Note presenting experimental results for transonic flow past cone-cylinder, axially symmetric bodies. The drag coefficient and surface Mach number are studied as the free-stream Mach number is varied and, wherever possible, the experimental results are compared with theoretical predictions. The theoretical problem of transonic flow past finite cones is discussed briefly and an approximate solution of the axially symmetric transonic equations, valid for a semi-infinite cone, is presented.

Note presenting a determination of the starting and operating pressure ratios for two supersonic wind tunnels which employed air injectors to supplement the primary pumping systems of the tunnels. Data are presented for tunnels operating at Mach numbers 3.85, 3.05, and 2.87 over a range of injector-to-tunnel mass-flow ratios of 0.5 to 1.35.

Report presenting an investigation of the low-speed two-dimensional aerodynamic characteristics of an NACA 0006 airfoil with area suction near the leading edge. For a given increase in lift coefficient, the minimum suction quantity required was related to the magnitude of the difference between the external pressure coefficients at the leading and trailing edges of the porous area. Results regarding the stall of the model, effect of chordwise extent of suction, chordwise variation of suction velocity, typical aerodynamic characteristics, effect of free-stream velocity, effect of a deflected flap, and some design considerations are provided.

"A method is presented for analyzing the stresses about a cutout in a circular cylinder of semimonocoque construction. The method involves the use of so-called perturbation solutions which are superposed on the stress distribution that would exist in the structure with no cutout in such a way as to give the effects of a cutout. The method can be used for any loading case for which the structure without the cutout can be analyzed and is sufficiently versatile to account for stringer and shear reinforcement about the cutout" (p. 1).

"The lift, drag, and pitching-moment characteristics of the AGARD Calibration Model B as determined in the Langley 9-inch supersonic tunnel are presented at Mach numbers of 1.62, 1.94, and 2.41 and at a Reynolds number, based on body length, of approximately 3.0 x 10(exp 6). The zero-lift drag data compared favorably with available data and were in the proper sequence for the effects of Reynolds number" (p. 1).

Report presenting turbulent-heat-transfer measurements obtained through the use of a three-dimensional axially symmetric annual nozzle consisting of an inner shaped plug and an outer cylindrical sleeve at Mach number 3.03. Results regarding the variation of wall temperature with longitudinal distance, variation of local Nusselt number with local Reynolds number, and variation of local temperature-recovery factor with local Reynolds number are provided.

Report presenting some information on landing-gear applied drag loads and on the nature of the wheel spin-up phenomenon in landing, based on a program of tests under controlled conditions in the impact basin. Results regarding fundamentals of the wheel spin-up process, comparison of the maximum loads in various types of tests, variation of coefficient of friction, and effect of prerotation are provided.

Report presenting an analog computer analysis of an acceleration restrictor which limits the elevator motion of an airplane. The device applies a mechanical brake to the elevator control system in response to signals. Results for a restrictor sensitive to combined signal of normal and pitching acceleration and a restrictor sensitive to normal and pitching acceleration and pitching velocity are provided.

Note presenting a theoretical investigation of a general method for predicting the flow field behind the wings of plane and cruciform wing and body combinations at transonic or supersonic speeds and slender configurations at supersonic speeds. The effects of certain geometric parameters as aspect ratio,angle of attack and incidence, ratio of body radius to wing semispan, and angle of bank on the vortex wake behidn wings of wing-body combinations were studied.

Average heat-transfer, temperature-recovery, and pressure-distribution measurements were obtained over the hemispherical nose of a body of revolution both with and without flow-separation spikes. The tests, conducted in a range of Reynolds numbers from 1.55 to 9.85 x 10(5) and from Mach number 0.12 to 5.04, indicated that in supersonic flow the addition of spikes approximately doubles the rate of heat transfer, regardless of spike length.