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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 37 records · Page 2
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NASA NTRS1992

High heat flux measurements and experimental calibrations/characterizations

Recent progress in techniques employed in the measurement of very high heat-transfer rates in reentry-type facilities at the Arnold Engineering Development Center (AEDC) is described. These advances include thermal analyses applied to transducer concepts used to make these measurements; improved heat-flux sensor fabrication methods, equipment, and procedures for determining the experimental time response of individual sensors; performance of absolute heat-flux calibrations at levels above 2,000 Btu/cu ft-sec (2.27 kW/cu cm); and innovative methods of performing in-situ run-to-run characterizations of heat-flux probes installed in the test facility. Graphical illustrations of the results of extensive thermal analyses of the null-point calorimeter and coaxial surface thermocouple concepts with application to measurements in aerothermal test environments are presented. Results of time response experiments and absolute calibrations of null-point calorimeters and coaxial thermocouples performed in the laboratory at intermediate to high heat-flux levels are shown. Typical AEDC high-enthalpy arc heater heat-flux data recently obtained with a Calspan-fabricated null-point probe model are included.

Kidd, Carl T.↗
NASA NTRS1995

Nonequilibrium Flow Expansion Experiment Around a Blunted Cone

Eleven tests were carried out in DLR's high enthalpy tunnel in Goettingen (HEG), Germany, for reservoir conditions ranging from 10 to 23 MJ/kg and a free stream Mach number of approximately 10. A blunted cone model with a cylindrical afterbody (sting) was investigated. To obtain information on the influence of defined parameters on the body back flow, especially of the reacting gas, the heat transfer rate along the body contour was measured with fast response surface thermocouples on the forebody and sensitive thin film heat transfer gauges on the base and sting of the model. Flow visualization with a holographic interferometry system was provided. Tests are described, and a preliminary interpretation of the observed flow effects are given.

Kastell, Dirk↗
NASA NTRS1995

User's manual for the one-dimensional hypersonic experimental aero-thermodynamic (1DHEAT) data reduction code

A FORTRAN computer code for the reduction and analysis of experimental heat transfer data has been developed. This code can be utilized to determine heat transfer rates from surface temperature measurements made using either thin-film resistance gages or coaxial surface thermocouples. Both an analytical and a numerical finite-volume heat transfer model are implemented in this code. The analytical solution is based on a one-dimensional, semi-infinite wall thickness model with the approximation of constant substrate thermal properties, which is empirically corrected for the effects of variable thermal properties. The finite-volume solution is based on a one-dimensional, implicit discretization. The finite-volume model directly incorporates the effects of variable substrate thermal properties and does not require the semi-finite wall thickness approximation used in the analytical model. This model also includes the option of a multiple-layer substrate. Fast, accurate results can be obtained using either method. This code has been used to reduce several sets of aerodynamic heating data, of which samples are included in this report.

Hollis, Brian R.↗
NASA NTRS2002

Fundamental Boiling and RP-1 Freezing Experiments

The prestart thermal conditioning of the hardware in LOX (liquid oxygen) systems involve heat transfer between LOX and metal where boiling plays a large role. Information is easily found on nucleate boiling, maximum heat flux, minimum heat flux and film boiling for common fluids like water. After looking at these standard correlations it was felt more data was needed for the cool down side transition boiling for the LN2 and LOX. In particular interest is the film boiling values, the temperature at which transition begins and the slope as peak heat flux is approached. The ultimate goal is an array of boiling heat transfer coefficient as a function of surface temperature which can be used in the chilldown model of the feed system, engine and bleed system for X-34. The first experiment consisted of an actual MC-1 LOX Impeller which had been machined backwards, that was instrumented with 17 surface thermocouples and submerged in liquid nitrogen. The thermocouples were installed on metal thicknesses varying from the thin inducer to the thick hub.

