have a similar trend with the pretest prediction. The test results will be reported in a future conference.

Description: A method has been developed to create an airfoil robust enough to operate satisfactorily in different environments. This method determines a robust, optimal, subsonic airfoil shape, beginning with an arbitrary initial airfoil shape, and imposes the necessary constraints on the design. Also, this method is flexible and extendible to a larger class of requirements and changes in constraints imposed.

Description: The Marshall Space Flight Center's Propulsion Systems Department has gained significant experience in the last year designing, building, and testing liquid engine components using additive manufacturing. The department has developed valve, duct, turbo-machinery, and combustion device components using this technology. Many valuable lessons were learned during this process. These lessons will be the focus of this presentation. We will present criteria for selecting part candidates for additive manufacturing. Some part characteristics are 'tailor made' for this process. Selecting the right parts for the process is the first step to maximizing productivity gains. We will also present specific lessons we learned about feature geometry that can and cannot be produced using additive manufacturing machines. Most liquid engine components were made using a two-step process. The base part was made using additive manufacturing and then traditional machining processes were used to produce the final part. The presentation will describe design accommodations needed to make the base part and lessons we learned about which features could be built directly and which require the final machine process. Tolerance capabilities, surface finish, and material thickness allowances will also be covered. Additive Manufacturing can produce internal passages that cannot be made using traditional approaches. It can also eliminate a significant amount of manpower by reducing part count and leveraging model-based design and analysis techniques. Information will be shared about performance enhancements and design efficiencies we experienced for certain categories of engine parts.

Description: No abstract available

Description: In this study, attributes are described for solid and liquid propulsion systems based on historical data. This study is not intended to compare liquid and solid propulsion system attributes, rather to present options for their use in various mission scenarios. US launch vehicle data from 1970 to 2008 was analyzed to assess solid and liquid propulsion development cost and schedule characteristics, performance features, and safety and mission success attributes. The study assessed historical trends for liquid and solid systems, and investigated implications of those trends. It was found that the two propulsion technologies have unique, common and complementary attributes that can be leveraged to meet mission requirements.

Description: No abstract available

Description: The Black Brant variation of the Standard Brant developed in the 1960's has been a workhorse motor of the NASA Sounding Rocket Project Office (SRPO) since the 1970's. In March 2012, the Black Brant Mk1 used on mission 36.277 experienced combustion instability during a flight at White Sands Missile Range, the third event in the last four years, the first occurring in November, 2009, the second in April 2010. After the 2010 event the program has been increasing the motor's throat diameter post-delivery with the goal of lowering the chamber pressure and increasing the margin against combustion instability. During the most recent combustion instability event, the vibrations exceeded the qualification levels for the Flight Termination System. The present study utilizes data generated from T-burner testing of multiple Black Brant propellants at the Naval Air Warfare Center at China Lake, to improve the combustion stability predictions for the Black Brant Mk1 and to generate new predictions for the Mk2. Three unique one dimensional (1-D) stability models were generated, representing distinct Black Brant flights, two of which experienced instabilities. The individual models allowed for comparison of stability characteristics between various nozzle configurations. A long standing "rule of thumb" states that increased stability margin is gained by increasing the throat diameter. In contradiction to this experience based rule, the analysis shows that little or no margin is gained from a larger throat diameter. The present analysis demonstrates competing effects resulting from an increased throat diameter accompanying a large response function. As is expected, more acoustic energy was expelled through the nozzle, but conversely more acoustic energy was generated due to larger gas velocities near the propellant surfaces.

Description: A device used in making differential measurements of a flow includes a flow obstruction and a support arm. The flow obstruction's forward portion is a nose cone. The flow obstruction's aft portion is coupled to the nose cone. The support arm's first end is coupled to an exterior wall of a conduit, and its second end is coupled to the forward portion of the flow obstruction. The support arm positions the flow obstruction in the conduit such that a flow region is defined around its nose cone, and such that the support arm's first and second end are separated from one another with respect to a length dimension of the conduit. Measurement ports are provided in the support arm and flow obstruction. Manifolds extending through the flow obstruction and support arm couple the ports to points at the exterior wall of the conduit.

