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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS2008

Laser-Induced Thermal Acoustic Measurements in a Highly Back-Pressured Scramjet Isolator Model: A Research Plan

Under the Propulsion Discipline of NASA s Fundamental Aeronautics Program s Hypersonics Project, a test apparatus, for testing a scramjet isolator model, is being constructed at NASA's Langley Research Center. The test apparatus will incorporate a 1-inch by 2-inch by 15-inch-long scramjet isolator model supplied with 2.1 lbm/sec of unheated dry air through a Mach 2.5 converging-diverging nozzle. The planned research will incorporate progressively more challenging measurement techniques to characterize the flow field within the isolator, concluding with the application of the Laser-Induced Thermal Acoustic (LITA) measurement technique. The primary goal of this research is to use the data acquired to validate Computational Fluid Dynamics (CFD) models employed to characterize the complex flow field of a scramjet isolator. This paper describes the test apparatus being constructed, pre-test CFD simulations, and the LITA measurement technique.

Middleton, Troy F.↗
NASA NTRS1992

Development of a hydrogen external burning flight test experiment on the NASA Dryden SR-71A aircraft

A captive-carry flight test experiment of the transonic external burning concept using gaseous hydrogen fuel has ben proposed for the SR-71A flight research test bed aircraft at the NASA Dryden Flight Research Facility, Edwards, California. This program will be the first large-scale, near-field flight test investigation of this base drag reduction technique at Mach numbers up to 3. Low-speed base drag reduction for supersonic and hypersonic flight vehicles is important in reducing vehicle drag and improving nozzle efficiency. Flight data will be correlated with a concurrent wind tunnel ground test of a similar test apparatus to confirm geometry and component performance scalability predictions and to investigate design parametrics under actual flight conditions. This paper describes the problem of transonic base drag for supersonic and hypersonic aircraft, a planned flight research program to address this problem, the test bed aircraft, and the flight experiment apparatus.

Hicks, John W.↗
NASA NTRS1993

Experimental research activities in dynamic response and sonic fatigue of hypersonic vehicle structures at NASA Langley Research Center

This paper presents an overview of experimental research activities being pursued at the NASA Langley Research Center for dynamic response and sonic fatigue of hypersonic vehicle structures. The capabilities of the principle test facility, the Thermal Acoustic Fatigue Apparatus, are first given. Results from recent dynamic response and sonic fatigue tests on candidate hypersonic vehicle structures are then presented.

Rizzi, Stephen A.↗
NASA NTRS2007

Fluorescence Visualization of Hypersonic Flow over Rapid Prototype Wind-Tunnel Models

Reentry models for use in hypersonic wind tunnel tests were fabricated using a stereolithography apparatus. These models were produced in one day or less, which is a significant time savings compared to the manufacture of ceramic or metal models. The models were tested in the NASA Langley Research Center 31-Inch Mach 10 Air Tunnel. Most of the models did not survive repeated tests in the tunnel, and several failure modes of the models were identified. Planar laser-induced fluorescence (PLIF) of nitric oxide (NO) was used to visualize the flowfields in the wakes of these models. Pure NO was either seeded through tubes plumbed into the model or via a tube attached to the strut holding the model, which provided localized addition of NO into the model s wake through a porous metal cylinder attached to the end of the tube. Models included several 2-inch diameter Inflatable Reentry Vehicle Experiment (IRVE) models and 5-inch diameter Crew Exploration Vehicle (CEV) models. Various configurations were studied including different sting placements relative to the models, different model orientations and attachment angles, and different NO seeding methods. The angle of attack of the models was also varied and the location of the laser sheet was scanned to provide three-dimensional flowfield information. Virtual Diagnostics Interface technology, developed at NASA Langley, was used to visualize the data sets in post processing. The use of calibration "dotcards" was investigated to correct for camera perspective and lens distortions in the PLIF images. Lessons learned and recommendations for future experiments are discussed.

Alderfer, D. W.↗
NASA NTRS2008

Fluorescence Imaging and Streamline Visualization of Hypersonic Flow over Rapid Prototype Wind-Tunnel Models

Reentry models for use in hypersonic wind tunnel tests were fabricated using a stereolithography apparatus. These models were produced in one day or less, which is a significant time savings compared to the manufacture of ceramic or metal models. The models were tested in the NASA Langley Research Center 31-Inch Mach 10 Air Tunnel. Only a few of the models survived repeated tests in the tunnel, and several failure modes of the models were identified. Planar laser-induced fluorescence (PLIF) of nitric oxide (NO) was used to visualize the flowfields in the wakes of these models. Pure NO was either seeded through tubes plumbed into the model or via a tube attached to the strut holding the model, which provided localized addition of NO into the model s wake through a porous metal cylinder attached to the end of the tube. Models included several 2- inch diameter Inflatable Reentry Vehicle Experiment (IRVE) models and 5-inch diameter Crew Exploration Vehicle (CEV) models. Various model configurations and NO seeding methods were used, including a new streamwise visualization method based on PLIF. Virtual Diagnostics Interface (ViDI) technology, developed at NASA Langley Research Center, was used to visualize the data sets in post processing. The use of calibration "dotcards" was investigated to correct for camera perspective and lens distortions in the PLIF images.

