Memorandum presenting an investigation of the effects of modifying a twin side-inlet duct system on an interceptor airplane as recorded and analyzed over an altitude range from about 25,000 to 51,000 feet throughout the transonic speed range to a Mach number of about 1.2. The modifications consisted primarily of redesigning the inlet lip, increasing the cross-sectional area of the inlet and diffuser, and adding a region of duct contraction ahead of the engine.

Report presenting comparative tests of the rolling power of plain constant-chord, full-delta, and half-delta ailerons on delta wings with varying degrees of sweepback were tested on rocket-propelled test vehicles. Results regarding the effect of leading-edge sweepback, a comparison of rolling effectiveness, and comparison with theory are provided.

Report presenting the pressure-recovery characteristics of a model with two scoops situated on the aft position of a long forebody and connected through diffusors to a common settling chamber at a range of Mach numbers. The effects of interaction between the flow in the two air-induction systems and of varying the mass flow through the boundary-layer scoops were studied. Results at several different angles of attack are provided.

From Introduction: "The data obtained for each configuration are presented in a manner to show the effects of fuel distribution and flame-holder design on net thrust, specific fuel consumption, exhaust-gas temperature, combustion efficiency, operable range of tail-pipe-burner fuel-air ratios, and maximum altitude limit."

A 24-inch-diameter shock-in-rotor supersonic compressor rotor designed to obtain simple radial equilibrium behind the normal shock obtained a maximum total-pressure ratio of 2.02 at an adiabatic efficiency of 0.74 and a weight flow of 61.5 pounds per second at design speed. Although this design method considerably reduced the radial forces, there was no apparent reduction in the radial forces, there was no apparent reduction in the radial redistribution of mass flow over the radial redistribution obtained in the previous investigation. This investigation indicates that the shock boundary-layer interaction pattern and associated effects have a very large effect on the radial redistribution of mass flow through the shock-in-rotor supersonic-compressor rotor.

Report presenting lateral and directional dynamic response characteristics of a 35 degree swept-wing fighter-type airplane determined from flight measurements and compared with predictions based on theoretical studies and wind-tunnel testing. Results regarding the frequency responses, theoretical and experimental transfer functions, stability derivatives, aeroelastic effects, response at low frequency, and effects of minor derivatives, product of inertia, and flight-path angle are provided.

Report presenting an investigation to evaluate the effects of an all-movable horizontal tail on the longitudinal characteristics of a sweptback wing and a fuselage of a type suitable for long-range high-speed airplanes. Tests were conducted at low speed and at high Mach numbers over a range of Reynolds numbers. Large reductions in longitudinal stability of the wing-fuselage-tail combination were found at lift coefficients well below the stall.

Report presenting an analysis of the force distribution on a slender pointed body of revolution at angle of attack utilizing pressure-distribution data and balance measurements. The data were obtained in the supersonic tunnel at a variety of Mach numbers and angles of attack. Results regarding normal forces and axial force are provided.

"An experimental investigation of a 0.8 hub-tip radius ratio, nontwisted-rotor-blade turbine designed for a stagnation-pressure ratio of 2.5 and an equivalent mean blade speed of 643 feet per second was made in a cold-air turbine with (a) nontwisted stator blades, and (b) twisted stator blades designed to maintain zero rotor-inlet incidence angles. Turbine efficiencies of the order of 0.85 at the design point were obtained with a nontwisted-rotor-blade turbine with a hub-tip radius ratio of 0.80. The turbine with the twisted stator blades gave higher efficiencies at the design point (of the order of 1.5 percentage points) than the turbine with the nontwisted stator blades" (p. 1).

"The pressure distribution and drag were determined for a spherical-nosed axially symmetric body with thin projecting rods at Mach numbers of 1.76, 1.93, and 2.10. The upstream projection distance of the rods was varied over a wide range to study changes in the character of the flow separation and to determine the variation of drag and pressure distribution with tip projection. Drag coefficients between 0.18 and 0.30 were obtained for most tip projections at each Mach number" (p. 1).

Report presenting the effects of Mach number between 0.50 and 0.95 and between 1.09 and 1.29 on the characteristics of the downwash and on the location of the wake behind a triangular wing of aspect ratio 4 as determined from semispan model tests.

