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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 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1965

Low heat leak connector for cryogenic system

Heat leak from the surrounding atmosphere during fluid transfer from a spaced shell-insulated vessel for storing liquified gas having an upper gaseous phase, in minimized by forming a relatively wide, shallow blister on the wall of the vessel at the point of transfer line connection. The shell and the opposed walls of the blister have aligned openings whose common axis passes centrally through the blister and is normal to the surfaces of the vessel and shell. A fluid transfer line conduit passing through the shell opening is in fluid-tight connection with the shell and blister wall. The fluid transfer line confines the fluid in a continuous stream. The blister is filled with a heat insulating material which provides a thermal break between the central wall portions of the blister. A connector at the bottom of the vessel comprises a tube extending between the openings in the blister which projects a short distance within the body of liquefied gas and terminates in a reverse bend to prevent backflow of liquid through the pipe.

Stelts, P. D.↗
NASA NTRS1979

A nozzle boundary layer model including the subsonic sublayer usable for determining boundary layer effects on plume flowfields

A model for the boundary layer at the exit plane of a rocket nozzle was developed which, unlike most previous models, includes the subsonic sublayer. The equations for the flow near the nozzle exit plane are presented and the method by which the subsonic sublayer transitions to supersonic flow in the plume is described. The resulting model describes the entire boundary layer and can be used to provide a startline for method-of-characteristics calculations of plume flowfields. The model was incorporated into a method of characteristics computer program and comparisons of computed results to experimental data show good agreement. The data used in the comparisons were obtained in tests in which mass fluxes from a 22.2-N (5 lbf) thrust engine were measured at angles off the nozzle centerline of up to 150 deg. Additional comparisons were made with data obtained during tests of a 0.89-N (0.2 lbr) monopropellant thruster and from the OH-64 space shuttle heating tests. The agreement with the data indicates that the model can be used for calculating plume backflow properties.

Cooper, B. P., Jr.↗
NASA NTRS1980

Analytical and experimental study of mean flow and turbulence characteristics inside the passages of an axial flow inducer

The inviscid and viscid effects existing within the passages of a three bladed axial flow inducer operating at a flow coefficient of 0.065 are investigated. The blade static pressure and blade limiting streamline angle distributions were determined and the three components of mean velocity, turbulence intensities, and turbulence stresses were measured at locations inside the inducer blade passage utilizing a rotating three sensor hotwire probe. Applicable equations were derived for the hotwire data reduction analysis and solved numerically to obtain the appropriate flow parameters. The three dimensional inviscid flow in the inducer was predicted by numerically solving the exact equations of motion, and the three dimensional viscid flow was predicted by incorporating the dominant viscous terms into the exact equations. The analytical results are compared with the experimental measurements and design values where appropriate. Radial velocities are found to be of the same order as axial velocities within the inducer passage, confirming the highly three dimensional characteristic of inducer flow. Total relative velocity distribution indicate a substantial velocity deficiency near the tip at mid-passage which expands significantly as the flow proceeds toward the inducer trailing edge. High turbulence intensities and turbulence stresses are concentrated within this core region. Considerable wake diffusion occurs immediately downstream of the inducer trailing edge to decay this loss core. Evidence of boundary layer interactions, blade blockage effects, radially inward flows, annulus wall effects, and backflows are all found to exist within the long, narrow passages of the inducer.

Gorton, C. A.↗
NASA NTRS1980

Three-layer interactive method for computing supersonic laminar separated flows

An interactive model for numerical computation of complicated two-dimensional flowfields including regions of reversed flow is proposed. The present approach is one of dividing the flowfield into three regions, in each of which a simplified mathematical model is applied: (1) outer, supersonic flow for which the full potential equation (hyperbolic) is used; (2) viscous, laminar layer in which the compressible boundary-layer model (parabolic) is used; and (3) recirculating flow modeled by the incompressible Navier-Stokes equations (elliptic). For matching of the numerical solutions in the three layers, two interaction models are developed: one for pressure interaction, the other for interaction between the shear layer and the recirculating flow. The uniform solution for the whole flowfield is then obtained by iteration of the local solutions under the constraints imposed by matching. The three-layer interactive model is used for solution of the flowfield past an asymmetric cavity. The method is shown to be capable of dealing with backflow without encountering problems at separation, characteristic to the boundary-layer approach.

Brandeis, J.↗
NASA NTRS1983

Eliminating Wind Tunnel Flow Breakdown

Undesirable vortexes near floor in small wind tunnels suppressed by simple device that alters flow pattern there. Air is injected along floor and interacts with backflow from wind-tunnel model. Results in smoother, more correct air-flow and to more-reliable wind-tunnel data.

