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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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NASA NTRS2009

Apparatus Measures Thermal Conductance Through a Thin Sample from Cryogenic to Room Temperature

An apparatus allows the measurement of the thermal conductance across a thin sample clamped between metal plates, including thermal boundary resistances. It allows in-situ variation of the clamping force from zero to 30 lb (133.4 N), and variation of the sample temperature between 40 and 300 K. It has a special design feature that minimizes the effect of thermal radiation on this measurement. The apparatus includes a heater plate sandwiched between two identical thin samples. On the side of each sample opposite the heater plate is a cold plate. In order to take data, the heater plate is controlled at a slightly higher temperature than the two cold plates, which are controlled at a single lower temperature. The steady-state controlling power supplied to the hot plate, the area and thickness of samples, and the temperature drop across the samples are then used in a simple calculation of the thermal conductance. The conductance measurements can be taken at arbitrary temperatures down to about 40 K, as the entire setup is cooled by a mechanical cryocooler. The specific geometry combined with the pneumatic clamping force control system and the steady-state temperature control approach make this a unique apparatus.

Tuttle, James G.↗
NASA NTRS2010

Fuel-Cell Power Source Based on Onboard Rocket Propellants

The use of onboard rocket propellants (dense liquids at room temperature) in place of conventional cryogenic fuel-cell reactants (hydrogen and oxygen) eliminates the mass penalties associated with cryocooling and boil-off. The high energy content and density of the rocket propellants will also require no additional chemical processing. For a 30-day mission on the Moon that requires a continuous 100 watts of power, the reactant mass and volume would be reduced by 15 and 50 percent, respectively, even without accounting for boiloff losses. The savings increase further with increasing transit times. A high-temperature, solid oxide, electrolyte-based fuel-cell configuration, that can rapidly combine rocket propellants - both monopropellant system with hydrazine and bi-propellant systems such as monomethyl hydrazine/ unsymmetrical dimethyl hydrazine (MMH/UDMH) and nitrogen tetroxide (NTO) to produce electrical energy - overcomes the severe drawbacks of earlier attempts in 1963-1967 of using fuel reforming and aqueous media. The electrical energy available from such a fuel cell operating at 60-percent efficiency is estimated to be 1,500 Wh/kg of reactants. The proposed use of zirconia-based oxide electrolyte at 800-1,000 C will permit continuous operation, very high power densities, and substantially increased efficiency of conversion over any of the earlier attempts. The solid oxide fuel cell is also tolerant to a wide range of environmental temperatures. Such a system is built for easy refueling for exploration missions and for the ability to turn on after several years of transit. Specific examples of future missions are in-situ landers on Europa and Titan that will face extreme radiation and temperature environments, flyby missions to Saturn, and landed missions on the Moon with 14 day/night cycles.

Ganapathi, Gani↗
NASA NTRS2010

Remaining Technical Challenges and Future Plans for Oil-Free Turbomachinery

The application of Oil-Free technologies (foil gas bearings, solid lubricants and advanced analysis and predictive modeling tools) to advanced turbomachinery has been underway for several decades. During that time, full commercialization has occurred in aircraft air cycle machines, turbocompressors and cryocoolers and ever-larger microturbines. Emerging products in the automotive sector (turbochargers and superchargers) indicate that high volume serial production of foil bearings is imminent. Demonstration of foil bearings in APU s and select locations in propulsion gas turbines illustrates that such technology also has a place in these future systems. Foil bearing designs, predictive tools and advanced solid lubricants have been reported that can satisfy anticipated requirements but a major question remains regarding the scalability of foil bearings to ever larger sizes to support heavier rotors. In this paper, the technological history, primary physics, engineering practicalities and existing experimental and experiential database for scaling foil bearings are reviewed and the major remaining technical challenges are identified.

