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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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NASA NTRS1987

Effect of flow oscillations on axial energy transport in a porous material

The effects of flow oscillations on axial energy diffusion in a porous medium, in which the flow is continuously disrupted by the irregularities of the porous structure, are analyzed. The formulation employs an internal heat transfer coefficient that couples the fluid and solid temperatures. The final relationship shows that the axial energy transport per unit cross-sectional area and time is directly proportional to the axial temperature gradient and the square of the maximum fluid displacement.

Siegel, R.↗

NASA NTRS1993

Analysis and design of an ultrahigh temperature hydrogen-fueled MHD generator

A coupled gas dynamics/radiative heat transfer analysis of partially ionized hydrogen, in local thermodynamic equilibrium, flowing through an ultrahigh temperature (10,000-20,000 K) magnetohydrodynamic (MHD) generator is performed. Gas dynamics are modeled by a set of quasi-one-dimensional, nonlinear differential equations which account for friction, convective and radiative heat transfer, and the interaction between the ionized gas and applied magnetic field. Radiative heat transfer is modeled using nongray, absorbing-emitting 2D and 3D P-1 approximations which permit an arbitrary variation of the spectral absorption coefficient with frequency. Gas dynamics and radiative heat transfer are coupled through the energy equation and through the temperature- and density-dependent absorption coefficient. The resulting nonlinear elliptic problem is solved by iterative methods. Design of such MHD generators as onboard, open-cycle, electric power supplies for a particular advanced airbreathing propulsion concept produced an efficient and compact 128-MWe generator characterized by an extraction ratio of 35.5 percent, a power density of 10,500 MWe/cu m, and a specific (extracted) energy of 324 MJe/kg of hydrogen. The maximum wall heat flux and total wall heat load were 453 MW/sq m and 62 MW, respectively.

Moder, Jeffrey P.↗

NASA NTRS1994

A Method for Thermal Analysis of Spiral Bevel Gears

A modeling method for analyzing the three-dimensional thermal behavior of spiral bevel gears has been developed. The model surfaces are generated through application of differential geometry to the manufacturing process for face-milled spiral bevel gears. Contact on the gear surface is found by combining tooth contact analysis with three-dimensional Hertzian theory. The tooth contact analysis provides the principle curvatures and orientations of the two surfaces. This information is then used directly in the Hertzian analysis to find the contact size and maximum pressure. Heat generation during meshing is determined as a function of the applied load, sliding velocity, and coefficient of friction. Each of these factors change as the point of contact changes during meshing. A nonlinear finite element program was used to conduct the heat transfer analysis. This program permitted the time- and position-varying boundary conditions, found in operation, to be applied to a one-tooth model. An example model and analytical results are presented.

Handschuh, Robert F.↗

NASA NTRSDate not supplied

Influence of Airfoil Angle of Attack on Ice Accretion Roughness

The influence of airfoil angle attack on roughness evolution and spatial variation was investigated in the Icing Research Tunnel (IRT) at the NASA Glenn Research Center. Two airfoil models were used for the study: a 53.34-cm (21-in.) NACA 0012 model and a 152.4-cm (60-in.) HAARP-II model. For the NACA 0012, the angle of attack was varied from 0° to 3°, and the ice accretion roughness was characterized while keeping the other icing parameters such as freestream velocity, freestream stagnation temperature, accumulation parameter, and median volumetric diameter (MVD) constant. For the HAARP-II, the angle of attack was varied from -1.9°, which is the non-lifting angle of attack for the airfoil, to 3°. A series of accumulation time progression cases for the HAARP-II at 2.3° angle of attack was also performed in Appendix C and Appendix O conditions and compared to the measurements presented by McClain et al. (2018) for the non-lifting angle of attack condition. The results demonstrate the strong influence of the changing local static pressure along the surface influencing the roughness characteristics. A correlation approach for the maximum equivalent sand-grain roughness along the surface of an airfoil is presented and considers the scaled accumulation time, the stagnation point freezing fraction, and the surface pressure relative to the static and stagnation pressures of the flow. Finally, an attempt to relate the maximum roughness quantities to local pressure coefficients and local icing quantities is presented and discussed.

