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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 505 records · Page 28
All sourcesNASA NTRSDOE OSTI
NASA NTRS1956

The Compressible Laminar Boundary Layer with Heat Transfer and Arbitrary Pressure Gradient

An approximate method for the calculation of the compressible laminar boundary layer with heat transfer and arbitrary pressure gradient, based on Thwaites' correlation concept, is presented. With the definition of dimensionless shear and heat-transfer parameters and an assumed correlation of these parameters in terms of a momentum parameter, a complete system of relations for calculating skin friction and heat transfer results. Knowledge of velocity or temperature profiles is not necessary in using this calculation method. When the method is applied to a convergent-divergent, axially symmetric rocket nozzle, it shows that high rates of heat transfer are obtained at the initial stagnation point and at the throat of the nozzle. Also indicated are negative displacement thicknesses in the convergent portion of the nozzle; these occur because of the high density within the lower portions of the cooled boundary layer. (author)

Cohen, Clarence B↗
NASA NTRS1941

The problem of cooling an air-cooled cylinder on an aircraft engine

An analysis of the cooling problem has been to show by what means the cooling of an air-cooled aircraft engine may be improved. Each means of improving cooling is analyzed on the basis of effectiveness in cooling with respect to power for cooling. The altitude problem is analyzed for both supercharged and unsupercharged engines. The case of ground cooling is also discussed. The heat-transfer process from the hot gases to the cylinder wall is discussed on the basis of the fundamentals of heat transfer and thermodynamics. Adiabatic air-temperature rise at a stagnation point in compressible flow is shown to depend only on the velocity of flow.

Brevoort, M J↗
NASA NTRS1953

Method for Calculation of Laminar Heat Transfer in Air Flow Around Cylinders of Arbitrary Cross Section (including Large Temperature Differences and Transpiration Cooling)

The solution of heat-transfer problems has become vital for many aeronautical applications. The shapes of objects to be cooled can often be approximated by cylinders of various cross sections with flow normal to the axis as, for instance heat transfer on gas-turbine blades and on air foils heated for deicing purposes. A laminar region always exists near the stagnation point of such objects. A method previously presented by E. R. G. Eckert permits the calculation of local heat transfer around the periphery of cylinders of arbitrary cross section in the laminar region for flow of a fluid with constant property values with an accuracy sufficient for engineering purposes. The method is based on exact solutions of the boundary-layer equations for incompressible wedge-type flow and on the postulate that at any point on the cylinder the boundary-layer growth is the same as that on a wedge with comparable flow conditions. This method is extended herein to take into account the influence of large temperature differences between the cylinder wall and the flow as well as the influence of transpiration cooling when the same medium as the outside flow is used as coolant.

Eckert, E R↗
NASA NTRS1956

Second-order subsonic airfoil theory including edge effects

Several recent advances in plane subsonic flow theory are combined into a unified second-order theory for airfoil sections of arbitrary shape. The solution is reached in three steps: the incompressible result is found by integration, it is converted into the corresponding subsonic compressible result by means of the second-order compressibility rule, and it is rendered uniformly valid near stagnation points by further rules. Solutions for a number of airfoils are given and are compared with the results of other theories and of experiment. A straight-forward computing scheme is outlined for calculating the surface velocities and pressures on any airfoil at any angle of attack

Van Dyke, Milton D↗
NASA NTRS1957

Investigation of the Laminar Aerodynamic Heat-transfer Characteristics of a Hemisphere-cylinder in the Langley 11-inch Hypersonic Tunnel at a Mach Number of 6.8

A program to investigate the aerodynamic heat transfer of a nonisothermal hemisphere-cylinder has been conducted in the Langley 11-inch hypersonic tunnel at a Mach number of 6.8 and a Reynolds number from approximately 0.14 x 10(6) to 1.06 x 10(6) based on diameter and free-stream conditions. The experimental heat-transfer coefficients were slightly less over the whole body than those predicted by the theory of Stine and Wanlass (NACA technical note 3344) for an isothermal surface. For stations within 45 degrees of the stagnation point the heat-transfer coefficients could be correlated by a single relation between local Stanton number and local Reynolds number. Pitot pressure profiles taken at a Mach number of 6.8 on a hemisphere-cylinder have verified that the local Mach number or velocity outside the boundary layer required in the theories may be computed from the surface pressures by using isentropic flow relations and conditions immediately behind a normal shock. The experimental pressure distribution at Mach number of 6.8 is closely predicted by the modified Newtonian theory.