DellaCorte, Christopher↗
NASA NTRS1995

NASA Tech Briefs, November 1995

The contents include: 1) Mission Accomplished; 2) Resource Report: Marshall Space Flight Center; 3) NASA 1995 Software of the Year Award; 4) Microbolometers Based on Epitaxial YBa2Cu3O(sub 7-x) Thin Films; 5) Garnet Random-Access Memory; 6) Fabrication of SNS Weak Links on SOS Substrates; 7) High-Voltage MOSFET Switching Circuit; 8) Asymmetric Switching for a PWM H-Bridge Power Circuit; 9) Better Ohmic Contacts for InP Semiconductor Devices; 10) Low-Bandgap Thermovoltaic Materials and Devices; 11) Digital Frequency-Differencing Circuit; 12) Imaging Magnetometer; 13) Computer-Assisted Monitoring of a Complex System; 14) Buffered Telemetry Demodulator; 15) Compact Multifunction Inspection Head; 16) Optical Detection of Fractures in Ceramic Diaphragms; 17) Eddy-Current Detection of Cracks in Reinforced Carbon/Carbon; 18) Apparent Thermal Conductivity of Multilayer Insulation; 19) Optimizing Misch-Metal Compositions in Metal Hydride Anodes; 20) Device for Sampling Surface Contamination; 21) Probabilistic Failure Assessment for Fatigue; 22) Probabilistic Fatigue and Flaw-Propagation Analysis; 23) Windows Program for Driving the TDU-850 Printer; 24) Subband/Transform MATLAB Functions for Processing Images; 25) Computing Equilibrium Chemical Compositions; 26) Program Processes Thermocouple Readings; 27) ICAN-Second-Generation Integrated Composite Analyzer; 28) Integrated Composite Analyzer with Damping Capabilities; 29) Computing Efficiency of Transfer of Microwave Power; 30) Program Calculates Power Demands of Electronic Designs; 31) Cost-Estimation Program; 32) Program Estimates Areas Required by Electronic Designs; 33) Program to Balance Mapped Turbopump Assemblies; 34) BiblioTech; 35) Controlling Mirror Tilt With a Bimorph Actuator; 36) Burst-Disk Device Simulates Effect of Pyrotechnic Device; 37) Bearing-Mounting Concept Accommodates Thermal Expansion; 38) Parallel-Plate Acoustic Absorbers for Hot Environments; 39) Adjustable-Length Strut Withstands Large Cyclic Loads; 40) Tool Indicates Contact Angles in Bearing Raceways; 41) Gravity Slides With Magnetic Braking; 42) High-Torque, Lightweight, Pneumatically Driven Wrench for Small Spaces; 43) Device for Testing Compatibility of an O-Ring; 44) Magnetic Heat Pump Containing Flow Diverters; 45) Variable-Tilt Helicopter Rotor Mast; 46) "Beach-Ball" Robotic Rovers; 47) Apparatus Would Measure Temperatures of Ball Bearings; 48) Flexible Borescope for Inspecting Ducts; 49) Texturing Copper To Reduce Secondary Emission of Electrons; 50) Automated Laser Cutting in Three Dimensions; 51) Algorithm Helps Monitor Engine Operation; 52) Flexible Revision of Data-Processing Communications; 53) Software for Managing the Use of Land; 54) Thermal Strap Increases Cryocooling Efficiency; 55) Reversible Nut With Engagement Indication; 56) Control Algorithms for Kinematically Redundant Manipulators; 57) Computed Hydrogen-Flow Splits in a Rocket Engine; 58) Pressure and Thermal Modeling of Rocket Launches; 59) Field of View of a Spacecraft Antenna: Analysis and Software; 60) Digital Controller for Laser-Beam-Steering Subsystem; 61) More About Beam-Steering Subsystem for Laser Communication; 62) Digital Controller for Laser-Beam-Steering Subsystem: Part 2; 63) Interface Circuit Board for Space-Shuttle Communications; 64) Automated Planning of Spacecraft Telecommunications; 65) Artifacts of Spectral Analysis of Instrument Readings; 66) Neural-Network Controller for Vibration Suppression; 67) Adaptive Finite-Element Computation in Fracture Mechanics; 68) Attitude Control for the Cassini Spacecraft; 69) Analytical Model for Fluid Dynamics in a Microgravity Environment; 70) Study of Rocket-Engine Joints Bonded by NVCU/NARloy-Z; 71) Improved Silicon Nitride for Advanced Heat Engines; 72) Parameters for Welding Aluminum/Lithium Alloys; 73) Lightweight Composite Intertank Structure; 74) Foil Patches Seal Small Vacuum Leaks; 75) Data Base on Cables and Connectors; 76) Effect of Clock Mode on Radiation Hardnessf an ADC; and 77) Fault-Tolerant Control for a Robotic Inspection System.

