Jet-plume-induced flow separation on a lifting entry body at Mach numbers from 4.00 to 6.00
Jet plume induced, flow separation on lifting entry model flat bottom by pressure distribution analysis
Mc Ghee, R. J.↗
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Computer simulation of hypersonic flow over the Space Shuttle Orbiter
Computer simulations of the flow field around the Space Shuttle Orbiter are described. Results of inviscid calculations are presented for the shock wave pattern and bottom centerline pressure distribution at 30 deg angle of attack. Results of viscous calculations are presented for wall pressure and heat transfer distributions for simple configurations representative of regions where shock wave-boundary layer interactions occur. The computer codes are verified by comparisons with wind-tunnel data and can be applied to flight conditions.
Inouye, M.↗
A semiempirical procedure for computing the water-pressure distribution on flat and v-bottom prismatic surfaces during impact or planing
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Smiley, Robert F↗
Three-dimensional fully spectral numerical method for mantle convection with depth-dependent properties
A semi-implicit fully spectral collocation method for the simulation of three-dimensional mantle convection with depth-dependent thermo-dynamic and transport properties is presented. The variable property Navier-Stokes equation expressed in terms of the primitive variable velocity and pressure is solved with the mass continuity and temperature equations. The periodic horizontal boundary conditions allow a Fourier expansion for the two horizontal directions. The stress-free, impermeable isothermal boundary conditions along with the depth dependent coefficients are handled with a Chebyshev expansion in the vertical direction. In the limit of an infinite Prandtl number appropriate to mantle convection, the inertial terms in the momentum equation are unimportant. In this case an explicit solution of a Poisson equation for pressure can be avoided; instead a fourth-order equation for vertical velocity can be solved. Simultaneous imposition of both impermeable and continuity boundary conditions during the vertical velocity evaluation is discussed. The pressure distributions on the top and bottom bounding planes were determined by means of an influence matrix technique. The numerical method employed here avoids time-splitting errors and enforces velocity boundary conditions and continuity over the entire domain, including the boundaries, to machine accuracy. Strongly time-dependent three-dimensional solutions up to a surface Rayleigh number of 1 x 10(exp 7) have been obtained. Strong upwellings, pulsating chaotically, are formed by the collective merging of cylindrical plumes.
Balachandar, S.↗
Distribution of pressure on fuselage of airplane model: communication from Rijks-Studiedienst voor de Luchtvaart of Amsterdam
In order to study the distribution of the pressure on the surfaces of a fuselage and the influence of the wing on the air flow along these surfaces, we have made tests pertaining to the bottom and one side.
PRESSURE DISTRIBUTION - FUSELAGES↗
Altitude Cooling Investigation of the R-2800-21 Engine in the P-47G Airplane. IV - Engine Cooling-Air Pressure Distribution
A study of the data obtained in a flight investigation of an R-2800-21 engine in a P-47G airplane was made to determine the effect of the flight variables on the engine cooling-air pressure distribution. The investigation consisted of level flights at altitudes from 5000 to 35,000 feet for the normal range of engine and airplane operation. The data showed that the average engine front pressures ranged from 0.73 to 0.82 of the impact pressure (velocity head). The average engine rear pressures ranged from 0.50 to 0.55 of the impact pressure for closed cowl flaps and from 0.10 to 0.20 for full-open cowl flaps. In general, the highest front pressures were obtained at the bottom of the engine. The rear pressures for the rear-row cylinders were .lower and the pressure drops correspondingly higher than for the front-row cylinders. The rear-pressure distribution was materially affected by cowl-flap position in that the differences between the rear pressures of the front-row and rear-row cylinders markedly increased as the cowl flaps were opened. For full-open cowl flaps, the pressure drops across the rear-row cylinders were in the order of 0.2 of the impact pressure greater than across the front-row cylinders. Propeller speed and altitude had little effect on the -coolingair pressure distribution, Increase in angle of inclination of the thrust axis decreased the front ?pressures for the cylinders at the top of the engine and increased them for the cylinders at the bottom of the engine. As more auxiliary air was taken from the engine cowling, the front pressures and, to a lesser extent, the rear pressures for the cylinders at the bottom of the engine decreased. No correlation existed between the cooling-air pressure-drop distribution and the cylinder-temperature distribution.
Kaufman, Samuel J.↗
Analysis of Experimental Investigations of the Planing Process of the Surface of Water
Pressure distribution and spray measurements were carried out on rectangular flat and V-bottom planing surfaces. Lift, resistance, and center of pressure data are analyzed and it is shown how these values may be computed for the pure planing procees of a flat or V-bottom suface of arbitrary beam, load and speed, the method being illustrated with the aid of an example.
