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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 847 records · Page 47
All sourcesNASA NTRSDOE OSTI
NASA NTRS1997

Modeling Cyclic Phase Change and Energy Storage in Solar Heat Receivers

Numerical results pertaining to cyclic melting and freezing of an encapsulated phase change material (PCM), integrated into a solar heat receiver, have been reported. The cyclic nature of the present melt/freeze problem is relevant to latent heat thermal energy storage (LHTES) systems used to power solar Brayton engines in microgravity environments. Specifically, a physical and numerical model of the solar heat receiver component of NASA Lewis Research Center's Ground Test Demonstration (GTD) project was developed and results compared with available experimental data. Multi-conjugate effects such as the convective fluid flow of a low-Prandtl-number fluid, coupled with thermal conduction in the phase change material, containment tube and working fluid conduit were accounted for in the model. A single-band thermal radiation model was also included to quantify reradiative energy exchange inside the receiver and losses through the aperture. The eutectic LiF-CaF2 was used as the phase change material (PCM) and a mixture of He/Xe was used as the working fluid coolant. A modified version of the computer code HOTTube was used to generate results for comparisons with GTD data for both the subcooled and two-phase regimes. While qualitative trends were in close agreement for the balanced orbit modes, excellent quantitative agreement was observed for steady-state modes.

Hall, Carsie A., III↗
NASA NTRS2001

Hypersonic Flows About a 25 degree Sharp Cone

This paper presents the results of a numerical study that examines the surface heating discrepancies observed between computed and measured values along a sharp cone. With Mach numbers of an order of 10 and the freestream length Reynolds number of an order of 10 000, the present computations have been made with the direct simulation Monte Carlo (DSMC) method by using the G2 code of Bird. The flow conditions are those specified for two experiments conducted in the Veridian 48-inch Hypersonic Shock Tunnel. Axisymmetric simulations are made since the test model was assumed to be at zero incidence. Details of the current calculations are presented, along with comparisons between the experimental data, for surface heating and pressure distributions. Results of the comparisons show major differences in measured and calculated results for heating distributions, with differences in excess of 25 percent for the two cases examined.

Moss, James N.↗
NASA NTRSDate not supplied

Design and Analysis of Battery Thermal Management Systems

Thermal management of battery cell packs is a critically needed technology. The purpose of this work is to design new and improved Battery Thermal Management Systems (BTMS) for use in electric airplanes. The BTMS should be 3D printable and hold twelve to sixteen 18650 batteries. To minimize aircraft’s weight, an actively air-cooled battery pack was chosen to avoid the excess weight of water-cooled and phase change material (PCM) cooled packs. As a result, significant changes were made to remove all metal in the pack and replace it with lower density polymer matrix composites (PMCs) which can be additively manufactured. Two air cooled battery pack configurations (traditional propeller fans & bladeless fans) were designed, modeled, and compared. These packs were first modeled in SolidWorks 2021 3D CAD, then imported into COMSOL MultiPhysics to be studied using the “Heat Transfer in Solids and Fluids” module. The design with the bladeless fans eliminated the need to use high conductivity heavy metal to remove unwanted thermal energy. These bladeless fans were designed entirely out of PMCs. This thermal pack design weighs 0.04 kg less than that of the traditional propeller fans design and has increased the battery pack energy density by 8.25 Wh/kg.

Thermal Management System for Battery Packs↗
NASA NTRS1975

Heat-transfer tests on the Rockwell International space shuttle orbiter with and without simulated protuberances

Aerothermodynamic tests on the forward half of the Rockwell International Space Shuttle Orbiter Configuration 140C were conducted at Mach number 8. The phase-change paint and thin-skin thermocouple techniques were used to determine the aerodynamic heating rates on the Orbiter models during simulated atmospheric reentry. Smooth 0.04-scale models and models with scaled protuberances and indentations which simulated the windshields, cargo bay door hinges, vents, and thruster nozzles were tested over an angle-of-attack range from 20 to 45 deg at yaw angles from -5 to 5 deg and at Reynolds numbers, based on the total Orbiter scaled length, from 2.15 to 15.9 million. Comparisons of the model heat-transfer rates obtained with a smooth surface and with scaled protuberances are presented.