Description: A multi-element airfoil system includes an airfoil element having a leading edge region and a skin element coupled to the airfoil element. A slat deployment system is coupled to the slat and the skin element, and is capable of deploying and retracting the slat and the skin element. The skin element substantially fills the lateral gap formed between the slat and the airfoil element when the slat is deployed. The system further includes an uncoupling device and a sensor to remove the skin element from the gap based on a critical angle-of-attack of the airfoil element. The system can alternatively comprise a trailing edge flap, where a skin element substantially fills the lateral gap between the flap and the trailing edge region of the airfoil element. In each case, the skin element fills a gap between the airfoil element and the deployed flap or slat to reduce airframe noise.

Description: A method for determining the optimum inlet geometry of a liquid rocket engine swirl injector includes obtaining a throttleable level phase value, volume flow rate, chamber pressure, liquid propellant density, inlet injector pressure, desired target spray angle and desired target optimum delta pressure value between an inlet and a chamber for a plurality of engine stages. The tangential inlet area for each throttleable stage is calculated. The correlation between the tangential inlet areas and delta pressure values is used to calculate the spring displacement and variable inlet geometry. An injector designed using the method includes a plurality of geometrically calculated tangential inlets in an injection tube; an injection tube cap with a plurality of inlet slots slidably engages the injection tube. A pressure differential across the injector element causes the cap to slide along the injection tube and variably align the inlet slots with the tangential inlets.

Description: The advanced modified high performance synthetic jet actuator with optimized curvature shape chamber (ASJA-M) is a synthetic jet actuator (SJA) with a lower volume reservoir or chamber. A curved chamber is used, instead of the conventional cylinder chamber, to reduce the dead volume of the jet chamber and increase the efficiency of the synthetic jet actuator. The shape of the curvature corresponds to the maximum displacement (deformation) profile of the electroactive diaphragm. The jet velocity and mass flow rate for the ASJA-M will be several times higher than conventional piezoelectric actuators.

Description: This project investigated a new type of small spacecraft propulsion for attitude control, specifically proximity and precision pointing control. Plasmonic force propulsion uses solar light focused on deep-subwavelength nanostructures to excite strong optical forces that accelerate and expel nanoparticle propellant. The goal of the project was to assess the feasibility of plasmonic force propulsion for nano/pico-satellite applications by evaluating key mission parameters for a nano/pico-satellite using plasmonic force propulsion in a NASA-relevant mission context. We achieved this goal and objective by evaluating plasmonic force propulsion within a NASA mission that required attitude control and precision pointing of a small satellite. We numerically simulated plasmonic force fields with asymmetric/gradient geometry and relevant solar light constraints, predicted nanoparticle velocity, mass flow rate, and resulting propulsion performance (thrust, specific impulse), and evaluated spacecraft position control resolution and pointing precision. Additionally we compared the precision pointing capabilities of plasmonic propulsion, as well as the mass, volume, and power requirements, with other state-of-the-art control techniques, such as reaction wheels and colloid/electrospray electric propulsion. The results are very exciting. Plasmonic force propulsion can significantly enhance the state-of-the-art in small spacecraft position and attitude control by 1-2 orders of magnitude. This is most succinctly shown in the figure below, which compares proximity and attitude control of plasmonic force propulsion (PFP) with other state-of-the-art thruster systems (CAT, VAT, electrospray). Additionally this figure also shows the proximity and attitude control required for different existing (James Webb Space Telescope, Hubble) and future (LISA and Stellar Imager) NASA missions. While some of these NASA missions are not small spacecraft missions, the requirements serve to illustrate the fact that more precise proximity and attitude control will be required for future NASA science missions. Stellar imager is a proposed NASA missions that requires an extremely high pointing precision of 0.1 milliarcseconds (2.7x10(exp -7) deg.) for an ultraviolet telescope that has over 200 the resolution of the Hubble Space Telescope, is able to take images showing details on the surfaces of other stars, consists of 20-30 small "mirror sats" flying in formation to produce a giant mirror, and requires each mirror-sat to be placed with nanometer precision and control its attitude with milliarcsecond precision.

