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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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At least 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS2020

SWIRP (Submm-Wave and Long Wave InfraRed Polarimeter); Development and Characterization of a Sub-Mm Polarimeter for Ice Cloud Investigations

A major source of uncertainty in climate models is the presence, shape and distribution of ice particles in the uppermost layers of the clouds. The effects of this component are poorly constrained, turning ice particles into an almost-free variable in many climate models.NASA-GSFC is developing a new instrument aimed at measuring the size and shape of ice particles. The instrument consists of two sub-mm polarimeters (at 220 and 670 GHz) coupled with a long-wave infrared polarimeter at 10 micron. Each polarimeter has identical V-pol and H-pol channels; the axes of polarization are defined geometrically by the orientation of the waveguide elements, and the purity has been measured in the lab. The instrument is configured as a conical scanner, suitable for deployment as a payload on a small satellite or on a high-altitude sub-orbital platform. From a 400 km orbit, the instrument has a 3dB spatial resolution of 20 (10) km at 220 (670) GHz and a swath of 600 km over 180 degrees of view.The BAPTA (Bearing And Power Transfer Assembly) carries heritage from the SSMIS design, now in its 22nd year of on-orbit operation, but with a much reduced SWaP (Size Weight and Power) footprint, suitable for a small satellite.The main components of the instrument have been fabricated and are undergoing final testing prior to their integration as a single unit. The sub-mm channels have dedicated secondary reflectors which illuminate a shared primary reflector. The receiving units are placed behind the focal point of the optical arrangement, so that all beams equally illuminate the primary reflector and are almost co-located on the ground (within a single 220 GHz footprint). Primary and secondary beam patterns have been measured and verified to match the as-designed expectations. A Zytex (TM) window is deployed to protect the secondary reflectors and the feed horns from debris and other contaminants, and to reduce the heat load from the active (hot) IR calibration unit. The insertion loss of Zytex has been measured and is accounted in the calibration equation of the sub-mm channels.The radiometric performance of the sub-mm receivers has been characterized in the lab and under operational conditions of temperature and pressure.This paper discusses the design constraints on the sub-mm components, details of the scientific goals and their flowdown, and describes the characterization of the polarimeters. Options to optimize the layout and distribution of the masses within the assembly, with the goal of making the instrument even more compact and fully-compatible with cubesat-class satellites will be presented.

De Amici, G.↗
NASA NTRSDate not supplied

Impact of Magnetohydrodynamic (MHD) Induced Drag on High Ballistic Coefficient Aerocapture at Neptune

This paper presents analyses illustrating the benefits of magnetohydrodynamic (MHD) drag force control for a capsule performing aerocapture at Neptune relative to the same capsule utilizing conventional aerodynamic-only control across a range of design and flight conditions. MHD technology utilizes an onboard mag-net to create a magnetic field that interacts with the ionized plasma surrounding the capsule at hypersonic velocities, thus creating an electron current and inducing a Lorentz force used to increase capsule drag. The study assumes a 4.57-meterMars Science Laboratory (MSL) 70-degree sphere-cone capsule aerodynamic shape arrives hyperbolically at Neptune and aerocaptures into a synchronous Tri-ton orbit (3986 km x 430,000 km orbit inclined 157 degrees) that offers repeated flyby opportunities. Hyperbolic arrival velocity and the capsule entry mass (i.e., ballistic coefficient) are varied over a wide range of values to study the impact of these variables on key flight performance and sizing parameters. Results show reduced heat rates, heat loads, and thermal protection system mass for MHD cases at all conditions, with the benefit of MHD relative to conventional aerodynamic-only options increasing as ballistic coefficient increases, suggesting the payoff for MHD is significantly better for high-mass entry system design concepts with high ballistic coefficients.

