Prediction of heat transfer in multiple glaze transparencies Final report
Techniques for evaluating thermal behavior of window systems for space vehicles exposed to environmental conditions attendant to planetary reentry
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Importance of skin temperature in the regulation of sweating
Human sweating regulation at rest, evaluating thermal inputs effects on thermoregulatory center and internal hypothalamic and skin temperatures
Bullard, R. W.↗
Services provided in support of the planetary quarantine requirements of the National Aeronautics and Space Administration
The project to evaluate thermal sterilization for unmanned landers is reported. A temperature controlled oven with a nitrogen gas supply containing a known concentration of water is discussed. The studies show that bacillus lentus, bacillus brevis, bacillus coagulans, atypical bacillus spp., and actinomycete are isolated heat survivors. The thermal resistance is given for naturally occurring airborne bacterial spores collected on exposed teflon ribbons.
Favero, M. S.↗
Evaluation of cermet materials suitable for lithium lubricated thrust bearings for high temperature operation
Cerment materials (HfC - 10 wt% W; HfC - 10 wt% TaC - 10 wt%W; HfC - 2 wt% CbC - 8 wt% Mo;Hfn - 10 wt% W; Hfn - 10 wt% TaN - 10 wt% W; and ZrC - 17 wt% W) were evaluated for possible use as lithium-lubricated bearings in the control system of a nuclear reactor. Tests of compatibility with lithium were made in T-111 (Ta-8W-2Hf) capsules at temperatures up to 1090 C. The tendencies of HfC-TaC-W, HfC-CbC-Mo, and HfN-W to bond to themselves and to the refractory alloys T-111 and TZM when enclosed in lithium-filled capsules under a pressure of 2000 psi at 980 and 1200 C for 1933 hours were evaluated. Thermal expansion characteristics were determined for the same three materials from room temperature to 1200 C. On the basis of these tests, HfC-10 TaC-10W and HfN-10W were selected as the best and second best candidates, respectively, of the materials tested for the bearing application.
Sinclair, J. H.↗
Automated space processing payloads study. Volume 2, book 2: Technical report, appendices A through E
Experiment hardware and operational requirements for space shuttle experiments are discussed along with payload and system concepts. Appendixes are included in which experiment data sheets, chamber environmental control and monitoring, method for collection and storage of electrophoretically-separated samples, preliminary thermal evaluation of electromagnetic levitation facilities L1, L2, and L3, and applicable industrial automation equipment are discussed.
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Tungsten wire-nickel base alloy composite development
Further development and evaluation of refractory wire reinforced nickel-base alloy composites is described. Emphasis was placed on evaluating thermal fatigue resistance as a function of matrix alloy composition, fabrication variables and reinforcement level and distribution. Tests for up to 1,000 cycles were performed and the best system identified in this current work was 50v/o W/NiCrAlY. Improved resistance to thermal fatigue damage would be anticipated for specimens fabricated via optimized processing schedules. Other properties investigated included 1,093 C (2,000 F) stress rupture strength, impact resistance and static air oxidation. A composite consisting of 30v/o W-Hf-C alloy fibers in a NiCrAlY alloy matrix was shown to have a 100-hour stress rupture strength at 1,093 C (2,000 F) of 365 MN/square meters (53 ksi) or a specific strength advantage of about 3:1 over typical D.S. eutectics.
Brentnall, W. D.↗
Outdoor tests of a liquid solar collector
Results of outdoor test program evaluating thermal performance of liquid solar collector are presented in report. Test article used is flatplate collector with antifreeze solution as working fluid.
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Generalized environmental control and life support system computer program (G1894), phase 3
The work performed during Phase 3 of the Generalized Environmental Control Life Support System (ECLSS) Computer Program is reported. Phase 3 of this program covered the period from December 1977 to September 1978. The computerized simulation of the Shuttle Orbiter ECLSS was upgraded in the following areas: (1) the payload loop of the Shuttle simulation was completely recoded and checked out; (2) the Shuttle simulation water and freon loop initialization logic was simplified to permit easier program input for the user; (3) the computerized simulation was modified to accept the WASP subroutine, which is a subroutine to evaluate thermal properties of water and freon; (4) the 1108 operating system was upgraded by LEC; (5) the Shuttle simulation was modified to permit failure cases which simulate zero component flow values; and (6) the Shuttle SEPS version was modified and secure files were setup on the 1108 and 1110 systems to permit simulation runs to be made from remote terminals.
Mcenulty, R. E.↗
Solid state SPS microwave generation and transmission study. Volume 2, phase 2: Appendices
The solid state sandwich concept for SPS was further defined. The design effort concentrated on the spacetenna, but did include some system analysis for parametric comparison reasons. Basic solid state microwave devices were defined and modeled. An initial conceptual subsystems and system design was performed as well as sidelobe control and system selection. The selected system concept and parametric solid state microwave power transmission system data were assessed relevant to the SPS concept. Although device efficiency was not a goal, the sensitivities to design of this efficiency were parametrically treated. Sidelobe control consisted of various single step tapers, multistep tapers and Gaussian tapers. A hybrid concept using tubes and solid state was evaluated. Thermal analyses are included with emphasis on sensitivities to waste heat radiator form factor, emissivity, absorptivity, amplifier efficiency, material and junction temperature.
