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NASA NTRS2003

Megawatt Electromagnetic Plasma Propulsion

The NASA Glenn Research Center program in megawatt level electric propulsion is centered on electromagnetic acceleration of quasi-neutral plasmas. Specific concepts currently being examined are the Magnetoplasmadynamic (MPD) thruster and the Pulsed Inductive Thruster (PIT). In the case of the MPD thruster, a multifaceted approach of experiments, computational modeling, and systems-level models of self field MPD thrusters is underway. The MPD thruster experimental research consists of a 1-10 MWe, 2 ms pulse-forming-network, a vacuum chamber with two 32 diffusion pumps, and voltage, current, mass flow rate, and thrust stand diagnostics. Current focus is on obtaining repeatable thrust measurements of a Princeton Benchmark type self field thruster operating at 0.5-1 gls of argon. Operation with hydrogen is the ultimate goal to realize the increased efficiency anticipated using the lighter gas. Computational modeling is done using the MACH2 MHD code, which can include real gas effects for propellants of interest to MPD operation. The MACH2 code has been benchmarked against other MPD thruster data, and has been used to create a point design for a 3000 second specific impulse (Isp) MPD thruster. This design is awaiting testing in the experimental facility. For the PIT, a computational investigation using MACH2 has been initiated, with experiments awaiting further funding. Although the calculated results have been found to be sensitive to the initial ionization assumptions, recent results have agreed well with experimental data. Finally, a systems level self-field MPD thruster model has been developed that allows for a mission planner or system designer to input Isp and power level into the model equations and obtain values for efficiency, mass flow rate, and input current and voltage. This model emphasizes algebraic simplicity to allow its incorporation into larger trajectory or system optimization codes. The systems level approach will be extended to the pulsed inductive thruster and other electrodeless thrusters at a future date.

Gilland, James↗

NASA NTRS2016

Photogrammetric Deflection Measurements for the Tiltrotor Test Rig (TTR) Multi-Component Rotor Balance Calibration

Calibrating the internal, multi-component balance mounted in the Tiltrotor Test Rig (TTR) required photogrammetric measurements to determine the location and orientation of forces applied to the balance. The TTR, with the balance and calibration hardware attached, was mounted in a custom calibration stand. Calibration loads were applied using eleven hydraulic actuators, operating in tension only, that were attached to the forward frame of the calibration stand and the TTR calibration hardware via linkages with in-line load cells. Before the linkages were installed, photogrammetry was used to determine the location of the linkage attachment points on the forward frame and on the TTR calibration hardware. Photogrammetric measurements were used to determine the displacement of the linkage attachment points on the TTR due to deflection of the hardware under applied loads. These measurements represent the first photogrammetric deflection measurements to be made to support 6-component rotor balance calibration. This paper describes the design of the TTR and the calibration hardware, and presents the development, set-up and use of the photogrammetry system, along with some selected measurement results.

Calibration↗

NASA NTRS2016

Photogrammetric Deflection Measurements for the Tiltrotor Test Rig (TTR) Multi-Component Rotor Balance Calibration

Calibration↗

NASA NTRS1993

A Software Package for Neural Network Applications Development

Original Backprop (Version 1.2) is an MS-DOS package of four stand-alone C-language programs that enable users to develop neural network solutions to a variety of practical problems. Original Backprop generates three-layer, feed-forward (series-coupled) networks which map fixed-length input vectors into fixed length output vectors through an intermediate (hidden) layer of binary threshold units. Version 1.2 can handle up to 200 input vectors at a time, each having up to 128 real-valued components. The first subprogram, TSET, appends a number (up to 16) of classification bits to each input, thus creating a training set of input output pairs. The second subprogram, BACKPROP, creates a trilayer network to do the prescribed mapping and modifies the weights of its connections incrementally until the training set is leaned. The learning algorithm is the 'back-propagating error correction procedures first described by F. Rosenblatt in 1961. The third subprogram, VIEWNET, lets the trained network be examined, tested, and 'pruned' (by the deletion of unnecessary hidden units). The fourth subprogram, DONET, makes a TSR routine by which the finished product of the neural net design-and-training exercise can be consulted under other MS-DOS applications.

