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At least 343 records · Page 19

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NASA NTRS1982

Relaxation solution of the full Euler equations

A numerical procedure for the relaxation solution of the full steady Euler equations is described. By embedding the Euler system in a second order surrogate system, central differencing may be used in subsonic regions while retaining matrix forms well suited to iterative solution procedures and convergence acceleration techniques. Hence, this method allows the development of stable, fully conservative differencing schemes for the solution of quite general inviscid flow problems. Results are presented for both subcritical and shocked supercritical internal flows. Comparisons are made with a standard time dependent solution algorithm.

Johnson, G. M.↗

NASA NTRS1982

Relaxation solution of the full Euler equations

Johnson, G. M.↗

NASA NTRS2000

Innovative Liner Concepts: Experiments and Impedance Modeling of Liners Including the Effect of Bias Flow

The study of normal impedance of perforated plate acoustic liners including the effect of bias flow was studied. Two impedance models were developed by modeling the internal flows of perforate orifices as infinite tubes with the inclusion of end corrections to handle finite length effects. These models assumed incompressible and compressible flows, respectively, between the far field and the perforate orifice. The incompressible model was used to predict impedance results for perforated plates with percent open areas ranging from 5% to 15%. The predicted resistance results showed better agreement with experiments for the higher percent open area samples. The agreement also tended to deteriorate as bias flow was increased. For perforated plates with percent open areas ranging from 1% to 5%, the compressible model was used to predict impedance results. The model predictions were closer to the experimental resistance results for the 2% to 3% open area samples. The predictions tended to deteriorate as bias flow was increased. The reactance results were well predicted by the models for the higher percent open area, but deteriorated as the percent open area was lowered (5%) and bias flow was increased. A fit was done on the incompressible model to the experimental database. The fit was performed using an optimization routine that found the optimal set of multiplication coefficients to the non-dimensional groups that minimized the least squares slope error between predictions and experiments. The result of the fit indicated that terms not associated with bias flow required a greater degree of correction than the terms associated with the bias flow. This model improved agreement with experiments by nearly 15% for the low percent open area (5%) samples when compared to the unfitted model. The fitted model and the unfitted model performed equally well for the higher percent open area (10% and 15%).

Kelly, Jeff↗

NASA NTRS2001

Thermal Analysis of a Carbon Fiber Rope Barrier for Use in the Reusable Solid Rocket Motor Nozzle Joint-2

This study provides for development and verification of analysis methods used to assess performance of a carbon fiber rope (CFR) thermal barrier system that is currently being qualified for use in Reusable Solid Rocket Motor (RSRM) nozzle joint-2. Modeled geometry for flow calculations considers the joint to be vented with the porous CFR barriers placed in the "open' assembly gap. Model development is based on a 1-D volume filling approach where flow resistances (assembly gap and CFRs) are defined by serially connected internal flow and the porous media "Darcy" relationships. Combustion gas flow rates are computed using the volume filling code by assuming a lumped distribution total joint fill volume on a per linear circumferential inch basis. Gas compressibility, friction and heat transfer are included in the modeling. Gas-to-wall heat transfer is simulated by concurrent solution of the compressible flow equations and a large thermal 2-D finite element (FE) conduction grid. The derived numerical technique loosely couples the FE conduction matrix with the compressible gas flow equations, Free constants that appear in the governing equations are calibrated by parametric model comparison to hot fire subscale test results. The calibrated model is then used to make full-scale motor predictions using RSRM aft dome environments. Model results indicate that CFR thermal barrier systems will provide a thermally benign and controlled pressurization environment for the RSRM nozzle joint-2 primary seal activation.

Clayton, J. Louie↗

NASA NTRS2002

Thermal Analysis of a Carbon Fiber Rope Barrier for Use in the Reusable Solid Rocket Motor Nozzle Joint-2

This study provides development and verification of analysis methods used to assess performance of a carbon fiber rope (CFR) thermal barrier system that is currently being qualified for use in Reusable Solid Rocket Motor (RSRM) nozzle joint-2. Modeled geometry for flow calculations considers the joint to be vented with the porous CFR barriers placed in the 'open' assembly gap. Model development is based on a 1-D volume filling approach where flow resistances (assembly gap and CFRs) are defined by serially connected internal flow and the porous media 'Darcy' relationships. Combustion gas flow rates are computed using the volume filling code by assuming a lumped distribution total joint fill volume on a per linear circumferential inch basis. Gas compressibility, friction and heat transfer are included in the modeling. Gas-to-wall heat transfer is simulated by concurrent solution of the compressible flow equations and a large thermal 2-D finite element (FE) conduction grid. The derived numerical technique loosely couples the FE conduction matrix with the compressible gas flow equations. Free constants that appear in the governing equations are calibrated by parametric model comparison to hot fire subscale test results. The calibrated model is then used to make full-scale motor predictions using RSRM aft dome environments. Model results indicate that CFR thermal barrier systems will provide a thermally benign and controlled pressurization environment for the RSRM nozzle joint-2 primary seal activation.