Entry Descent and Landing↗
NASA NTRS1992

Modeling and analysis of the Space Station Freedom Active Thermal Control System radiators using SINDA/FLUINT

Detailed SINDA/FLUINT models have been developed to assess the adequacy of the Space Station Freedom External Thermal Control System's radiator design, as well as to simulate the effects of various operational scenarios. Design margins can be deduced by quantifying such real-world effects as flow maldistribution due to line-pressure drop and uneven heat loads. The factor which most reduced the heat-rejection capacity is an uneven heat load split between the two low temperature loops.

Andish, Kambiz K.↗
NASA NTRS2011

Thermal Vacuum Testing of a Proto-flight Miniature Loop Heat Pipe with Two Evaporators and Two Condensers

This paper describes thermal vacuum testing of a proto-flight miniature loop heat pipe (MLHP) with two evaporators and two condensers designed for future small systems applications requiring low mass, low power and compactness. Each evaporator contains a wick with an outer diameter of 6.35 mm, and each has its own integral compensation chamber (CC). Miniaturization of the loop components reduces the volume and mass of the thermal system. Multiple evaporators provide flexibility for placement of instruments that need to be maintained at the same temperature, and facilitate heat load sharing among instruments, reducing the auxiliary heater power requirement. A flow regulator is used to regulate heat dissipations between the two condensers, allowing flexible placement of radiators on the spacecraft. A thermoelectric converter (TEC) is attached to each CC for control of the operating temperature and enhancement of start-up success. Tests performed include start-up, power cycle, sink temperature cycle, high power and low power operation, heat load sharing, and operating temperature control. The proto-flight MLHP demonstrated excellent performance in the thermal vacuum test. The loop started successfully and operated stably under various evaporator heat loads and condenser sink temperatures. The TECs were able to maintain the loop operating temperature within b1K of the desired set point temperature at all power levels and all sink temperatures. The un-powered evaporator would automatically share heat from the other powered evaporator. The flow regulator was able to regulate the heat dissipation among the radiators and prevent vapor from flowing into the liquid line.

Ku, Jentung↗
NASA NTRSDate not supplied

Vapor Cooling of a Structural Skirt for a Large-Scale Hydrogen Tank

The demonstration of vapor cooling on a structural skirt was one of the main objectives of the Structural Heat Intercept, Insulation, and Vibration Evaluation Rig (SHIIVER) testing. SHIIVER consisted of a 4-meter diameter stainless steel tank with an aluminum forward skirt. The forward skirt was cooled by directing effluent vapor from the tank through two flow channels that each went 360 degrees around the interior of the skirt in a helical pattern. Flow rates, temperatures, and pressures in the system were measured allowing for the calculation of heat load removed via cooling stream and heat load reduction into the tank. Testing occurred at fill levels between 25% and 90% full using both liquid hydrogen and liquid nitrogen. Boil-off rate was varied independent of the skirt performance by adding multilayer insulation over the spray-on-foam insulation on the tank domes, while leaving the barrel insulated only with spray-on-foam. The results from the testing, which demonstrate vapor cooling reversed heat flow from the skirt into the tank and reduces total propellant heat load by as much as 19%, are analyzed and discussed.

Cryogenic Fluid Management↗
NASA NTRS2011

Study of Unsteady, Sphere-Driven, Shock-Induced Combustion for Application to Hypervelocity Airbreathing Propulsion

