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At least 163 records · Page 9

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NASA NTRS1972

Analytical investigation of two-phase heat transfer in a porous matrix.

One-dimensional two-phase flow transpiration cooling through porous metals is studied analytically. The differential energy equations for the solid and for the liquid with temperature dependent properties are solved simultaneously with a semi-empirical pressure drop correlation. Formulation of the problem both assuming thermal equilibrium (assumes Clausius-Claperyon equation pertains) and nonequilibrium (defines a parameter representing fraction of energy transferred to mixture which contributes to evaporation) is given. Representative solutions showing the influences and significance of the controlling dimensionless parameters are presented.

Frost, W.↗

NASA NTRS1972

Experimental investigation of two-phase heat transfer in a porous matrix.

One-dimensional two-phase flow transpiration cooling through porous metal is studied experimentally. The experimental data is compared with a previous one-dimensional analysis. Good agreement with calculated temperature distribution is obtained as long as the basic assumptions of the analytical model are satisfied. Deviations from the basic assumptions are caused by nonhomogeneous and oscillating flow conditions. Preliminary derivation of nondimensional parameters which characterize the stable and unstable flow conditions is given. Superheated liquid droplets observed sputtering from the heated surface indicated incomplete evaporation at heat fluxes well in access of the latent energy transport. A parameter is developed to account for the nonequilibrium thermodynamic effects. Measured and calculated pressure drops show contradicting trends which are attributed to capillary forces.

Von Reth, R.↗

NASA NTRS1973

Induction simulation of gas core nuclear engine

The design, construction and operation of an induction heated plasma device known as a combined principles simulator is discussed. This device incorporates the major design features of the gas core nuclear rocket engine such as solid feed, propellant seeding, propellant injection through the walls, and a transpiration cooled, choked flow nozzle. Both argon and nitrogen were used as propellant simulating material, and sodium was used for fuel simulating material. In addition, a number of experiments were conducted utilizing depleted uranium as the fuel. The test program revealed that satisfactory operation of this device can be accomplished over a range of operating conditions and provided additional data to confirm the validity of the gas core concept.

Poole, J. W.↗

NASA NTRS1972

Investigation of extendable nozzle concepts

Preliminary research and technology were performed toward the development of a transpiration cooled inflatable nozzle extension. The J-2 engine of the Saturn-Apollo launch vehicle was selected as a design application test bed to provide realistic operational parameters and a possible hot firing demonstration. The nozzle extension attached at the 27.5 to 1 area ratio of the basic J-2 nozzle exit and extended to an area ratio of 48 to 1. The final nozzle extension design as evolved during this effort utilized conically woven stainless steel airmat as the major construction material. Model and full scale nozzle extensions were fabricated and tested although an actual J-2 hot firing demonstration was not possible due to the lack of a proper altitude test opportunity.

Scott, C. N.↗

NASA NTRS1973

Cavity temperature and flow characteristics in a gas-core test reactor

A test reactor concept for conducting basic studies on a fissioning uranium plasma and for testing various gas-core reactor concepts is analyzed. The test reactor consists of a conventional fuel-element region surrounding a 61-cm-(2-ft-) diameter cavity region which contains the plasma experiment. The fuel elements provide the neutron flux for the cavity region. The design operating conditions include 60-MW reactor power, 2.7-MW cavity power, 200-atm cavity pressure, and an average uranium plasma temperature of 15,000 K. The analytical results are given for cavity radiant heat transfer, hydrogen transpiration cooling, and uranium wire or powder injection.

Putre, H. A.↗

NASA NTRS1973

Conformal mapping technique for two-dimensional porous media and jet impingement heat transfer

Transpiration cooling and liquid metals both provide highly effective heat transfer. Using Darcy's law in porous media, and the inviscid approximation for liquid metals, the local fluid velocity in these flows equals the gradient of a potential, The energy equation and flow region are simplified when transformed into potential plane coordinates. In these coordinates the present problems are reduced to heat conduction solutions which are mapped into the physical geometry. Results are obtained for a porous region with simultaneously prescribed surface temperature and heat flux, heat transfer in a two-dimensional porous bed, and heat transfer for two liquid metal slot jets impinging on a heated plate.

Siegel, R.↗

NASA NTRS1974

Method of electroforming a rocket chamber

A transpiration cooled rocket chamber is made by forming a porous metal wall on a suitably shaped mandrel. The porous wall may be made of sintered powdered metal, metal fibers sintered on the mandrel or wires woven onto the mandrel and then sintered to bond the interfaces of the wires. Intersecting annular and longitudinal ribs are then electroformed on the porous wall. An interchamber wall having orifices therein is then electroformed over the annular and longitudinal ribs. Parallel longitudinal ribs are then formed on the outside surface of the interchamber wall after which an annular jacket is electroformed over the parallel ribs to form distribution passages therewith. A feed manifold communicating with the distribution passages may be fabricated and welded to the rocket chamber or the feed manifold may be electroformed in place.