Carter, L. D.↗
NASA NTRS1986

Functional and performance tests of two capillary pumped loop engineering models

This paper presents the results of the functional and performance tests for two capillary pumped loop (CPL) engineering models. Both CPL systems are aluminum/ammonia transport systems which contain eight parallel evaporators and six parallel condensers in a single loop. Tests conducted include the transport limit, heat load shearing between evaporators, liquid inventory/temperature control by the reservoir, pressure priming under heat load, diode function of condensers, and isolation of a single deprimed evaporator. Consistent performance results were obtained for both systems. Transport capabilities of up to 70 kw-m with individual evaporators managing up to 1.7 kw, with a corresponding input heat flux of 15w/sq cm, were demonstrated. These tests demonstrated the ability of a CPL system to operate over a wide range of conditions and thus established the viability of these systems for high power thermal management of large spacecraft, such as the NASA Space Station.

Ku, J.↗
NASA NTRS1986

Post-perihelion brightening of Halley's Comet: A case of nuclear summer

Increased brightness and gas production rates for Halley's comet after perihelion are explained as a result of seasonal effects on an obliquely rotating nucleus. The highly eccentric cometary orbit causes a rapid change in solar declination as the comet rounds perihelion, resulting in drastic changes in the insolation reaching the Northern and Southern Hemispheres of the nucleus. The rapid heating of the Northern Hemisphere post-perihelion likely results in substantial cracking of the non-volatile surface crust due to thermal stresses, exposing areas of fresh volatile ices. The orientation of the triaxial ellipsoid nucleus may also play a role in exposing more surface area to continuous sunlight and sublimation after perihelion. Post-perihelion brightening models based on heat flow and storage in sub-surface layers of the nucleus pre-perihelion are likely not viable because of the low thermal conductivities of porous, low density cometary surface materials.

Weissman, Paul R.↗
NASA NTRS1994

Cable Discharge System for fundamental detonator studies

Sandia National Laboratories has recently completed the modification and installation of a cable discharge system (CDS) which will be used to study the physics of exploding bridgewire (EBW) detonators and exploding foil initiators (EFI or slapper). Of primary interest are the burst characteristics of these devices when subjected to the constant current pulse delivered by this system. The burst process involves the heating of the bridge material to a conductive plasma and is essential in describing the electrical properties of the bridgewire foil for use in diagnostics or computer models. The CDS described herein is capable of delivering up to an 8000 A pulse of 3 micron duration. Experiments conducted with the CDS to characterize the EBW and EFI burst behavior are also described. In addition, the CDS simultaneous VISAR capability permits updating the EFI electrical Gurney analysis parameters used in our computer simulation codes. Examples of CDS generated data for a typical EFI and EBW detonator are provided.

Peevy, Gregg R.↗
NASA NTRS2017

Altitude Scaling of Thermal Ice Protection Systems in Running Wet Operation

A study into the effects of altitude on an aircraft thermal Ice Protection System (IPS) performance has been conducted by the National Research Council Canada (NRC) in collaboration with the NASA Glenn Icing Branch. The study included tests of an airfoil model, with a heated-air IPS, installed in the NRCs Altitude Icing Wind Tunnel (AIWT) at altitude and ground level conditions.

Running wet mode↗
NASA NTRSDate not supplied

CFD Simulations of the IHF 13-Inch Nozzle Flow: 55-deg Sphere-Cone Model, Manufactured Fences and Gaps

This paper reports computational analyses of testsin a high enthalpy arc-jet facility at NASA Ames Research Center. These tests were conducted using 55-deg sphere-cone models placed in a free jet downstream of the 13-inch diameter conical nozzle in the Ames 60-MW Interaction Heating Facility. Some of the sphere-cone models include surface features such as manufactured fences and gaps intended to simulate effects of differential recession, all of which disturb the flow, producing augmented heating locally and downstream.Test calibration data were obtained using slug and Gardon gage stagnation calorimeters, and a sphere-cone calorimeter model with six Gardon gages and five pressure tabs.The present analysis comprises computational fluid dynamics simulations of the nonequilibrium flowfield in the facility nozzle and test box, including the models tested, and comparisons with the experimental measurements. These simulations take into account nonuniform total enthalpy and mass flux profiles at the nozzle inlet as well as the expansion waves emanating from the nozzle exit and their effects on the model flowfields.

Arc-jets↗
NASA NTRSDate not supplied

CFD Simulations of the IHF 13-Inch Nozzle Flow: 55-deg Sphere-Cone Model, Manufactured Fences and Gaps

This paper reports computational analyses of tests in a high enthalpy arc-jet facility at NASA Ames Research Center. These tests were conducted using 55-deg sphere-cone models placed in a free jet downstream of the 13-inch diameter conical nozzle in the Ames 60-MW Interaction Heating Facility. Some of the sphere-cone models include surface features such as manufactured fences and gaps intended to simulate effects of differential recession, all of which disturb the flow, producing augmented heating locally and downstream. Test calibration data were obtained using slug and Gardon gage stagnation calorimeters, and a sphere-cone calorimeter model with six Gardon gages and five pressure tabs. The present analysis comprises computational fluid dynamics simulations of the nonequilibrium flowfield in the facility nozzle and test box, including the models tested, and comparisons with the experimental measurements. These simulations take into account nonuniform total enthalpy and mass flux profiles at the nozzle inlet as well as the expansion waves emanating from the nozzle exit and their effects on the model flowfields.