Baran, Robert H.↗

NASA NTRSDate not supplied

Creep Testing of Vectran Yarn

NASA and industry are developing inflatable space structures to provide deployable habitable volumes for future exploration missions in space and on the surface of the Moon and Mars. These structures must withstand the constant pressure loading of multi-year missions. These missions therefore require excellent creep resistance in the constituent cordage and webbings that provide the primary structure of the vessel, which is typically constructed using high-strength synthetic materials such as Vectran. Real-time and accelerated creep testing of Vectran webbings has been performed at NASA Langley Research Center (LaRC), however, acceptable correlation between the results of the two approaches has not yet been achieved. Given the additional complexity of the webbing architecture, creep testing at the yarn level is being studied to reduce the number of interacting properties and focus on the core creep behavior to see if better correlation can be achieved between real-time and accelerated tests. To achieve this objective a creep test stand that is able to support simultaneous testing of five Vectran yarns had been previously designed and built. The test stand utilizes two clamp-wrap grips to mount the yarn and has a 10:1 load ratio, via a lever-arm, henceforth, known as the lever-arm test stand. Real-time creep testing (without measuring the strain of the yarn) in this fixture produces yarn ruptures in the gauge section. However, a large scatter in the times to failure was seen in the results across multiple loading conditions. For example, a test at 70% of the ultimate tensile strength (UTS) was found to have specimens that failed within a matter of hours and others that took over 6 weeks to fail. Prior test data on Vectran webbings had indicated high variability in the material strength, leading to higher than average scatter in the creep testing of webbings, which could translate to the scatter in creep behavior at the yarn level as well. However, there was still the possibility that the lever-arm test stand was causing the high scatter in data; specifically, due to potential variability in the loading of each specimen. Therefore, an additional creep test setup was designed and built to attempt to determine if the large scatter in the results was due to inherent material properties, the lever-arm test stand, or both. The newly designed creep test stand, henceforth known as the conventional test stand, utilized the more typical method for creep testing which involved hanging a weight from the yarn with no additional load amplification. The conventional test stand was designed and fabricated in-house at LaRC. High speed cameras were used to validate that the specimens in the conventional test stand failed in the gauge section. Once it was determined that the conventional test stand would provide valid failures in the gauge section, creep tests on the Vectran yarns were run to compare to the lever-arm test stand results. Preliminary results for a 70% UTS creep test have some specimens failing within a number of hours and some failing in excess of 6 weeks. These results are similar to those seen on the lever-arm test stand, pointing towards a large variability inherent in the material. The poster/presentation will focus on the design, fabrication, and validation of the conventional test stand as well as present preliminary results comparing the creep data from the conventional and lever-arm test stands

creep testing↗

DOE OSTI2025

A harmonized protocol to assess the single-cell performance of proton exchange membrane water electrolyzers

The harmonization of testing protocols for proton exchange membrane (PEM) electrolyzers is essential for ensuring accurate and reliable performance assessments and accelerating the development of hydrogen production technologies. This protocol provides a structured approach to PEM electrolyzer setup and testing, incorporating key considerations for test station design and single-cell characterization techniques. Polarization curves and electrochemical impedance spectroscopy (EIS) are detailed, along with best practices from academic and industry research groups to enhance data accuracy and comparability. By addressing material variability and harmonizing testing methodologies, this framework enables more precise evaluations of membrane electrode assemblies and electrolyzer components. Harmonized protocols not only streamline development efforts but also foster collaboration across institutions, ultimately supporting the commercialization of hydrogen solutions through improved stack efficiency and durability.