A premixed, shock-induced combustion engine has been proposed in the past as a viable option for operating in the Mach 10 to 15 range in a single stage to orbit vehicle. In this approach, a shock is used to initiate combustion in a premixed fuel/air mixture. Apparent advantages over a conventional scramjet engine include a shorter combustor that, in turn, results in reduced weight and heating loads. There are a number of technical challenges that must be understood and resolved for a practical system: premixing of fuel and air upstream of the combustor without premature combustion, understanding and control of instabilities of the shock-induced combustion front, ability to produce sufficient thrust, and the ability to operate over a range of Mach numbers. This study evaluated the stability of the shock-induced combustion front in a model problem of a sphere traveling in a fuel/air mixture at high Mach numbers. A new, rapid analysis method was developed and applied to study such flows. In this method the axisymmetric, body-centric Navier-Stokes equations were expanded about the stagnation streamline of a sphere using the local similarity hypothesis in order to reduce the axisymmetric equations to a quasi-1D set of equations. These reduced sets of equations were solved in the stagnation region for a number of flow conditions in a premixed, hydrogen/air mixture. Predictions from the quasi-1D analysis showed very similar stable or unstable behavior of the shock-induced combustion front as compared to experimental studies and higher-fidelity computational results. This rapid analysis tool could be used in parametric studies to investigate effects of fuel rich/lean mixtures, non-uniformity in mixing, contaminants in the mixture, and different chemistry models.

Axdahl, Erik↗
NASA NTRS1978

Preliminary design of the cryogenic cooled limb scanning interferometer radiometer (CLIR), phase 2

A cryogenic system to be utilized with a mounted pointing system is designed. The analysis indicates that the resulting design is significantly heavier than previous design due to the increased diameter of the system, resulting in greater tank and vacuum enclosure weights. Additional weight is also required in the area where the payload is attached to the pointing system in order to provide additional stiffness to the payload. The vacuum shell required to enclose the instrument is shorter and therefore an attendant reduction in instrument package weight would be expected. The primary resonance of the cooler instrument package occurs at 21 Hz, a mode in which the flexibility of the vacuum shell allows motion of the cryogen tank and instrument, and which can be increased by additional stiffening in the region of the pointing system vacuum shell connection at the cost of additional weight. The second resonance at 59 Hz is due to vibration of the helium tank on the fiberglass support tubes. The parasitic heat load can be reduced to 0.28 W by the utilization of a sandwich support tube structure in place of the monocoque structure assumed in the design.

Source record↗
NASA NTRS1988

Manipulator For A Vacuum Chamber

Rotary seal provides for external actuators. Positioner in vacuum employs four actuators; up/down, horizontal rotary, horizontal, and vertical rotary. External up/down and azimuthal actuators drives shaft entering vacuum chamber through sliding and rotating seal. Placement of actuator motors outside vacuum chamber reduces heat-removal load on system. Atop 2.5-cm-diameter shaft is stage supporting masses as great as 15 kg. Small motors on stage provide horizontal translation and rotation in vertical plane. Manipulator developed for calibrating plasma detectors by subjecting them to particle beams in vacuum. Standard, commercially available parts used.

Hunsaker, F.↗
NASA NTRS1992

Low gravity thermal stratification of liquid helium on SHOOT

Estimates of the extent and impact of thermal stratification are presented as well as predictions of the behavior of the HeI/HeII boundary. Although thermal stratification of cryogens can be problematic and lead to their inefficient use in low gravity, for SHOOT the occurrence is beneficial both during ground hold and in orbit and presents no hazards. On the ground the parasitic heat load is both reduced and more efficiently removed. In orbit the pumpdown proceeds at a much more rapid rate, allowing orbital operations to begin earlier. The thermal conductivity of the aluminum tank and the normal liquid plus cooling at the liquid/vapor interface as the vapor bubble grows are sufficient to prevent undesirably high vapor pressures in the tank.

Shirron, P. J.↗
NASA NTRS1991

Design of the solid cryogen dewar for the Near-Infrared Camera and Multi-Object Spectrometer

A multipurpose, second-generation HST instrument for imaging and spectroscopy in the 1 to 2.5 microns wavelength region is being developed. The Near-Infrared Camera Multi-Object Spectrometer (NICMOS) is unique since it is the only HST instrument operating in the NIR and cryogenically cooled. The NICMOS detector arrays are cooled to 58 K by a solid-nitrogen (SN2) dewar with a predicted lifetime of nearly five years. To obtain this long lifetime, a hybrid cooling approach using thermoelectric coolers (TECs) is employed to reduce the parasitic heat load on the SN2. The design features used to promote long life, the predicted lifetime improvements provided by the TECs, and the performance degradation in the event of TEC failure(s) are discussed.