Goode, Brian↗
NASA NTRS1996

Radiative Ignition and the Transition to Flame Spread Investigated in the Japan Microgravity Center's 10-sec Drop Shaft

The Radiative Ignition and Transition to Spread Investigation (RITSI) is a shuttle middeck Glovebox combustion experiment developed by the NASA Lewis Research Center, the National Institute for Standards and Technology (NIST), and Aerospace Design and Fabrication (ADF). It is scheduled to fly on the third United States Microgravity Payload (USMP-3) mission in February 1996. The objective of RITSI is to experimentally study radiative ignition and the subsequent transition to flame spread in low gravity in the presence of very low speed air flows in two- and three-dimensional configurations. Toward this objective, a unique collaboration between NASA, NIST, and the University of Hokkaido was established to conduct 15 science and engineering tests in Japan's 10-sec drop shaft. For these tests, the RITSI engineering hardware was mounted in a sealed chamber with a variable oxygen atmosphere. Ashless filter paper was ignited during each drop by a tungsten-halogen heat lamp focused on a small spot in the center of the paper. The flame spread outward from that point. Data recorded included fan voltage (a measure of air flow), radiant heater voltage (a measure of radiative ignition energy), and surface temperatures (measured by up to three surface thermocouples) during ignition and flame spread.

Source record↗
NASA NTRS2007

Shuttle Entry Imaging Using Infrared Thermography

During the Columbia Accident Investigation, imaging teams supporting debris shedding analysis were hampered by poor entry image quality and the general lack of information on optical signatures associated with a nominal Shuttle entry. After the accident, recommendations were made to NASA management to develop and maintain a state-of-the-art imagery database for Shuttle engineering performance assessments and to improve entry imaging capability to support anomaly and contingency analysis during a mission. As a result, the Space Shuttle Program sponsored an observation campaign to qualitatively characterize a nominal Shuttle entry over the widest possible Mach number range. The initial objectives focused on an assessment of capability to identify/resolve debris liberated from the Shuttle during entry, characterization of potential anomalous events associated with RCS jet firings and unusual phenomenon associated with the plasma trail. The aeroheating technical community viewed the Space Shuttle Program sponsored activity as an opportunity to influence the observation objectives and incrementally demonstrate key elements of a quantitative spatially resolved temperature measurement capability over a series of flights. One long-term desire of the Shuttle engineering community is to calibrate boundary layer transition prediction methodologies that are presently part of the Shuttle damage assessment process using flight data provided by a controlled Shuttle flight experiment. Quantitative global imaging may offer a complementary method of data collection to more traditional methods such as surface thermocouples. This paper reviews the process used by the engineering community to influence data collection methods and analysis of global infrared images of the Shuttle obtained during hypersonic entry. Emphasis is placed upon airborne imaging assets sponsored by the Shuttle program during Return to Flight. Visual and IR entry imagery were obtained with available airborne imaging platforms used within DoD along with agency assets developed and optimized for use during Shuttle ascent to demonstrate capability (i.e., tracking, acquisition of multispectral data, spatial resolution) and identify system limitations (i.e., radiance modeling, saturation) using state-of-the-art imaging instrumentation and communication systems. Global infrared intensity data have been transformed to temperature by comparison to Shuttle flight thermocouple data. Reasonable agreement is found between the flight thermography images and numerical prediction. A discussion of lessons learned and potential application to a potential Shuttle boundary layer transition flight test is presented.

Horvath, Thomas↗
NASA NTRS2010

The HYTHIRM Project: Flight Thermography of the Space Shuttle During the Hypersonic Re-entry