Description: For more than five decades, Radioisotope Power Systems (RPS) have played a critical role in the exploration of space, enabling missions of scientific discovery to destinations across the solar system by providing electrical power to explore remote and challenging environments - some of the hardest to reach, darkest, and coldest locations in the solar system. In particular, RPS has met the demand of many long-duration mission concepts for continuous power to conduct science investigations independent of change in sunlight or variations in surface conditions like shadows, thick clouds, or dust.

Description: The use of tow steered composites, where fibers follow prescribed curvilinear paths within a laminate, can improve upon existing capabilities related to aeroelastic tailoring of wing structures, though this tailoring method has received relatively little attention in the literature. This paper demonstrates the technique for both a simple cantilevered plate in low-speed flow, as well as the wing box of a full-scale high aspect ratio transport configuration. Static aeroelastic stresses and dynamic flutter boundaries are obtained for both cases. The impact of various tailoring choices upon the aeroelastic performance is quantified: curvilinear fiber steering versus straight fiber steering, certifiable versus noncertifiable stacking sequences, a single uniform laminate per wing skin versus multiple laminates, and identical upper and lower wing skins structures versus individual tailoring.

Description: NASA needs a method of not only propelling and rotating small satellites, but also to track their position and orientation. We propose a concept that will, for the first time, demonstrate both tracking and propulsion simultaneously in the same system.

Description: This is part 2 of a two part document. Part 1 is titled: "Aeroservoelastic Testing of Free Flying Wind Tunnel Models Part 1: A Sidewall Supported Semispan Model Tested for Gust Load Alleviation and Flutter Suppression." A team comprised of the Air Force Research Laboratory (AFRL), Boeing, and the NASA Langley Research Center conducted three aeroservoelastic wind tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, flexible vehicles. In the first of these three tests, a full-span, aeroelastically scaled, wind tunnel model of a joined wing SensorCraft vehicle was mounted to a force balance to acquire a basic aerodynamic data set. In the second and third tests, the same wind tunnel model was mated to a new, two degree of freedom, beam mount. This mount allowed the full-span model to translate vertically and pitch. Trimmed flight at10 percent static margin and gust load alleviation were successfully demonstrated. The rigid body degrees of freedom required that the model be flown in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free-flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort. The balance and free flying wind tunnel tests will be summarized. The design of the trim and gust load alleviation control laws along with the associated results will also be discussed.

Description: The dual-bell rocket nozzle was first proposed in 1949, offering a potential improvement in rocket nozzle performance over the conventional-bell nozzle. Despite the performance advantages that have been predicted, both analytically and through static test data, the dual-bell nozzle has still not been adequately tested in a relevant flight environment. In 2013 a proposal was constructed that offered a National Aeronautics and Space Administration (NASA) F-15 airplane as the flight testbed, with the plan to operate a dual-bell rocket nozzle during captive-carried flight. If implemented, this capability will permit nozzle operation into an external flow field similar to that of a launch vehicle, and facilitate an improved understanding of dual-bell nozzle plume sensitivity to external flow-field effects. More importantly, this flight testbed can be utilized to help quantify the performance benefit with the dual-bell nozzle, as well as to advance its technology readiness level. Toward this ultimate goal, this report provides plans for future flights to quantify the external flow field of the airplane near the nozzle experiment, as well as details on the conceptual design for the dual-bell nozzle cold-flow propellant feed system integration within the NASA F-15 Propulsion Flight Test Fixture. The current study shows that this concept of flight research is feasible, and could result in valuable flight data for the dual-bell nozzle.

