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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 NTRS1938

Pressure distribution over an NACA 23012 airfoil with an NACA 23012 external-airfoil flap

Report presents the results of pressure-distribution tests of an NACA 23012 airfoil with an NACA 23012 external airfoil flap made in the 7 by 10-foot wind tunnel. The pressures were measured on the upper and lower surfaces at one chord section on both the main airfoil and on the flap for several different flap deflections and at several angles of attack. A test installation was used in which the airfoil was mounted horizontally in the wind tunnel between vertical end planes so that two-dimensional flow was approximated. The data are presented in the form of pressure-distribution diagrams and as graphs of calculated coefficients for the airfoil-and-flap combination and for the flap alone.

Wenzinger, Carl J↗
NASA NTRS1938

Pressure distribution over an NACA 23012 airfoil with a slotted and a plain flap

Report presents the results of pressure-distribution of an NACA 23012 airfoil equipped with a slotted flap and with a plain flap conducted in the 7 by 10-foot wind tunnel. A test installation was used in which the 7-foot-span airfoil was mounted vertically between the upper and lower sides of the closed test section so that two-dimensional flow was approximated. The pressures were measured on the upper and lower surfaces at one chord section both on the main airfoil and on the flaps for several different flap deflections and at several angles of attack. The data are presented in the form of pressure-distribution diagrams and as graphs of calculated section coefficients for the airfoil-and-flap combinations and also for the flaps alone. The results are useful for application to rib and flap structural design; in addition, the plain-flap data furnish considerable information applicable to the structural design of plain ailerons.

Wenzinger, Carl J↗
NASA NTRS1977

An investigation to determine the static pressure distribution of the 0.00548 scale shuttle solid rocket booster (MSFC model number 468) during reentry in the NASA/MSFC 14 inch trisonic wind tunnel (SA28F)

The results of a pressure test of a .00548 scale 146 inch Space Shuttle Solid Rocket Booster (SRB) with and without protuberances, conducted in a 14 x 14 inch trisonic wind tunnel are presented. Static pressure distributions for the SRB at reentry attitudes and flight conditions were obtained. Local longitudinal and ring pressure distributions are presented in tabulated form. Integration of the pressure data was performed. The test was conducted at Mach numbers of 0.40 to 4.45 over an angle of attack range from 60 to 185 degrees. Roll angles of 0, 45, 90 and 315 degrees were investigated. Reynolds numbers per foot varied for selected Mach numbers.

Braddock, W. F.↗
NASA NTRS1987

A fast approach to designing airfoils from given pressure distribution in compressible flows

A new inverse method for aerodynamic design of airfols is presented for subcritical flows. The pressure distribution in this method can be prescribed as a function of the arc length of the as-yet unknown body. This inverse problem is shown to be mathematically equivalent to solving only one nonlinear boundary value problem subject to known Dirichlet data on the boundary. The solution to this problem determines the airfoil, the freestream Mach number, and the upstream flow direction. The existence of a solution to a given pressure distribution is discussed. The method is easy to implement and extremely efficient. A series of results for which comparisons are made with the known airfoils is presented.

Daripa, Prabir↗
NASA NTRS1974

Heat-transfer and pressure distributions for laminar separated flows downstream of rearward-facing steps with and without mass suction

Heat-transfer and pressure distributions were measured for laminar separated flows downstream of rearward-facing steps with and without mass suction. The flow conditions were such that the boundary-layer thickness was comparable to or larger than the step height. For both suction and no-suction cases, an increase in the step height resulted in a sharp decrease in the initial heat-transfer rates behind the step. Downstream, however, the heat transfer gradually recovered back to less than or near attached-flow values. Mass suction from the step base area increased the local heat-transfer rates; however, this effect was relatively weak for the laminar flows considered. Even removal of the entire approaching boundary layer raised the post-step heat-transfer rates only about 10 percent above the flatplate values. Post-step pressure distributions were found to depend on the entrainment conditions at separation. In the case of the solid-faced step, a sharp pressure drop behind the step was followed by a very short plateau and relatively fast recompression. For the slotted-step connected to a large plenum but without suction, the pressure drop at the base was much smaller and the downstream recompression more gradual than that for solid-faced step.

Brown, R. D.↗
NASA NTRS1972

Heat transfer and pressure distributions at M equals 8 on 0.029 scale models of the Viking entry vehicle

An investigation in the Langley Research Center Mach-8 Variable Density Hypersonic Tunnel was made of the pressure distributions and heat transfer rate distributions on two 0.029 scale Viking Entry Vehicle models. Comparable ranges of test Reynolds number were exercised for the two tests between run conditions around 4 million and conditions of about 1.6 million. At angles of attack less than 20 degrees the pressure ratio distribution referenced to stagnation pressure appeared invariant with Reynolds number. Increasing angle of attack results in a flatter distribution of both the windward and leeward pressure distributions; in addition, the stagnation point shifted into the windward plane. A subsequent rise in the heating rate profile on the leeward side with further increase in angle of attack is attributed to boundary layer natural transition to turbulent flow. Schlieren photographs were taken for flow field visualization and to correct model angle of attack.

Faye-Petersen, R.↗
NASA NTRS1953

Method of Estimating the Incompressible-flow Pressure Distribution of Compressor Blade Sections at Design Angle of Attack

A method was devised for estimating the incompressible-flow pressure distribution over compressor blade sections at design angle of attack. The theoretical incremental velocities due to camber and thickness of the section as an isolated airfoil are assumed proportional to the average passage velocity and are modified by empirically determined interference factors. Comparisons were made between estimated and test pressure distributions of NACA 65-series sections for typical conditions. Good agreement was obtained.

COMPRESSOR FLOW THEORY AND EXPERIMENT↗
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↗