Oonk, Rodney L.↗
NASA NTRS1997

High Resolution X-Ray Spectroscopy Using Microcalorimeters

During the past 13 years high resolution X ray spectrometers have been developed that use cryogenically cooled microcalorimeters. These devices have inherently high signal-to-noise by operating at temperatures below 0.1 K and can achieve an energy resolution of < 10 eV over the 0.1-10 keV band. Existing devices use doped semiconductor thermometers and typically employ HgTe absorbers. The energy resolution depends on achieving a low heat capacity for the device. For soft X ray applications a relatively thin absorber (approximately 1 micrometer) may be used and an energy resolution of approximately 7 eV has been achieved. For applications up to approximately 10 keV an absorber thickness of approximately 10 micrometer is required and the energy resolution is typically approximately 12 eV. Improvements to the energy resolution in this energy band could be achieved if the problems of thermalizing X rays in low heat capacity superconductors can be overcome. The recent work on transition edge thermometers by Irwin et nl. looks particularly promising because of the higher sensitivity achievable from a sharp superconducting transition. The relatively low impedance of such a device permits the use of a low noise SQUID amplifier for readout. This would also significantly reduce the cryogen heat load compared with JFETs required by higher impedance semiconductor thermometers.

Kelley, R. L.↗
NASA NTRS2007

Numerical Study of Flow Augmented Thermal Management for Entry and Re-Entry Environments

The use of a flow augmented thermal management system for entry and re-entr environments is one method for reducing heat and drag loads. This concept relies on jet penetration from supersonic and hypersonic counterflowing jets that could significantly weaken and disperse the shock-wave system of the spacecraft flow field. The objective of this research effort is to conduct parametric studies of the supersonic flow over a 2.6% scale model of the Apollo capsule, with and without the counterflowing jet, using time-accurate and steady-state computational fluid dynamics simulations. The numerical studies, including different freestream Mach number angle of attack counterflowing jet mass flow rate, and nozzle configurations, were performed to examine their effect on the drag and beat loads and to explore the counternowing jet condition. The numerical results were compared with the test data obtained from transonic blow-down wind-tunnel experiments conducted independently at NASA MSFC.

Cheng, Gary C.↗
NASA NTRS2015

The Structural Heat Intercept-Insulation-Vibration Evaluation Rig (SHIVER)

NASA is currently investigating methods to reduce the boil-off rate on large cryogenic upper stages. Two such methods to reduce the total heat load on existing upper stages are vapor cooling of the cryogenic tank support structure and integration of thick multilayer insulation systems to the upper stage of a launch vehicle. Previous efforts have flown a 2-layer MLI blanket and shown an improved thermal performance, and other efforts have ground-tested blankets up to 70 layers thick on tanks with diameters between 2 3 meters. However, thick multilayer insulation installation and testing in both thermal and structural modes has not been completed on a large scale tank. Similarly, multiple vapor cooled shields are common place on science payload helium dewars; however, minimal effort has gone into intercepting heat on large structural surfaces associated with rocket stages. A majority of the vapor cooling effort focuses on metallic cylinders called skirts, which are the most common structural components for launch vehicles. In order to provide test data for comparison with analytical models, a representative test tank is currently being designed to include skirt structural systems with integral vapor cooling. The tank is 4 m in diameter and 6.8 m tall to contain 5000 kg of liquid hydrogen. A multilayer insulation system will be designed to insulate the tank and structure while being installed in a representative manner that can be extended to tanks up to 10 meters in diameter. In order to prove that the insulation system and vapor cooling attachment methods are structurally sound, acoustic testing will also be performed on the system. The test tank with insulation and vapor cooled shield installed will be tested thermally in the B2 test facility at NASAs Plumbrook Station both before and after being vibration tested at Plumbrooks Space Power Facility.

