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Results for “Freestream Mach Number”

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 163 records · Page 9
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NASA NTRS1973

High-speed leading edge problem.

The sharp leading edge problem has been studied for both monatomic and diatomic gases using the Boltzmann equation with the Bhatnagar-Gross-Krook type models as the governing equation and the discrete ordinate method with a closed-boundary value approach as a tool. Plate length relative to the freestream mean free path is taken to be 52. The gas-surface interaction law is assumed to be diffuse reflection. The local distribution functions of molecular velocities and internal energies (for the diatomic gas) for the entire flowfield have been calculated for a freestream Mach number of 6.1. Comparisons are made between the calculated results and experimental data.

Huang, A. B.↗
NASA NTRS1974

Transonic conical flow

The problem of inviscid, steady transonic conical flow, formulated in terms of the small disturbance theory, is studied. The small disturbance equation and similarity rules are presented, and a boundary value problem is formulated for the case of a supersonic freestream Mach number. The equation for the perturbation potential is solved numerically using an elliptic finite difference system. The difference equations are solved with a point relaxation algorithm that is also capable of capturing the shock wave during the iteration procedure by using the boundary conditions at the shock. Numerical calculations, for shock location, pressure distribution and drag coefficient, are presented for a family of nonlifting conical wings. The theory of slender wings is also presented and analytical results for pressure and drag coefficients are obtained.

Agopian, K. G.↗
NASA NTRS1974

Rearward-facing steps in laminar supersonic flows with and without suction

An experimental investigation of heat-transfer and pressure distributions within regions of laminar separated flows produced by two-dimensional rearward-facing steps has been carried out at freestream Mach numbers of around 4 in the range of step height-to-boundary layer thickness varying from 0.1 to 2.4. With no suction from the separated area, the ratio of the maximum post-step heat transfer to the attached-flow values was less than unity. The maximum heating-rate region was located far downstream of the reattachment plate stagnation point. Mass suction from the separated area increased the local heating rates, this effect was however relatively weak for purely laminar flow conditions and the competing effect of the step height clearly predominated. At step heights comparable with boundary-layer thickness, even removing the entire approaching boundary layer was not sufficient to raise the post-step heating rates above the flat-plate values.

Jakubowski, A. K.↗
NASA NTRS1974

Low-speed test of translating lip axisymmetric inlets for subsonic transports

Translating lip, axisymmetric inlets have been tested at freestream Mach numbers between 0.0 and 0.30 and at angles of attack between 0 and 90 degrees. Three isolated inlet models were tested. One model represented the inlet geometry at cruise, that is, with no lip translation. The other two models had forward translating lips with different contours. The low forward speed inlet angle-of-attack capability with the translating lip was increased more than twenty degrees over the basic untranslated lip configuration at the higher inlet airflows. The static inlet performance with the translating lip was also greatly improved. The design translation distance, defined by potential flow analysis, was near optimum.

Henne, P. A.↗
NASA NTRS1973

Relaxation methods for transonic flow about wing-cylinder combinations and lifting swept wings

The mixed elliptic-hyperbolic relaxation method for obtaining steady-state solutions to two-dimensional transonic potential equations is extended to the transonic small disturbance equation in three dimensions. In particular, transonic flow is considered both about nonlifting wing-cylinder combinations and over thin lifting wings with sweep and taper. The treatment is restricted to freestream Mach numbers less than one and to wings with subsonic trailing edges.

Bailey, F. R.↗
NASA NTRS1975

Design and analysis of a supersonic penetration/maneuvering fighter

The design of three candidate air combat fighters which would cruise effectively at freestream Mach numbers of 1.6, 2.0, and 2.5 while maintaining good transonic maneuvering capability, is considered. These fighters were designed to deliver aerodynamically controlled dogfight missiles at the design Mach numbers. Studies performed by Rockwell International in May 1974 and guidance from NASA determined the shape and size of these missiles. The principle objective of this study is the aerodynamic design of the vehicles; however, configurations are sized to have realistic structures, mass properties, and propulsion systems. The results of this study show that air combat fighters in the 15,000 to 23,000 pound class would cruise supersonically on dry power and still maintain good transonic maneuvering performance.

