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At least 145 records · Page 8

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NASA NTRS1985

A theoretical study of non-adiabatic surface effects for a model in the NTF cryogenic wind tunnel

A theoretical analysis was made of the severity and effect of nonadiabatic surface conditions for a model in the NTF cryogenic wind tunnel. The nonadiabatic condition arises from heaters that are used to maintain a constant thermal environment for instrumentation internal to the model. The analysis was made for several axi-symmetric representations of a fuselage cavity, using a finite element heat conduction code. Potential flow and boundary layer codes were used to calculate the convection condition for the exterior surface of the model. The results of the steady state analysis show that it is possible to maintain the surface temperature very near the adiabatic value, with the judicious use of insulating material. Even for the most severe nonadiabatic condition studied, the effects on skin friction drag and displacement thickness were only marginally significant. The thermal analysis also provided an estimate of the power required to maintain a specified cavity temperature.

Macha, J. M.↗

NASA NTRS1985

Numerical investigation of internal high-speed viscous flows using a parabolic technique

A feasibility study has been conducted to assess the applicability of an existing parabolic analysis (ADD-Axisymmetric Diffuser Duct), developed previously for subsonic viscous internal flows, to mixed supersonic/subsonic flows with heat addition simulating a SCRAMJET combustor. A study was conducted with the ADD code modified to include additional convection effects in the normal momentum equation when supersonic expansion and compression waves are present. A set of test problems with weak shock and expansion waves have been analyzed with this modified ADD method and stable and accurate solutions were demonstrated provided the streamwise step size was maintained at levels larger than the boundary layer displacement thickness. Calculations made with further reductions in step size encountered departure solutions consistent with strong interaction theory. Calculations were also performed for a flow field with a flame front in which a specific heat release was imposed to simulate a SCRAMJET combustor. In this case the flame front generated relatively thick shear layers which aggravated the departure solution problem. Qualitatively correct results were obtained for these cases using a marching technique with the convective terms in the normal momentum equation suppressed. It is concluded from the present study that for the class of problems where strong viscous/inviscid interactions are present a global iteration procedure is required.

Anderson, O. L.↗

NASA NTRS1985

Flow field survey near the rotational plane of an advanced design propeller on a JetStar airplane

An investigation was conducted to obtain upper fuselage surface static pressures and boundary layer velocity profiles below the centerline of an advanced design propeller. This investigation documents the upper fuselage velocity flow field in support of the in-flight acoustic tests conducted on a JetStar airplane. Initial results of the boundary layer survey show evidence of an unusual flow disturbance, which is attributed to the two windshield wiper assemblies on the aircraft. The assemblies were removed, eliminating the disturbances from the flow field. This report presents boundary layer velocity profiles at altitudes of 6096 and 9144 m (20,000 and 30,000 ft) and Mach numbers from 0.6 to 0.8, and it investigated the effects of windshield wiper assemblies on these profiles. Because of the unconventional velocity profiles that were obtained with the assemblies mounted, classical boundary layer parameters, such as momentum and displacement thicknesses, are not presented. The effects of flight test variables (Mach number and angles of attack and sideslip) and an advanced design propeller on boundary layer profiles - with the wiper assemblies mounted and removed - are presented.

Walsh, K. R.↗

NASA NTRS1986

The application to airfoils of a technique for reducing orifice-induced pressure error at high Reynolds numbers

A wind tunnel investigation was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel to study the effects of porous (sintered metal) plug orifices on orifice-induced static-pressure measurement error at high Reynolds numbers. A NACA airfoil was tested at Mach numbers from 0.60 to 0.80 and at Reynolds numbers from 6 x 1,000,000 to 40 x 1,000,000. Data are included which compare pressure measurements obtained from porous plug orifices and from conventional orifices with diameters of 0.025 cm (0.010 in.) and 0.102 cm (0.040 in.). The two dimensional airfoil code GRUMFOIL was used to calculate boundary layer displacement thickness. The response time and the downstream effect of the porous plug orifice were considered in this investigation. The results showed that the porous plug orifice could be a viable method of reducing pressure error. The data also showed that the pressure measurements obtained with a 0.102-cm-diameter orifice were very close to the measurements obtained with 0.025-cm-diameter orifice over such of the airfoil and that downstream of a shock the orifice size was not critical.