Danehy, Paul M.↗
NASA NTRS1992

Thermal control/oxidation resistant coatings for titanium-based alloys

Extensive research and development efforts have been expended toward development of thermal control and environmental protection coatings for NASP and generic hypersonic vehicle applications. The objective of the coatings development activities summarized here was to develop light-weight coatings for protecting advanced titanium alloys from oxidation in hypersonic vehicle applications. A number of new coating concepts have been evaluated. Coated samples were exposed to static oxidation tests at temperatures up to 1000 C using a thermogravimetric apparatus. Samples were also exposed to simulated hypersonic flight conditions for up to 10 hr to determine their thermal and chemical stability and catalytic efficiency. The emittance of samples was determined before and after exposure to simulated hypersonic flight conditions.

Clark, Ronald K.↗
NASA NTRS2012

Improved Finite-Volume Method for Radiative Hydrodynamics

Fully coupled simulations of hydrodynamics and radiative transfer are essential to a number of fields ranging from astrophysics to engineering applications. Of particular interest in this work are hypersonic atmospheric entries and associated experimental apparatus, e.g., shock tubes and high enthalpy testing facilities. The radiative transfer calculations must supply to the CFD a heating term in the energy equation in the form of the divergence of the radiative heat flux and the radiative heat fluxes to bounding surfaces. It is most efficient to solve the radiative transfer equation on the same grid as the CFD solution, and this work presents an algorithm with improved accuracy for such simulations on structured and unstructured grids compared to more conventional approaches. Results will be shown for shock radiation during hypersonic reentry. Issues of parallelization within a radiation sweep will also be discussed.

Wray, Alan↗
NASA NTRS1994

Design of a variable area diffuser for a 15-inch Mach 6 open-jet tunnel

The Langley 15-inch Mach 6 High Temperature Tunnel was recently converted from a Mach 10 Hypersonic Flow Apparatus. This conversion was effected to improve the capability of testing in Mach 6 air at relatively high reservoir temperatures not previously possible at Langley. Elevated temperatures allow the matching of the Mach numbers, Reynolds numbers, and ratio of wall-to-adiabatic-wall temperatures (TW/Taw) between this and the Langley 20-inch Mach 6 CF4 Tunnel. This ratio is also matched for Langley's 31-inch Mach 10 Tunnel and is an important parameter useful in the simulation of slender bodies such as National Aerospace Plane (NASP) configurations currently being studied. Having established the nozzle's operating characteristics, the decision was made to install another test section to provide model injection capability. This test section is an open-jet type, with an injection system capable of injecting a model from retracted position to nozzle centerline between 0.5 and 2 seconds. Preliminary calibrations with the new test section resulted in Tunnel blockage. This blockage phenomenon was eliminated when the conical center body in the diffuser was replaced. The issue then, is to provide a new and more efficient variable area diffuser configuration with the capability to withstand testing of larger models without sending the Tunnel into an unstart condition. Use of the 1-dimensional steady flow equation with due regard to friction and heat transfer was employed to estimate the required area ratios (exit area / throat area) in a variable area diffuser. Correlations between diffuser exit Mach number and area ratios, relative to the stagnation pressure ratios and diffuser inlet Mach number were derived. From these correlations, one can set upper and lower operating pressures and temperatures for a given diffuser throat area. In addition, they will provide appropriate input conditions for the full 3-dimensional computational fluid dynamics (CFD) code for further simulation studies.

Loney, Norman W.↗
NASA NTRS1973

Apparatus and method for generating large mass flow of high temperature air at hypersonic speeds

A description is given of tests performed in the Mach number 8 to 10 regime to generate high temperature air at hypersonic speeds. The test flow was obtained by pressuring a ceramic lined storage tank with air to a pressure of about 100 to 200 atm. The air is heated to temperatures of 7,000 R to 8,000 R prior to introduction into the tank by passing air through an electric arc heater. The air cools to 5,500 R to 6,000 R while in the tank and then it is rapidly released through a Mach number 8 to 10 nozzle. Cold air or hot driver gas under pressure is injected simultaneously with the release of pressurized hot air and regulated to enter the tank at the same rate the hot air is leaving. This holds pressure in the tank constant during the test. Alternatively, the driver gas is introduced into the tank prior to the discharge of the hot air to increase the pressure in the tank and permit tests at higher pressures than can be handled by the arc heater.