Child, R. D.↗
NASA NTRS1975

Experimental evaluation of a Mach 3.5 axisymmetric inlet

Wind tunnel test results for a large scale inlet model designed for Mach 3.5 are presented and compared with analytical predictions. The inlet is an axisymmetric mixed-compression type with a lip diameter of 49.723 cm. The inlet design was developed using analytical procedures. Data are shown for freestream Mach numbers from 0.6 to 3.5. The test results indicate that boundary layer bleed requirements can be accurately predicted. Good agreement was obtained with analytical predictions of the flowfield structure and boundary layer development in the supersonic diffuser yielding high performance at the design Mach number. The highest engine face total pressure recovery at Mach 3.5 was 85.8%; this was obtained at 0.05 Mach tolerance with only 2.8% total pressure distortion and 13.4% bleed. In the started Mach number range from 1.6 to 3.5, the total pressure recovery in the throat, downstream of the terminal normal shock, ranged between 91% and 95%. Total pressure losses in the subsonic diffuser varied from 3% to 13%. The higher losses occuring between Mach 2.5 and 3.2 were believed to be caused by the rapid rate of increase in the area of the diffuser just downstream of the throat, possibly coupled with inadequate centerbody throat bleed. In the unstarted mode at transonic speeds, the maximum inlet flow was over 99% of the theoretical maximum capture mass-flow.

Syberg, J.↗
NASA NTRS1975

Hypersonic engine airplane integration

A summary of the research activities is presented along with the design and analysis of an integrated scramjet engine which utilizes heat conduction flames. The scramjet engine model described was tested at a freestream Mach number of 4.0 by personnel at the General Applied Science Laboratory. The research effort was primarily concerned with the interpretation and analysis of the experimental data. The mixing tests were conducted with cold ambient air, while the combustion tests were conducted with hot vitiated air. The purpose of the cold mixing tests was to determine the degree of mixedness of the fuel at the ignitor station so as to certify if the fuel injection conditions and fuel injectors locations were properly selected to give a uniform fuel concentration profile at the ignitor station. A second objective was to assess the influence of the fuel injection on the inlet flow. Several fuel injection configurations were tested.

Source record↗
NASA NTRS1975

Internal flow calculations for axisymmetric supersonic inlets at angle of attack

Several solutions for the internal flow in axisymmetric supersonic inlets at angle of attack were obtained. A shock-capturing finite-difference technique was used to obtain the solutions. Good overall agreement between the shock-capturing solutions and experimental data was obtained, except in regions of strong viscous effects or boundary-layer removal. The required centerbody translation with changes in freestream Mach number and angle of attack was obtained and agrees qualitatively with experimental data.

Presley, L. L.↗
NASA NTRS1976

Investigation of slot cooling at high subsonic speeds

The paper discusses the results of an experimental and numerical investigation of tangential slot injection film cooling with zero pressure gradients in subsonic boundary layers at freestream Mach numbers of 0.4, 0.6, and 0.8. The results are compared with the predictions obtained from a finite-difference boundary-layer program developed by NASA for slot injection into turbulent supersonic boundary layers. The two sets of results are found to compare favorably. The numerical results point to the existence of a unique relation between isothermal effectiveness and adiabatic effectiveness, thereby confirming the existence of similarity conditions for temperature and velocity profiles.

Ferri, A.↗
NASA NTRS1976

Influence of nonconservative differencing on transonic streamline shapes

A computer program recently developed by South and Brandt (1976) which contained the Murman (1973) conservative finite-difference scheme is easily modified to use the Garabedian and Korn (1971) nonconservative finite difference scheme. This program solves the transonic small disturbance equation for only symmetric flow, but incorporates several iterative solution techniques. Results are presented for the case where the equally spaced computational grid extended to infinity in both the streamwise and normal directions. Streamline shapes are obtained along several grid lines by a streamwise integration of the normal component of the perturbation velocity. Comparison cases are run for a 10% thick parabolic arc airfoil at zero incidence for freestream Mach numbers of 0, 0.70, 0.84, and 0.95. It is shown that the use of a nonconservative finite-difference scheme in transonic flow calculations destroys the global mass balance when shocks are present. This lack of mass balance may prove to be more crucial in the case of an unconfined external flow.

Newman, P. A.↗
NASA NTRS1976

Computation of viscous transonic flow about a lifting airfoil

The viscous transonic flow about a stationary body in free air was numerically investigated. The geometry chosen was a symmetric NACA 64A010 airfoil at a freestream Mach number of 0.8, a Reynolds number of 4 million based on chord, and angles of attack of 0 and 2 degrees. These conditions were such that, at 2 degrees incidence unsteady periodic motion was calculated along the aft portion of the airfoil and in its wake. Although no unsteady measurements were made for the NACA 64A010 airfoil at these flow conditions, interpolated steady measurements of lift, drag, and surface static pressures compared favorably with corresponding computed time-averaged lift, drag, and surface static pressures.