Plentovich, E. B.↗

NASA NTRS1986

Extensional tectonics during the igneous emplacement of the mafic-ultramafic rocks of the Barberton greenstone belt

The simatic rocks (Onverwacht Group) of the Barberton greenstone belt are part of the Jamestown ophiolite complex. This ophiolite, together with its thick sedimentary cover occupies a complex thrust belt. Field studies have identified two types of early faults which are entirely confined to the simatic rocks and are deformed by the later thrusts and associated folds. The first type of fault (F1a) is regional and always occurs in the simatic rocks along and parallel to the lower contacts of the ophiolite-related cherts (Middle Marker and equivalent layers). These fault zones have previously been referred to both as flaser-banded gneisses and as weathering horizons. In general the zones range between 1-30m in thickness. Displacements along these zones are difficult to estimate, but may be in the order of 1-100 km. The structures indicate that the faults formed close to horizontal, during extensional shear and were therefore low angle normal faults. F1a zones overlap in age with the formation of the ophiolite complex. The second type of faults (F1b) are vertical brittle-ductile shear zones, which crosscut the complex at variable angles and cannot always be traced from plutonic to overlying extrusive (pillowed) simatic rocks. F1b zones are also apparently of penecontemporaneous origin with the intrusive-extrusive igneous processs. F1b zones may either represent transform fault-type activity or represent root zones (steepened extensions) of F1a zones. Both fault types indicate extensive deformation in the rocks of the greenstone belt prior to compressional overthrust tectonics.

Dewit, M. J.↗

NASA NTRS1986

ALESEP. Part 2: A computer program for the analysis of leading edge separation bubbles on infinite swept wings

A program called ALESEP is presented for the analysis of the inviscid-viscous interaction which occurs due to the presence of a closed laminar-transitional separation bubble on an airfoil or infinite swept wing. The ALESEP code provides an iterative solution of the boundary layer equations expressed in an inverse formulation coupled to a Cauchy integral representation of the inviscid flow. This interaction analysis is treated as a local perturbation to a known solution obtained from a global airfoil analysis; hence, part of the required input to the ALESEP code are the reference displacement thickness and tangential velocity distributions. Special windward differencing may be used in the reversed flow regions of the separation bubble to accurately account for the flow direction in the discretization of the streamwise convection of momentum. The ALESEP code contains a forced transition model based on a streamwise intermittency function, a natural transition model based on a solution of the integral form of the turbulent kinetic energy equation, and an empirical natural transition model.

Davis, R. L.↗

NASA NTRS1986

Tables for correcting airfoil data obtained in the Langley 0.3-meter transonic cryogenic tunnel for sidewall boundary-layer effects

Tables for correcting airfoil data taken in the Langley 0.3-meter Transonic Cryogenic Tunnel for the presence of sidewall boundary layer are presented. The corrected Mach number and the correction factor are minutely altered by a 20 percent change in the boundary layer virtual origin distance. The sidewall boundary layer displacement thicknesses measured for perforated sidewall inserts and without boundary layer removal agree with the values calculated for solid sidewalls.

Jenkins, R. V.↗

NASA NTRS1986

The acoustic field of a point source in a uniform boundary layer over an impedance plane

The acoustic field of a point source in a boundary layer above an impedance plane is investigated anatytically using Obukhov quasi-potential functions, extending the normal-mode theory of Chunchuzov (1984) to account for the effects of finite ground-plane impedance and source height. The solution is found to be asymptotic to the surface-wave term studies by Wenzel (1974) in the limit of vanishing wind speed, suggesting that normal-mode theory can be used to model the effects of an atmospheric boundary layer on infrasonic sound radiation. Model predictions are derived for noise-generation data obtained by Willshire (1985) at the Medicine Bow wind-turbine facility. Long-range downwind propagation is found to behave as a cylindrical wave, with attention proportional to the wind speed, the boundary-layer displacement thickness, the real part of the ground admittance, and the square of the frequency.

Zorumski, W. E.↗

NASA NTRS1986

Comparison of theoretical and flight-measured local flow aerodynamics for a low-aspect-ratio fin

Flight test and theoretical aerodynamic data were obtained for a flight test fixture mounted on the underside of an F-104G aircraft. The theoretical data were generated using two codes: a two-dimensional transonic code called code H, and a three-dimensional subsonic and supersonic code called wing-body. Pressure distributions generated by the codes for the flight test fixture, as well as compared with the flight-measured data. The two-dimensional code pressure distributions compared well except at the minimum pressure point and the trailing edge. Shock locations compared well except at high transonic speeds. However, the two-dimensional code did not adequately predict the displacement thickness of the flight test fixture. The three-dimensional code pressure distributions compared well except at the trailing edge of the flight test fixture.

Johnson, J. Blair↗

NASA NTRS1987

Calculation of sidewall boundary-layer parameters from rake measurements for the Langley 0.3-meter transonic cryogenic tunnel

Correction of airfoil data for sidewall boundary-layer effects requires a knowledge of the boundary-layer displacement thickness and the shape factor with the tunnel empty. To facilitate calculation of these quantities under various test conditions for the Langley 0.3 m Transonic Cryogenic Tunnel, a computer program was written. This program reads the various tunnel parameters and the boundary-layer rake total head pressure measurements directly from the Engineering Unit tapes to calculate the required sidewall boundary-layer parameters. Details of the method along with the results for a sample case are presented.