Description: No abstract available

Description: Computational Fluid Dynamics (CFD) simulations were completed for a series of convergent nozzles in participation of the American Institute of Aeronautics and Astronautics (AIAA) 1st Propulsion Aerodynamics Workshop. The simulations were performed using the Wind-US flow solver. Discharge and thrust coefficients were computed for four axisymmetric nozzles with nozzle pressure ratios (NPR) ranging from 1.4 to 7.0. The computed discharge coefficients showed excellent agreement with available experimental data; the computed thrust coefficients captured trends observed in the experimental data, but over-predicted the thrust coefficient by 0.25 to 1.0 percent. Sonic lines were computed for cases with NPR 〉= 2.0 and agreed well with experimental data for NPR 〉= 2.5. Simulations were also performed for a 25 deg. conic nozzle bifurcated by a flat plate at NPR = 4.0. The jet plume shock structure was compared with and without the splitter plate to the experimental data. The Wind-US simulations predicted the shock structure well, though lack of grid resolution in the plume reduced the sharpness of the shock waves. Unsteady Reynolds-Averaged Navier-Stokes (URANS) simulations and Detached Eddy Simulations (DES) were performed at NPR = 1.6 for the 25 deg conic nozzle with splitter plate. The simulations predicted vortex shedding from the trailing edge of the splitter plate. However, the vortices of URANS and DES solutions appeared to dissipate earlier than observed experimentally. It is believed that a lack of grid resolution in the region of the vortex shedding may have caused the vortices to break down too soon

Description: The primary objective of this trade study report is to explore the potential of using Active Flow Control (AFC) for achieving lighter and mechanically simpler high-lift systems for transonic commercial transport aircraft. This assessment was conducted in four steps. First, based on the Common Research Model (CRM) outer mold line (OML) definition, two high-lift concepts were developed. One concept, representative of current production-type commercial transonic transports, features leading edge slats and slotted trailing edge flaps with Fowler motion. The other CRM-based design relies on drooped leading edges and simply hinged trailing edge flaps for high-lift generation. The relative high-lift performance of these two high-lift CRM variants is established using Computational Fluid Dynamics (CFD) solutions to the Reynolds-Averaged Navier-Stokes (RANS) equations for steady flow. These CFD assessments identify the high-lift performance that needs to be recovered through AFC to have the CRM variant with the lighter and mechanically simpler high-lift system match the performance of the conventional high-lift system. Conceptual design integration studies for the AFC-enhanced high-lift systems were conducted with a NASA Environmentally Responsible Aircraft (ERA) reference configuration, the so-called ERA-0003 concept. These design trades identify AFC performance targets that need to be met to produce economically feasible ERA-0003-like concepts with lighter and mechanically simpler high-lift designs that match the performance of conventional high-lift systems. Finally, technical challenges are identified associated with the application of AFC-enabled highlift systems to modern transonic commercial transports for future technology maturation efforts.

Description: Exploration of our solar system has brought great knowledge to our nation's scientific and engineering community over the past several decades. As we expand our visions to explore new, more challenging destinations, we must also expand our technology base to support these new missions. NASA's Space Technology Mission Directorate is tasked with developing these technologies for future mission infusion and continues to seek answers to many existing technology gaps. One such technology gap is related to compact power systems (greater than 1 kWe) that provide abundant power for several years where solar energy is unavailable or inadequate. Below 1 kWe, Radioisotope Power Systems have been the workhorse for NASA and will continue, assuming its availability, to be used for lower power applications similar to the successful missions of Voyager, Ulysses, New Horizons, Cassini, and Curiosity. Above 1 kWe, fission power systems become an attractive technology offering a scalable modular design of the reactor, shield, power conversion, and heat transport subsystems. Near term emphasis has been placed in the 1-10kWe range that lies outside realistic radioisotope power levels and fills a promising technology gap capable of enabling both science and human exploration missions. History has shown that development of space reactors is technically, politically, and financially challenging and requires a new approach to their design and development. A small team of NASA and DOE experts are providing a solution to these enabling FPS technologies starting with the lowest power and most cost effective reactor series named "Kilopower" that is scalable from approximately 1-10 kWe.

Description: Plume impingement analyses were performed for the European Service Module (ESM) propulsion system Orbital Maneuvering System engine (OMS-E), auxiliary engines, and reaction control system (RCS) engines. The heat flux from plume impingement on the solar arrays and other surfaces are evaluated. This information is used to provide inputs for the ESM thermal analyses and help determine the optimal configuration for the RCS engines.