This report describes a NASA Langley led endeavor sponsored by the NASA Engineering Safety Center, the Space Shuttle Program Office and the NASA Aeronautics Research Mission Directorate to demonstrate a quantitative thermal imaging capability. A background and an overview of several multidisciplinary efforts that culminated in the acquisition of high resolution calibrated infrared imagery of the Space Shuttle during hypervelocity atmospheric entry is presented. The successful collection of thermal data has demonstrated the feasibility of obtaining remote high-resolution infrared imagery during hypersonic flight for the accurate measurement of surface temperature. To maximize science and engineering return, the acquisition of quantitative thermal imagery and capability demonstration was targeted towards three recent Shuttle flights - two of which involved flight experiments flown on Discovery. In coordination with these two Shuttle flight experiments, a US Navy NP-3D aircraft was flown between 26-41 nautical miles below Discovery and remotely monitored surface temperature of the Orbiter at Mach 8.4 (STS-119) and Mach 14.7 (STS-128) using a long-range infrared optical package referred to as Cast Glance. This same Navy aircraft successfully monitored the Orbiter Atlantis traveling at approximately Mach 14.3 during its return from the successful Hubble repair mission (STS-125). The purpose of this paper is to describe the systematic approach used by the Hypersonic Thermodynamic Infrared Measurements team to develop and implement a set of mission planning tools designed to establish confidence in the ability of an imaging platform to reliably acquire, track and return global quantitative surface temperatures of the Shuttle during entry. The mission planning tools included a pre-flight capability to predict the infrared signature of the Shuttle. Such tools permitted optimization of the hardware configuration to increase signal-to-noise and to maximize the available dynamic range while mitigating the potential for saturation. Post flight, analysis tools were used to assess atmospheric effects and to convert the 2-D intensity images to 3-D temperature maps of the windward surface. Comparison of the spatially resolved global thermal measurements to surface thermocouples and CFD prediction is made. Successful demonstration of a quantitative, spatially resolved, global temperature measurement on the Shuttle suggests future applications towards hypersonic flight test programs within NASA, DoD and DARPA along with flight test opportunities supporting NASA's project Constellation.

Horvath, Thomas J.↗
NASA NTRS2011

Venting of a Water/Inhibited Propylene Glycol Mixture in a Vacuum Environment-Characterization and Representative Test Results

A planned use of the Orion space vehicle involves its residence at the International Space Station for six months at a time. One concept of operations involves temporarily venting portions of the idle Orion active thermal control system (ATCS) during the docked phase, preventing freezing. The venting would have to be reasonably complete with few, if any, completely filled pockets of frozen liquid. Even if pockets of frozen liquid did not damage the hardware during the freezing process, they could prevent the system from filling completely prior to its reactivation. The venting of single component systems in a space environment has been performed numerous times and is well understood. Local nucleation occurs at warm, relatively massive parts of the system, which creates vapor and forces the bulk liquid out of the system. The remnants of the liquid will freeze, then evaporate over time through local heating. Because the Orion ATCS working fluid is a 50/50 mixture of water and inhibited propylene glycol, its boiling behavior was expected to differ from that of a pure fluid. It was thought that the relatively high vapor pressure water might evaporate preferentially, leaving behind a mixture enriched with the low vapor pressure propylene glycol, which would be vaporization ]resistant. Owing to this concern, a test was developed to compare the evaporation behavior of pure water, a 50/50 mixture of water and inhibited propylene glycol, and inhibited propylene glycol. The test was performed using room temperature fluids in an instrumented thin walled stainless steel vertical tube. The 1 in x 0.035 in wall tube was instrumented with surface thermocouples and encased in closed cell polyurethane foam. Reticulated polyurethane foam was placed inside the tube to reduce the convection currents. A vacuum system connected to the top of the tube set the pressure boundary condition. Tests were run for the three fluids at back pressures ranging from 1 to 18 torr. During each test, the mass of the test article was measured as it changed over time, as was its temperature and backpressure. The tests were successful. Somewhat surprisingly, the results showed that the evaporation behavior of the three fluids had more similarities than differences. The 50/50 mixture evaporated similarly to the pure water - albeit at a slower rate. The test results indicate that our extensive space - based experience with venting of single component fluids can be applied to the problem of Orion ATCS venting as long as the appropriate puts, takes, and caveats are applied.

Ungar, Eugene K.↗
NASA NTRS2012

Thermographic Imaging of the Space Shuttle During Re-Entry Using a Near Infrared Sensor