08 HYDROGEN↗

NASA NTRSDate not supplied

Flight Performance and Stability of Space Launch System Core Stage Thrust Vector Control

The Space Launch System (SLS) Core Stage (CS) Thrust Vector Control (TVC) system is comprised of eight mechanical feedback Shuttle heritage Type III TVC actuators and four RS-25 engines, each attached to a Shuttle heritage gimbal block/bearing. The Core Stage TVC shares vehicle control authority with the SLS 5-segment Solid Rocket Boosters (SRBs) during boost phase flight, and is the sole means of vehicle flight control during in exoatmospheric flight following SRB separation. TVC responses during Green Run Hot Fire (GRHF) testing revealed that the TVC did not meet its performance specifications. Step and frequency responses exhibited unexpected departures from prior laboratory data and modeled behavior. Post-test analysis determined that the characteristics of the structure and gimbal friction are significantly influenced by the thrust-loaded conditions, and the command avionics exhibited a small but important gain nonlinearity. Using the available test data, the design team augmented the flight control TVC models to bound the observed results and include the additional fidelity needed for vehicle flight control analysis so as to build sufficient rationale for flight certification. Prior to the Green Run tests, “simplex” linear models typically used for flight control analysis did not include gimbal friction and other nonlinearities owing to long-standing assumptions that these effects were negligible in the Shuttle Orbiter TVC system. Following the Green Run findings, simulation analysis of the flight dynamics in the time and frequency domain revealed the propensity for a flight control limit cycle oscillation (LCO) if friction and structural compliances fell near the edges of test-predicted bounds. While the “most probable” models did not predict an in-flight LCO, the SLS Program conservatively proceeded with a system-wide evaluation and ultimate acceptance of the possibility for a small amplitude, low-frequency TVC LCO in flight. A final validation of the extensive test and modeling effort occurred when the first flight of SLS successfully demonstrated the fully integrated performance of the vehicle’s TVC system This paper is the final installment in a seven-paper series surveying the design, engineering, test validation, and flight performance of the Core Stage Thrust Vector Control system. In this paper, the development of flight rationale in light of the TVC responses observed in Green Run is discussed, along with a review of the flight telemetry illustrating the correlation of the preflight predictions with the observed performance.

John H. Wall↗

DOE OSTI2024

Development of Superconducting RF Cavity in Traveling-Wave Regime at Fermilab

Niobium Superconducting RF (SRF) cavities have a theoretical peak magnetic field which limits the accelerating field to 50-60 MV/m. Presently, all SRF cavities operate in a Standing Wave (SW) resonance field in which particles experience an accelerating force alternating from zero to peak. In contrast, a resonance field in Traveling Wave (TW) mode propagates along with a structure, so particles in such field can experience a constant acceleration force and could have higher energy gain than that of SW mode. This phenomenon is defined by the cavity’s transit time factor, T. A TW structure proposed in an early study achieves T ~0.9, suggesting an increase in acceleration per structure by more than 20% compared to a SW structure (T ~0.7). The early stages of developments had been funded by several SBIR grants to Euclid Techlabs and completed in collaboration with Fermilab through a 1-cell prototype and a proof-of-principle 3-cell TW cavity. It demonstrated the TW resonance excitation at room temperature in the “as-fabricated” 3-cell structure. Here we report recent progresses and the first cryogenic testing of the 3-cell TW cavity in 2 K liquid helium at Fermilab.

72 PHYSICS OF ELEMENTARY PARTICLES AND FIELDS↗

NASA NTRS1990

Transatmospheric vehicle research

Research was conducted into the alternatives to the supersonic combustion ramjet (scramjet) engine for hypersonic flight. A new engine concept, the Oblique Detonation Wave Engine (ODWE) was proposed and explored analytically and experimentally. Codes were developed which can couple the fluid dynamics of supersonic flow with strong shock waves, with the finite rate chemistry necessary to model the detonation process. An additional study was conducted which compared the performance of a hypersonic vehicle powered by a scramjet or an ODWE. Engineering models of the overall performances of the two engines are included. This information was fed into a trajectory program which optimized the flight path to orbit. A third code calculated the vehicle size, weight, and aerodynamic characteristics. The experimental work was carried out in the Ames 20MW arc-jet wind tunnel, focusing on mixing and combustion of fuel injected into a supersonic airstream. Several injector designs were evaluated by sampling the stream behind the injectors and analyzing the mixture with an on-line mass spectrometer. In addition, an attempt was made to create a standing oblique detonation wave in the wind tunnel using hydrogen fuel. It appeared that the conditions in the test chamber were marginal for the generation of oblique detonation waves.

