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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 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

Wind-tunnel measurements of the chordwise pressure distribution and profile drag of a research airplane model incorporating a 17-percent-thick supercritical wing

The Langley 8-foot transonic pressure tunnel to determine the wing chordwise pressure distribution for a 0.09-scale model of a research airplane incorporating a 17-percent-thick supercritical wing. Airfoil profile drag was determined from wake pressure measurements at the 42-percent-semispan wing station. The investigation was conducted at Mach numbers from 0.30 to 0.80 over an angle-of-attack range sufficient to include buffet onset. The Reynolds number based on the mean geometric chord varied from 2 x 10 to the 6th power at Mach number 0.30 to 3.33 x 10 to the 6th power at Mach number 0.65 and was maintained at a constant value of 3.86 x 10 to the 6th power at Mach numbers from 0.70 to 0.80. Pressure coefficients for four wing semispan stations and wing-section normal-force and pitching-moment coefficients for two semispan stations are presented in tabular form over the Mach number range from 0.30 to 0.80. Plotted chordwise pressure distributions and wake profiles are given for a selected range of section normal-force coefficients over the same Mach number range.

Ferris, J. C.↗
NASA NTRS1929

The Pressure Distribution over the Horizontal and Vertical Tail Surfaces of the F6C-4 Pursuit Airplane in Violent Maneuvers

This investigation of the pressure distribution on the tail surfaces of a pursuit airplane in violent maneuvers was conducted for the purpose of determining the maximum loads likely to be encountered on these surfaces in flight. The information is a part of that needed for a revision of existing loading specifications to bring these into closer agreement with the actual flight conditions. A standard F6C-4 airplane was used and the pressure distribution over the right horizontal and complete vertical tail surfaces was recorded throughout violent maneuvers. The results show that the existing loading specifications do not conform satisfactorily to the loadings existent in critical conditions, and in some cases were exceeded by the loads obtained. An acceleration of 10.5 G. Was recorded in one maneuver in which the pilot suffered severely; it is therefore indicated that the limits of the physical resistance of the pilot to violent maneuvers are being approached. Navy specifications for the structural design of tail surfaces are included as an appendix. (author)

Rhode, R V↗
NASA NTRS1990

Unsteady surface pressure distributions on a delta wing undergoing large amplitude pitching motions

Wind tunnel experiments were performed on a 70-deg-sweep delta wing to determine the effect of a sinusoidal pitching motion on the pressure field on the suction side of the wing. Pressure taps were placed from 35-90 percent of the chord, at 60 percent of the local semi-span. Pressure coefficients were measured as functions of Reynolds number and pitch rate. The surface pressure distribution was seen to vary at the same frequency as the pitching frequency, though distortion due to the vortex breakdown was observed. Comparing the upstroke (angle of attack increasing) and downstroke (angle of attack decreasing) pressures for a specific angle of attack, a time lag in the pressure distribution was observed. The downstroke pressures were slightly larger at the forward chord locations. Vortex breakdown was seen to have the most significant effect at the 40-45-percent chord location, where an increase in local pressure was apparent, as well as a distortion of the periodic pressure fluctuation.

Thompson, S. A.↗
NASA NTRS1952

Pressure distribution and aerodynamic coefficients associated with heat addition to supersonic air stream adjacent to two-dimensional supersonic wing

The modifications in the pressure distributions and the aerodynamic coefficients associated with additions of heat to the two-dimensional supersonic in viscid flow field adjacetnt to the lower surface of of a 5-percent-thickness symmetrical circular-arc wing are presented in this report. The pressure distributions are obtained by the use of graphical method which gives the two-dimensional supersonic inviscid flow field obtained with moderate heat addition. The variation is given of the lift-drag ratio and of the aerodynamic coefficients of lift, drag, and moment with free stream Mach number, angle of attack, and parameters defining extent and amount of heat addition. The six graphical solutions used in this study included Mach numbers of 3.0 and 5.0 and angles of attack of 0 degrees and 2 degrees.

Pinkel, I Irving↗
NASA NTRS1953

Theoretical pressure distribution and wave drags for conical boattails

Afterbody pressure distributions and wave drag were calculated using a second-order theory for a variety of conical boattails at zero angle of attack. Results are presented for Mach numbers from 1.5 to 4.5, area ratios from 0.200 to 0.800, and boattail angle from 3 degrees to 11 degrees. The results indicate that for a given boattail angle, the wave drag decreases with increasing Mach number and area ratio. The wave drag, for a constant area ratio, increases with increasing boattail angle. For a specific Mach number, area ratio, and fineness ratio, a comparison of the wave-drag coefficients for conical, tangent-parabolic, and secant-parabolic boattails showed the conical boattail to have the smallest wave drag.

Jack, John R↗
NASA NTRS1976

A graphical method of determining pressure distribution in two-dimensional flow

By a generalization of the Joukowski method, a procedure is developed for effecting localized modifications of airfoil shapes and for determining graphically the resultant changes in the pressure distribution. The application of the procedure to the determination of the pressure distribution over airfoils of original design is demonstrated. Formulas for the lift, the moment, and the aerodynamic center are also given.

Jones, R. T.↗
NASA NTRS1941

A graphical method of determining pressure distribution in two-dimensional flow

By a generalization of the Joukowski method, a procedure is developed for effecting localized modifications of airfoil shapes and for determining graphically the resultant changes in the pressure distribution. The application of the procedure to the determination of the pressure distribution over airfoils of original design is demonstrated. Formulas for the lift, the moment, and the aerodynamic center are also given.