Murthy, A. V.↗

NASA NTRS1987

Analysis of viscous transonic flow over airfoil sections

A full Navier-Stokes solver has been used to model transonic flow over three airfoil sections. The method uses a two-dimensional, implicit, conservative finite difference scheme for solving the compressible Navier-Stokes equations. Results are presented as prescribed for the Viscous Transonic Airfoil Workshop to be held at the AIAA 25th Aerospace Sciences Meeting. The NACA 0012, RAE 2822 and Jones airfoils have been investigated for both attached and separated transonic flows. Predictions for pressure distributions, loads, skin friction coefficients, boundary layer displacement thickness and velocity profiles are included and compared with experimental data when possible. Overall, the results are in good agreement with experimental data.

Huff, Dennis L.↗

NASA NTRS1986

Heat and mass transfer rates during flow of dissociated hydrogen gas over graphite surface

To improve upon the performance of chemical rockets, the nuclear reactor has been applied to a rocket propulsion system using hydrogen gas as working fluid and a graphite-composite forming a part of the structure. Under the boundary layer approximation, theoretical predictions of skin friction coefficient, surface heat transfer rate and surface regression rate have been made for laminar/turbulent dissociated hydrogen gas flowing over a flat graphite surface. The external stream is assumed to be frozen. The analysis is restricted to Mach numbers low enough to deal with the situation of only surface-reaction between hydrogen and graphite. Empirical correlations of displacement thickness, local skin friction coefficient, local Nusselt number and local non-dimensional heat transfer rate have been obtained. The magnitude of the surface regression rate is found low enough to ensure the use of graphite as a linear or a component of the system over an extended period without loss of performance.

Nema, V. K.↗

NASA NTRS1987

Analysis of viscous transonic flow over airfoil sections

Huff, Dennis L.↗

NASA NTRS1987

A simultaneous viscous-inviscid interaction calculation procedure for transonic turbulent flows

A new simultaneous viscous-inviscid interaction scheme has been developed for the analysis of steady turbulent transonic separated flows. The viscous and inviscid solutions are coupled through the displacement concept using a transpiration velocity. The local solutions of the full potential and boundary-layer equations are treated simultaneously using the finite-difference method. The displacement thickness is treated as an unknown and is obtained as a part of the solution through a global iteration procedure of the space marching scheme. The Cebeci-Smith and Johnson-King models are used to simulate the turbulence. The computational examples showed that the simultaneous method is more efficient and robust than the semi-inverse method for transonic flows with a strong interaction.

Lee, D.↗

NASA NTRS1987

An experimental mapping of the flow field behind a glaze ice shape on a NACA 0012 airfoil

The flow field about a NACA 0012 airfoil with a simulated glaze ice shape was studied. Split hot-film anemometry was used to measure the streamwise velocity component in the upper and lower separation bubbles aft of the glaze ice horns. Velocity profiles were presented as well as the boundary layer momentum and displacement thickness distributions through the bubbles. Data were presented at angles of attack of 0, 2 and 4 deg and clearly showed the large region of reverse flow. A detailed discussion of split hot-film acquisition and data reduction, and possible sources of error has been included.

Spring, Samuel A.↗

NASA NTRS1988

Investigation of viscous/inviscid interaction in transonic flow over airfoils with suction

The viscous/inviscid interaction over transonic airfoils with and without suction is studied. The streamline angle at the edge of the boundary layer is used to couple the viscous and inviscid flows. The potential flow equations are solved for the inviscid flow field. In the shock region, the Euler equations are solved using the method of integral relations. For this, the potential flow solution is used as the initial and boundary conditions. An integral method is used to solve the laminar boundary-layer equations. Since both methods are integral methods, a continuous interaction is allowed between the outer inviscid flow region and the inner viscous flow region. To avoid the Goldstein singularity near the separation point the laminar boundary-layer equations are derived in an inverse form to obtain solution for the flows with small separations. The displacement thickness distribution is specified instead of the usual pressure distribution to solve the boundry-layer equations. The Euler equations are solved for the inviscid flow using the finite volume technique and the coupling is achieved by a surface transpiration model. A method is developed to apply a minimum amount of suction that is required to have an attached flow on the airfoil. The turbulent boundary layer equations are derived using the bi-logarithmic wall law for mass transfer. The results are found to be in good agreement with available experimental data and with the results of other computational methods.

Vemuru, C. S.↗

NASA NTRS1988

Effect of stage loading on endwall flows in an axial flow compressor rotor

This paper reports results from investigations conducted to determine the effect of stage loading on endwall flows in a low speed axial flow compressor. These investigations consisted of two sets of measurements. The first set consisted of radial transverse of flow properties at the rotor inlet and exit, at five flow coefficients. These measurements are used to determine the boundary layer integral parameters. The displacement thicknesses at the rotor hub and tip agree reasonably well with Smith's (1970) correlation for multistage axial compressors. The second set consisted of measurements of static pressures on the rotor blade at four flow coefficients. From these measurements, lift coefficient is determined. Also, loss of lift coefficient near the tip is calculated, and is attributed mainly to the tip leakage flows.

Sitaram, N.↗