Description: A high efficiency radioisotope power system is being developed for long-duration NASA space science missions. The U.S. Department of Energy (DOE) managed a flight contract with Lockheed Martin Space Systems Company (LMSSC) to build Advanced Stirling Radioisotope Generators (ASRGs), with support from NASA Glenn Research Center (GRC). Sunpower Inc. held two parallel contracts to produce Advanced Stirling Convertors (ASCs), one with DOELockheed Martin to produce ASC-F flight units, and one with GRC for the production of ASC-E3 engineering unit pathfinders that are built to the flight design. In support of those contracts, GRC provided testing, materials expertise, government furnished equipment, inspections, and related data products to DOELockheed Martin and Sunpower. The technical support includes material evaluations, component tests, convertor characterization, and technology transfer. Material evaluations and component tests have been performed on various ASC components in order to assess potential life-limiting mechanisms and provide data for reliability models. Convertor level tests have been used to characterize performance under operating conditions that are representative of various mission conditions. Technology transfers enhanced contractor capabilities for specialized production processes and tests. Despite termination of flight ASRG contract, NASA continues to develop the high efficiency ASC conversion technology under the ASC-E3 contract. This paper describes key government furnished services performed for ASRG and future tests used to provide data for ongoing reliability assessments.

Description: The measured aerodynamic performance of a compact, high work factor, single-stage centrifugal compressor, comprising an impeller, diffuser, 90-bend, and exit guide vane (EGV), is reported. Performance levels are based on steady-state total-pressure and total-temperature rake and angularity-probe data acquired at key machine rating planes during recent testing at NASA Glenn Research Center. Aerodynamic performance at the stage level are reported for operation between 70 to 105 of design corrected speed, with subcomponent (impeller, diffuser, and exitguide-vane) detailed flow field measurements presented and discussed at the 100 design-speed condition. Individual component losses from measurements are compared with pre-test predictions on a limited basis.

Description: NASA is developing mission concepts for a solar electric propulsion technology demonstration mission. A number of mission concepts are being evaluated including ambitious missions to near Earth objects. The demonstration of a high-power solar electric propulsion capability is one of the objectives of the candidate missions under consideration. In support of NASAs exploration goals, a number of projects are developing extensible technologies to support NASAs near and long term mission needs. Specifically, the Space Technology Mission Directorate Solar Electric Propulsion Technology Demonstration Mission project is funding the development of a 12.5-kW magnetically shielded Hall thruster system to support future NASA missions. This paper presents the design attributes of the thruster that was collaboratively developed by the NASA Glenn Research Center and the Jet Propulsion Laboratory. The paper provides an overview of the magnetic, plasma, thermal, and structural modeling activities that were carried out in support of the thruster design. The paper also summarizes the results of the functional tests that have been carried out to date. The planned thruster performance, plasma diagnostics (internal and in the plume), thermal, wear, and mechanical tests are outlined.

Description: The NASA Space Technology mission Directorate's (STMD) Green Propellant Infusion Mission (GPIM) Technology Demonstration Mission (TDM) will demonstrate an operational AF-M315E green propellant propulsion system. Aerojet-Rocketdyne is responsible for the development of the propulsion system payload. This paper statuses the propulsion system module development, including thruster design and system design; Initial test results for the 1N engineering model thruster are presented. The culmination of this program will be high-performance, green AF-M315E propulsion system technology at TRL 7+, with components demonstrated to TRL 9, ready for direct infusion to a wide range of applications for the space user community.

Description: The components required for an in-space iodine vapor-fed Hall effect thruster propellant management system are described. A laboratory apparatus was assembled and used to produce iodine vapor and control the flow through the application of heating to the propellant reservoir and through the adjustment of the opening in a proportional flow control valve. Changing of the reservoir temperature altered the flowrate on the timescale of minutes while adjustment of the proportional flow control valve changed the flowrate immediately without an overshoot or undershoot in flowrate with the requisite recovery time associated with thermal control systems. The flowrates tested spanned a range from 0-1.5 mg/s of iodine, which is sufficient to feed a 200-W Hall effect thruster.

