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Results for “PRESSURE DISTRIBUTION - AIRFOILS”

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 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1937

Pressure-Distribution Measurements on O-2H Airplane in Flight

Results are given of pressure-distribution measurements made over two different horizontal tail surfaces and the right wing cellule, including the slipstream area, of an observation-type biplane. Measurements were also taken of air speed, control-surface positions, control-stick forces, angular velocities, and accelerations during various abrupt maneuvers. These maneuvers consisted of push-downs and pull-ups from level flight, dive pull-outs, and aileron rolls with various thrust conditions. The results from the pressure-distribution measurements over the wing cellule are given on charts showing the variation of individual rib coefficients with wing coefficients; the data from the tail-surface pressure-distribution measurements are given mainly as total loads and moments. These data are supplemented by time histories of the measured quantities and isometric views of the rib pressure distributions occurring in abrupt maneuvers.

Pearson, H A↗
NASA NTRS1989

Large scale advanced propeller blade pressure distributions - Prediction and data

Two Euler analysis techniques, finite difference and finite volume, are employed to predict the blade surface pressure distributions of a large scale advanced propeller. The predicted pressure distributions are compared with wind tunnel data. Both techniques produced blade pressure distributions which are in fairly good agreement with the data over the range of test Mach numbers of 0.2 to 0.78. However, the numerical simulations fail to predict correctly the measured pressure distributions for the low Mach number, high power case which seem to have a leading edge vortex. A discussion of the compressibility effects is also presented.

Nallasamy, M.↗
NASA NTRS1989

Large-scale advanced propeller blade pressure distributions: Prediction and data

Two Euler analyses techniques, finite difference and finite volume, were employed to predict the blade surface pressure distributions of a large scale advanced propeller. The predicted pressure distributions were compared with wind tunnel data. Both techniques produced blade pressure distributions that are in fairly good agreement with the data over the range of test Mach numbers of 0.2 to 0.78. However, the numerical simulations fail to predict correctly the measured pressure distributions for the low Mach number, high power case. The data indicate the presence of a leading edge vortex for this case. A discussion of the compressibility effects is also presented.

Nallasamy, M.↗
NASA NTRS1982

KC-135 wing and winglet flight pressure distributions, loads, and wing deflection results with some wind tunnel comparisons

A full-scale winglet flight test on a KC-135 airplane with an upper winglet was conducted. Data were taken at Mach numbers from 0.70 to 0.82 at altitudes from 34,000 feet to 39,000 feet at stabilized flight conditions for wing/winglet configurations of basic wing tip, 15/-4 deg, 15/-2 deg, and 0/-4 deg winglet cant/incidence. An analysis of selected pressure distribution and data showed that with the basic wing tip, the flight and wind tunnel wing pressure distribution data showed good agreement. With winglets installed, the effects on the wing pressure distribution were mainly near the tip. Also, the flight and wind tunnel winglet pressure distributions had some significant differences primarily due to the oilcanning in flight. However, in general, the agreement was good. For the winglet cant and incidence configuration presented, the incidence had the largest effect on the winglet pressure distributions. The incremental flight wing deflection data showed that the semispan wind tunnel model did a reasonable job of simulating the aeroelastic effects at the wing tip. The flight loads data showed good agreement with predictions at the design point and also substantiated the predicted structural penalty (load increase) of the 15 deg cant/-2 deg incidence winglet configuration.

Montoya, L. C.↗
NASA NTRS1982

Circumferential pressure distributions in a model labyrinth seal

A research program to isolate and study leakage flow through labyrinth glands was initiated. Circumferential pressure distributions were measured in the labyrinth glands with geometry appropriate to the high pressure labyrinths in large steam turbines. Knowledge of this pressure distribution is essential as it is this unequal pressure field that results in the destabilizing force. Parameters that are likely to affect the pressure distributions are incorporated into the test rig. Some preliminary pressure profiles are presented.

Leong, Y. M. M. S.↗
NASA NTRS1975

A Method for Calculating the Induced Pressure Distribution Associated with a Jet in a Crossflow

A model is presented which can be used to study the loss of lift during hovering and horizontal flight of the VTOL aircraft. The model numerically predicts the pressure distribution induced by a round, turbulent, unheated, subsonic jet exhausting normally through a flat plate into a subsonic crossflow. The complete model assumes that the predominant features of the flow are jet entrainment and a pair of contrarotating vortices which form downstream of the jet. Experimentally determined vortex properties and a reasonable assumption concerning jet entrainment were used. Potential flow considerations were used except in the wake region, where a simple method for approximating the pressure distribution was suggested. The calculated pressure distribution, lift, and pitching moments on the flat plate are presented for a jet to crossflow velocity ratio of 8 and were compared with experimental results. A computer program is given which was used to calculate the pressure distribution across the flat plate.