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NASA NTRS2011

Broad Area Cooler Concepts for Cryogenic Propellant Tanks

Numerous studies and ground tests have shown that broad area cooling (also known as distributed cooling) can reduce or eliminate cryogenic propellant boil-off and enable long duration storage in space. Various combinations of cryocoolers, circulators, heat exchangers and other hardware could be used to build the system. In this study, several configurations of broad area cooling systems were compared by weighing hardware combinations, input power requirements, component availability, and Technical Readiness Level (TRL). The preferred system has a high TRL and can be scaled up to provide cooling capacities on the order of 150W at 90K

Christie, R. J.↗
NASA NTRS2011

Modeling a Transient Pressurization with Active Cooling Sizing Tool

As interest in the area of in-space zero boil-off cryogenic propellant storage develops, the need to visualize and quantify cryogen behavior during ventless tank self-pressurization and subsequent cool-down with active thermal control has become apparent. During the course of a mission, such as the launch ascent phase, there are periods that power to the active cooling system will be unavailable. In addition, because it is not feasible to install vacuum jackets on large propellant tanks, as is typically done for in-space cryogenic applications for science payloads, instances like the launch ascent heating phase are important to study. Numerous efforts have been made to characterize cryogenic tank pressurization during ventless cryogen storage without active cooling, but few tools exist to model this behavior in a user-friendly environment for general use, and none exist that quantify the marginal active cooling system size needed for power down periods to manage tank pressure response once active cooling is resumed. This paper describes the Transient pressurization with Active Cooling Tool (TACT), which is based on a ventless three-lump homogeneous thermodynamic self-pressurization model1 coupled with an active cooling system estimator. TACT has been designed to estimate the pressurization of a heated but unvented cryogenic tank, assuming an unavailable power period followed by a given cryocooler heat removal rate. By receiving input data on the tank material and geometry, propellant initial conditions, and passive and transient heating rates, a pressurization and recovery profile can be found, which establishes the time needed to return to a designated pressure. This provides the ability to understand the effect that launch ascent and unpowered mission segments have on the size of an active cooling system. A sample of the trends found show that an active cooling system sized for twice the steady state heating rate would results in a reasonable time for tank pressure recovery with ZBO of a liquid oxygen propellant tank.