Crawford, Davis H↗
NASA NTRS1948

Surface-pressure Distributions on a Systematic Group of NACA 1-series Cowlings with and Without Spinners

A method for calculating the flow fields of axially symmetric bodies from their pressure distributions is reported in NACA RM No. L8I17. In order to facilitate application of this method to the important case of the cowling-spinner combination, for use in the design of propellers, the present paper presents static-pressure distributions on the tops of 79 high-critical-speed NACA 1-series cowling-spinner combinations over wide ranges of inlet-velocity ratio at angles of attack of 0 degrees, 2 degrees, 4 degrees, and 6 degrees. Static-pressure distributions around the nose sections of several cowlings are given in greater detail to aid in estimating the pressures near the stagnation points and to show the effect of changes in the internal lip shape. The effects of the operation of a typical propeller on the surface pressures on the cowling are shown for one configuration. The pressure distributions over the nine NACA 1-series nose inlets used as the basic components of these combinations are also presented ro supplement the existing open-nose-cowling data of NACA ACR No. L5F30a which are applicable to the case of the rotating cowling.

Boswinkle, Robert W JR↗
NASA NTRS1994

Measured and calculated spectral radiation from a blunt body shock layer in an arc-jet wind tunnel

Spectra of the shock layer radiation incident on the stagnation point of a blunt body placed in an arc-jet wind tunnel were measured over the wavelength range from 600 nm to 880 nm. The test gas was a mixture of 80 percent air and 20 percent argon by mass, and the run was made in a highly nonequilibrium environment. The observed spectra contained contributions from atomic lines of nitrogen, oxygen, and argon, of bound-free and free-free continua, and band systems of N2 and N2(+). The measured spectra were compared with the synthetic spectra, which were obtained through four steps: the calculation of the arc-heater characteristics, of the nozzle flow, of the blunt-body flow, and the nonequilibrium radiation processes. The results show that the atomic lines are predicted approximately correctly, but all other sources are underpredicted by orders of magnitude. A possible explanation for the discrepancy is presented.

Babikian, Dikran S.↗
NASA NTRS1994

Turbulence modeling for separated flow

Two projects are described in this report. The first involves assessing turbulence models in separated flow. The second addresses the anomalous behavior of certain turbulence models in stagnation point flow. The primary motivation for developing turbulent transport models is to provide tools for computing non-equilibrium, or complex, turbulent flows. Simple flows can be analyzed using data correlations or algebraic eddy viscosities, but in more complicated flows such as a massively separated boundary layer, a more elaborate level of modeling is required. It is widely believed that at least a two-equation transport model is required in such cases. The transport equations determine the evolution of suitable velocity and time-scales of the turbulence. The present study included assessment of second-moment closures in several separated flows, including sharp edge separation; smooth wall, pressure driven separation; and unsteady vortex shedding. Flows with mean swirl are of interest for their role in enhancing mixing both by turbulent and mean motion. The swirl can have a stabilizing effect on the turbulence. An axi-symmetric extension to the INS-2D computer program was written adding the capability of computing swirling flow. High swirl can produce vortex breakdown on the centerline of the jet and it occurs in various combustors.

Durbin, Paul A.↗
NASA NTRS1994

Review of chemical-kinetic problems of future NASA missions, II: Mars entries

The present work aims to derive a set of thermomechanical relaxation rate parameters and chemical reaction rate coefficients relevant to future interplanetary missions. It also attempts to assess the impact of thermochemical nonequilibrium phenomena on radiative heating rates for the stagnation point of the Martian entry vehicle.

TRT-EXPERIMENTAL↗
NASA NTRS1994

Aerothermodynamic methods for a Mars environmental survey Mars entry

Computational fluid dynamics models for the thermodynamics and transport properties used in an equilibrium version of the Langley aerothermodynamics upwind relaxation algorithm (LAURA) for Mars atmospheric entries are described. In addition, the physical models used in a nonequilibrium version of LAURA for Mars-entry flows are described. Uncertainties exist in defining constants used in the transport properties for the equilibrium model and in many of the physical models for the nonequilibrium version. Solutions from the two codes using the best available constants are examined at the Mars-entry conditions characteristics of the Mars environmental survey mission. While the flowfields are near thermal equilibrium, chemical nonequilibrium effects are present in the entry cases examined. Convective heating at the stagnation point for these flows (assuming fully catalytic wall boundary conditions) is approximately 100 W/cm(exp 2). Radiative heating is negligible.