Jones, Robert T↗
NASA NTRS1949

Experimental Pressure Distributions over Wing Tips at Mach Number 1.9 I : Wing Tip with Subsonic Leading Edge

An investigation was conducted at a Mach number of 1.91 to determine spanwise pressure distribution over a wing tip in a region influenced by a sharp subsonic leading edge swept back at 70 degrees. Except for pressure distribution on the top surface in the immediate vicinity of the subsonic leading edge, the maximum difference between linearized theory and experimental data was 2 1/2 percent (of free-stream dynamic pressure) for angles of attack up to 4 degrees and 7 percent for angles of attack up to 8 degrees. Pressures on the top surface nearest the subsonic edge indicated local expansions beyond values predicted by linearized theory.

FLOW, SUPERSONIC↗
NASA NTRS1989

Wind-tunnel investigation of aerodynamic characteristics and wing pressure distributions of an airplane with variable-sweep wings modified for laminar flow

A wind tunnel test was conducted to evaluate the aerodynamic characteristics and wing pressure distributions of a variable wing sweep aircraft having wing panels that are modified to promote laminar flow. The modified wing section shapes were incorporated over most of the exposed outer wing panel span and were obtained by extending the leading edge and adding thickness to the existing wing upper surface forward of 60 percent chord. Two different wing configurations, one each for Mach numbers 0.7 and 0.8, were tested on the model simultaneously, with one wing configuration on the left side and the other on the right. The tests were conducted at Mach numbers 0.20 to 0.90 for wing sweep angles of 20, 25, 30, and 35 degrees. Longitudinal, lateral and directional aerodynamic characteristics of the modified and baseline configurations, and selected pressure distributions for the modified configurations, are presented in graphical form without analysis. A tabulation of the pressure data for the modified configuration is available as microfiche.

Hallissy, James B.↗
NASA NTRS1959

Chordwise Pressure Distributions over Several NACA 16-Series Airfoils at Transonic Mach Numbers up to 1.25

A two-dimensional wind-tunnel investigation of the pressure distributions over several NACA 16-series airfoils with thicknesses of 4, 6, 9, and 12 percent of the chord and design lift coefficients of 0, 0.2, 1 and 0.5 has been conducted in the Langley airfoil test apparatus at transonic Mach numbers from 0.7 to 1.25. The tests ranged in Reynolds number from 2.4 x 10 (exp 6) to 2.8 x 10 (exp 6) and in angle of attack from -10 to 12 degrees. Chordwise pressure distributions and schlieren flow photographs are presented without analysis.

Ladson, Charles L.↗
NASA NTRS1974

Pressure distribution of a twin-engine upper-surface blown jet-flap model

An investigation has been made to determine the chordwise and spanwise pressure distributions of a small-scale upper-surface blown jet-augmented flap STOL model. The model was powered by two simulated high-bypass-ratio turbofan engines mounted ahead of and above an unswept-untapered wing in a nacelle having a rectangular nozzle. The results of the investigation are presented as tabulated and plotted chordwise pressure distribution coefficients for nine spanwise stations.

Smith, C. C., Jr.↗
NASA NTRS1928

Pressure Distribution over a Wing and Tail Rib of a VE-7 and of a TS Airplane in Flight

This investigation was made to determine the pressure distribution over a rib of the wing and over a rib of the horizontal tail surface of an airplane in flight and to obtain information as to the time correlation of the loads occurring on these ribs. Two airplanes, VE-7 and TS, were selected in order to obtain the information for a thin and a thick wing section. In each case the pressure distribution was recorded for the full range of angle of attack in level flight and throughout violent maneuvers. The results show: (a) that the present rib load specifications in use by the Army Air Corps and the Bureau of Aeronautics, Navy Department, are in fair agreement with the loads actually occurring in flight, but could be slightly improved; (b) that there appears to be no definite sequence in which wing and tail surface ribs reach their respective maximum loads in different maneuvers; (c) that in accelerated flight, at air speeds less than or equal to 60 per cent of the maximum speed, the accelerations measured agree very closely with the theoretically possible maximum accelerations. In maneuvers at higher air speeds the observed accelerations were smaller than those theoretically possible. (author)

Crowley, J W , Jr↗
NASA NTRS1957

Comparison of NACA 65-series compressor-blade pressure distributions and performance in a rotor and in cascade

An investigation has been conducted to compare the performance of NACA 65-series compressor blades in two-dimensional cascade with that in an axial flow compressor. Blade pressure distributions were obtained by the use of a mercury-seal pressure-transfer device. The comparison indicated that cascade data accurately predicted the turning angle and blade pressure distribution obtained in the compressor at design conditions.

CASCADES↗
NASA NTRS1928

Pressure distribution on wing ribs of the VE-7 and TS airplanes in flight Part II : pull-ups

This paper is the second of a series of notes, each of which presents the complete results of pressure distribution tests made at Langley Field by the National Advisory Committee for Aeronautics, on wing and tail ribs of the VE-7 and TS airplanes for a particular maneuver of flight. The results for pull-ups are presented in the form of curves which show the variation of pressure distribution, total loads, normal acceleration and center of pressure with respect to time.

Rhode, R V↗