Guzik, Monica C.↗
NASA NTRS2011

NASA Tech Briefs, May 2011

Topics covered include: 1) Method to Estimate the Dissolved Air Content in Hydraulic Fluid; 2) Method for Measuring Collimator-Pointing Sensitivity to Temperature Changes; 3) High-Temperature Thermometer Using Cr-Doped GdAlO3 Broadband Luminescence; 4)Metrology Arrangement for Measuring the Positions of Mirrors of a Submillimeter Telescope; 5) On-Wafer S-Parameter Measurements in the 325-508-GHz Band; 6) Reconfigurable Microwave Phase Delay Element for Frequency Reference and Phase-Shifter Applications; 7) High-Speed Isolation Board for Flight Hardware Testing; 8) High-Throughput, Adaptive FFT Architecture for FPGA-Based Spaceborne Data Processors; 9) 3D Orbit Visualization for Earth-Observing Missions; 10) MaROS: Web Visualization of Mars Orbiting and Landed Assets; 11) RAPID: Collaborative Commanding and Monitoring of Lunar Assets; 12) Image Segmentation, Registration, Compression, and Matching; 13) Image Calibration; 14) Rapid ISS Power Availability Simulator; 15) A Method of Strengthening Composite/Metal Joints; 16) Pre-Finishing of SiC for Optical Applications; 17) Optimization of Indium Bump Morphology for Improved Flip Chip Devices; 18) Measuring Moisture Levels in Graphite Epoxy Composite Sandwich Structures; 19) Marshall Convergent Spray Formulation Improvement for High Temperatures; 20) Real-Time Deposition Monitor for Ultrathin Conductive Films; 21) Optimized Li-Ion Electrolytes Containing Triphenyl Phosphate as a Flame-Retardant Additive; 22) Radiation-Resistant Hybrid Lotus Effect for Achieving Photoelectrocatalytic Self-Cleaning Anticontamination Coatings; 23) Improved, Low-Stress Economical Submerged Pipeline; 24) Optical Fiber Array Assemblies for Space Flight on the Lunar Reconnaissance Orbiter; 25) Local Leak Detection and Health Monitoring of Pressurized Tanks; 26) Dielectric Covered Planar Antennas at Submillimeter Wavelengths for Terahertz Imaging; 27) Automated Cryocooler Monitor and Control System; 28) Broadband Achromatic Phase Shifter for a Nulling Interferometer; 29) Super Dwarf Wheat for Growth in Confined Spaces; 30) Fine Guidance Sensing for Coronagraphic Observatories; 31) Single-Antenna Temperature- and Humidity-Sounding Microwave Receiver; 32) Multi-Wavelength, Multi-Beam, and Polarization-Sensitive Laser Transmitter for Surface Mapping; 33) Optical Communications Link to Airborne Transceiver; 34) Ascent Heating Thermal Analysis on Spacecraft Adaptor Fairings; 35) Entanglement in Self-Supervised Dynamics; 36) Prioritized LT Codes; 37) Fast Image Texture Classification Using Decision Trees; 38) Constraint Embedding Technique for Multibody System Dynamics; 39) Improved Systematic Pointing Error Model for the DSN Antennas; 40) Observability and Estimation of Distributed Space Systems via Local Information-Exchange Networks; 41) More-Accurate Model of Flows in Rocket Injectors; 42) In-Orbit Instrument-Pointing Calibration Using the Moon as a Target; 43) Reliability of Ceramic Column Grid Array Interconnect Packages Under Extreme Temperatures; 44) Six Degrees-of-Freedom Ascent Control for Small-Body Touch and Go; and 45) Optical-Path-Difference Linear Mechanism for the Panchromatic Fourier Transform Spectrometer.

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NASA NTRS2011

Landsat Data Continuity Mission

The Landsat Data Continuity Mission (LDCM) is currently under development and is on schedule to launch the 8th satellite in the Landsat series in December of 2012. LDCM is a joint project between the National Aeronautics and Space Administration (NASA) and the United States Geological Survey (USGS). NASA is responsible for developing and launching the flight hardware and on-orbit commissioning and USGS is responsible for developing the ground system and operating the system onorbit after commissioning. Key components of the flight hardware are the Operational Land Imager (OLI), nearing completion by Ball Aerospace & Technologies Corp in Boulder, CO, the Thermal Infrared Sensor (TIRS), being built by NASA's Goddard Space Flight Center and the spacecraft, undergoing integration at Orbital Sciences Corp in Gilbert, Arizona. The launch vehicle will be an Atlas-5 with launch services provided by NASA's Kennedy Space Center. Key ground systems elements are the Mission Operations Element, being developed by the Hammers Corporation, and the Collection Activity Planning Element, Ground Network Element, and Data Processing and Archive System, being developed internally by the USGS Earth Resources Observations and Science (EROS) Center. The primary measurement goal of LDCM is to continue the global coverage of moderate spatial resolution imagery providing continuity with the existing Landsat record. The science goal for this imagery is to monitor land use and land cover, particularly as it relates to global climate change. Together the OLI and TIRS instruments on LDCM replace the ETM+ instrument on Landsat-7 with significant enhancements. The OLI is a pushbroom design instrument where the scanning mechanism of the ETM+ is effectively replaced by a long line of detectors. The OLI has 9 spectral bands with similar spatial resolution to ETM+: 7 of them similar to the reflective spectral bands on ETM+ and two new bands. The two new bands cover (1) the shorter wavelength blue part of the spectrum to help with coastal studies and aerosol analyses/atmospheric correction and (2) an atmospheric water absorption band, where the Earth surface is generally not visible, but Cirrus clouds are, to aid in cloud detection and screening. The radiometry of OLI benefits from improved SNR, dynamic range and quantization. OLI is undergoing system testing with a delivery scheduled for Spring 2011. The TIRS is also a pushbroom design and used QWIPS detectors that require cooling to 43K using a cryocooler. It.has two spectral bands, effectively splitting the ETM+ band 6 in half, that can be used as a split window to aid in atmospheric correction. It has nominally 100 m spatial resolution as opposed to the 60 m of Landsat-7 ETM+: TIRS has commenced integration and test, with a delivery to the spacecraft vendor scheduled for Winter 2011-2012. The Orbital spacecraft currently being integrated for LDCM will have improved capabilities for pointing over previous missions. These capabilities will allow the OLI and TIRS instruments to point off-nadir the equivalent of one WRS-2 path to increase the chances of coverage for high priority targets, particularly in the event of natural disasters. Also, the pointing capability will allow the calibration of the OLI using the sun (roughly weekly), the moon (monthly), stars (during commissioning) and the Earth (at 90 deg from normal orientation, a.k.a., side slither) quarterly. The solar calibration will be used for OLI absolute and relative calibration, the moon for trending the stability of the OLI response, the stars will be used for Line of Sight determination and the side slither will be an alternate OLI and relative gain determination methodology. The spacecraft is scheduled to begin integration with the OLI instrument in Summer 2011. The LDCM data processing and archive system (DPAS), located at USGS EROS, generates the products for distribution to users. Like Landsat-7 this includes an image assessment system for characrizing instrument performance and updating calibration parameters. Products will be generated that include the spectral bands from both instruments, terrain corrected and registered to the geoid. Also, like Landsat-7, data products will be distributed at no charge to the user. The current status and plans of the space and ground segments of the LDCM project will be presented along with performance predictions as available. More detailed information on the two instruments is intended to be presented in separate papers.