AEROTHERMODYNAMIC METHODS↗
NASA NTRS1993

Interactions between gravity waves and cold air outflows in a stably stratified uniform flow

Interactions between gravity waves and cold air outflows in a stably stratified uniform flow forced by various combinations of prescribed heat sinks and sources are studied using a hydrostatic two-dimensional nonlinear numerical model. The formation time for the development of a stagnation point or reversed flow at the surface is not always directly proportional to the Froude number when wave reflections exist from upper levels. A density current is able to form by the wave-otuflow interaction, even though the Froude number is greater than a critical value. This is the result of the wave-outflow interaction shifting the flow response to a different location in the characteristic parameter space. A density current is able to form or be destroyed due to the wave-outflow interaction between a traveling gravity wave and cold air outflow. This is proved by performing experiments with a steady-state heat sink and an additional transient heat source. In a quiescent fluid, a region of cold air, convergence, and upward motion is formed after the collision between two outflows produced by two prescribed heat sinks. After the collision, the individual cold air outflows lose their own identity and merge into a single, stationary, cold air outflow region. Gravity waves tend to suppress this new stationary cold air outflow after the collision. The region of upward motion associated with the collision is confined to a very shallow layer. In a moving airstream, a density current produced by a heat sink may be suppressed or enhanced nonlinearly by an adjacent heat sink due to the wave-outflow interaction.

Lin, Yuh-Lang↗
NASA NTRS1995

Analytical studies of hypersonic viscous dissociated flows

This project primarily dealt with integral boundary-layer solution techniques that are directly applicable to the problem of determining aerodynamic heating rates of hypersonic vehicles like X-33 in the vicinity of stagnation points, windward centerlines, and swept-wing leading edges. The analyses include effects of finite-rate gas chemistry across the boundary layer and finite-rate catalysis of atom recombination at the surface. A new approach for combining the insight afforded by integral boundary-layer analysis with comprehensive (and expensive) computational fluid dynamic (CFD) flowfield solutions of the thin-layer Navier-Stokes equations was developed. The approach extracts CFD derived quantities at the wall and at the boundary layer edge for inclusion in a post-processing boundary-layer analysis. The post-processed data base allows a designer at a workstation to ask and answer the following questions: (1) How much does the heating change if one uses a thermal protection system (TPS) with different catalytic properties than was used in the original CFD solution? (2) How does the heating change when one moves the interface of two different TPS materials with different catalytic efficiencies for the purpose of reducing vehicle weight and expense? The answer to the second question is particularly critical, because abrupt changes from low catalytic efficiency to high catalytic efficiency can lead to localized increase in heating which exceeds the usually conservative estimate provided by a fully catalytic wall assumption. A secondary issue that was addressed involves the prediction of heating levels in the vicinity of sharp corners that are transverse to or aligned with the flow. An example of the first case is heating at the edge of the COMET reentry module. An example of the second case is heating along the side edge of a deflected body flap on an SSV. The difficulty of putting grids in the vicinity of such corners with continuously varying metric coefficients causes problems in CFD predictions. A preliminary theory for prediction that says the heating at the corner is X percent of the heating N boundary-layer thicknesses inboard was developed. This will prove useful to analytically evaluate the possible benefits of rounding the edges of these configurations and defining how much rounding is sufficient.

Inger, George R.↗
NASA NTRS1994

Viscous Shock-Layer Analysis of Two-Dimensional and Axisymmetric Flows

Results are obtained for cylindrical leading edges of proposed transatmospheric vehicles by employing a two-dimensional viscous shock-layer code for nonequilibrium gas flows. The accuracy and efficiency of the planar code is verified through detailed comparisons with other predictions. This study includes results for 6-deg half-angle bodies with nose radii ranging from 0.01 to 2.0 ft for both cylindrically blunted wedges and spherically blunted cones (included for comparison). Some results are presented as a ratio of the noncatalytic to the corresponding fully catalytic heating value to illustrate the maximum potential for a heating reduction in dissociated nonequilibrium flows. Generally, this ratio and the individual heating rates are smaller for cylindrically blunted wedges with small nose radii as compared to the spherically blunted cones (for the same nose radius). Therefore, a larger potential exists for heating reduction in cylindrically blunted as compared with the spherically blunted surfaces. However, the results presented at higher altitudes (where the slip effects become important) show that the spherically, blunted nose gives lower stagnation-point heating due to stronger merged shock-layer effects as compared with a cylindrically blunted nose.