Walitt, L.↗
NASA NTRS1977

An experimental investigation of shock wave-turbulent boundary layer interactions with and without boundary layer suction: A data summary report

Tabulated data from a series of experimental studies of the interaction of a shock wave with a turbulent boundary layer in axisymmetric flow configurations is presented. The studies were conducted at the walls of circular wind tunnels and on the cylindrical centerbody of an annular flow channel. Detailed pitot pressure profiles and wall static pressure profiles upstream of, within and downstream of the interaction region are given. Results are presented for flows at nominal freestream Mach Numbers of 2, 3 and 4. For studies at the tunnel sidewalls, the shock waves were produced by conical shock generators mounted on the centerline of the wind tunnel at zero angle of attack. The annular ring generator was used to produce the shock wave at the centerbody of the annular flow channel. The effects of boundary layer bleed were examined in the investigation. Both bleed rate and bleed location were studied. Most of the bleed studies were conducted with bleed holes drilled normal to the wall surface but the effects of slot suction were also examined. A summary of the principal results and conclusions is given.

Sun, C. C.↗
NASA NTRS1978

Transonic lifting line theory - Numerical procedure for shock-free flows

An algorithm for computing transonic lifting-line theory without shocks is presented; the numerical procedure relies on two-dimensional analyses developed to solve inviscid flow equations for slender airfoils. From a series of solutions to two-dimensional problems in which span effects appear parametrically, the three-dimensional potential field characterization is obtained. Numerical results are given for a lifting wing with an elliptic (spanwise) distribution of chord and a NACA-0012 cross section (a freestream Mach number of 0.63 and an angle of attack of 2 deg are assumed).

Small, R. D.↗
NASA NTRS1979

Base pressure and heat transfer tests of the 0.0225-scale space shuttle plume simulation model (19-OTS) in yawed flight conditions in the NASA-Lewis 10x10-foot supersonic wind tunnel (test IH83)

Wind tunnel tests were performed to determine pressures, heat transfer rates, and gas recovery temperatures in the base region of a rocket firing model of the space shuttle integrated vehicle during simulated yawed flight conditions. First and second stage flight of the space shuttle were simulated by firing the main engines in conjunction with the SRB rocket motors or only the SSME's into the continuous tunnel airstream. For the correct rocket plume environment, the simulated altitude pressures were halved to maintain the rocket chamber/altitude pressure ratio. Tunnel freestream Mach numbers from 2.2 to 3.5 were simulated over an altitude range of 60 to 130 thousand feet with varying angle of attack, yaw angle, nozzle gimbal angle and SRB chamber pressure. Gas recovery temperature data derived from nine gas temperature probe runs are presented. The model configuration, instrumentation, test procedures, and data reduction are described.

Foust, J. W.↗
NASA NTRS1979

Unsteady thin airfoil theory for transonic flows with embedded shocks

Classical unsteady thin airfoil theory fails for low frequencies at the subsonic freestream Mach number, because of the formation of a shock wave that shields the forward region of the airfoil from aft generated disturbances. In the present paper, the classical thin airfoil theory is modified to account for the presence and induced motion of such shocks. The modification consists of taking the steady local Mach number to be a simple step discontinuity, normal to the undisturbed flow, separating two uniform regions. Predicted regions are shown to correlate well both with the experiment and finite difference calculations.

Williams, M. H.↗
NASA NTRS1979

Refraction of sound by a shear layer - Experimental assessment

An experimental study was conducted to determine the refraction angle and amplitude changes associated with sound transmission through a circular, open jet shear layer. Both on-axis and off-axis acoustic source locations were used. Source frequency varied from 1 kHz to 10 kHz while freestream Mach number varied from 0.1 to 0.4. The experimental results were compared with an existing refraction theory which was extended to account for off-axis source positions. A simple experiment was also conducted to assess the importance of turbulence scattering between 1 kHz and 25 kHz.

Schlinker, R. H.↗
NASA NTRS1980

Experimental investigation of the asymmetric body vortex wake

An experimental investigation of the asymmetric body vortex wake of a circular cylinder in high subsonic flow is presented. Laser velocimeter, force and moment, and surface hot wire measurements were obtained for a freestream Mach number of 0.6 and Reynolds number (based on body diameter) of 0.62 x 10 to the 6th. Two component laser velocimeter measurements were made at three body cross-flow planes, x/d = 4, 8, and 12, and angles of attack of 25, 35, and 45 deg. Laser vapor screen photographs were also obtained at these body stations and angles of attack. Surface hot wire measurements were used to determine if any vortex switching occurred at various angles of attack of the body. The laser velocimeter measurements are related to the vapor screen photographs and side force measurements. These results show that more than one asymmetric body vortex wake configuration can exist for the same angle of attack and body roll angle.

Oberkampf, W. L.↗