Description: The use of iodine propellant for Hall thrusters has been studied and proposed by multiple organizations due to the potential mission benefits over xenon. In 2013, NASA Marshall Space Flight Center competitively selected a project for the maturation of an iodine flight operational feed system through the Technology Investment Program. Multiple partnerships and collaborations have allowed the team to expand the scope to include additional mission concept development and risk reduction to support a flight system demonstration, the iodine Satellite (iSAT). The iSAT project was initiated and is progressing towards a technology demonstration mission preliminary design review. The current status of the mission concept development and risk reduction efforts in support of this project is presented.

Description: This paper makes no attempt to comprehensively review erosive burning models or the data collected in pursuit of them; the interested reader could begin with Landsbaum for a historical summary. However, a discussion and comparison to recent work by McDonald and Rettenmaier and Heister will be included, along with data generated by Strand, et. al. Suffice it to say that the search for a way to predict erosive burning in any size motor with formulas cleanly applicable to a typical 1D ballistics analysis has been long thwarted. Some models were based on testing that failed to adequately simulate the solid rocket motor environment. In most cases, no realtime burn rate measurement was available. Two popular models, even when calibrated to recent motorlike realtime burn rate data obtained by Furfaro, were shown by McMillin to be inadequate at modeling erosive burning in the Space Shuttle Reusable Solid Rocket Motor (RSRM), the Space Launch Systems' FiveSegment RSRM (RSRMV), and the fivesegment Engineering Test Motor (ETM)3. Subsequently to the data cited from Strand and Furfaro, additional motors of the same kind as Furfaro's were tested with RSRMV propellant, utilizing 7 segments per motor and 3 throat sizes. By measuring propellant web thickness with ultrasonic gages, the burn rate was determined at crossflow Mach numbers up to Mach 0.8. Furthermore, because of the different throat sizes in otherwise identical motors, this provides a unique look at the effect of pressure and base burn rate on the erosive response. Figure 1 shows example of the data pertaining to the high Mach motor, where the port area is initially less than the throat area. The burn rate data was processed using a smoothing method developed to reduce the noise without too severely introducing end effects that limit the range of useful data. Then, an empirical ballistics scheme was used to estimate the flow condition based on the burn rate measurements and pressure measured between each segment.

Description: Conical shell theory and piston theory aerodynamics are used to study the aeroelastic stability of the thermal protection system (TPS) on the NASA Hypersonic Inflatable Aerodynamic Decelerator (HIAD). Structural models of the TPS consist of single or multiple orthotropic conical shell systems resting on several circumferential linear elastic supports. The shells in each model may have pinned (simply-supported) or elastically-supported edges. The Lagrangian is formulated in terms of the generalized coordinates for all displacements and the Rayleigh-Ritz method is used to derive the equations of motion. The natural modes of vibration and aeroelastic stability boundaries are found by calculating the eigenvalues and eigenvectors of a large coefficient matrix. When the in-flight configuration of the TPS is approximated as a single shell without elastic supports, asymmetric flutter in many circumferential waves is observed. When the elastic supports are included, the shell flutters symmetrically in zero circumferential waves. Structural damping is found to be important in this case. Aeroelastic models that consider the individual TPS layers as separate shells tend to flutter asymmetrically at high dynamic pressures relative to the single shell models. Several parameter studies also examine the effects of tension, orthotropicity, and elastic support stiffness.

Description: The primary objective of this work was to develop and demonstrate a process for accurate and efficient uncertainty quantification and certification prediction of low-boom, supersonic, transport aircraft. High-fidelity computational fluid dynamics models of multiple low-boom configurations were investigated including the Lockheed Martin SEEB-ALR body of revolution, the NASA 69 Delta Wing, and the Lockheed Martin 1021-01 configuration. A nonintrusive polynomial chaos surrogate modeling approach was used for reduced computational cost of propagating mixed, inherent (aleatory) and model-form (epistemic) uncertainty from both the computation fluid dynamics model and the near-field to ground level propagation model. A methodology has also been introduced to quantify the plausibility of a design to pass a certification under uncertainty. Results of this study include the analysis of each of the three configurations of interest under inviscid and fully turbulent flow assumptions. A comparison of the uncertainty outputs and sensitivity analyses between the configurations is also given. The results of this study illustrate the flexibility and robustness of the developed framework as a tool for uncertainty quantification and certification prediction of low-boom, supersonic aircraft.