Gupta, Roop N.↗
NASA NTRS1996

Blunt body near wake flow field at Mach 6

Tests were conducted in a Mach 6 flow to examine the reattachment process of an axisymmetric free shear layer associated with the near wake of a 70 deg. half angle, spherically blunted cone with a cylindrical after body. Model angle of incidence was fixed at 0 deg. and free-stream Reynolds numbers based on body diameter ranged from 0.5 x 10(exp 6) to 4 x 10(exp 6). The sensitivity of wake shear layer transition on reattachment heating was investigated. The present perfect gas study was designed to compliment results obtained previously in facilities capable of producing real gas effects. The instrumented blunted cone model was designed primarily for testing in high enthalpy hypervelocity shock tunnels in both this country and abroad but was amenable for testing in conventional hypersonic blowdown wind tunnels as well. Surface heating rates were inferred from temperature - time histories from coaxial surface thermocouples on the model forebody and thin film resistance gages along the model base and cylindrical after body. General flow feature (bow shock, wake shear layer, and recompression shock) locations were visually identified by schlieren photography. Mean shear layer position and growth were determined from intrusive pitot pressure surveys. In addition, wake surveys with a constant temperature hot-wire anemometer were utilized to qualitatively characterize the state of the shear layer prior to reattachment. Experimental results were compared to laminar perfect gas predictions provided by a 3-D Navier Stokes code (NSHYP). Shear layer impingement on the instrumented cylindrical after body resulted in a localized heating maximum that was 21 to 29 percent of the forebody stagnation point heating. Peak heating resulting from the reattaching shear layer was found to be a factor of 2 higher than laminar predictions, which suggested a transitional shear layer. Schlieren flow visualization and fluctuating voltage time histories and spectra from the hot wire surveys across the shear layer substantiate this observation. The sensitivity of surface heating to forebody roughness was characterized for a reattaching shear layer. For example, at R(sub infinity), d = 4 x 10(exp 6), when the shear layer was transitional, the magnitude of peak heating from shear layer impingement was reduced by approximately 24 percent when transition grit was applied to the forebody. The spatial location of the local peak, however, remained unchanged.

Horvath, Thomas J.↗
NASA NTRS1996

Analytic corrections to CFD heating predictions accounting for changes in surface catalysis

Integral boundary-layer solution techniques applicable to the problem of determining aerodynamic heating rates of hypersonic vehicles in the vicinity of stagnation points and windward centerlines are briefly summarized. A new approach for combining the insight afforded by integral boundary-layer analysis with comprehensive (but time intensive) computational fluid dynamic (CFD) flowfield solutions of the thin-layer Navier-Stokes equations is described. The approach extracts CFD derived quantities at the wall and at the boundary layer edge for inclusion in a post-processing boundary-layer analysis. It allows a designer at a workstation to address two questions, given a single CFD solution. (1) How much does the heating change for a thermal protection system with different catalytic properties than was used in the original CFD solution? (2) How does the heating change at the interface of two different TPS materials with an abrupt change in catalytic efficiency? The answer to the second question is particularly important, because abrupt changes from low to high catalytic efficiency can lead to localized increase in heating which exceeds the usually conservative estimate provided by a fully catalytic wall assumption.

Gnoffo, Peter A.↗
NASA NTRS1996

DSMC simulations of OREX entry conditions

Results of direct simulation Monte Carlo (DSMC) solutions are presented for the Japanese Orbital Reentry Experiment (OREX) vehicle, a 50 deg half-angle spherically blunted cone with a nose radius of 1.35 m and a base diameter of 3.4 m. The flow conditions simulated are those for entry into the Earth's atmosphere at a nominal velocity of about 7.4 km/s and zero incidence. Calculations are made for the higher altitude portion of entry, encompassing the transitional flow regime (altitudes of 200 to 80 km). Comparisons with flight measured values are made for axial acceleration, surface pressure, and stagnation point heating.

Moss, James N.↗
NASA NTRS1995

Theory of the Effects of Small Gravitational Levels on Droplet Gasification

A mathematical model taking into account small (and constant) gravitational levels is developed for vaporization of an isolated liquid droplet suspended in a stagnant atmosphere. A goal of the present analysis is to see how small gravitational levels affect droplet gasification characteristics. Attention is focused upon determining the effects on gas-phase phenomena. The conservation equations arc normalized and nondimensionalized, and a small parameter that accounts for the effects of gravity is identified. This parameter is the square of the inverse of a Froude number based on the gravitational acceleration, the droplet radius, and a characteristic gas-phase velocity at the droplet surface. Asymptotic analyses are developed in terms of this parameter. In the analyses, different spatial regions are identified. Near a droplet, gravitational effects are negligible in the first approximation, and the flowfield is spherically symmetric to the leading order. Analysis shows, however, that outer zones exist where gravitational effects cannot be neglected; it is expected that a stagnation point will be present in an outer zone that is not present when gravity is totally absent. The leading order and higher-order differential equations for each zone are derived and solved. The solutions allow the effects of gravity on vaporization rates and temperature, velocity and species fields to be determined.

Beitelmal, A.↗
NASA NTRS1996

From Chondrules to Planetesimals: Solids and Turbulence in the Solar Nebula

The usual theory of planetesimal formation is untenable because turbulence inhibits gravitational instability. However, turbulence can actually concentrate chondrule-sized particles by factors up to a million near stagnation points. The implications for accretion may be profound.

Dobrovolskis, Anthony R.↗