Dietz, W. E., Jr.↗
NASA NTRS1973

Measured pressure distributions of large-angle cones in hypersonic flows of tetrafluoromethane, air, and helium

An experimental study of surface pressure distributions on a family of blunt and sharp large angle cones was made in hypersonic flows of helium, air, and tetrafluoromethane. The effective isentropic exponents of these flows were 1.67, 1.40, and 1.12. Thus, the effect of large shock density ratios such as might be encountered during planetary entry because of real-gas effects could be studied by comparing results in tetrafluoromethane with those in air and helium. It was found that shock density ratio had a large effect on both shock shape and pressure distribution. The differences in pressure distribution indicate that for atmospheric flight at high speed where real-gas effects produce large shock density ratios, large-angle cone vehicles can be expected to experience different trim angles of attack, drag coefficient, and lift-drag ratios than those for ground tests in air wind tunnels.

Jones, R. A.↗
NASA NTRS1948

Surface-pressure Distributions on a Systematic Group of NACA 1-series Cowlings with and Without Spinners

A method for calculating the flow fields of axially symmetric bodies from their pressure distributions is reported in NACA RM No. L8I17. In order to facilitate application of this method to the important case of the cowling-spinner combination, for use in the design of propellers, the present paper presents static-pressure distributions on the tops of 79 high-critical-speed NACA 1-series cowling-spinner combinations over wide ranges of inlet-velocity ratio at angles of attack of 0 degrees, 2 degrees, 4 degrees, and 6 degrees. Static-pressure distributions around the nose sections of several cowlings are given in greater detail to aid in estimating the pressures near the stagnation points and to show the effect of changes in the internal lip shape. The effects of the operation of a typical propeller on the surface pressures on the cowling are shown for one configuration. The pressure distributions over the nine NACA 1-series nose inlets used as the basic components of these combinations are also presented ro supplement the existing open-nose-cowling data of NACA ACR No. L5F30a which are applicable to the case of the rotating cowling.

Boswinkle, Robert W JR↗
NASA NTRS1924

Remarks on the Pressure Distribution over the Surface of an Ellipsoid, Moving Translationally Through a Perfect Fluid

This note, prepared for the National Advisory Committee for Aeronautics, contains a discussion of the pressure distribution over ellipsoids when in translatory motion through a perfect fluid. An easy and convenient way to determine the magnitude of the velocity and of the pressure at each point of the surface of an ellipsoid of rotation is described. The knowledge of such pressure distribution is of great practical value for the airship designer. The pressure distribution over the nose of an airship hull is known to be in such good agreement with the theoretical distribution as to permit basing the computation of the nose stiffening structure on the theoretical distribution of pressure.

Munk, Max M↗
NASA NTRS1985

Detailed pressure distribution measurements obtained on several configurations of an aspect-ratio-7 variable twist wing

Detailed pressure distribution measurements were made for 11 twist configurations of a unique, multisegmented wing model having an aspect ratio of 7 and a taper ratio of 1. These configurations encompassed span loads ranging from that of an untwisted wing to simple flapped wings both with and without upper-surface spoilers attached. For each of the wing twist configurations, electronic scanning pressure transducers were used to obtain 580 surface pressure measurements over the wing in about 0.1 sec. Integrated pressure distribution measurements compared favorably with force-balance measurements of lift on the model when the model centerbody lift was included. Complete plots and tabulations of the pressure distribution data for each wing twist configuration are provided.

Holbrook, G. T.↗
NASA NTRS1987

Experimental cavity pressure distributions at supersonic speeds

An investigation was conducted to define pressure distributions for rectangular cavities over a range of free-stream Mach numbers and cavity dimensions. These pressure distributions together with schlieren photographs are used to define the critical values of cavity length-to-depth ratio that separate open type cavity flows from closed type cavity flows. For closed type cavity flow, the shear layer expands over the cavity leading edge and impinges on the cavity floor, whereas for open type cavity flow, the shear layer bridges the cavity. The tests were conducted by using a flat-plate model permitting the cavity length to be remotely varied from 0.5 to 12 in. Cavity depths and widths were varied from 0.5 to 2.5 in. The flat-plate boundary layer approaching the cavity was turbulent and had a thickness of approximately 0.2 in. at the cavity front face for the range of test Mach numbers from 1.5 to 2.86. Presented are a discussion of the results and a complete tabulation of the experimental data.

