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.


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    Title :

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


    Contributors:

    Publication date :

    1959-06-01


    Type of media :

    Miscellaneous


    Type of material :

    No indication


    Language :

    English


    Keywords :



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