Adelman, Henry G.↗

NASA NTRS2004

Lunar and Planetary Science XXXV: Missions and Instruments: Hopes and Hope Fulfilled

The titles in this section include: 1) Mars Global Surveyor Mars Orbiter Camera in the Extended Mission: The MOC Toolkit; 2) Mars Odyssey THEMIS-VIS Calibration; 3) Early Science Operations and Results from the ESA Mars Express Mission: Focus on Imaging and Spectral Mapping; 4) The Mars Express/NASA Project at JPL; 5) Beagle 2: Mission to Mars - Current Status; 6) The Beagle 2 Microscope; 7) Mars Environmental Chamber for Dynamic Dust Deposition and Statics Analysis; 8) Locating Targets for CRISM Based on Surface Morphology and Interpretation of THEMIS Data; 9) The Phoenix Mission to Mars; 10) First Studies of Possible Landing Sites for the Phoenix Mars Scout Mission Using the BMST; 11) The 2009 Mars Telecommunications Orbiter; 12) The Aurora Exploration Program - The ExoMars Mission; 13) Electron-induced Luminescence and X-Ray Spectrometer (ELXS) System Development; 14) Remote-Raman and Micro-Raman Studies of Solid CO2, CH4, Gas Hydrates and Ice; 15) The Compact Microimaging Spectrometer (CMIS): A New Tool for In-Situ Planetary Science; 16) Preliminary Results of a New Type of Surface Property Measurement Ideal for a Future Mars Rover Mission; 17) Electrodynamic Dust Shield for Solar Panels on Mars; 18) Sensor Web for Spatio-Temporal Monitoring of a Hydrological Environment; 19) Field Testing of an In-Situ Neutron Spectrometer for Planetary Exploration: First Results; 20) A Miniature Solid-State Spectrometer for Space Applications - Field Tests; 21) Application of Laser Induced Breakdown Spectroscopy (LIBS) to Mars Polar Exploration: LIBS Analysis of Water Ice and Water Ice/Soil Mixtures; 22) LIBS Analysis of Geological Samples at Low Pressures: Application to Mars, the Moon, and Asteroids; 23) In-Situ 1-D and 2-D Mapping of Soil Core and Rock Samples Using the LIBS Long Spark; 24) Rocks Analysis at Stand Off Distance by LIBS in Martian Conditions; 25) Evaluation of a Compact Spectrograph/Detection System for a LIBS Instrument for In-Situ and Stand-Off Detection; 26) Analysis of Organic Compounds in Mars Analog Samples; 27) Report of the Organic Contamination Science Steering Group; 28) The Water-Wheel IR (WIR) - A Contact Survey Experiment for Water and Carbonates on Mars; 29) Mid-IR Fiber Optic Probe for In Situ Water Detection and Characterization; 30) Effects of Subsurface Sampling & Processing on Martian Simulant Containing Varying Quantities of Water; 31) The Subsurface Ice Probe (SIPR): A Low-Power Thermal Probe for the Martian Polar Layered Deposits; 32) Deploying Ground Penetrating Radar in Planetary Analog Sites to Evaluate Potential Instrument Capabilities on Future Mars Missions; 33) Evaluation of Rock Powdering Methods to Obtain Fine-grained Samples for CHEMIN, a Combined XRD/XRF Instrument; 34) Novel Sample-handling Approach for XRD Analysis with Minimal Sample Preparation; 35) A New Celestial Navigation Method for Mars Landers; 36) Mars Mineral Spectroscopy Web Site: A Resource for Remote Planetary Spectroscopy.