Cryogenic Fluid Management↗
NASA NTRS2017

Structural Heat Intercept, Insulation, and Vibration Evaluation Rig (SHIIVER)

Abstract: NASA is currently investigating methods to reduce the boil-off rate on large cryogenic upper stages. Two such methods to reduce the total heat load on existing upper stages are vapor cooling of the cryogenic tank support structure and integration of thick multilayer insulation systems to the upper stage of a launch vehicle. Previous efforts have flown a 2-layer MLI blanket and shown an improved thermal performance, and other efforts have ground-tested blankets up to 70 layers thick on tanks with diameters between 2 3 meters. However, thick multilayer insulation installation and testing in both thermal and structural modes has not been completed on a large scale tank. Similarly, multiple vapor cooled shields are common place on science payload helium dewars; however, minimal effort has gone into using boil-off to intercept heat on large structural surfaces associated with rocket stages. A majority of the vapor cooling effort focuses on metallic cylinders called skirts, which are the most common structural components for launch vehicles. In order to provide test data for comparison with analytical models, a 4 meter diameter test tank and assembly is currently being designed to include skirt structural systems with integral vapor cooling. In order to give a more representative result for actual applications, the technologies will be designed for an 8.4 m and scaled down to a 4 meter tank. In order to prove that the insulation system and vapor cooling attachment methods are structurally sound, acoustic testing will also be performed on the system. The test tank with insulation and vapor cooled shield installed will be tested thermally in the B2 test facility at NASAs Plum Brook Station both before and after being vibration tested at Plum Brooks Space Power Facility.

Johnson, Wesley↗
NASA NTRS1979

Flight prototype CO2 and humidity control system

A regenerable CO2 and humidity control system is presently being developed for potential use on shuttle as an alternative to the baseline lithium hydroxide system. The system utilizes a sorbent material (designated HS-C) to adsorb CO2 and the latent heat load from the cabin atmosphere and desorb the CO2 and water vapor overboard when exposed to a space vacuum, thus reducing the overall vehicle heat rejection load. Continuous operation is achieved by utilizing two beds which are alternatively cycled between adsorption and desorption. The HS-C material process was verified. Design concepts for the auxiliary components for the HS-C prototype system were generated. Performance testing verified system effectiveness in controlling CO2 partial pressure and humidity.

Rudy, K. M.↗
NASA NTRS1980

Baffle design for earth radiation rejection in the Cryogenic Limb-Scanning Interferometer/Radiometer

The Cryogenic Limb-Scanning Interferometer/Radiometer (CLIR) is being developed to observe infrared emissions of the earth's upper atmosphere from space. The earth's surface is an extended source of intense background radiation with a small angular separation from the desired scene. The CLIR employs an off-axis Gregorian Telescope whose primary mirror and baffles are cooled by an open-cycle cryogen system. A system of specular annular baffles has been developed to minimize both stray light problems and cryogen consumption by retro-mapping the aperture into itself. Each off-axis ray which enters the aperture and strikes the specular baffle surface is reflected so that it passes out of the aperture and is not absorbed on a cryogenic surface. The specular baffle which lies closest to the aperture is an ellipsoid whose foci trace out the circular aperture on revolution about the axis. Its theoretical 'ray trace' efficiency is 100 percent. A subsequent baffle has an elliptical cross section whose near focus traces out the central hole in the ellipsoidal baffle and whose far focus traces out the aperture. Its theoretical efficiency is about 90 percent. These baffles reduce the earth radiation heat load on the cryogenic cooler by an order of magnitude, changing it from the dominant cause of cryogen consumption to a relatively small effect. An aperture shield is also desirable to reduce cryogen consumption, stray light, and contamination.