High resolution calibrated near infrared (NIR) imagery of the Space Shuttle Orbiter was obtained during hypervelocity atmospheric re-entry of the STS-119, STS-125, STS-128, STS-131, STS-132, STS-133, and STS-134 missions. This data has provided information on the distribution of surface temperature and the state of the airflow over the windward surface of the Orbiter during descent. The thermal imagery complemented data collected with onboard surface thermocouple instrumentation. The spatially resolved global thermal measurements made during the Orbiter s hypersonic re-entry will provide critical flight data for reducing the uncertainty associated with present day ground-to-flight extrapolation techniques and current state-of-the-art empirical boundary-layer transition or turbulent heating prediction methods. Laminar and turbulent flight data is critical for the validation of physics-based, semi-empirical boundary-layer transition prediction methods as well as stimulating the validation of laminar numerical chemistry models and the development of turbulence models supporting NASA s next-generation spacecraft. In this paper we provide details of the NIR imaging system used on both air and land-based imaging assets. The paper will discuss calibrations performed on the NIR imaging systems that permitted conversion of captured radiant intensity (counts) to temperature values. Image processing techniques are presented to analyze the NIR data for vignetting distortion, best resolution, and image sharpness. Keywords: HYTHIRM, Space Shuttle thermography, hypersonic imaging, near infrared imaging, histogram analysis, singular value decomposition, eigenvalue image sharpness

Zalameda, Joseph N.↗
NASA NTRS2016

Orion EFT-1 Cavity Heating Tile Experiments and Environment Reconstruction

Developing aerothermodynamic environments for deep cavities, such as those produced by micrometeoroids and orbital debris impacts, poses a great challenge for engineers. In order to assess existing cavity heating models, two one-inch diameter cavities were flown on the Orion Multi-Purpose Crew Vehicle during Exploration Flight Test 1 (EFT1). These cavities were manufactured with depths of 1.0 in and 1.4 in, and they were both instrumented. Instrumentation included surface thermocouples upstream, downstream and within the cavities, and additional thermocouples at the TPS-structure interface. This paper will present the data obtained, and comparisons with computational predictions will be shown. Additionally, the development of a 3D material thermal model will be described, which will be used to account for the three-dimensionality of the problem when interpreting the data. Furthermore, using a multi-dimensional inverse heat conduction approach, a reconstruction of a time- and space-dependent flight heating distribution during EFT1 will be presented. Additional discussions will focus on instrumentation challenges and calibration techniques specific to these experiments. The analysis shown will highlight the accuracies and/or deficiencies of current computational techniques to model cavity flows during hypersonic re-entry.

Salazar, Giovanni↗
NASA NTRS2016

Exploration Flight Test 1 Afterbody Aerothermal Environment Reconstruction

The Exploration Flight Test 1 vehicle included roughly 100 near surface thermocouples on the after body of the vehicle. The temperature traces at each of these instruments have been used to perform inverse environment reconstruction to determine the aerothermal environment experienced during re-entry of the vehicle. This paper provides an overview of the reconstructed environments and identifies critical aspects of the environment. These critical aspects include transition and reaction control system jet influence. A blind test of the process and reconstruction tool was also performed to build confidence in the reconstructed environments. Finally, an uncertainty quantification analysis was also performed to identify the impact of each of the uncertainties on the reconstructed environments.

Hyatt, Andrew J.↗
NASA NTRSDate not supplied

MEDLI2: MISP Inferred Aerothermal Environment and Flow Transition Assessment

The Mars Entry, Descent, and Landing Instrumentation 2 (MEDLI2) sensor suite on the Mars2020 mission contained multiple sensors on the aeroshell to measure the aerothermal environment during entry into the Martian atmosphere. These sensors performed superbly and successfully returned forebody and aftbody heating measurements. Analysis of MEDLI2 data indicated flow transitioning from a laminar to turbulent state on the heatshield. No evidence of flow transition was observed on the backshell. Two methods were used to estimate flow transition times on the heatshield: (1) temperature gradient of near-surface thermocouple data and (2) heat flux gradient from an inverse reconstruction approach using thermocouple data and material response modeling. Both methods produced similar transition times with an estimated accuracy of ±1 s. To assess various transition criteria, transition parameters were evaluated at each sensor location using flow field solutions from computational fluid dynamics (CFD) simulations. The idea was to use conservative values inferred from MEDLI2 data as transition criteria for other Mars missions. To test this hypothesis, MEDLI data from the Mars Science Laboratory (MSL) mission was used to compare predicted vs. actual flow transition times. The comparisons suggest smooth wall transition criteria are not well-suited in modeling the rapid progression of a turbulent transition front. Transition criteria containing a roughness element parameter agreed better with the flight data. In summary, critical transition values derived from MEDLI2 data may be used as a starting point in constructing a flow transition model for future Mars missions.

