"A method of analysis based on slender-wing theory is developed to investigate the characteristics in roll of slender cruciform wings and wing-body combinations. The method makes use of the conformal mapping processes of classical hydrodynamics which transform the region outside a circle and the region outside an arbitrary arrangement of line segments intersecting at the origin. The method of analysis may be utilized to solve other slender cruciform wing-body problems involving arbitrarily assigned boundary conditions" (p. 1).

The effect of Mach number and structural damping on single-degree-of-freedom pitching of a wing is presented. Some experimental results are compared with theory and good agreement is found for certain ranges of an inertia parameter.

Report presents the results of an investigation conducted in the Langley stability tunnel to determine the effects of horizontal tails of various sizes and at various tail lengths (when loaded on the fuselage center line) and also the effects of vertical location of the horizontal tail relative to the wing on the low-speed static longitudinal stability and on the steady-state rotary damping in pitch for a complete-model configuration. The wing and tail surfaces had the quarter-chord lines swept back 45 degrees and had aspect ratios of 4.

Report presents the results of measurements of the longitudinal stability and control characteristics of a large airplane using a mechanical feel device in combination with a booster incorporated in the elevator-control system. Tests were made to investigate the feasibility of eliminating the aerodynamic control forces through use of a booster and of providing control-feel forces mechanically. The feel device consisted of a centering spring which restrained the control stick through a linkage which was changed as a function of the dynamic pressure. Provisions were made for trimming and for manual adjustment of the force gradient. The system was designed to approximate the control-force characteristics that would result with a conventional elevator control with linear hinge-moment characteristics.

"A supersonic inlet with supersonic deceleration of the flow entirely outside of the inlet is considered a particular arrangement with fixed geometry having a central body with a circular annular intake is analyzed, and it is shown theoretically that this arrangement gives high pressure recovery for a large range of Mach number and mass flow and, therefore, is practical for use on supersonic airplanes and missiles. Experimental results confirming the theoretical analysis give pressure recoveries which vary from 95 percent for Mach number 1.33 to 86 percent for number 2.00. These results were originally presented in a classified document of the NACA in 1946" (p. 1).

From Summary: "The longitudinal and lateral-directional dynamic-response characteristics of a 35 degree swept-wing fighter-type airplane determined from flight measurements are presented and compared with predictions based on theoretical studies and wind-tunnel data. Flights were made at an altitude of 35,000 feet covering the Mach number range of 0.50 to 1.04. A limited amount of lateral-directional data were also obtained at 10,000 feet. The flight consisted essentially of recording transient responses to pilot-applied pulsed motions of each of the three primary control surfaces. These transient data were converted into frequency-response form by means of the Fourier transformation and compared with predicted responses calculated from the basic equations. Experimentally determined transfer functions were used for the evaluation of the stability derivatives that have the greatest effect on the dynamic response of the airplane. The values of these derivatives, in most cases, agreed favorably with predictions over the Mach number range of the test."

"The general equations of motion for an airplane with a number of spherical fuel tanks are derived. The motion of the fuel is approximated by the motion of solid pendulums. The same type of derivation and equations are shown to apply to any type of fuel tank where the motion of the fuel may be represented in terms of undamped harmonic oscillators. Motions are calculated for a present-day high-speed airplane and a free-flying airplane model with two spherical tanks in the symmetry plane" (p. 1).

Report presents the results of interferometer measurements of the flow field near two-dimensional wedge and circular-arc sections of zero angle of attack at high-subsonic and low-supersonic velocities. Both subsonic flow with local supersonic zone and supersonic flow with detached shock wave have been investigated. Pressure distributions and drag coefficients as a function of Mach number have been obtained.

"A theoretical study has been made of the aerodynamic characteristics at zero angle of attack of a thin, doubly symmetrical double-wedge profile in the range of supersonic flight speed in which the bow wave is detached. The analysis utilizes the equations of the transonic small-disturbance theory and involves no assumptions beyond those implicit in this theory. The mixed flow about the front half of the profile is calculated by relaxation solution of boundary conditions along the shock polar and sonic line" (p. 1).

"An iterative transformation procedure suggested by H. Wielandt for numerical solution of flutter and similar characteristic-value problems is presented. Application of this procedure to ordinary natural-vibration problems and to flutter problems is shown by numerical examples. Comparisons of computed results with experimental values and with results obtained by other methods of analysis are made" (p. 1).

"Experimental measurements of the aerodynamic reactions on a symmetrical airfoil oscillating harmonically in a two-dimensional flow are presented and analyzed. Harmonic motions include pure pitch and pure translation, for several amplitudes and superimposed on an initial angle of attack, as well as combined pitch and translation. The apparatus and testing program are described briefly and the necessary theoretical background is presented" (p. 1).

"In accordance with the act of Congress, approved March 3,1915 (U.S. C. title 50, sec. 151), which established the National Advisory Committee for Aeronautics the Committee submits its thirty-seventh annual report for the fiscal year 1951. The United States is engaged in expanding military aviation to levels never before reached except in the midst of a major war. In Korea, our military aircraft are engaged in combat with airplanes of an unfriendly nation evidently able to build military aircraft of increasing capabilities. In this environment, the NACA is responsible for conducting an adequate program of scientific research to open the way for the design of aircraft and missile of superior performance" (p. VII).