Icing↗

NASA NTRS1951

Analysis of turbulent free-convection boundary layer on flat plate

With the use of Karman's integrated momentum equation for the boundary layer and data on the wall-shearing stress and heat transfer in forced-convection flow, a calculation was carried out for the flow and heat transfer in the turbulent free-convection boundary layer on a vertical flat plate. The calculation is for a fluid with a Prandtl number that is close to 1. A formula was derived for the heat-transfer coefficient that was in good agreement with experimental data in the range of Grashof numbers from 10sup10 to 10sup12. Because of the good agreement between the theoretical formula and the experimental data, the formula may be used to obtain data for high Grashof numbers. The calculation also yielded formulas for the maximum velocity in the boundary layer and for boundary-layer thickness.

Eckert, E R G↗

NASA NTRS1956

Investigation of Heat Transfer From

The convective heat transfer from the surface of an ellipsoidal forebody of fineness ratio 3 and 20-inch maximum diameter was investigated in clear air for both stationary and rotating operation over a range of conditions including air speeds up to 240 knots, rotational speeds up to 1200 rpm, and angles of attack of 0 deg, 3 deg, and 6 deg. The results are presented in the form of heat-transfer coefficients and the correlation of Nusselt and Reynolds numbers. Both a uniform surface temperature and a uniform input heater density distribution were used. The experimental results agree well with theoretical predictions for uniform surface temperature distribution. Complete agreement was not obtained with uniform input heat density in the laminar-flow region because of conduction effects. No significant effects of rotation were obtained over the range of airstream and rotational speeds investigated. Operation at angle of attack had only minor effects on the local heat transfer. Transition from laminar to turbulent heat transfer occurred over a wide range of Reynolds numbers. The location of transition depended primarily on surface roughness and pressure and temperature gradients. Limited transient heating data indicate that the variation of surface temperature with time followed closely an exponential relation.

Lewis, James P.↗

NASA NTRS1984

Multiple-pass trajectories for an aeroassisted orbital transfer vehicle

This paper investigates the use of multiple-pass trajectories for aeroassisted orbital transfer vehicles (AOTV) as a means of reducing the severe aeroheating environment and the extreme sensitivity to off-nominal trajectory and atmospheric conditions. One-, two-, and three-pass trajectories were calculated for AOTV's with ballistic coefficients ranging from 5 to 150 psf. Compared to the single-pass case, the maximum heat rate was reduced by 30 percent for two passes and 45 percent for three passes, while the maximum acceleration was reduced by 40 percent and 55 percent, respectively. The sensitivity of the trajectories to variations in atmospheric and orbital parameters was not significantly reduced.

Rehder, J. J.↗

NASA NTRS1941

The design of fins for air-cooled cylinders

An analysis was made to determine the proportions of fins made of aluminum, copper, magnesium, and steel necessary to dissipate maximum quantities of heat for different fin widths, fin weights, and air-flow conditions. The analysis also concerns the determination of the optimum fin proportions when specified limits are placed on the fin dimensions. The calculation of the heat flow in the fins is based on experimentally verified, theoretical equations. The surface heat-transfer coefficients used with this equation were taken from previously reported experiments. In addition to the presentation of fin-design information, this investigation shows that optimum fin dimensions are inappreciably affected by the differences in air flow that are obtained with different air-flow arrangements or by small changes in the length of the air-flow path.

Biermann, Arnold E↗

NASA NTRS1985

Effects of cooling system parameters on heat transfer in PAFC stack

Analytical and experimental study for the effects of cooling system parameters on the heat transfer and temperature distribution in the electrode plates of a phosphoric acid fuel-cell has been conducted. An experimental set-up that simulates the operating conditions prevailing in a phosphoric-acid fuel-cell stack was designed and constructed. The set-up was then used to measure the overall heat transfer coefficient, the thermal contact resistance, and the electrode temperature distribution for two different cooling plate configurations. Two types of cooling plate configurations, serpentine and straight, were tested. Air, water, and oil were used as coolants. Measurements for the heat transfer coefficient and the thermal contact resistance were made for various flow rates ranging from 16 to 88 Kg/hr, and stack clamping pressure ranging from O to 3448 Kpa. The experimental results for the overall heat transfer coefficient were utilized to derive mathematical relations for the overall heat transfer coefficient as a function of stack clamping pressure and Reynolds number for the three coolants. The empirically derived formulas were incorporated in a previously developed computer program to predict electrodes temperature distribution and the performance of the stack cooling system. The results obtained were then compared with those available in the literature. The comparison showed maximum deviation of +/- 11%.