Chun Y Tang↗
NASA NTRS2001

Thermocouple Rakes for Measuring Boundary Layer Flows Extremely Close to Surface

Of vital interest to aerodynamic researchers is precise knowledge of the flow velocity profile next to the surface. This information is needed for turbulence model development and the calculation of viscous shear force. Though many instruments can determine the flow velocity profile near the surface, none of them can make measurements closer than approximately 0.01 in. from the surface. The thermocouple boundary-layer rake can measure much closer to the surface than conventional instruments can, such as a total pressure boundary layer rake, hot wire, or hot film. By embedding the sensors (thermocouples) in the region where the velocity is equivalent to the velocity ahead of a constant thickness strut, the boundary-layer flow profile can be obtained. The present device fabricated at the NASA Glenn Research Center microsystem clean room has a heater made of platinum and thermocouples made of platinum and gold. Equal numbers of thermocouples are placed both upstream and downstream of the heater, so that the voltage generated by each pair at the same distance from the surface is indicative of the difference in temperature between the upstream and downstream thermocouple locations. This voltage differential is a function of the flow velocity, and like the conventional total pressure rake, it can provide the velocity profile. In order to measure flow extremely close to the surface, the strut is made of fused quartz with extremely low heat conductivity. A large size thermocouple boundary layer rake is shown in the following photo. The latest medium size sensors already provide smooth velocity profiles well into the boundary layer, as close as 0.0025 in. from the surface. This is about 4 times closer to the surface than the previously used total pressure rakes. This device also has the advantage of providing the flow profile of separated flow and also it is possible to measure simultaneous turbulence levels within the boundary layer.

Hwang, Danny P.↗
NASA NTRS1987

Design, calibration and error analysis of instrumentation for heat transfer measurements in internal combustion engines

The paper reports the methodology and uncertainty analyses of instrumentation for heat transfer measurements in internal combustion engines. Results are presented for determining the local wall heat flux in an internal combustion engine (using a surface thermocouple-type heat flux gage) and the apparent flame-temperature and soot volume fraction path length product in a diesel engine (using two-color pyrometry). It is shown that a surface thermocouple heat transfer gage suitably constructed and calibrated will have an accuracy of 5 to 10 percent. It is also shown that, when applying two-color pyrometry to measure the apparent flame temperature and soot volume fraction-path length, it is important to choose at least one of the two wavelengths to lie in the range of 1.3 to 2.3 micrometers. Carefully calibrated two-color pyrometer can ensure that random errors in the apparent flame temperature and in the soot volume fraction path length will remain small (within about 1 percent and 10-percent, respectively).

Ferguson, C. R.↗
NASA NTRS1987

Heat-Conducting Anchors for Thermocouples

Metal particles in adhesive aid heat transfer. Aluminum caps containing silver-filled epoxy used as high-thermal-conductance anchors for thermocouples, epoxy providing thermal path between mounting surfaces and thermocouple measuring junctions. Normally, epoxy-filled aluminum caps used when measuring steady-state temperatures. Silver-filled epoxy used when thermocouple not isolated electrically from surface measured.

Macdavid, Kenton S.↗
NASA NTRS2017

Decoupled Method for Reconstruction of Surface Conditions From Internal Temperatures On Ablative Materials With Uncertain Recession Model

Obtaining measurements of flight environments on ablative heat shields is both critical for spacecraft development and extremely challenging due to the harsh heating environment and surface recession. Thermocouples installed several millimeters below the surface are commonly used to measure the heat shield temperature response, but an ill-posed inverse heat conduction problem must be solved to reconstruct the surface heating environment from these measurements. Ablation can contribute substantially to the measurement response making solutions to the inverse problem strongly dependent on the recession model, which is often poorly characterized. To enable efficient surface reconstruction for recession model sensitivity analysis, a method for decoupling the surface recession evaluation from the inverse heat conduction problem is presented. The decoupled method is shown to provide reconstructions of equivalent accuracy to the traditional coupled method but with substantially reduced computational effort. These methods are applied to reconstruct the environments on the Mars Science Laboratory heat shield using diffusion limit and kinetically limited recession models.

Oliver, A. Brandon↗