Sottorf, W.↗
TWINTAN: A program for transonic wall interference assessment in two-dimensional wind tunnels
A method for assessing the wall interference in transonic two dimensional wind tunnel test was developed and implemented in a computer program. The method involves three successive solutions of the transonic small disturbance potential equation to define the wind tunnel flow, the perturbation attriburable to the model, and the equivalent free air flow around the model. Input includes pressure distributions on the model and along the top and bottom tunnel walls which are used as boundary conditions for the wind tunnel flow. The wall induced perturbation fields is determined as the difference between the perturbation in the tunnel flow solution and the perturbation attributable to the model. The methodology used in the program is described and detailed descriptions of the computer program input and output are presented. Input and output for a sample case are given.
Kemp, W. B., Jr.↗
Passive shock wave/boundary layer control for transonic airfoil drag reduction
An investigation of the passive shock wave/boundary layer interaction control for reducing the drag in 12 percent thick circular arc and 14 percent thick supercritical airfoils was conducted in a 3 in. x 15.4 in. transonic wind tunnel at transonic Mach numbers. A porous surface with a cavity beneath it was positioned on the area of the airfoils, mounted on the test section bottom wall, where the shock wave occurs. The static pressure distributions over the airfoil, the wake impact pressure data for determining the profile drag, and the Schlieren photographs for porous surface airfoils are presented and compared with the results for solid surface airfoils. With the porous surface the normal shock wave for solid surface was changed to a lambda shock wave system, and the wake impact pressure data indicated an appreciable drag reduction for both airfoils with the porous surface at transonic speeds while causing little or no loss of lift. The effect of porosity and cavity size is investigated and off-design performance is discussed.
Bahi, L.↗
TWINTN4: A program for transonic four-wall interference assessment in two-dimensional wind tunnels
A method for assessing the wall interference in transonic two-dimensional wind tunnel tests including the effects of the tunnel sidewall boundary layer was developed and implemented in a computer program named TWINTN4. The method involves three successive solutions of the transonic small disturbance potential equation to define the wind tunnel flow, the equivalent free air flow around the model, and the perturbation attributable to the model. Required input includes pressure distributions on the model and along the top and bottom tunnel walls which are used as boundary conditions for the wind tunnel flow. The wall-induced perturbation field is determined as the difference between the perturbation in the tunnel flow solution and the perturbation attributable to the model. The methodology used in the program is described and detailed descriptions of the computer program input and output are presented. Input and output for a sample case are given.
Kemp, W. B., Jr.↗
Porosity effect on supercritical airfoil drag reduction by shock wave/boundary layer control
An investigation of the passive shock wave/boundary layer control for reducing the drag of 14 percent-thick supercritical airfoil was conducted in the 3 in. x 15.4 in. RPI Transonic Wind Tunnel at transonic Mach numbers. Various porous surfaces with a cavity beneath it was positioned on the area of the airfoil, mounted on the test section bottom wall, where the shock wave occurs. The static pressure distributions over the airfoil, the wake impact pressure survey for determining the profile drag and the Schlieren photographs for porous surfaces are presented and compared with the results for solid surface airfoil. With a uniform porosity surface the normal shock wave for solid surface was changed to a lambda shock wave, and the wake impact pressure data indicated an appreciable drag reduction at transonic Mach numbers. For a free stream Mach number of 0.81 the profile drag coefficient for the airfoil top surface with uniform porosity was 46 percent lower than for the solid surface airfoil.
Nagamatsu, H. T.↗
Calibration of the 13- by 13-inch adaptive wall test section for the Langley 0.3-meter transonic cryogenic tunnel
A 13 by 13 inch adaptive wall test section was installed in the 0.3 Meter Transonic Cryogenic Tunnel circuit. This new test section is configured for 2-D airfoil testing. It has four solid walls. The top and bottom walls are flexible and movable whereas the sidewalls are rigid and fixed. The wall adaptation strategy employed requires the test section wall shapes associated with uniform test section Mach number distributions. Calibration tests with the test section empty were conducted with the top and bottom walls linearly diverged to approach a uniform Mach number distribution. Pressure distributions were measured in the contraction cone, the test section, and the high speed diffuser at Mach numbers from 0.20 to 0.95 and Reynolds numbers from 10 to 100 x 10 (exp 6)/per foot.