Clayton, J. Louie↗

NASA NTRS1991

Modeling flow at the nozzle of a solid rocket motor

The mechanical behavior of a rocket motor internal flow field results in a system of nonlinear partial differential equations which can be solved numerically. The accuracy and the convergence of the solution of the system of equations depends largely on how precisely the sharp gradients can be resolved. An adaptive grid generation scheme is incorporated into the computer algorithm to enhance the capability of numerical modeling. With this scheme, the grid is refined as the solution evolves. This scheme significantly improves the methodology of solving flow problems in rocket nozzle by putting the refinement part of grid generation into the computer algorithm.

Chow, Alan S.↗

NASA NTRS1988

Development of a thermal and structural analysis procedure for cooled radial turbines

A procedure for computing the rotor temperature and stress distributions in a cooled radial turbine are considered. Existing codes for modeling the external mainstream flow and the internal cooling flow are used to compute boundary conditions for the heat transfer and stress analysis. The inviscid, quasi three dimensional code computes the external free stream velocity. The external velocity is then used in a boundary layer analysis to compute the external heat transfer coefficients. Coolant temperatures are computed by a viscous three dimensional internal flow cade for the momentum and energy equation. These boundary conditions are input to a three dimensional heat conduction code for the calculation of rotor temperatures. The rotor stress distribution may be determined for the given thermal, pressure and centrifugal loading. The procedure is applied to a cooled radial turbine which will be tested at the NASA Lewis Research Center. Representative results are given.

Kumar, Ganesh N.↗

NASA NTRS1988

Development of a thermal and structural analysis procedure for cooled radial turbines

A procedure for computing the rotor temperature and stress distributions in a cooled radial turbine is considered. Existing codes for modeling the external mainstream flow and the internal cooling flow are used to compute boundary conditions for the heat transfer and stress analyses. An inviscid, quasi three-dimensional code computes the external free stream velocity. The external velocity is then used in a boundary layer analysis to compute the external heat transfer coefficients. Coolant temperatures are computed by a viscous one-dimensional internal flow code for the momentum and energy equation. These boundary conditions are input to a three-dimensional heat conduction code for calculation of rotor temperatures. The rotor stress distribution may be determined for the given thermal, pressure and centrifugal loading. The procedure is applied to a cooled radial turbine which will be tested at the NASA Lewis Research Center. Representative results from this case are included.

Kumar, Ganesh N.↗

NASA NTRS1983

Computation of viscous flow in curved ducts and comparison with experimental data

A three dimensional analysis for fully viscous subsonic internal flow is evaluated. The analysis, designated PEPSIG, solves an approximate form of the Navier-Stokes equations by an implicit spatial marching procedure. Results of calculations are presented for laminar flow through two different circular cross-sectioned 180 degree bends, and for laminar and turbulent flow through circular and square cross-sectioned 22.5 to 22.5 degree S-ducts. Quantitative comparisons with experimental data are shown for all cases. Special emphasis is placed on verifying the ability of the analysis to accurately predict the distorted flow fields resulting from pressure-driven secondary flows.

Towne, C. E.↗

NASA NTRS1984

Computation of viscous flow in curved ducts and comparison with experimental data

A three dimensional analysis for fully viscous subsonic internal flow is evaluated. The analysis, designated PEPSIG, solves an approximate form of the Navier-Stokes equations by an implicit spatial marching procedure. Results of calculations are presented for laminar flow through two different circular cross-sectioned 180 degree bends, and for laminar and turbulent flow through circular and square cross-sectioned 22.5 to 22.5 degree S-ducts. Quantitative comparisons with experimental data are shown for all cases. Special emphasis is placed on verifying the ability of the analysis to accurately predict the distored flow fields resulting from pressure-driven secondary flows. Previously announced in STAR as N84-13404

Towne, C. E.↗

NASA NTRS1975

Aerodynamic Analyses Requiring Advanced Computers, Part 1

Papers are presented which deal with results of theoretical research on aerodynamic flow problems requiring the use of advanced computers. Topics discussed include: viscous flows, boundary layer equations, turbulence modeling and Navier-Stokes equations, and internal flows.

Source record↗

NASA NTRS1978

Aerothermodynamic Assessment of Corrugated Panel Thermal Protection Systems

The feasibility of using corrugated panels as a thermal protection system for an advanced space transportation vehicle was investigated. The study consisted of two major tasks: development of improved correlations for wind tunnel heat transfer and pressure data to yield design techniques, and application of the design techniques to determine if corrugated panels have application future aerospace vehicles. A single-stage-to-orbit vehicle was used to assess advantages and aerothermodynamic penalties associated with use of such panels. In the correlation task, experimental turbulent heat transfer and pressure data obtained on corrugation roughened surfaces during wind tunnel testing were analyzed and compared with flat plate data. The correlations and data comparisons included the effects of a large range of geometric, inviscid flow, internal boundary layer, and bulk boundary layer parameters in supersonic and hypersonic flow.