Description: Direct numerical simulations (DNS) of Mach 6 turbulent boundary layer with nominal freestream Mach number of 6 and Reynolds number of Re(sub T) approximately 460 are conducted at two wall temperatures (Tw/Tr = 0.25, 0.76) to investigate the generated pressure fluctuations and their dependence on wall temperature. Simulations indicate that the influence of wall temperature on pressure fluctuations is largely limited to the near-wall region, with the characteristics of wall-pressure fluctuations showing a strong temperature dependence. Wall temperature has little influence on the propagation speed of the freestream pressure signal. The freestream radiation intensity compares well between wall-temperature cases when normalized by the local wall shear; the propagation speed of the freestream pressure signal and the orientation of the radiation wave front show little dependence on the wall temperature.

Description: Solar and drag sail technology is entering the mainstream for space propulsion applications within NASA and around the world. Solar sails derive propulsion by reflecting sunlight from a large, mirror- like sail made of a lightweight, reflective material. The continuous sunlight pressure provides efficient primary propulsion, without the expenditure of propellant or any other consumable, allowing for very high V maneuvers and long-duration deep space exploration. Drag sails increase the aerodynamic drag on Low Earth Orbit (LEO) spacecraft, providing a lightweight and relatively inexpensive approach for end-of-life deorbit and reentry. Since NASA began investing in the technology in the late 1990's, significant progress has been made toward their demonstration and implementation in space. NASA's Marshall Space Flight Center (MSFC) managed the development and testing of two different 20-m solar sail systems and rigorously tested them under simulated space conditions in the Glenn Research Center's Space Power Facility at Plum Brook Station, Ohio. One of these systems, developed by L'Garde, Inc., is planned for flight in 2015. Called Sunjammer, the 38m sailcraft will unfurl in deep space and demonstrate solar sail propulsion and navigation as it flies to Earth-Sun L1. In the Flight Center (MSFC) managed the development and testing of two different 20-m solar sail systems and rigorously tested them under simulated space conditions in the Glenn Research Center's Space Power Facility at Plum Brook Station, Ohio. One of these systems, developed by L'Garde, Inc., is planned for flight in 2015. Called Sunjammer, the 38m sailcraft will unfurl in deep space and demonstrate solar sail propulsion and navigation as it flies to Earth-Sun L1. In the interim, NASA MSFC funded the NanoSail-D, a subscale drag sail system designed for small spacecraft applications. The NanoSail-D flew aboard the Fast Affordable Science and Technology SATellite (FASTSAT) in 2010, also developed by MSFC, and began its mission after it was ejected from the FASTSAT into Earth orbit, where it remained for several weeks before deorbiting as planned. NASA recently selected two small satellite missions for study as part of the Advanced Exploration Systems (AES) Program, both of which will use solar sails to enable their scientific objectives. Lunar Flashlight, managed by JPL, will search for and map volatiles in permanently shadowed Lunar craters using a solar sail as a gigantic mirror to steer sunlight into the shaded craters. The Near Earth Asteroid (NEA) Scout mission will use the sail as primary propulsion allowing it to survey and image one or more NEA's of interests for possible future human exploration. Both are being studied for possible launch in 2017. The Planetary Society's privately funded LightSail-A and -B cubesat-class spacecraft are nearly complete and scheduled for launch in 2015 and 2016, respectively. MMA Design launched their DragNet deorbit system in November 2013, which will deploy from the STPSat-3 spacecraft as an end of life deorbit system. The University of Surrey is building a suite of cubesat class drag and solar sail systems that will be launched beginning in 2015. As the technology matures, solar sails will increasingly be used to enable science and exploration missions that are currently impossible or prohibitively expensive using traditional chemical and electric rockets. For example, the NASA Heliophysics Decadal Survey identifies no less than three such missions for possible flight before the mid-2020's. Solar and drag sail propulsion technology is no longer merely an interesting theoretical possibility; it has been demonstrated in space and is now a critical technology for science and solar system exploration.