Mineck, Raymond E.↗
Distribution of wave drag and lift in the vicinity of wing tips at supersonic speeds
The point-source-distribution method of calculating the aerodynamic coefficients of thin wings at supersonic speeds was extended to include the effect of the region between the wing boundary and the foremost Mach wave from the wing leading edge. The effect of this region on the surface velocity potential has been determined by an equivalent function, which is evaluated over a portion of the wing surface. In this manner, the effect of angles of attack and yaw as well as the asymmetry of top and bottom wing surfaces may be calculated. As examples of the method, the pressure distribution on a thin plate wing of rectangular plan form as well as the lift and the drag coefficients as a function of Mach number, angle of attack, and aspect ratio are calculated. The equations for the surface velocity potential of several other plan forms are also included. (author)
AIRFOILS, FLAT-PLATE↗
A simplified fourwall interference assessment procedure for airfoil data obtained in the Langley 0.3-meter transonic cryogenic tunnel
A simplified fourwall interference assessment method has been described, and a computer program developed to facilitate correction of the airfoil data obtained in the Langley 0.3-m Transonic Cryogenic Tunnel (TCT). The procedure adopted is to first apply a blockage correction due to sidewall boundary-layer effects by various methods. The sidewall boundary-layer corrected data are then used to calculate the top and bottom wall interference effects by the method of Capallier, Chevallier and Bouinol, using the measured wall pressure distribution and the model force coefficients. The interference corrections obtained by the present method have been compared with other methods and found to give good agreement for the experimental data obtained in the TCT with slotted top and bottom walls.
Murthy, A. V.↗
An inerference assessment approach for a three-dimensional slotted tunnel with sparse wall pressure data
The various procedures referred to as wall interference assessment and correction procedures presume the existence of a surface distribution of data (usually static pressure) measured over a surface on or near the tunnel walls for each test point to be assessed. An alternative approach in which a reasonably sophisticated computer model of the test section flow would be fitted parametrically to a sparse set of measured data is presented. The measurements provides line distributions of static pressure near the center lines of the top, side and bottom walls. The development of a test section model incorporating explicit recognition of discrete slots of finite length with controlled flow reentry into the solid wall downstream portion of the tunnel is shown.
Kemp, W. B., Jr.↗
Atmospheric tides on Venus. III - The planetary boundary layer
Diurnal solar heating of Venus' surface produces variable temperatures, winds, and pressure gradients within a shallow layer at the bottom of the atmosphere. The corresponding asymmetric mass distribution experiences a tidal torque tending to maintain Venus' slow retrograde rotation. It is shown that including viscosity in the boundary layer does not materially affect the balance of torques. On the other hand, friction between the air and ground can reduce the predicted wind speeds from about 5 to about 1 m/sec in the lower atmosphere, more consistent with the observations from Venus landers and descent probes. Implications for aeolian activity on Venus' surface and for future missions are discussed.
Dobrovolskis, A. R.↗
Computation of imaginary-side pressure distributions over the flexible walls of the test section insert for the 0.3-M transonic cryogenic tunnel
Two dimensional airfoil testing in an adaptive wall test-section wind tunnel requires the computation of the imaginary flow fields extending outward from the top and bottom test section walls. A computer program was developed to compute the flow field which would be associated with an arbitrary test section wall shape. The program is based on incompressible flow theory with a Prandtl-Glauert compressibility correction. The program was validated by comparing the streamline and the pressure field generated by a source in uniform flow with the results from the computer program. A listing of the program, the validation test results, and a sample program are included.
Goodyer, M. J.↗
Interpretation of waverider performance data using computational fluid dynamics
A computational study was conducted to better understand experimental results obtained from wind tunnel tests of a Mach 4 waverider model and a comparative reference configuration. The experimental results showed that the performance of the reference configuration was slightly better than that of the waverider model. These results contradict waverider design theory, which suggests that a waverider optimized for maximum lift-to-drag should provide better performance than any other non-waverider configuration at a given design point, especially at hypersonic speeds. The computational results showed that the predicted surface pressure values and the integrated lift and drag coefficients from the pressure distributions were much lower for the reference model than for the flat-top model, due to the reference model bottom surface having a slight expansion. The lift-to-drag ratios for the flat-top model were higher due to a relatively low drag for the same amount of lift. These results indicate that the performance advantage of the reference model was due to the shape of the bottom surface and not due to the flat top surface. The results also showed that the reference model exhibited the same shock attachment characteristics as the waverider because the planform shapes were identical. CFD predictions show that the planform shape gives the waverider an advantage in performance over conventional hypersonic vehicles and that altering the bottom surface of a waverider does not cause significant performance degradation.
Cockrell, Charles E., Jr.↗