Source record↗

NASA NTRS2022

Design of Hypervelocity Flow Generator (HFG) and Its Diagnostics

Ground facilities for hypersonic research are a key element for successful development of aerodynamically proven hypervelocity vehicles. Design concepts and diagnostics of a hypervelocity flow generator (HFG) were made as a test platform for hypersonic/hypervelocity spacecraft models at the NASA Langley Research Center. The HFG is a hypersonic flow field generator using optically heated gas which is blown into an 80 cu m vacuum chamber. The vacuum chamber is kept at a stable vacuum pressure with a combination of three large vacuum pumps, while the HFG is in the test mode. The HFG provides a relatively small test section with approximately a 20 cm window. This facility was designed to generate 2.45 km/sec of flow speed, and potentially generate a continuous flow with the nozzle and vacuum system. The window in test section provides a direct view of the shock wave around a model in order to measure temperature, pressure, and density profile within the shock layer. One of the key test goals under this project is to understand why the emission spectra from the standing shockwave plasma predicted by Lora-Loran codes are significantly different from the measured emission spectra from the Flight Investigation Reentry Environment (FIRE II) Flight. The correct estimate of the thermal loading on the leading edge of hypersonic vehicles greatly affects the aerodynamic design, the material selection for the vehicle, and the cooling requirement and can be obtained by the precise modeling of emission spectra from shockwave plasma. However, the estimation of thermal loading is not an easy task due to complex non-equilibrium radiative process within high temperature shock layers that still falls into a category of cold plasma. Direct flight experiments are the most desirable, but not a cost-effective approach. Analysis by computational fluid dynamics (CFD) offers many test flexibilities. However, the CFD codes must be fully tested and validated with experimental data before the codes are effectively used for practical design. Large discrepancies between experiments and codes appear in hypersonic/hypervelocity flow regimes at high altitudes of 60 km ~ 90 km. This HFG facility offers some important parameters for CFD code validation, such as collision cross-sections, relaxation times, reaction rate coefficients and transportation coefficients. The HFG test facility is based on the ejection flow of high temperature gas heated over to 3500 K through a nozzle. The tungsten gas chamber of the HFG is heated up to a desired temperature by a 60 kW optical power beam source. This system consists of an optical power source, a thermal chamber, an expansion nozzle, a test section, and an 80 cu m vacuum tank. 60 kW optical input power is obtained from the 150 kW Vortek arc lamp system (by Vortek Industries, Vancouver, Canada). This optical beam is focused to heat the gas chamber within which a flow media is heated. The maximum achievable temperature of the flow medium reached approximately 3500 K or even higher but is limited by the melting point of the chamber material used. The exhaust velocity through the nozzle was determined by the stagnation temperature and the molecular weight of the working medium at the test section. To provide design parameters, a NASA Chemical Equilibrium with Application (CEA) computer program is used for the simulation of aerothermal data. This CEA program can calculate chemical equilibrium and properties of complex mixtures using shock tube parameters. For nitrogen gas at 2666 K stagnation temperature, the maximum achievable velocity at the test section is approximately 2.45 km/sec which is within the range of the thermal velocity of 8000 K shock layer. Based on the calculation through the CEA program, the design parameters of a HFG were determined and implemented for the test section that includes an expansion nozzle. The installed test section of the experimental facility can sustain a condition of a re-entry vehicle from the Space at an altitude of 60-90 km. After installation of the HFG, the system was fully tested and its operational parameters were measured. An 80 cu m vacuum chamber of HFG was set at 1 torr level to keep a stable downstream condition. The pumping time to reach the minimum vacuum pressure (~ 1 torr) at the test chamber from atmospheric pressure was approximately an hour using Kinney (MBV-14000/MB -1600/KT-300) and Stokes (Model 1772 and 412) pumps. Such a setup condition allowed a continued stable operation of the HFG experiment with flowrates through 1-, 3-, or 5-mm diameters of nozzle throat. The flow characteristics of the HFG for various operating conditions were performed using a focused Schlieren method. At 0.8 torr chamber pressure, a barrel shock was observed at the test section. With a sphere obstacle of 12 mm diameter in the flow, a bow shock (~ 2-mm thickness) was observed by a focused Schlieren visualization method.