Markham, Brian↗
NASA NTRS2011

Technique for Configuring an Actively Cooled Thermal Shield in a Flight System

Broad area cooling shields are a mass-efficient alternative to conductively cooled thermal radiation shielding. The shield would actively intercept a large portion of incident thermal radiation and transport the heat away using cryogenic helium gas. The design concept consists of a conductive and conformable surface that maximizes heat transfer and formability. Broad Area Cooled (BAC) shields could potentially provide considerable mass savings for spaceflight applications by eliminating the need for a rigid thermal radiation shield for cryogen tanks. The BAC consists of a network of capillary tubes that are thermally connected to a conductive shield material. Chilled helium gas is circulated through the network and transports unwanted heat away from the cryogen tanks. The cryogenic helium gas is pumped and chilled simultaneously using a specialized pulse-tube cryocooler, which further improves the mass efficiency of the system. By reducing the thermal environment temperature from 300 to 100 K, the radiative heat load on a cryogen tank could be reduced by an order of magnitude. For a cryogenic liquid propellant scenario of oxygen and hydrogen, the boiloff of hydrogen would be significantly reduced and completely eliminated for oxygen. A major challenge in implementing this technology on large tanks is that the BAC system must be easily scalable from lab demonstrations to full-scale missions. Also, the BAC shield must be conformable to complex shapes like spheres without losing the ability to maintain constant temperature throughout. The initial design maximizes thermal conductivity between the capillary tube and the conductive radiation shielding by using thin, corrugated aluminum foil with the tube running transverse to the folds. This configuration has the added benefit of enabling the foil to stretch and contract longitudinally. This allows the BAC to conform to the complex curvature of a cryogen tank, which is key to its success. To demonstrate a BAC shield system with minimal impact to current cryogen tank designs, the shielding must be applied after the final assembly of the tank and supporting structure. One method is to pre-fabricate the shield in long strips. A spool of corrugated aluminum foil with a thermally sunk aluminum capillary running through the center could then be simply wound around the cryogen tanks and encapsulated within the multi-layer insulation (MLI) blanket. Then, on orbit, the BAC would intercept thermal radiation coming in through the MLI and transport it away from the cryogen tanks. An optimization of the design could be done to take into account mass savings from thinner MLI blankets, eliminating solid thermal shields, and ultimately, a reduction in the required cryogen tank size.