Maynard, O. E.↗
IRAS observations of the Pluto-Charon system
High-signal-to-noise-ratio observations of the Pluto-Charon system at 25, 60, and 100 microns using IRAS are combined with visual-magnitude and mutual-eclipse constraints to evaluate thermal models of Pluto and Charon. These models are consistent with eclipse observation by Dunbar and Tedesco (1986) but not with Reinsch and Pakull (1987). The most likely model for Charon is the standard asteroid model, typical for the icy Galilean and Saturnian satellites. Charon models with a significant atmosphere can be ruled out. Based on currently available radius and albedo constraints, no significant numerical distinction is possible between Pluto models ranging from isothermal spheres with surface emissivity between 0.4 and 0.9. Concerns regarding the viability of an emissivity as low as 0.4 favor the higher-emissivity models. The globally uniform surface temperature of Pluto may thus at present be as low as 45 K, with a methane column abundance of 6.7 cm atm. The most likely models are centered on radii of 1180 and 747 km and albedos of 0.47 and 0.26 for Pluto and Charon, respectively.
Aumann, H. H.↗
High temperature NASP engine seal development
Key to the development of advanced hypersonic engines such as those being considered for the National Aerospace Plane (NASP) is the development and evaluation of high temperature, flexible seals that must seal the many feet of gaps between the articulating and stationary engine panels. Recent seal progress made at NASA-Lewis is reviewed in the areas of seal concept maturation, test rig development, and performance tests. A test fixture was built at NASA capable of subjecting candidate 3 ft long seals to engine simulated temperatures (up to 1500 F), pressures (up to 100 psi), and engine wall distortions (up to 0.15 in only 18 in span). Leakage performance test results at high temperatures are presented for an innovative high temperature, flexible ceramic wafer seal. Also described is a joint Pratt and Whitney/NASA planned test program to evaluate thermal performance of a braided rope seal under engine simulated heat flux rates (up to 400 Btu/sq ft s), and supersonic flow conditions. These conditions are produced by subjecting the seal specimen to hydrogen oxygen rocket exhaust that flows tangent to the specimen.
Steinetz, Bruce M.↗
Aerothermodynamic environments of aerobraking vehicles for manned Mars missions
The aerothermodynamic environments of manned spacraft aerobraking in the Martian and earth atmospheres are evaluated. Thermal performance of aerobrake concepts are examined for current cryogenic-aerobrake and advanced propulsion missions entailing three different modes of aerobraking: (1) aerocapture into an orbit about Mars, (2) descent and landing at Mars, and (3) Mars return direct entry at earth. Analyses for these vehicles and modes included both radiative and convective heating, where radiative heating is shown to be a significant portion of the total stagnation point heating induced on the vehicle. A comprehensive parametric study of the effects of ballistic coefficient, nose radius, entry velocity, and L/D on stagnation point heating is described. Optimal nose radii for ranges of ballistic coefficient and entry velocity are determined. The peak heating rates are shown to be 83 W/sq cm and 90 W/sq cm for a low and high L/D Mars transfer vehicle configuration, respectively. Heating profiles for these vehicles using boundary layer techniques show that a high L/D shape will result in a smaller high-temperature region provided the flow is laminar. An examination of a crew return vehicle for a Mars return direct entry trajectory shows that the thermal protection for this aerobrake will require an ablative material for heat rejection due to the large heating rates (about 1 kW/sq cm).
Ledoux, Stephen T.↗
Numerical simulation of the reactive flow in advanced (HSR) combustors using KIVA-2
Recent work has been done with the goal of establishing ultralow emission aircraft gas turbine combustors. A significant portion of the effort is the development of three dimensional computational combustor models. The KIVA-II computer code which is based on the Implicit Continuous Eulerian Difference mesh Arbitrary Lagrangian Eulerian (ICED-ALE) numerical scheme is one of the codes selected by NASA to achieve these goals. This report involves a simulation of jet injection through slanted slots within the Rich burn/Quick quench/Lean burn (RQL) baseline experimental rig. The RQL combustor distinguishes three regions of combustion. This work specifically focuses on modeling the quick quench mixer region in which secondary injection air is introduced radially through 12 equally spaced slots around the mixer circumference. Steady state solutions are achieved with modifications to the KIVA-II program. Work currently underway will evaluate thermal mixing as a function of injection air velocity and angle of inclination of the slots.
Winowich, Nicholas S.↗
Interdisciplinary analysis from the spacecraft thermal analyst's viewpoint
Interdisciplinary analyses performed at the Marshall Space Flight Center based on multidiscipline data flows are reviewed. Particular attention is given to global software packages for allowing efficient interdisciplinary data exchange, the use of a number of smaller individual packages, specific software programs utilized by the various interfacing disciplines to perform calculations in terms of interface compatibility, and tasks based on thermal evaluations integral. Recommendations for improvements in the MSFC equipment, software, and organization are proposed.