Brandon, H. J.↗

NASA NTRS1979

Influence of corrugation-roughness shape on aerodynamic heating in thick turbulent boundary layers

The results of an experimental program to evaluate the effects of wave shape on heat transfer and pressure distributions of corrugation roughened flat plates in thick turbulent boundary layers are presented. The wave shaping effects studied were wavelength, wave amplitude, and wave radius. The experimental program consisted of tests in the sidewall boundary layer of the Langley Continuous Flow Hypersonic Tunnel at a Mach number of 10.3 and Reynolds numbers from 1.3 to 8.2 x 10 to the 6th per meter. The effects of wave shape are analyzed, correlated and compared with previous data obtained in both thick and thin boundary layers. The correlations include the effects of a large range of geometric, inviscid flow, internal boundary layer, and bulk boundary layer parameters in supersonic and hypersonic flow.

Brandon, H. J.↗

NASA NTRS1981

Dynamic phenomena in sunspots

A detailed summary of observed dynamic phenomena associated with sunspots is presented, together with a description of the observational techniques and available analytical formulations for the processes under study. The phenomena detected thus far are grouped into aperiodic events and oscillations and waves. Aperiodic phenomena comprise umbral flares, the superpenumbra, and inverse Evershed flow. Internal, aperiodic manifestations include penumbral grains, the photospheric penumbral dark fibrils, Evershed flow, umbral dots, the inhomogeneity of the umbral magnetic field, and umbral turbulence. Oscillations and flashes are seen in the umbra, while running waves and dark puffs have been detected in the penumbra, and oscillations are located in the photosphere. All the observed features are evidence of mass motion and change on time scales of less than an hour.

Moore, R. L.↗

NASA NTRS1983

PAN AIR modeling studies

PAN AIR is a computer program that predicts subsonic or supersonic linear potential flow about arbitrary configurations. The code's versatility and generality afford numerous possibilities for modeling flow problems. Although this generality provides great flexibility, it also means that studies are required to establish the dos and don'ts of modeling. The purpose of this paper is to describe and evaluate a variety of methods for modeling flows with PAN AIR. The areas discussed are effects of panel density, internal flow modeling, forebody modeling in subsonic flow, propeller slipstream modeling, effect of wake length, wing-tail-wake interaction, effect of trailing-edge paneling on the Kutta condition, well- and ill-posed boundary-value problems, and induced-drag calculations. These nine topics address problems that are of practical interest to the users of PAN AIR.

Towne, M. C.↗

NASA NTRS1992

SRMAFTE facility checkout model flow field analysis

The Solid Rocket Motor Air Flow Equipment (SRMAFTE) facility was constructed for the purpose of evaluating the internal propellant, insulation, and nozzle configurations of solid propellant rocket motor designs. This makes the characterization of the facility internal flow field very important in assuring that no facility induced flow field features exist which would corrupt the model related measurements. In order to verify the design and operation of the facility, a three-dimensional computational flow field analysis was performed on the facility checkout model setup. The checkout model measurement data, one-dimensional and three-dimensional estimates were compared, and the design and proper operation of the facility was verified. The proper operation of the metering nozzles, adapter chamber transition, model nozzle, and diffuser were verified. The one-dimensional and three-dimensional flow field estimates along with the available measurement data are compared.

Dill, Richard A.↗

NASA NTRS1998

Application of Pressure Sensitive Paint to Confined Flow at Mach Number 2.5

Pressure sensitive paint (PSP) is a novel technology that is being used frequently in external aerodynamics. For internal flows in narrow channels, and applications at elevated nonuniform temperatures, however, there are still unresolved problems that complicate the procedures for calibrating PSP signals. To address some of these problems, investigations were carried out in a narrow channel with supersonic flows of Mach 2.5. The first set of tests focused on the distribution of the wall pressure in the diverging section of the test channel downstream of the nozzle throat. The second set dealt with the distribution of wall static pressure due to the shock/wall interaction caused by a 25 deg. wedge in the constant Mach number part of the test section. In addition, the total temperature of the flow was varied to assess the effects of temperature on the PSP signal. Finally, contamination of the pressure field data, caused by internal reflection of the PSP signal in a narrow channel, was demonstrated. The local wall pressures were measured with static taps, and the wall pressure distributions were acquired by using PSP. The PSP results gave excellent qualitative impressions of the pressure field investigated. However, the quantitative results, specifically the accuracy of the PSP data in narrow channels, show that improvements need to be made in the calibration procedures, particularly for heated flows. In the cases investigated, the experimental error had a standard deviation of +/- 8.0% for the unheated flow, and +/- 16.0% for the heated flow, at an average pressure of 11 kpa.

Lepicovsky, J.↗

NASA NTRS1983

Design of Multistage Axial-Flow Compressors

Program developed for computing aerodynamic design of multistage axialflow compressor and associated blading geometry input for internal flow analysis. Aerodynamic solution gives velocity diagrams on selected streamlines of revolution at blade row edges. Program written in FORTRAN IV.

Crouse, J. E.↗