Fortini, A.↗

NASA NTRS1974

Conformal mapping technique for two-dimensional porous media and jet impingement heat transfer

Transpiration cooling and liquid metals both provide highly effective heat transfer. Using Darcy's law in porous media and the inviscid approximation for liquid metals, the local fluid velocity in these flows equals the gradient of a potential. The energy equation and flow region are simplified when transformed into potential plane coordinates. In these coordinates, the present problems are reduced to heat conduction solutions which are mapped into the physical geometry. Results are obtained for a porous region with simultaneously prescribed surface temperature and heat flux, heat transfer in a two-dimensional porous bed, and heat transfer for two liquid metal slot jets impinging on a heated plate.

Siegel, R.↗

NASA NTRS1976

Radiative behavior of a gas layer seeded with soot

Gaseous film or transpiration cooling may be used to reduce the heat flux reaching the wall of a container or other structures. Such a protective film, however, is usually not effective for reducing radiative heat transfer as most gases are transparent in the temperature range for which solid walls can exist. Therefore, heat transfer was examined for a gaseous layer seeded with radiation-absorbing carbon particles (soot) and flowing along a surface. The layer was subjected to an external high temperature source of blackbody radiation. The radiative behavior was found to depend on a parameter containing particle concentration, layer thickness and source temperature. Only a very small particle volume concentration, in the range of .0001, was required to obtain high absorption in a 1-cm-thick layer for typical conditions. The results provide the distance along the surface for which the heat transfer to the wall remains within an acceptable limit and the particles remain below a temperature at which they will melt or vaporize. The wall protection by the layer lasts only until the particles vaporize or the layer becomes so hot that it reradiates substantially to the wall. Depending on the layer mass velocity the protection may be effective for a distance along the wall of only a few layer thickness. Hence, to protect greater wall lengths, it will be necessary to introduce the suspension through multiple slots or holes along the wall.

Siegel, R.↗

NASA NTRS1984

Comparisons of rational engineering correlations of thermophoretically-augmented particle mass transfer with STAN5-predictions for developing boundary layers

Modification of the code STAN5 to properly include thermophoretic mass transport, and examination of selected test cases developing boundary layers which include variable properties, viscous dissipation, transition to turbulence and transpiration cooling. Under conditions representative of current and projected GT operation, local application of St(M)/St(M),o correlations evidently provides accurate and economical engineering design predictions, especially for suspended particles characterized by Schmidt numbers outside of the heavy vapor range.

Gokoglu, S. A.↗

NASA NTRS1984

Materials for Advanced Turbine Engines (MATE): Project 3: Design, fabrication and evaluation of an oxide dispersion strengthened sheet alloy combustor liner, volume 1

The suitability of wrought oxide dispersion strengthened (ODS) superalloy sheet for gas turbine engine combustor applications was evaluated. Incoloy MA 956 (FeCrAl base) and Haynes Developmental Alloy (HDA) 8077 (NiCrAl base) were evaluated. Preliminary tests showed both alloys to be potentially viable combustor materials, with neither alloy exhibiting a significant advantage over the other. Both alloys demonstrated a +167C (300 F) advantage of creep and oxidation resistance with no improvement in thermal fatigue capability compared to a current generation combustor alloy (Hastelloy X). MA956 alloy was selected for further demonstration because it exhibited better manufacturing reproducibility than HDA8077. Additional property tests were conducted on MA956. To accommodate the limited thermal fatigue capability of ODS alloys, two segmented, mechanically attached, low strain ODS combustor design concepts having predicted fatigue lives or = 10,000 engine cycles were identified. One of these was a relatively conventional louvered geometry, while the other involved a transpiration cooled configuration. A series of 10,000 cycle combustor rig tests on subscale MA956 and Hastelloy X combustor components showed no cracking, thereby confirming the beneficial effect of the segmented design on thermal fatigue capability. These tests also confirmed the superior oxidation and thermal distortion resistance of the ODS alloy. A hybrid PW2037 inner burner liner containing MA956 and Hastelloy X components was designed and constructed.

Henricks, R. J.↗

NASA NTRS1984

Ceramic composite liner material for gas turbine combustors

The application of ceramics to gas turbine combustor liners to reduce liner metal temperature was studied in an experiment in which yttria-stabilized zirconia plasma was sprayed on compliant metal substrates exposed to near stoichiometric combustion. The strain isolation pad materials chosen were Hoskins Alloy 875 and BRUNSLLOY 534 Fiber Metal of 0.25 and 0.38 cm thicknesses and 35 and 45 percent density levels. Combustor screening tests of all specimens showed no evidence of deterioration or failure. Specimens exposed to flame temperatures in excess of 2100 K were convectively or convective-transpiration cooled and were evaluated in a 10 cm sq flame tube at inlet air temperature of 533 K and pressure of 0.5 MPa. The results suggest the superiority of a system composed of the Hoskins Alloy 875 compliant pad with 0.25 cm thickness and 35 percent density coupled with a NiCrAlY bond coat and a 8 percent Y2O3-ZrO2 ceramic top coat of 0.19 cm thickness.