Hypervelocity flow↗

NASA NTRS2014

Evaluation of the First Transport Rotorcraft Airframe Crash Testbed (TRACT 1) Full-Scale Crash Test

In 2012, the NASA Rotary Wing Crashworthiness Program initiated the Transport Rotorcraft Airframe Crash Testbed (TRACT) research program by obtaining two CH-46E helicopters from the Navy CH-46E Program Office (PMA-226) at the Navy Flight Readiness Center in Cherry Point, North Carolina. Full-scale crash tests were planned to assess dynamic responses of transport-category rotorcraft under combined horizontal and vertical impact loading. The first crash test (TRACT 1) was performed at NASA Langley Research Center's Landing and Impact Research Facility (LandIR), which enables the study of critical interactions between the airframe, seat, and occupant during a controlled crash environment. The CH-46E fuselage is categorized as a medium-lift rotorcraft with fuselage dimensions comparable to a regional jet or business jet. The first TRACT test (TRACT 1) was conducted in August 2013. The primary objectives for TRACT 1 were to: (1) assess improvements to occupant loads and displacement with the use of crashworthy features such as pre-tensioning active restraints and energy absorbing seats, (2) develop novel techniques for photogrammetric data acquisition to measure occupant and airframe kinematics, and (3) provide baseline data for future comparison with a retrofitted airframe configuration. Crash test conditions for TRACT 1 were 33-ft/s forward and 25-ft/s vertical combined velocity onto soft soil, which represent a severe, but potentially survivable impact scenario. The extraordinary value of the TRACT 1 test was reflected by the breadth of meaningful experiments. A total of 8 unique experiments were conducted to evaluate ATD responses, seat and restraint performance, cargo restraint effectiveness, patient litter behavior, and photogrammetric techniques. A combination of Hybrid II, Hybrid III, and ES-2 Anthropomorphic Test Devices (ATDs) were placed in forward and side facing seats and occupant results were compared against injury criteria. Loads from ATDs in energy absorbing seats and restraints were within injury limits. Severe injury was likely for ATDs in forward facing passenger seats, legacy troop bench seats, and a three-tiered patient litter. In addition, two standing ATDs were used to evaluate the benefit of Mobile Aircrew Restraint Systems (MARS) versus a standard gunner's belt. The ATD with the MARS survived the impact, while fatal head blunt trauma occurred for the standing ATD held by the legacy gunner's belt. In addition to occupant loading, the structural response of the airframe was assessed based on accelerometers located throughout the airframe and using three-dimensional photogrammetric techniques. Analysis of the photogrammetric data indicated regions of maximum deflection and permanent deformation.

Annett, Martin S.↗

DOE OSTI2026

Semi-Transparent Perovskite Solar Cells in a Stacked Tandem Module: Cooperative Research and Development Final Report, CRADA Number CRD-19-00810

This project seeks to develop device design, materials composition, and processing tools and parameters to fabricate semi-transparent perovskite solar cells and modules for application in stand-alone products or added to other solar cells in a mechanically stacked tandem configuration. This technology presents significant advanced manufacturing challenges and opportunities in getting to scale, including development of perovskite inks, scalable perovskite and heterojunction deposition and annealing processes, heterojunction composition, transparent electrode composition and deposition process, anti-reflection layer composition and deposition process, and cell to module integration processes. Modification 5: The proposed project seeks to develop device design, materials composition, and processing tools and parameters to fabricate semi-transparent perovskite solar cells and modules for application in stand-alone products or added to other solar cells in a mechanically-stacked tandem configuration. This technology presents significant advanced manufacturing challenges and opportunities in getting to scale, including development of perovskite inks, scalable perovskite and heterojunction deposition and annealing processes, heterojunction composition, transparent electrode composition and deposition process, passivation layers including in module scribes, anti-reflection layer composition and deposition process, and cell to module integration processes. Advanced metrology and characterization will be performed on perovskite films, cells and module. Furthermore, we will examine module or materials recycling for circular economy considerations. Modifcation 6: Gigahertz frequency microwave pump-probe spectroscopies are highly sensitive to thin film semiconductor photoconductivity of individual and stacks of layers that comprise perovskite solar cells. As such, these techniques will be used to qualify reproducibility and quality correlations during the manufacturing process. Modification 7: Mechanical adhesion of top contacts within perovskite modules significantly impacts the durability of the module when exposed to accelerated degradation testing. The adhesion between the perovskite/transport layer interface and the transport layer/top contact interface are both very sensitive small changes in processing. ALD processing conditions of the transport layer will be tuned to optimize the mechanical adhesion within the perovskite module stack.