Stallings, Robert L., Jr.↗
NASA NTRS1979

Experimental studies on the effects of a sting support on the pressure distribution around a spherical object

Experiments were conducted on a spherical object, 2.5 inches in diameter, to obtain the pressure distribution around its meridan plane. In most of the tests the sphere was provided with a tail consisting of a circular cylinder that was attached directly to the rear with its axis in alignment with the center of the sphere. In some tests the tail was removed and the sphere alone was tested for comparison purposes. The main object of the tests was to obtain information on tail interference with the pressure distribution. The results of the tests show that the pressure distribution was affected by the presence of the tail to a minor extent only, while major differences occurred with the variation of the Reynolds number. The experiments were performed both in an open as well as inside a closed wind tunnel under steady flow conditions at Reynolds numbers ranging from 0.91 to 2.6 x 100000.

Barna, P. S.↗
NASA NTRS1983

Subsonic panel method for designing wing surfaces from pressure distribution

An iterative method has been developed for designing wing section contours corresponding to a prescribed subcritical distribution of pressure. The calculations are initialized by using a surface panel method to analyze a baseline wing or wing-fuselage configuration. A first-order expansion to the baseline panel method equations is then used to calculate a matrix containing the partial derivative of potential at each control point with respect to each unknown geometry parameter. In every iteration cycle, the matrix is used both to calculate the geometry perturbation and to analyze the perturbed geometry. The distribution of potential on the perturbed geometry is established by simple linear extrapolation from the baseline solution. The extrapolated potential is converted to pressure by Bernoulli's equation. Not only is the accuracy of the approach good for very large perturbations, but the computing cost of each complete iteration cycle is substantially less than one analysis solution by a conventional panel method.

Bristow, D. R.↗
NASA NTRS1972

Subsonic pressure distributions around a solid model of an inflatable decelerator attached to the base of an ogive-cylinder

A wind-tunnel investigation was conducted at free-stream Mach numbers from 0.20 to 1.00 and corresponding Reynolds numbers, based on maximum afterbody diameter, from 2.25 x one million to 6.90 x one million on a solid model of an attached inflatable decelerator (AID) connected to the base of an ogive-cylinder. Tests were conducted to obtain ram-air and surface pressure distributions about the AID. AID shapes derived for subsonic deployment are dependent on the pressure distributions used in their derivation, and the different shapes obtained are dependent on the Mach number for which the design is made. The resulting pressure distributions were used in a design program to obtain new shapes which were compared with the original pressure-distribution shape.

Sawyer, J. W.↗
NASA NTRS1989

Verification of spatial and temporal pressure distributions in segmented solid rocket motors

A wide variety of analytical tools are in use today to predict the history and spatial distributions of pressure in the combustion chambers of solid rocket motors (SRMs). Experimental and analytical methods are presented here that allow the verification of many of these predictions. These methods are applied to the redesigned space shuttle booster (RSRM). Girth strain-gage data is compared to the predictions of various one-dimensional quasisteady analyses in order to verify the axial drop in motor static pressure during ignition transients as well as quasisteady motor operation. The results of previous modeling of radial flows in the bore, slots, and around grain overhangs are supported by approximate analytical and empirical techniques presented here. The predictions of circumferential flows induced by inhibitor asymmetries, nozzle vectoring, and propellant slump are compared to each other and to subscale cold air and water tunnel measurements to ascertain their validity.

Salita, Mark↗
NASA NTRS1947

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.↗
NASA NTRS1928

The Measurement of Pressure Through Tubes in Pressure Distribution Tests

The tests described in this report were made to determine the error caused by using small tubes to connect orifices on the surface of aircraft to central pressure capsules in making pressure distribution tests. Aluminum tubes of 3/16-inch inside diameter were used to determine this error. Lengths from 20 feet to 226 feet and pressures whose maxima varied from 2 inches to 140 inches of water were used. Single-pressure impulses for which the time of rise of pressure from zero to a maximum varied from 0.25 second to 3 seconds were investigated. The results show that the pressure recorded at the capsule on the far end of the tube lags behind the pressure at the orifice end and experiences also a change in magnitude. For the values used in these tests the time lag and pressure change vary principally with the time of rise of pressure from zero to a maximum and the tube length. Curves are constructed showing the time lag and pressure change. Empirical formulas are also given for computing the time lag. Analysis of pressure distribution tests made on airplanes in flight shows that the recorded pressures are slightly higher than the pressures at the orifice and that the time lag is negligible. The apparent increase in pressure is usually within the experimental error, but in the case of the modern pursuit type of airplane the pressure increase may be 5 per cent. For pressure-distribution tests on airships the analysis shows that the time lag and pressure change may be neglected.

Hemke, Paul E↗