Arc-jets↗
NASA NTRSDate not supplied

CFD Simulations of the IHF 13-Inch Nozzle Flow: 55-deg Sphere-Cone Model, Manufactured Fences and Gaps

This paper reports computational analyses of tests in a high enthalpy arc-jet facility at NASA Ames Research Center. These tests were conducted using 55-deg sphere-cone models placed in a free jet downstream of the 13-inch diameter conical nozzle in the Ames 60-MW Interaction Heating Facility. Some of the sphere-cone models include surface features such as manufactured fences and gaps intended to simulate effects of differential recession, all of which disturb the flow, producing augmented heating locally and downstream. Test calibration data were obtained using slug and Gardon gage stagnation calorimeters, and a sphere-cone calorimeter model with six Gardon gages and five pressure tabs. The present analysis comprises computational fluid dynamics simulations of the nonequilibrium flowfield in the facility nozzle and test box, including the models tested, and comparisons with the experimental measurements. These simulations take into account nonuniform total enthalpy and mass flux profiles at the nozzle inlet as well as the expansion waves emanating from the nozzle exit and their effects on the model flowfields.

Arc-jets↗
NASA NTRS1982

Computer modeling of thermoelectric generator performance

Features of the DEGRA 2 computer code for simulating the operations of a spacecraft thermoelectric generator are described. The code models the physical processes occurring during operation. Input variables include the thermoelectric couple geometry and composition, the thermoelectric materials' properties, interfaces and insulation in the thermopile, the heat source characteristics, mission trajectory, and generator electrical requirements. Time steps can be specified and sublimation of the leg and hot shoe is accounted for, as are shorts between legs. Calculations are performed for conduction, Peltier, Thomson, and Joule heating, the cold junction can be adjusted for solar radition, and the legs of the thermoelectric couple are segmented to enhance the approximation accuracy. A trial run covering 18 couple modules yielded data with 0.3% accuracy with regard to test data. The model has been successful with selenide materials, SiGe, and SiN4, with output of all critical operational variables.

Chmielewski, A. B.↗
NASA NTRS1986

Methods for computing comet core temperatures

The temperature profile within the comet nucleus provides the key to an understanding of the history of the volatiles within a comet. Certain difficulties arise in connection with current cometary temperature models. It is shown that the constraint of zero net heat flow can be used to derive general analytical expressions which will allow for the determination of comet core temperature for a spherically symmetric comet, taking into account information about the surface temperature and the thermal conductivity. The obtained results are compared with the expression for comet core temperatures considered by Klinger (1981). Attention is given to analytical results, an example case, and numerical models. The formalization developed makes it possible to determine the core temperature on the basis of the numerical models of the surface temperature.

Mckay, C. P.↗
NASA NTRSDate not supplied

Thermomechanical Modeling of Woven Materials With Particle-Based, Explicit-Fiber Simulations

Fiber-based materials are extensively used to protect spacecraft during entry. Insulative fibers, often in a fiber network or woven, provide rigidity, strength, and control of material anisotropy and density. Woven thermal protection materials, such as ADEPT (Adaptable, Deployable Entry and Placement Technology), 3D-MAT (3-Dimensional Multifunctional Ablative Thermal Protection), and 3MDCP (3D Woven Mid-Density Carbon Phenolic), enable missions with stronger and denser materials for entry profiles with high shear and heat flux. Vulnerabilities to woven thermal protection materials include manufacturing-induced material property variation, and impact from micrometeoroids. Simulating woven materials under these conditions require models that can resolve hierarchal structures, thermomechanical behavior, and failure. To address this, we simulate weave thermal conduction and mechanical deformation. We simulate the full weave with a coarse-grained yarn model is presented. The model combines a validated, high-resolution single 3MDCP yarn model and phenolic resin model. Instead of modeling every fiber, each yarn ply with order 10, instead of order 1000, fibers. The discrete element bonded particle model (DEM-BPM) of fibers captures the thermal and mechanical behavior within and between fibers. We study the proportion of heat transfer and stress via the contact network, fiber bonds, and overall weave geometry.