Barkfknecht, Peter↗
NASA NTRS2011

Issues of Long-Term Cryogenic Propellant Storage in Microgravity

Modern multi-layer insulation (MLI) allows to sharply reduce the heat leak into cryogenic propellant storage tanks through the tank surface and, as a consequence, significantly extend the storage duration. In this situation the MLI penetrations, such as support struts, feed lines, etc., become one of the most significant challenges of the tanks heat management. This problem is especially acute for liquid hydrogen (LH2) storage, since currently no efficient cryocoolers exist that operate at very low LH2 temperatures (20K). Even small heat leaks under microgravity conditions and over the period of many months give rise to a complex slowly-developing, large-scale spatiotemporal physical phenomena in a multi-phase liquid-vapor mixture. These phenomena are not well-understood nor can be easily controlled. They can be of a potentially hazardous nature for long-term on-orbital cryogenic torage, propellant loading, tank chilldown, engine restart, and other in-space cryogenic fluid management operations. To support the engineering design solutions that would mitigate these effects a detailed physics-based analysis of heat transfer, vapor bubble formation, growth, motion, coalescence and collapse is required in the presence of stirring jets of different configurations and passive cooling devices such as MLI, thermodynamic vent system, and vapor-cooled shield. To develop physics-based models and correlations reliable for microgravity conditions and long-time scales there is a need for new fundamental data to be collected from on-orbit cryogenic storage experiments. Our report discusses some of these physical phenomena and the design requirements and future studies necessary for their mitigation. Special attention is payed to the phenomena occurring near MLI penetrations.

Muratov, C. B.↗
NASA NTRS2013

CheMin Instrument Performance and Calibration on Mars

The CheMin (Chemistry and Mineralogy) instrument on the Mars Science Laboratory rover Curiosity uses a CCD detector and a Co-anode X-ray tube source to acquire both mineralogy (from the pattern of Co diffraction) and chemical information (from energies of fluoresced X-rays). A key component of the CheMin instrument is the ability to move grains within sample cells during analysis, providing multiple, random grain orientations that disperse diffracted X-ray photons along Debye rings rather than producing discrete Laue spots. This movement is accomplished by piezoelectric vibration of the sample cells. A cryocooler is used to maintain the CCD at a temperature at about -50 C in order to obtain energy resolution better than 250 eV, allowing discrimination of diffracted Co K X-rays from Fe K and other fluorescent X-rays. A detailed description of CheMin is provided in [1]. The CheMin flight model (FM) is mounted within the body of Curiosity and has been operating on Mars since August 6, 2012. An essentially identical sister instrument, the CheMin demonstration model (DM), is operated in a Mars environment chamber at JPL.

Vaniman, D. T.↗
NASA NTRS2012

Sensitivity of Mission Energy Consumption to Turboelectric Distributed Propulsion Design Assumptions on the N3-X Hybrid Wing Body Aircraft

In a previous study by the authors it was shown that the N3-X, a 300 passenger hybrid wing body (HWB) aircraft with a turboelectric distributed propulsion (TeDP) system, was able to meet the NASA Subsonic Fixed Wing (SFW) project goal for N+3 generation aircraft of at least a 60% reduction in total energy consumption as compared to the best in class current generation aircraft. This previous study combined technology assumptions that represented the highest anticipated values that could be matured to technology readiness level (TRL) 4-6 by 2030. This paper presents the results of a sensitivity analysis of the total mission energy consumption to reductions in each key technology assumption. Of the parameters examined, the mission total energy consumption was most sensitive to changes to total pressure loss in the propulsor inlet. The baseline inlet internal pressure loss is assumed to be an optimistic 0.5%. An inlet pressure loss of 3% increases the total energy consumption 9%. However changes to reduce inlet pressure loss can result in additional distortion to the fan which can reduce fan efficiency or vice versa. It is very important that the inlet and fan be analyzed and optimized as a single unit. The turboshaft hot section is assumed to be made of ceramic matrix composite (CMC) with a 3000 F maximum material temperature. Reducing the maximum material temperature to 2700 F increases the mission energy consumption by only 1.5%. Thus achieving a 3000 F temperature in CMCs is important but not central to achieving the energy consumption objective of the N3-X/TeDP. A key parameter in the efficiency of superconducting motors and generators is the size of the superconducting filaments in the stator. The size of the superconducting filaments in the baseline model is assumed to be 10 microns. A 40 micron filament, which represents current technology, results in a 200% increase in AC losses in the motor and generator stators. This analysis shows that for a system with 40 micron filaments the higher stator losses plus the added weight and power of larger cryocoolers results in a 4% increase in mission energy consumption. If liquid hydrogen is used to cool the superconductors the 40 micron fibers results in a 200% increase in hydrogen required for cooling. Each pound of hydrogen used as fuel displaces 3 pounds of jet fuel. For the N3-X on the reference mission the additional hydrogen due to the increase stator losses reduces the total fuel weight 10%. The lighter fuel load and attendant vehicle resizing reduces the total energy consumption more than the higher stator losses increase it. As a result with hydrogen cooling there is a slight reduction in mission energy consumption with increasing stator losses. This counter intuitive result highlights the need to consider the full system impact of changes rather than just at the component or subsystem level.