Hackett, J. E.↗
NASA NTRS1981

Electric thruster research

The multipole discharge chamber of an electrostatic ion thruster is discussed. No reductions in discharge losses were obtained, despite repeated demonstration of anode potentials more positive than the bulk of the discharge plasma. The penalty associated with biased anode operation was reduced as the magnetic integral above the biased anodes was increased. The hollow cathode is discussed. The experimental configuration of the Hall current thruster had a uniform field throughout the ion generation and acceleration regions. To obtain reliable ion generation, it was necessary to reduce the magnetic field strength, to the point where excessive electron backflow was required to establish ion acceleration. The theoretical study of ion acceleration with closed electron drift paths resulted in two classes of solutions. One class has the continuous potential variation in the acceleration region that is normally associated with a Hall current accelerator. The other class has an almost discontinuous potential step near the anode end of the acceleration region. This step includes a significant fraction of the total acceleration potential difference.

Kaufman, H. R.↗
NASA NTRS1986

Space Shuttle Main Engine fuel preburner augmented spark igniter shutdown detonations

Detonations were experienced in the Space Shuttle Main Engine fuel preburner (FPB) augmented spark igniter (ASI) during engine cutoff. Several of these resulted in over pressures sufficient to damage the FPB ASI oxidizer system. The detonations initiated in the FPB ASI oxidizer line when residual oxidizer (oxygen) in the line mixed with backflowing fuel (hydrogen) and detonated. This paper reviews the damage history to the FPB ASI oxidizer system, an engineering assessment of the problem cause, a verification of the mechanisms, the hazards associated with the detonations, and the solution implemented.

Dexter, C. E.↗
NASA NTRS1986

Thermodynamic evaluation of transonic compressor rotors using the finite volume approach

The development of a computational capability to handle viscous flow with an explicit time-marching method based on the finite volume approach is summarized. Emphasis is placed on the extensions to the computational procedure which allow the handling of shock induced separation and large regions of strong backflow. Appendices contain abstracts of papers and whole reports generated during the contract period.

Moore, John↗
NASA NTRS1987

Experimental study of two separating turbulent boundary layers

A detailed study of two strong adverse pressure gradient flows, one with a free-stream velocity of 35 m/sec, at throat (producing a Re sub theta of 27000 at detachment) and another with free-stream velocity of 22 m/sec, at throat (producing a Re sub theta of 19000 at detachment) is presented. In these examples flows separate slowly and reattach very rapidly over a very short distance in a streamwise direction. In the backflow region, there appears to be a semi-logarithmically flat region in the streamwise fluctuating velocity component, u', which spreads over a definite range of y/delta. In power spectra, the flow variables phi sub upsilon upsilon (kappa sub 1 delta)/ -uv bar sub max vs. kappa sub 1 delta forms a unique set of scaling parameters for adverse pressure gradient flows. Experimental results show good agreement with previous studies.

Nagabushana, K. A.↗
NASA NTRS1988

Evaluation and Improvement of Liquid Propellant Rocket Chugging Analysis Techniques. Part 2: a Study of Low Frequency Combustion Instability in Rocket Engine Preburners Using a Heterogeneous Stirred Tank Reactor Model

During the shutdown of the space shuttle main engine, oxygen flow is shut off from the fuel preburner and helium is used to push the residual oxygen into the combustion chamber. During this process a low frequency combustion instability, or chug, occurs. This chug has resulted in damage to the engine's augmented spark igniter due to backflow of the contents of the preburner combustion chamber into the oxidizer feed system. To determine possible causes and fixes for the chug, the fuel preburner was modeled as a heterogeneous stirred tank combustion chamber, a variable mass flow rate oxidizer feed system, a constant mass flow rate fuel feed system and an exit turbine. Within the combustion chamber gases were assumed perfectly mixed. To account for liquid in the combustion chamber, a uniform droplet distribution was assumed to exist in the chamber, with mean droplet diameter determined from an empirical relation. A computer program was written to integrate the resulting differential equations. Because chamber contents were assumed perfectly mixed, the fuel preburner model erroneously predicted that combustion would not take place during shutdown. The combustion rate model was modified to assume that all liquid oxygen that vaporized instantaneously combusted with fuel. Using this combustion model, the effect of engine parameters on chamber pressure oscillations during the SSME shutdown was calculated.

Bartrand, Timothy A.↗
NASA NTRS1987

Absolute instability of the Gaussian wake profile

Linear parallel-flow stability theory has been used to investigate the effect of viscosity on the local absolute instability of a family of wake profiles with a Gaussian velocity distribution. The type of local instability, i.e., convective or absolute, is determined by the location of a branch-point singularity with zero group velocity of the complex dispersion relation for the instability waves. The effects of viscosity were found to be weak for values of the wake Reynolds number, based on the center-line velocity defect and the wake half-width, larger than about 400. Absolute instability occurs only for sufficiently large values of the center-line wake defect. The critical value of this parameter increases with decreasing wake Reynolds number, thereby indicating a shrinking region of absolute instability with decreasing wake Reynolds number. If backflow is not allowed, absolute instability does not occur for wake Reynolds numbers smaller than about 38.