Abdul-Aziz, Ali A.↗

NASA NTRS1978

A numerical simulation of seasonal stratospheric climate. I - Zonal temperatures and winds

Two numerical experiments were designed to isolate the effects of seasonal and latitudinal variations in longwave radiative transfer processes. The first experiment aimed at computing the rate of temperature change in the stratosphere due to longwave radiative transfer by CO2, H2O and O3 from Ramanathan's (1976) radiative-convective model applied to the troposphere-stratosphere circulation. The second experiment employs the Newtonian cooling approximation in which the rate of temperature change in the stratosphere by longwave radiative transfer is set equal to the product of the Newtonian cooling coefficient 'h' and the departure of the local temperature from a reference temperature. It is shown that the latitudinal temperature distribution of the lower stratosphere is maintained by the combined effects of dynamics, O3 solar heating and the longwave radiative coupling between troposphere and stratosphere. The latitudinal gradient in the troposphere-stratosphere longwave radiative coupling is maximum during winter and spring. The radiative response time (1/h) of the middle and upper stratosphere undergoes significant latitudinal and seasonal variations, largely due to the temperature dependence of h.

Ramanathan, V.↗

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 NTRS1990

Thermal/structural analyses of several hydrogen-cooled leading-edge concepts for hypersonic flight vehicles

The aerodynamic heating at high flight Mach numbers, when shock interference heating is included, can be extremely high and can exceed the capability of most conventional metallic and potential ceramic materials available. Numerical analyses of the heat transfer and thermal stresses are performed on three actively cooled leading-edge geometries (models) made of three different materials to address the issue of survivability in a hostile environment. These analyses show a mixture of results from one configuration to the next. Results for each configuration are presented and discussed. Combinations of enhanced internal film coefficients and high material thermal conductivity of copper and tungsten are predicted to maintain the maximum wall temperature for each concept within acceptable operating limits. The exception is the TD nickel material which is predicted to melt for most cases. The wide range of internal impingement film coefficients (based on correlations) for these conditions can lead to a significant uncertainty in expected leading-edge wall temperatures. The equivalent plastic strain, inherent in each configuration which results from the high thermal gradients, indicates a need for further cyclic analysis to determine component life.

Gladden, Herbert J.↗

NASA NTRS1990

Thermal/structural analyses of several hydrogen-cooled leading-edge concepts for hypersonic flight vehicles

Gladden, Herbert J.↗

NASA NTRS1992

Variable Property Flow in Rectangular Ducts with Repeated Rectangular Rib Roughness

PHOENICS was used to simulate turbulent flow of supercritical hydrogen through asymmetrically heated rectangular ducts with one side rib-roughened, in order to provide design data for active cooling systems of hypersonic aircraft where the hydrogen fuel is used as coolant. Variable property and conjugate heat transfer effects were considered. The computations using PHOENICS required provision of property sub-routines for supercritical hydrogen, a low-Reynolds number form of the k-epsilon turbulence model, a rough-wall wall function and a procedure for calculating proper diffusion coefficients at the fluid-solid interface. The effect of each design variable such as width and height of duct and wall thickness, on pressure drop and maximum wall temperature was investigated. PHOENICS version 1.4 was used on an IBM 3090 computer.

Youn, B.↗

NASA NTRS1984

Measurements of local convective heat transfer coefficients on ice accretion shapes

The thin-skin heat rate technique was used to determine local convective heat transfer coefficients for four representative ice accretion shapes. The shapes represented three stages of glaze ice formation and one rime ice formation; the ice models had varying degrees of surface roughness. In general, convective heat transfer was higher in regions where the model's surfaces were convex and lower in regions where the surfaces were concave. The effect of roughness was different for the glaze and rime ice shapes. On the glaze ice shapes, roughness increased the maximum Nu by 80 percent, but the other Nu values were virtually unchanged. On the rime ice shape, the Nu numbers near the stagnation point were unchanged. The maximum Nu value increased by 45 percent, and the Nu number downstream of the peak increased by approximately 150 percent.

Arimilli, R. V.↗

NASA NTRS1997

Experimental and Computational Studies of the Flow Over a Sting Mounted Planetary Probe Configuration