Humphries, W. R.↗
Radiant Energy Measurements from a Scaled Jet Engine Axisymmetric Exhaust Nozzle for a Baseline Code Validation Case
A non-flowing, electrically heated test rig was developed to verify computer codes that calculate radiant energy propagation from nozzle geometries that represent aircraft propulsion nozzle systems. Since there are a variety of analysis tools used to evaluate thermal radiation propagation from partially enclosed nozzle surfaces, an experimental benchmark test case was developed for code comparison. This paper briefly describes the nozzle test rig and the developed analytical nozzle geometry used to compare the experimental and predicted thermal radiation results. A major objective of this effort was to make available the experimental results and the analytical model in a format to facilitate conversion to existing computer code formats. For code validation purposes this nozzle geometry represents one validation case for one set of analysis conditions. Since each computer code has advantages and disadvantages based on scope, requirements, and desired accuracy, the usefulness of this single nozzle baseline validation case can be limited for some code comparisons.
Baumeister, Joseph F.↗
Temperature Distributions in Semitransparent Coatings: A Special Two-Flux Solution
Radiative transfer is analyzed in a semitransparent coating on an opaque substrate and in a semitransparent layer for evaluating thermal protection behavior and ceramic component performance in high temperature applications. Some ceramics are partially transparent for radiative transfer, and at high temperatures internal emission and reflections affect their thermal performance. The behavior is examined for a ceramic component for which interior cooling is not provided. Two conditions are considered: (1) the layer is heated by penetration of radiation from hot surroundings while its external surface is simultaneously film cooled by convection, and (2) the surface is heated by convection while the semitransparent material cools from within by radiant emission leaving through the surface. By using the two-flux method, which has been found to yield good accuracy in previous studies, a special solution is obtained for these conditions. The analytical result includes isotropic scattering and requires only an integration to obtain the temperature distribution within the semitransparent material. Illustrative results are given to demonstrate the nature of the thermal behavior.
Siegel, Robert↗
Emission Spectral Measurements in the Plenum of an Arc Jet Wind Tunnel
Arc jet wind tunnel facilities are used to evaluate thermal protection system materials for re-entry vehicles. The high speed, high temperature flowfield generated by the arc jet can simulate the extreme aerodynamic heating environment experienced during re-entry so that the survivability of heat shield materials and performance of various designs options can be tested. Although the re-entry heating environment can be approximated in the arc jet facility, the flowfield only partially simulates the actual re-entry flight conditions. Reynolds numbers are not matched so that surface shear stress distributions and mass transfer rates due to ablation or other mechanisms are not modeled correctly. Unlike flight conditions the arc freestream air is in non-equilibrium because of the rapid expansion that occurs in the supersonic nozzle. To properly study the actual re-entry flow environment, computational fluid dynamics, computational chemistry and radiation models must be used. Arc jet tunnel tests serve to validate these models. To perform accurate simulations inlet and boundary-conditions are needed, which come from measurements of the flowfield. The present study is concerned with measurements in the plenum region of an arc heater. In the past, conditions in the arc heater flowfield have been predicted using simulations since conventional measurement techniques could not be used in the harsh extremely high temperature environment. The present study is part of a recent push to utilize optical techniques to help better characterize the arc jet flowfields. Emission measurements have been made in the shock layer and the constrictor section of the arc heater to determine temperatures and species number densities. LIF measurements have been made in the free stream to determine temperature and velocity.
Donohue, Jim↗
Comparison of High Aspect Ratio Cooling Channel Designs for a Rocket Combustion Chamber
An analytical investigation on the effect of high aspect ratio (height/width) cooling channels, considering different coolant channel designs, on hot-gas-side wall temperature and coolant pressure drop for a liquid hydrogen cooled rocket combustion chamber, was performed. Coolant channel design elements considered were: length of combustion chamber in which high aspect ratio cooling was applied, number of coolant channels, and coolant channel shape. Seven coolant channel designs were investigated using a coupling of the Rocket Thermal Evaluation code and the Two-Dimensional Kinetics code. Initially, each coolant channel design was developed, without consideration for fabrication, to reduce the hot-gas-side wall temperature from a given conventional cooling channel baseline. These designs produced hot-gas-side wall temperature reductions up to 22 percent, with coolant pressure drop increases as low as 7.5 percent from the baseline. Fabrication constraints for milled channels were applied to the seven designs. These produced hot-gas-side wall temperature reductions of up to 20 percent, with coolant pressure drop increases as low as 2 percent. Using high aspect ratio cooling channels for the entire length of the combustion chamber had no additional benefit on hot-gas-side wall temperature over using high aspect ratio cooling channels only in the throat region, but increased coolant pressure drop 33 percent. Independent of coolant channel shape, high aspect ratio cooling was able to reduce the hot-gas-side wall temperature by at least 8 percent, with as low as a 2 percent increase in coolant pressure drop. The design with the highest overall benefit to hot-gas-side wall temperature and minimal coolant pressure drop cooling can now be done in relatively short periods of time with multiple iterations.
Wadel, Mary F.↗