14 SOLAR ENERGY↗

NASA NTRS2006

Self-Passivating Lithium/Solid Electrolyte/Iodine Cells

Robust lithium/solid electrolyte/iodine electrochemical cells that offer significant advantages over commercial lithium/ iodine cells have been developed. At room temperature, these cells can be discharged at current densities 10 to 30 times those of commercial lithium/iodine cells. Moreover, from room temperature up to 80 C, the maximum discharge-current densities of these cells exceed those of all other solid-electrolyte-based cells. A cell of this type includes a metallic lithium anode in contact with a commercial flexible solid electrolyte film that, in turn, is in contact with an iodine/ graphite cathode. The solid electrolyte (the chemical composition of which has not been reported) offers the high ionic conductivity needed for high cell performance. However, the solid electrolyte exhibits an undesirable chemical reactivity to lithium that, if not mitigated, would render the solid electrolyte unsuitable for use in a lithium cell. In this cell, such mitigation is affected by the formation of a thin passivating layer of lithium iodide at the anode/electrolyte interface. Test cells of this type were fabricated from iodine/graphite cathode pellets, free-standing solid-electrolyte films, and lithium-foil anodes. The cathode mixtures were made by grinding together blends of nominally 10 weight percent graphite and 90 weight percent iodine. The cathode mixtures were then pressed into pellets at 36 kpsi (248 MPa) and inserted into coin-shaped stainless-steel cell cases that were coated with graphite paste to minimize corrosion. The solid-electrolyte film material was stamped to form circular pieces to fit in the coin cell cases, inserted in the cases, and pressed against the cathode pellets with polyethylene gaskets. Lithium-foil anodes were placed directly onto the electrolyte films. The layers described thus far were pressed and held together by stainless- steel shims, wave springs, and coin cell caps. The assembled cells were then crimped to form hermetic seals. It was found that the solid electrolyte films became discolored within seconds after they were placed in contact with the cathodes - a result of facile diffusion of iodine through the solid electrolyte material (see figure).

Bugga, Ratnakumar↗

NASA NTRS2000

Effects of Melt Convection and Solid Transport on Macrosegregation and Grain Structure in Equiaxed Al-Cu Alloys

Macrosegregation in metal casting can be caused by thermal and solutal melt convection, and the transport of unattached solid crystals resulting from nucleation in the bulk liquid or dendrite fragmentation. To develop a comprehensive numerical model for the casting of alloys, an experimental study has been conducted to generate benchmark data with which such a solidification model could be tested. The objectives were: (1) experimentally study the effects of solid transport and thermosolutal convection on macrosegregation and grain size; and (2) provide a complete set of boundary conditions temperature data, segregation data, and grain size data - to validate numerical models. Through the control of end cooling and side wall heating, radial temperature gradients in the sample and furnace were minimized. Thus the vertical crucible wall was adiabatic. Samples at room temperature were 24 cc and 95 mm long. The alloys used were Al-1 wt. pct. Cu, and Al- 10 wt. pct. Cu; the starting point for solidification was isothermal at 710 and 685 C respectively. To induce an equiaxed structure various amounts of the grain refiner TiB2 were added. Samples were either cooled from the top, or the bottom. Several trends in the data stand out. In attempting to model these experiments, concentrating on these trends or differences may be beneficial.