Bremer, J. C.↗
NASA NTRSDate not supplied

Calorimeter Heat Flux Trends in NASA’s Subscale Rotating Detonation Rocket Engine

A calorimeter-style rotating detonation rocket engine (RDRE) outer chamber body was designed, additively manufactured in GRCop-42, and hot-fire tested in test cell 105 at NASA Marshall Space Flight Center (MSFC) in Huntsville, AL in August 2024. Steady-state, spatially resolved heat flux and chamber pressure were measured at 13 and 9 locations from the injector face, respectively, for a variety of test conditions and hardware configurations with gaseous methane and oxygen propellants. The 2-in. diameter highly instrumented design, enabled by metal additive manufacturing (AM), allowed for simultaneous measurement of heat flux and chamber pressure through 44 long duration hot-fire tests without failure. Of the 44 tests, only 3 of the tests featured chamber pressures below 100 psia and are discussed in this work. Heat flux trends for this subscale hardware are explored with respect to local propellant mass flux for several different injectors, chamber lengths, and chamber “subsonic” area ratios (contraction ratios). Results show that for increasing chamber pressure, a reduced amount of energy per kilogram of propellant is transferred to the wall. Switching injectors affects the near-face heat flux and chamber pressure profiles. Chamber length does not noticeably affect the heat flux profile but increases bulk heat load. A subsonic area ratio greater than 1 reduces heat losses to the wall compared to a straight annulus case. Equivalence ratio shift did not significantly affect heat flux. Computational fluid dynamics (CFD) and conjugate heat transfer (CHT) analyses were performed to delineate 3-D heat transfer and coolant mass flow maldistribution effects and create a calorimeter transfer function to transform heat flux data and mitigate the profile distortions. Additional CFD/CHT simulations were performed using experimental data as boundary conditions to assess local nucleate boiling propensity during testing and to assess projected hot wall temperatures for future fatigue assessment. Vapor fractions were observed locally for several cases. It is anticipated that this effort will provide a common framework for coolant system design/analysis in calorimeter systems for rocket-based and air-breathing combustors, ensuring greater repeatability in calorimetry measurements across the research community.

Joseph Hernandez-McCloskey↗
NASA NTRS2015

Numerical Experiments of Counterflowiing Jet Effects on Supersonic Slender-Body Configurations

Previous studies have demonstrated that the use of counterflowing jets can greatly reduce the drag and heat loads on blunt-body geometries, especially when the long penetration mode jet condition can be established. Previously, the authors had done some preliminary numerical studies to determine the ability to establish long penetration mode jets on a typical Mach 1.6 slender configuration, and study its impact on the boom signature. The results indicated that a jet with a longer penetration length was required to achieve any impact on the boom signature of a typical Mach 1.6 slender configuration. This paper focuses on an in-depth parametric study, done using the space-time conservation element solution element Navier-Stokes flow solver, for investigating the effect of various counterflowing jet conditions/configurations on two supersonic slender-body models (cone-cylinder and quartic body of revolution). The study is aimed at gaining a better understanding of the relationship between the shock penetration length and reduction of drag and boom signature for these two supersonic slender-body configurations. Different jet flow rates, Mach numbers, nozzle jet exit diameters and jet-to-base diameter ratios were examined. The results show the characteristics of a short-to-long-to-short penetration-mode pattern with the increase of jet mass flow rates, observed across various counterflowing jet nozzle configurations. Though the optimal shock penetration length for potential boom-signature mitigation is tied to the long penetration mode, it often results in a very unsteady flow and leads to large oscillations of surface pressure and drag. Furthermore, depending on the geometry of the slender body, longer jet penetration did not always result in maximum drag reduction. For the quartic geometry, the maximum drag reduction corresponds well to the longest shock penetration length, while this was not the case for the cone-cylinder-as the geometry was already optimized for drag. Numerical results and assessments obtained from this parametric study along with the recommendation for future implementation of counterflowing jets as a means for drag and noise reduction are detailed in this paper.

Venkatachari, Balaji Shankar↗