Ercegovic, D. B.↗

NASA NTRS1984

High temperature performance of flexible thermal protection materials

Aero convective tests of several flexible thermal protection system (FTPS) concepts were conducted in the NASA Ames Research Center 20 MW arcjet aero heating wind tunnel. The concepts consisted of quilted insulation blankets with nextel AB312 felt insulation stitched between cover cloths with AB312 thread. The cover cloths were commercially available nextel AB312 and nicalon fabrics. The specimens were subjected to convective heat fluxes ranging from 7 to 35 Btu/per sq ft per sec at stagnation pressures of .005 to .02 atm. Specimens were tested both with and without transpiration cooling. Results indicated that both the nextel and nicalon fabrics offer the potential for higher temperature applications than current FTPS, and nicalon appears to be capable of withstanding temperatures well above 2500 degrees F with minimal degradation.

Savage, R. T.↗

NASA NTRS1984

Comparisons of rational engineering correlations of thermophoretically-augmented particle mass transfer with STAN5-predictions for developing boundary layers

Gokoglu, S. A.↗

NASA NTRS1988

Quick actuating closure and handling system

A quick activating closure and handling system, which utilizes conical sections for locking, was developed to allow quick access to the combustor internal components of the 8 ft High Temperature Tunnel. These critical components include the existing methane spraybar, a transpiration cooled nozzle and the new liquid oxygen (LOX) injection system housed within the combustor. A substantial cost savings will be realized once the mechanism is installed since it will substantially reduce the access time and increase the time available for conducting wind tunnel tests. A need exists for more frequent inspections when the wind tunnel operates at the more severe conditions generated by using LOX in the combustor. A loads analysis and a structural (finite element) analysis were conducted to verify that the new closure system is compatible with the existing pressure shell. In addition, strain gages were placed on the pressure vessel to verify how the pressure shell reacts to transient pressure loads. A scale model of the new closure system was built to verify the operation of the conical sections in the locking mechanisms.

Allred, Johnny W.↗

NASA NTRS1990

A unique high heat flux facility for testing hypersonic engine components

A major concern in advancing the state-of-the-art technologies for hypersonic vehicles is the development of an aeropropulsion system capable of withstanding high thermal loads expected during hypersonic flights. Consequently, there is a need for experimental facilities capable of providing a high heat flux environment for testing compound concepts and verifying analyses. A hydrogen/oxygen rocket engine was developed to provide a high enthalpy/high heat flux environment for component evaluation. This Hot Gas Facility is capable of providing heat fluxes ranging from 200 (on flat surfaces) up to 8000 Btu per sq ft per sec (at a leading edge stagnation point). Gas temperatures up to 5500 R can be attained as well as Reynolds numbers up to 360,000 per ft. Test articles such as cowl leading edges, transpiration-cooled seals, fuel injectors, and cooled panel concepts can be evaluated with gaseous hydrogen as coolant. This facility and its configuration and test capabilities are discussed. Results from flow characterization experiments are also shown and their implications considered.

Melis, Matthew E.↗

NASA NTRS1992

Ram accelerator direct space launch system - New concepts

The ram accelerator, a chemically driven ramjet-in-tube device is a new option for direct launch of acceleration-insensitive payloads into earth orbit. The projectile is the centerbody of a ramjet and travels through a tube filled with a premixed fuel-oxidizer mixture. The tube acts as the cowl of the ramjet. A number of new concepts for a ram accelerator space launch system are presented. The velocity and acceleration capabilities of a number of ram accelerator drive modes, including several new modes, are given. Passive (fin) stabilization during atmospheric transit is investigated and found to be promising. Gasdynamic heating in-tube and during atmospheric transit is studied; the former is found to be severe, but may be alleviated by the selection of the most suitable drive modes, transpiration cooling, or a hydrogen gas core in the launch tube. To place the payload in earth orbit, scenarios using one impulse and three impulses (with an aeropass) and a new scenario involving an auxiliary vehicle are studied. The auxiliary vehicle scenario is found to be competitive regarding payload, and requires a much simpler projectile, but has the disadvantage of requiring the auxiliary vehicle.

Bogdanoff, David W.↗

NASA NTRS1992

A database of aerothermal measurements in hypersonic flow for CFD validation

This paper presents an experimental database selected and compiled from aerothermal measurements obtained on basic model configurations on which fundamental flow phenomena could be most easily examined. The experimental studies were conducted in hypersonic flows in 48-inch, 96-inch, and 6-foot shock tunnels. A special computer program was constructed to provide easy access to the measurements in the database as well as the means to plot the measurements and compare them with imported data. The database contains tabulations of model configurations, freestream conditions, and measurements of heat transfer, pressure, and skin friction for each of the studies selected for inclusion. The first segment contains measurements in laminar flow emphasizing shock-wave boundary-layer interaction. In the second segment, measurements in transitional flows over flat plates and cones are given. The third segment comprises measurements in regions of shock-wave/turbulent-boundary-layer interactions. Studies of the effects of surface roughness of nosetips and conical afterbodies are presented in the fourth segment of the database. Detailed measurements in regions of shock/shock boundary layer interaction are contained in the fifth segment. Measurements in regions of wall jet and transpiration cooling are presented in the final two segments.

Holden, M. S.↗