This paper summarizes the results of a series of experimental studies in the LENS shock tunnel and computations with DSMC and Navier Stokes codes which have been made to examine the aerothermal and flowfield characteristics of the flow over a sting-supported planetary probe configuration in hypervelocity air and nitrogen flows. The experimental program was conducted in the LENS hypervelocity shock tunnel at total enthalpies of 5and 10 MJkg for a range of reservoir pressure conditions from 70 to 500 bars. Heat transfer and pressure measurements were made on the front and rear face of the probe and along the supporting sting. High-speed and single shot schlieren photography were also employed to examine the flow over the model and the time to establish the flow in the base recirculation region. Predictions of the flowfield characteristics and the distributions of heat transfer and pressure were made with DSMC codes for rarefied flow conditions and with the Navier-Stokes solvers for the higher pressure conditions where the flows were assumed to be laminar. Analysis of the time history records from the heat transfer and pressure instrumentation on the face of the probe and in the base region indicated that the base flow was fully established in under 4 milliseconds from flow initiation or between 35 and 50 flow lengths based on base height. The measurements made in three different tunnel entries with two models of identical geometries but with different instrumentation packages, one prepared by NASA Langley and the second prepared by CUBRC, demonstrated good agreement between heat transfer measurements made with two different types of thin film and coaxial gage instrumentation. The measurements of heat transfer and pressure to the front face of the probe were in good agreement with theoretical predictions from both the DSMC and Navier Stokes codes. For the measurements made in low density flows, computations with the DSMC code were found to compare well with the pressure and heat transfer measurements on the sting, although the computed heat transfer rates in the recirculation region did not exhibit the same characteristics as the measurements. For the 10MJkg and 500 bar reservoir match point condition, the measurements and heat transfer along the sting from the first group of studies were in agreement with the Navier Stokes solutions for laminar conditions. A similar set of measurements made in later tests where the model was moved to a slightly different position in the test section indicated that the boundary layer in the reattachment compression region was close to transition or transitional where small changes in the test environment can result in larger than laminar heating rates. The maximum heating coefficients on the sting observed in the present studies was a small fraction of similar measurements obtained at nominally the same conditions in the HEG shock tunnel, where it is possible for transition to occur in the base flow, and in the low enthalpy studies conducted in the NASA Langley high Reynolds number Mach 10 tunnel where the base flow was shown to be turbulent. While the hybrid Navier- StokedDMSC calculations by Gochberg et al. (Reference 1) suggested that employing the Navier- Stokes calculations for the entire flowfield could be seriously in error in the base region for the 10 MJkg, 500 bar test case, similar calculations performed by Cornell, presented here, do not.

Holden, Michael S.↗

NASA NTRS1947

Temperatures and Stresses on Hollow Blades For Gas Turbines

The present treatise reports on theoretical investigations and test-stand measurements which were carried out in the BMW Flugmotoren GMbH in developing the hollow blade for exhaust gas turbines. As an introduction the temperature variation and the stress on a turbine blade for a gas temperature of 900 degrees and circumferential velocities of 600 meters per second are discussed. The assumptions onthe heat transfer coefficients at the blade profile are supported by tests on an electrically heated blade model. The temperature distribution in the cross section of a blade Is thoroughly investigated and the temperature field determined for a special case. A method for calculation of the thermal stresses in turbine blades for a given temperature distribution is indicated. The effect of the heat radiation on the blade temperature also is dealt with. Test-stand experiments on turbine blades are evaluated, particularly with respect to temperature distribution in the cross section; maximum and minimum temperature in the cross section are ascertained. Finally, the application of the hollow blade for a stationary gas turbine is investigated. Starting from a setup for 550 C gas temperature the improvement of the thermal efficiency and the fuel consumption are considered as well as the increase of the useful power by use of high temperatures. The power required for blade cooling is taken into account.

Pollmann, Erich↗

NASA NTRS1995

Comparisons of the Maxwell and CLL gas/surface interaction models using DSMC

The behavior of two different models of gas-surface interactions is studied using the Direct Simulation Monte Carlo (DSMC) method. The DSMC calculations examine differences in predictions of aerodynamic forces and heat transfer between the Maxwell and the Cercignani-Lampis-Lord (CLL) models for flat plate configurations at freestream conditions corresponding to a 140 km orbit around Venus. The size of the flat plate represents one of the solar panels on the Magellan spacecraft, and the freestream conditions correspond to those experienced during aerobraking maneuvers. Results are presented for both a single flat plate and a two-plate configuration as a function of angle of attack and gas-surface accommodation coefficients. The two-plate system is not representative of the Magellan geometry but is studied to explore possible experiments that might be used to differentiate between the two gas-surface interaction models. The Maxwell and CLL models produce qualitatively similar results for the aerodynamic forces and heat transfer on a single flat plate. However, the flow fields produced with the two models are qualitatively different for both the single-plate and two-plate calculations. These differences in the flowfield lead to predictions of the angle of attack for maximum heat transfer in a two plate configuration that are distinctly different for the two gas-surface interactions models.

Hedahl, Marc O.↗