Felder, James L.↗
NASA NTRS2013

Performance Testing of the Engineering Model Astro-H 3-stage ADR

The Japanese Astro-H mission will include the Soft X-ray Spectrometer (SXS) instrument provided by NASA/GSFC. The SXS will perform imaging spectroscopy in the soft x-ray band using a 6x6 array of silicon microcalorimeters operated at 50 mK. The detectors are cooled by a 3-stage adiabatic demagnetization refrigerator (ADR), which is configured to use either a 1.3 K superfluid helium tank or a 4.5 K Joule-Thomson cryocooler as a heat sink. At present, the engineering model SXS, including the detectors and ADR, has been performance tested at GSFC and integrated with the EM dewar in Japan. The flight model SXS is currently being fabricated. This paper presents test results of the EM ADR and changes that will be implemented in the flight version.

Shirron, Peter J.↗
NASA NTRS2009

Experimental Results of Integrated Refrigeration and Storage System Testing

Launch operations engineers at the Kennedy Space Center have identified an Integrated Refrigeration and Storage system as a promising technology to reduce launch costs and enable advanced cryogenic operations. This system uses a close cycle Brayton refrigerator to remove energy from the stored cryogenic propellant. This allows for the potential of a zero loss storage and transfer system, as well and control of the state of the propellant through densification or re-liquefaction. However, the behavior of the fluid in this type of system is different than typical cryogenic behavior, and there will be a learning curve associated with its use. A 400 liter research cryostat has been designed, fabricated and delivered to KSC to test the thermo fluid behavior of liquid oxygen as energy is removed from the cryogen by a simulated DC cycle cryocooler. Results of the initial testing phase focusing on heat exchanger characterization and zero loss storage operations using liquid oxygen are presented in this paper. Future plans for testing of oxygen densification tests and oxygen liquefaction tests will also be discussed. KEYWORDS: Liquid Oxygen, Refrigeration, Storage

Notardonato, W. U.↗
NASA NTRS2012

Design of the PIXIE Adiabatic Demagnetization Refrigerators

The Primordial Inflation Explorer (PIXIE) is a proposed mission to densely map the polarization of the cosmic microwave background. It will operate in a scanning mode from a sun-synchronous orbit, using low temperature detectors (at 0.1 K) and located inside a teslescope that is cooled to approximately 2.73 K - to match the background temperature. A mechanical cryocooler operating at 4.5 K establishes a low base temperature from which two adiabatic demagnetization refrigerator (ADR) assemblies will cool the telescope and detectors. To achieve continuous scanning capability, the ADRs must operate continuously. Complicating the design are two factors: 1) the need to systematically vary the temperature of various telescope components in order to separate the small polarization signal variations from those that may arise from temperature drifts and changing gradients within the telescope, and 2) the orbital and monthly variations in lunar irradiance into the telescope barrels. These factors require the telescope ADR to reject quasi-continuous heat loads of 2-3 millwatts, while maintaining a peak heat reject rate of less than 12 milliwatts. The detector heat load at 0.1 K is comparatively small at 1-2 microwatts. This paper will describe the 3-stage and 2-stage continuous ADRs that will be used to meet the cooling power and temperature stability requirements of the PIXIE detectors and telescope.