Hultgren, Lennart S.↗
NASA NTRS1988

Features of separating turbulent boundary layers

In the present study of two strong adverse pressure gradient flows, mean flow and turbulence characteristics are measured, together with frequency spectra, using hot-wire and laser anemometry. In these separating flows, reattachment occurs over a shorter distance than separation. It is noted that the outer flow variables form a unique set of scaling parameters for streamwise power spectra in adverse pressure gradient flows. The inner flow scaling of Perry et al. (1985) for streamwise spectra does not hold in the backflow region unless the value of the downstream-upstream intermittency in the flow is unity.

Nagabushana, K. A.↗
NASA NTRS1988

Thermal and pressure diffusion effects in high altitude flows

An apparently anomalous change in composition adjacent to a non-catalytic surface in an earlier DSMC calculation of a non-reacting gas mixture is shown to have been due to thermal diffusion. Both thermal and pressure diffusion can lead to species separation effects in the flow of gas mixtures, but are not generally included in Navier-Stokes formulations. Additional DSMC calculations are made for blunt-body and plume expansion flows using both a simple gas and an otherwise identical gas mixture. These test calculations indicate that the pressure distribution on the forward face of a blunt body is largely unaffected by the species separation in the stagnation region. The gas density in this region is significantly affected and there is a slight change in the heat transfer to the surface. The pressure distribution around the shoulder of the body is affected. Very large separations occur in the plume expansion of the gas mixtures and there is an order of magnitude increase in density in the backflow region for the mixture.

Bird, G. A.↗
NASA NTRS1989

A prediction of 3-D viscous flow and performance of the NASA low-speed centrifugal compressor

A prediction of the 3-D turbulent flow in the NASA Low-Speed Centrifugal Compressor Impeller has been made. The calculation was made for the compressor design conditions with the specified uniform tip clearance gap. The predicted performance is significantly worse than that predicted in the NASA design study. This is explained by the high tip leakage flow in the present calculation and by the different model adopted for tip leakage flow mixing. The calculation gives an accumulation for high losses in the shroud/pressure-side quadrant near the exit of the impeller. It also predicts a region of meridional backflow near the shroud wall. Both of these flow features should be extensive enough in the NASA impeller to allow detailed flow measurements, leading to improved flow modelling. Recommendations are made for future flow studies in the NASA impeller.

Moore, John↗
NASA NTRS1989

Absolute/convective instabilities and the convective Mach number in a compressible mixing layer

Two aspects of the stability of a compressible mixing layer: Absolute/Convective instability and the convective Mach number were considered. It was shown that, for Mach numbers less than one, the compressible mixing layer is convectively unstable unless there is an appreciable amount of backflow. Also presented was a rigorous derivation of a convective Mach number based on linear stability theory for the flow of a multi-species gas in a mixing layer. The result is compared with the heuristic definitions of others and to selected experimental results.

Jackson, T. L.↗
NASA NTRS1989

Numerical simulations of magnetized jets

The present axisymmetric numerical simulations of light hypersonic jets allow unmagnetized jets and jets carrying a dynamically important magnetic field to be contrasted. After decelerating a weakly magnetized jet through a series of weak, oblique shocks, a Mach disk and a strong annular shock are encountered near the outer edges of the contact discontinuity separating the shocked fluid from the shocked ambient gas. Upon passing the annular shock, the gas quickly expands and enters a backflowing cocoon surrounding the jet. The overall speed of advance of the jet is reduced; matter near the jet axis which passes through the terminal Mach disk accumulates in a plug, and gas is discharged into the cocoon by the intermittent shedding of vortices. When magnetic stresses dominate, however, the jet is rapidly decelerated via a Mach disk and strong annular shock.

Lind, Kevin R.↗
NASA NTRS1989

The measurement of boundary layers on a compressor blade in cascade. Volume 2: Data tables

Measurements were made of the boundary layers and wakes about a highly loaded, double-circular-arc compressor blade in cascade. These laser Doppler velocimetry measurements have yielded a very detailed and precise data base with which to test the application of viscous computational codes to turbomachinery. In order to test the computational codes at off-design conditions, the data have been acquired at a chord Reynolds number of 500,000 and at three incidence angles. Average values and 95 percent confidence bands were tabularized for the velocity, local turbulence intensity, skewness, kurtosis, and percent backflow. Tables also exist for the blade static-pressure distributions and boundary layer velocity profiles reconstructed to account for the normal pressure gradient.

Zierke, William C.↗
NASA NTRS1989

Embedded function methods for compressible high speed turbulent flow

Fundamental issues relating to compressible turbulent flow are addressed. The focus has been on developing methods and testing concepts for attached flows rather than trying to force a conventional law of the wall into a zone of backflow. Although the dynamics of the near-wall flow in an attached turbulent boundary layer are relatively well documented, the dynamical features of a zone of reversed turbulent flow are not, nor are they well understood. Incompressibility introduces effects and issues that have been dealt with only marginally in the literature, therefore, the present work has been focussed on attached high-speed flows. The wall function method has been extended up through the supersonic to hypersonic speeds. Algorithms have been successfully introduced into the code that calculates the flow all the way to the wall, and testing is being carried out for progressively more complex flow situations.

Walker, J. D. A.↗