Boatright, W. B.↗
NASA NTRS1992

The Langley 15-inch Mach 6 High Temperature Tunnel

The Langley 15-Inch Mach 6 High Temperature Tunnel was recently converted from the former Mach 10 Hypersonic Flow Apparatus. This conversion was made primarily to provide the capability of testing in Mach 6 air at higher reservoir temperatures than previously possible at Langley. These elevated temperatures allow the matching of the Mach number, Reynolds number, and ratio of wall-to-adiabatic-wall temperatures (Tw/Taw) between this facility and the Langley 20-Inch Mach 6 CF4 Tunnel; Tw/Taw is also matched for Langley's 31-Inch Mach 10 Tunnel. A brief history and general description of this facility are presented along with a discussion of the recently completed modifications and upgrades.

Hodge, Jeffrey S.↗
NASA NTRSDate not supplied

Testing of ACC-6 Carbon-Carbon at High Strain Rates and High Temperatures

Some materials, such as steel, titanium, and tungsten, have a yield stress that depends on the strain rate, and actually are stronger at high rates of load or deformation. The increase in strength is relevant; and if not taken into account, the computer models will not match the experimental behavior observed in typical shock loading, ordnance, or ballistic tests. For higher velocity (hypervelocity impacts), an equation of state that accounts for non-linearities and temperature in the pressure-volume relation, may also be needed. This rate-dependent behavior of the strength has also been observed in some composite materials. Southwest Research Institute® (SwRI®) has started an internal research effort to measure and better model the behavior of materials, in particular carbon/carbon composites like ACC-6, at ultra-high temperatures for both static and high rates of deformation. The ultimate objective is to obtain reliable computer models of carbon/carbon material under hypersonic flight conditions, hypervelocity impacts, shock-loading situations such as booster-vehicle separation, and munition operation events. SwRI® has successfully modeled in the past ballistic impact on other composites, namely S-2 glass/epoxy, Kevlar/resin, polyethylene fibers/resin, and carbon fiber composites. This presentation will show test results relevant to the strength properties of the ACC-6 C/C material which were measured at quasistatic, medium, and the moderately high strain rates (~103 s—1) achievable in a conventional Split Hopkinson bar apparatus. Some of the tests were performed at both high-rates and high temperatures typical of hypersonic flight, something that, to the authors’ knowledge, has never been published in the past for this particular material. The test method will be presented, as well as the comparison of the stress-strain curves for the different temperatures and strain-rates of interest. A discussion of how the results can be used in the typical hydrocodes of interest will be included.

carbon/carbon↗
NASA NTRS1973

Apparatus and method for generating large mass flow of high temperature air at hypersonic speeds

High temperature, high mass air flow and a high Reynolds number test air flow in the Mach number 8-10 regime of adequate test flow duration is attained by pressurizing a ceramic-lined storage tank with air to a pressure of about 100 to 200 atmospheres. The air is heated to temperatures of 7,000 to 8,000 R prior to introduction into the tank by passing the air over an electric arc heater means. The air cools to 5,500 to 6,000 R while in the tank. A decomposable gas such as nitrous oxide or a combustible gas such as propane is injected into the tank after pressurization and the heated pressurized air in the tank is rapidly released through a Mach number 8-10 nozzle. The injected gas medium upon contact with the heated pressurized air effects an exothermic reaction which maintains the pressure and temperature of the pressurized air during the rapid release.

Sabol, A. P.↗
NASA NTRS1981

Hot-film static-pressure probe for surveying flow fields

A static pressure probe employing hot-film sensors has been developed for the rapid measurement of the static pressure fields surrounding analytic shapes in hypersonic flows. The hot-film probe is a modification of the standard static pressure probe, consisting of a front hot-film sensor operated as a resistance thermometer, a rear sensor operated at an overheat ratio of 1.5 to 1.8 and a small sonic orifice installed inside the tubing of a conventional device. The probe has been calibrated in helium and air over a range of temperatures and pressures in a bell jar apparatus, with a repeatability of the data to within + or - 0.015 mm Hg. Comparative tests of the hot-film and conventional static pressure probes in a hypersonic helium wind tunnel at Mach 20 and various Reynolds numbers have indicated the settling time of the hot-film probe to be on the order of milliseconds, as compared with 30 sec for the conventional probe. The pressures measured by the two probes were found to be within 10% of each other. Although the hot-film probe makes flow-field static pressure surveys more practical in blowdown hypersonic wind tunnels, viscous and flow angle effects still must be assessed under the conditions of use.

Ashby, G. C., Jr.↗