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Results for “MACH NUMBER EFFECTS - PROPELLERS”

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 361 records · Page 20
All sourcesNASA NTRSDOE OSTI
NASA NTRS1957

Base Pressure at Supersonic Speeds on Two-dimensional Airfoils and on Bodies of Revolution with and Without Fins Having Turbulent Boundary Layers

An analysis has been made of available experimental data to show the effects of most of the variables that are more predominant in determining base pressure at supersonic speeds. The analysis covers base pressures for two-dimensional airfoils and for bodies of revolution with and without stabilizing fins and is restricted to turbulent boundary layers. The present status of available experimental information is summarized as are the existing methods for predicting base pressure. A simple semiempirical method is presented for estimating base pressure. For two-dimensional bases, this method stems from an analogy established between the base-pressure phenomena and the peak pressure rise associated with the separation of the boundary layer. An analysis made for axially symmetric flow indicates that the base pressure for bodies of revolution is subject to the same analogy. Based upon the methods presented, estimations are made of such effects as Mach number, angle of attack, boattailing, fineness ratio, and fins. These estimations give fair predictions of experimental results. (author)

LOVE EUGENE S↗
NASA NTRS1956

A Factor Affecting Transonic Leading-edge Flow Separation

A change in flow pattern that was observed as the free-stream Mach number was increased in the vicinity of 0.8 was described in NACA Technical Note 1211 by Lindsey, Daley, and Humphreys. The flow on the upper surface behind the leading edge of an airfoil at an angle of attack changed abruptly from detached flow with an extensive region of separation to attached supersonic flow terminated by a shock wave. In the present paper, the consequences of shock-wave - boundary layer interaction are proposed as a factor that may be important in determining the conditions under which the change in flow pattern occurs. Some experimental evidence in support of the importance of this factor is presented.

FLOW, TURBULENT↗
NASA NTRS1957

Thin airfoil theory based on approximate solution of the transonic flow equation

A method is presented for the approximate solution of the nonlinear equations transonic flow theory. Solutions are found for two-dimensional flows at a Mach number of 1 and for purely subsonic and purely supersonic flows. Results are obtained in closed analytic form for a large and significant class of nonlifting airfoils. At a Mach number of 1 general expressions are given for the pressure distribution on an airfoil of specified geometry and for the shape of an airfoil having a prescribed pressure distribution. Extensive comparisons are made with available data, particularly for a Mach number of 1, and with existing solutions.

WING-SECTION THEORY↗
NASA NTRS1957

Comparison of calculated and experimental load distributions on thin wings at high subsonic and sonic speeds

A method for calculating the aerodynamic loading on a wing in combination with a body is presented. Calculated results are compared with experimentally measured data for two wing-body configurations throughout a range of Mach number up to 1.0. The magnitude and the distribution of spanwise loading of the calculated data are generally in good agreement with the experimental data.

WINGS, COMPLETE - THEORY↗
NASA NTRS1958

Compilation of Information on the Transonic Attachment of Flows at the Leading Edges of Airfoils

Schlieren photographs have been compiled of the two-dimensional flow at transonic speeds past 37 airfoils. These airfoils have variously shaped profiles, and some are related in thickness and camber. The data for these airfoils were analyzed to provide basic information on the flow changes involved and to determine factors affecting transonic-flow attachment, which is a transition from separated to unseparated flow at the leading edges of two-dimensional airfoils at fixed angles as the subsonic Mach number is increased.

FLOW, MIXED↗
NASA NTRS1949

Investigation of a Thin Straight Wing of Aspect Ratio 4 by the NACA Wing-flow Method : Lift and Pitching-moment Characteristics of the Wing Alone

This report presents measurements of the lift and pitching-moment characteristics of a straight wing of aspect ratio 4, taper ratio 0.5, having a symmetrical double-wedge airfoil section with a maximum thickness of 4.4-percent chord. The tests were conducted in the Mach number range 0.51 to 1.20, Reynolds numbers 380,000 to 660,000, by the NACA wing-flow method. The results are compared with theory and with wind-tunnel tests of a similar model. It is indicated that in the Mach number range 0.82 to 1.00 the model surface, profile, and test Reynolds number all would be very important considerations in any attempt to study or predict full-scale characteristics from small-model tests.

MACH NUMBER EFFECTS - COMPLETE WINGS↗
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 NTRS1948

An Investigation of Aileron Oscillations at Transonic Speeds on NACA 23012 and NACA 65-212 Airfoils by the Wing-flow Method

An investigation is being conducted to determine the feasibility of studying aileron buzz by means of the wing-flow method. Two semispan models which had an aspect ratio of 6 and a taper ratio of 2 with quarter-chord half-span mass-balanced ailerons have been used. One had an NACA 23012 airfoil section and the second, an NACA 65-212 airfoil section. The ailerons on both models were subject to buzz over a small range of Mach number near 0.9. Data obtained by wing-flow testing agreed reasonably well with full-scale flight results.

MACH NUMBER EFFECTS - COMPLETE WINGS↗