bonded particle↗
NASA NTRS1961

Boundary-Layer Transition on a Group of Blunt Nose Shapes at a Mach Number of 2.20

An investigation has been made to study boundary-layer transition on six axisymmetrical blunt bodies of revolution. Model shapes were selected with respect to the degree of favorable pressure gradient over the model surface. Tests were conducted at a Mach number of 2.20 and over a range of free-stream Reynolds number per foot of about 1.4 x 10(exp 6) to 6.5 x 10(exp 6). The tests were made at an angle of attack of 0 deg. with zero heat transfer. For the hemisphere, the flow remained essentially laminar over the model surface length for the entire pressure range of the tests. For a strong favorable pressure gradient followed by any weak favorable, neutral, or adverse gradient, the tendency was for transition to occur at or immediately behind the shoulder. A single strip of three-dimensional roughness in the region of strong favorable pressure gradient did not fix transition on the models at the roughness location except at the maximum test pressures, whereas a second roughness strip added in a region of neutral or adverse pressure gradient did fix transition. Experimental pressure coefficients agreed closely with modified Newtonian theory except in the shoulder region.

Jackson, Mary W.↗
NASA NTRS1977

Prediction of in-depth gap heating ratios from wing glove model test data

In-depth gap heating ratios were predicted down RSI tile sidewalls based on temperature measurements obtained from the JSC arc-jet Wing Glove model tests in order to develop gap heating ratios which resulted in the best possible fit of test data and to produce a set of engineering verification heating ratios similar in shape to one another which could be used at various body points on the Orbiter during reentry. The Rockwell TPS Multidimensional heat conduction program was used to perform 3-D thermal analyses using a 3.0 in. thick section of a curved RSI tile with 283 nodal points. Correlation with test data shows that the predicted heating ratios were significantly higher down in the gap than the zero pressure values for T/C stacks 39 and 38 on the Wing Glove model. For stack 37 (in a low pressure region), the baseline heating ratio overpredicted the temperature data. This analysis, which showed that the heating ratios were a strong function of the product of pressure and pressure gradient, will be used to compare with recent Gap/Step and Ames Double Wedge test/analysis results in the effort to identify the Orbiter gap response to high delta P flight environment.

Source record↗
NASA NTRS1995

Coupled Ablation, Heat Conduction, Pyrolysis, Shape Change and Spallation of the Galileo Probe

The Galileo probe enters the atmosphere of Jupiter in December 1995. This paper presents numerical methodology and detailed results of our final pre-impact calculations for the heat shield response. The calculations are performed using a highly modified version of a viscous shock layer code with massive radiation coupled with a surface thermochemical ablation and spallation model and with the transient in-depth thermal response of the charring and ablating heat shield. The flowfield is quasi-steady along the trajectory, but the heat shield thermal response is dynamic. Each surface node of the VSL grid is coupled with a one-dimensional thermal response calculation. The thermal solver includes heat conduction, pyrolysis, and grid movement owing to surface recession. Initial conditions for the heat shield temperature and density were obtained from the high altitude rarefied-flow calculations of Haas and Milos. Galileo probe surface temperature, shape, mass flux, and element flux are all determined as functions of time along the trajectory with spallation varied parametrically. The calculations also estimate the in-depth density and temperature profiles for the heat shield. All this information is required to determine the time-dependent vehicle mass and drag coefficient which are necessary inputs for the atmospheric reconstruction experiment on board the probe.

Milos, Frank S.↗
NASA NTRS1989

Investigation of heat transfer in zirconium potassium perchlorate at low temperature: A study of the failure mechanism of the NASA standard initiator

The objective of this work was to study the reasons for the failure of pyrotechnic initiators at very low temperatures (10 to 100 K). A two-dimensional model of the NASA standard initiator was constructed to model heat transfer from the electrically heated stainless steel bridgewire to the zirconium potassium perchlorate explosive charge and the alumina charge cup. Temperature dependent properties were used in the model to simulate initiator performance over a wide range of initial temperatures (10 to 500 K). A search of the thermophysical property data base showed that pure alumina has a very high thermal conductivity at low temperatures. It had been assumed to act as a thermal insulator in all previous analyses. Rapid heat transfer from the bridgewire to the alumina at low initial temperatures was shown to cause failure of the initiators if the wire did not also make good contact with the zirconium potassium perchlorate charge. The mode is able to reproduce the results of the tests that had been conducted to investigate the cause for failure. It also provides an explanation for previously puzzling results and suggests simple design changes that will increase reliability at very low initial temperatures.

Varghese, Philip L.↗