adiabatic demagnetization refrigerator↗
NASA NTRS2014

An Overview of SBIR Phase 2 Airbreathing Propulsion Technologies

Technological innovation is the overall focus of NASA's Small Business Innovation Research (SBIR) program. The program invests in the development of innovative concepts and technologies to help NASA's mission directorates address critical research and development needs for agency projects. This report highlights innovative SBIR Phase II projects from 2007-2012 specifically addressing areas in Airbreathing Propulsion which is one of six core competencies at NASA Glenn Research Center. There are twenty technologies featured with emphasis on a wide spectrum of applications such as with a Turbo-Brayton cryocooler for aircraft superconducting systems, braided composite rotorcraft structures, engine air brake, combustion control valve, flexible composite driveshaft, and much more. Each article in this booklet describes an innovation, technical objective, and highlights NASA commercial and industrial applications. This report serves as an opportunity for NASA personnel including engineers, researchers, and program managers to learn of NASA SBIR's capabilities that might be crosscutting into this technology area. As the result, it would cause collaborations and partnerships between the small companies and NASA Programs and Projects resulting in benefit to both SBIR companies and NASA.

Nguyen, Hung D.↗
NASA NTRS2014

Operation of an ADR Using Helium Exchange Gas as a Substitute for a Failed Heat Switch

The Soft X-ray Spectrometer (SXS) is one of four instruments on the Japanese Astro-H mission, which is currently planned for launch in late 2015. The SXS will perform imaging spectroscopy in the soft X-ray band (0.3-12 keV) using a 6 6 pixel array of microcalorimeters cooled to 50 mK. The detectors are cooled by a 3-stage adiabatic demagnetization refrigerator (ADR) that rejects heat to either a superfluid helium tank (at 1.2 K) or to a 4.5 K Joule-Thomson (JT) cryocooler. Four gas-gap heat switches are used in the assembly to manage heat flow between the ADR stages and the heat sinks. The engineering model (EM) ADR was assembled and performance tested at NASA/GSFC in November 2011, and subsequently installed in the EM dewar at Sumitomo Heavy Industries, Japan. During the first cooldown in July 2012, a failure of the heat switch that linked the two colder stages of the ADR to the helium tank was observed. Operation of the ADR requires some mechanism for thermally linking the salt pills to the heat sink, and then thermally isolating them. With the failed heat switch unable to perform this function, an alternate plan was devised which used carefully controlled amounts of exchange gas in the dewar's guard vacuum to facilitate heat exchange. The process was successfully demonstrated in November 2012, allowing the ADR to cool the detectors to 50 mK for hold times in excess of 10 h. This paper describes the exchange-gas-assisted recycling process, and the strategies used to avoid helium contamination of the detectors at low temperature.

Adiabatic demagnetization↗
NASA NTRS2015

Architecture, Voltage, and Components for a Turboelectric Distributed Propulsion Electric Grid

The development of a wholly superconducting turboelectric distributed propulsion system presents unique opportunities for the aerospace industry. However, this transition from normally conducting systems to superconducting systems significantly increases the equipment complexity necessary to manage the electrical power systems. Due to the low technology readiness level (TRL) nature of all components and systems, current Turboelectric Distributed Propulsion (TeDP) technology developments are driven by an ambiguous set of system-level electrical integration standards for an airborne microgrid system (Figure 1). While multiple decades' worth of advancements are still required for concept realization, current system-level studies are necessary to focus the technology development, target specific technological shortcomings, and enable accurate prediction of concept feasibility and viability. An understanding of the performance sensitivity to operating voltages and an early definition of advantageous voltage regulation standards for unconventional airborne microgrids will allow for more accurate targeting of technology development. Propulsive power-rated microgrid systems necessitate the introduction of new aircraft distribution system voltage standards. All protection, distribution, control, power conversion, generation, and cryocooling equipment are affected by voltage regulation standards. Information on the desired operating voltage and voltage regulation is required to determine nominal and maximum currents for sizing distribution and fault isolation equipment, developing machine topologies and machine controls, and the physical attributes of all component shielding and insulation. Voltage impacts many components and system performance.

electric faults↗