Rerko, Rodney S.↗

NASA NTRS2007

Compact and Integrated Liquid Bismuth Propellant Feed System

Operation of Hall thrusters with bismuth propellant has been shown to be a promising path toward high-power, high-performance, long-lifetime electric propulsion for spaceflight missions [1]. There has been considerable effort in the past three years aimed at resuscitating this promising technology and validating earlier experimental results indicating the advantages of a bismuth-fed Hall thruster. A critical element of the present effort is the precise metering of propellant to the thruster, since performance cannot be accurately assessed without an accurate accounting of mass flow rate. Earlier work used a pre./post-test propellant weighing scheme that did not provide any real-time measurement of mass flow rate while the thruster was firing, and makes subsequent performance calculations difficult. The motivation of the present work is to develop a precision liquid bismuth Propellant Management System (PMS) that provides hot, molten bismuth to the thruster while simultaneously monitoring in real-time the propellant mass flow rate. The system is a derivative of our previous propellant feed system [2], but the present system represents a more compact design. In addition, all control electronics are integrated into a single unit and designed to reside on a thrust stand and operate in the relevant vacuum environment where the thruster is operating, significantly increasing the present technology readiness level of liquid metal propellant feed systems. The design of various critical components in a bismuth PMS are described. These include the bismuth reservoir and pressurization system, 'hotspot' flow sensor, power system and integrated control system. Particular emphasis is given to selection of the electronics employed in this system and the methods that were used to isolate the power and control systems from the high-temperature portions of the feed system and thruster. Open loop calibration test results from the 'hotspot' flow sensor are reported, and results of integrated thruster/PMS tests demonstrate operation of the feed system in the relevant environment.

Polzin, Kurt A.↗

NASA NTRS2017

Development of an IVE/EVA Compatible Prototype Cold-Gas Cubesat Propulsion System at NASA/JSC

Cold-gas propulsion systems are well suited for some applications because they are simple to design and build, have low operating costs, and are non-toxic. The inherent tradeoff, however, is their relatively low impulse density. Nevertheless, a modest propulsion system, sized for Cubesats and designed for affordability, presents an attractive system solution for some missions, such as an on-orbit inspection free-flyer. NASA has a long-standing effort to develop propulsion systems appropriate for very high delta-V cubesat missions, such as geo transfer orbits, and there are commercially available Cubesat propulsion systems with considerably more impulse capability, but, these are both prohibitively expensive for some development customers and face compatibility constraints for crewed applications, such as operation within ISS. A relatively conventional cold-gas system has been developed at NASA/JSC taking advantage of existing miniature industrial components, additive manufacturing techniques and in-house qualification of the system. The result is a nearly modular system with a 1U form factor. Compressed nitrogen is stored in a small high-pressure tank, then regulated and distributed to 12 thrusters. Maneuvering thrust can be adjusted, with a typical value of 40 mN, and the delta-V delivered to a 3U Cubesat would be approximately 7 m/s. These values correspond to the performance parameters for an inspection mission previously established at JSC for inspection of the orbiter prior to reentry. Environmental testing was performed to meet ISS launch and workmanship standards, along with the expected thermal environment for an inspection mission. Functionality has been demonstrated, and performance in both vacuum and relevant blow down scenarios was completed. Several avenues for further improvement are also explored. Details of the system, components, integration, tests, and test data are presented in this paper.

Radke, Christopher↗

NASA NTRS1995

Development of an L-, C-, and X-band radar for backscattering studies over vegetation

With the recent surge of interest in global change, the impact of different ecosystems on the Earth's carbon budget has become the focus of many scientific studies. Studies have been launched by NASA and other agencies to address this issue. One such study is the Boreal Ecosystem-Atmosphere Study (BOREAS). BOREAS focuses on the boreal ecosystem in Northern Canada. As a part of the BOREAS study, we have developed a helicopter-borne three-band radar system for measuring the scattering coefficient of various stands within the boreal forest. During the summer of 1994 the radar was used at the southern study area (SSA) in Saskatchewan over the young jack pine (YJP), old jack pine (OJP), old black spruce (OBS) and old aspen (OA) sites. The data collected will be used to study the interaction of microwaves with forest canopy. By making use of three different frequency bands the contribution to the backscatter from each of the layers within the canopy can be determined. Using the knowledge gained from these studies, we will develop algorithms to enable more accurate interpretation of SAR images of the boreal region. This report describes in detail the development of the L-, C- and X-band radar system. The first section provides background information and explains the objectives of the boreal forest experiment. The second section describes the design and implementation of the radar system. All of the subsystems of the radar are explained in this section. Next, problems that were encountered during system testing and the summer experiments are discussed. System performance and results are then presented followed by a section on conclusions and further work.

Lockhart, G. Lance↗