"The flow of a compressible fluid past a sphere fixed in a uniform stream is calculated to the third order of approximation by means of the Janzen-Rayleigh method. The velocity and the pressure distribution over the surface of the sphere are computed and the terms involving the fourth power of the Mach number, neglected in Rayleigh's calculation, are shown to be of considerable importance as the local velocity of sound is approached on the sphere. The critical Mach number, that is, the value of the Mach number at which the maximum velocity of the fluid past the sphere is just equal to the local velocity of sound, is calculated for both the second and the third approximation and is found to be, respectively, Mcr=0.587 and Mcr=0.573" (p. 1).

Pressure-distribution tests of an NACA 0009 airfoil with an 80-percent-chord plain flap and three plain tabs, having chord of 10, 20, and 30 percent of the flap chord, were made. Section data suitable for application to the design of horizontal and vertical tail surfaces were obtained. Resultant-pressure diagrams for the airfoil with the flap and the 20-percent-chord tab are presented. Plots are also given of increments of normal-force and hinge-moment coefficients for the airfoil, the flap, and the three tabs. A comparison of some characteristic slopes for the 30-, the 50-, and the 80-percent-chord flaps, tested in the general investigation of plain flaps for control surfaces, is included. Section aerodynamic and load data have been made available for a wide range of flap and a tab chords to be used on an NACA 0009 airfoil or on other conventional sections.

From Summary: "Pressure-distribution tests of an NACA 0009 airfoil with a 30-percent-chord plain flap and three plain tabs, having chords 10, 20, and 30 percent of the flap chord, were made. The purpose of these tests was to continue an investigation to supply structural and aerodynamic section data that may be applied to the design of horizontal and vertical tail surfaces. The results are presented as diagrams of resultant pressures and of resultant-pressure increments for the airfoil with the flap and the 20-percent-chord tab. Increments of normal-force and hinge-moment coefficients for the airfoil, the flap, and the three tabs are also given. At all unstalled flap and tab deflections, the experimental distributions agree well with those calculated by an analytical method. The agreement is poor, however, then the stalled or the unstalled condition of the flap or tab deflected alone was changed to an unstalled or stalled condition by the simultaneous deflection of both the flap and the tab."

Aerodynamic force tests of an NACA 23018 airfoil with a Gwinn flap having a chord 25 percent of the overall chord and of an NACA 23015 airfoil with a plain flap having a 25-percent chord were conducted to determine the relative merits of the Gwinn and the plain flaps. The tests indicated that, based on speed-range ratios, the plain flap was more effective than the Gwinn flap. At small flap deflections, the plain flap had lower drag coefficients at lift-coefficient values less than 0.70. For lift coefficients greater than 0.70, however, the Gwinn flap at all downward flap deflections had the lower drag coefficients.