A method is derived for calculating the damping coefficients in pitch and roll for a series of triangular wings and a restricted series of sweptback wings at supersonic speeds. The elementary "supersonic source" solution of the linearized equation of motion is used to find the potential function of a line of doublets, and the flows are obtained by surface distributions of these doublet lines. The damping derivatives for triangular wings are found to be a function of the ratio of the tangent of the apex angle to the tangent of the Mach angle. As this ratio becomes equal to and greater than 1.0 for triangular wings, the damping derivatives, in pitch and in roll, become constant. The damping derivative in roll becomes equal to one-half the value calculated for an infinite rectangular wing, and the damping derivative in pitch for pitching about the apex becomes equal to 3.375 times that of an infinite rectangular wing.

From Introduction: "The present report presents the results of such an analysis as regards the longitudinal-stability and control characteristics of the various airplanes tested."

The performance of the 11-stage axial-flow compressor in the X24C-4B turbojet engine was analyzed on the basis of results obtained from an investigation of the complete engine in the NACA Cleveland altitude wind tunnel. The engine was operated with four, exhaust nozzles of different outlet area over a range of engine speeds from 6000 to 12,500 rpm, corrected engine speeds from approximately 6100 to 13,600 rpm, and compressor Mach numbers from 0.45 to 1.00. Data are presented for engine operation over a range of simulated altitudes from 15,000 to 45,000 feet and simulated flight Mach numbers from 0.24 to 1.08.

The magnitude and the effects of boundary-layer separation on normal-force-curve slopes, centers of pressure, pressure distributions, and lift and drag coefficients were determined for various bodies of revolution with conical tail flares at Mach numbers from 3.0 to 6.3. Some of the experimental results are compared to theoretical predictions of the aerodynamic characteristics of the bodies.

Report presenting the effects of modifying a twin side-inlet duct system on an interceptor airplane over a range of altitudes throughout the transonic speed range to Mach number 1.2. The primary modifications were redesigning the inlet lip, increasing the cross-sectional area of the inlet and diffuser, and adding a region of duct contraction ahead of the engine.

Windmilling characteristics of several turbojet and turbine-propeller engines were investigated individually over a wide range of flight conditions in the NACA Lewis altitude wind tunnel. A study was made of all these data and windmilling performance of gas turbine engines was generalized. Although internal-drag, air-flow, and total-pressure-drop parameters were generalized to a single curve for both the axial-flow type engines and another for the centrifugal-flow engine.

Report presents a comparison of the theoretical and experimental expressions of lift, drag, and center of pressure of supercavitating hydrofoils. The results obtained from both methods appear to agree well with one another. The application of cavitation number to a variety of situations and speeds is also provided.

Report presenting a small-scale apparatus for simulating the motion and heating of ballistic missiles and its elements of design and operation. Experiments demonstrate that conditions for simulation are fulfilled according to theoretical requirements. Results regarding the application to full-scale missiles, shadowgraph observations, observations of recovered models, and spectroscopic observations are provided.

Memorandum presenting a wind-tunnel investigation of the three types of lateral control devices on a low-aspect-ratio trapezoidal wing. Tip ailerons, a plain spoiler, and a vented spoiler were investigated to ascertain the relative characteristics of each system. The results indicated the superiority in roll power of the tip ailerons over the plain and vented ailerons.

From Summary :"An investigation was conducted to evaluate the possibilities of utilizing the high-performance characteristics of triptane and xylidines blended with 28-R fuel in order to increase fuel economy by the use of high compression ratios and maximum-economy spark setting. Full-scale single-cylinder knock tests were run with 20 deg B.T.C. and maximum-economy spark settings at compression ratios of 6.9, 8.0, and 10.0, and with two inlet-air temperatures. The fuels tested consisted of triptane, four triptane and one xylidines blend with 28-R, and 28-R fuel alone."

The partial differential equation for the perturbation velocity potential is examined for free-stream Mach numbers close to and equal to one. It is found that, under the assumptions of linearized theory, solutions can be found consistent with the theory for lifting-surface problems both in stationary three-dimensional flow and in unsteady two-dimensional flow. Several examples are solved including a three dimensional swept-back wing and two dimensional harmonically-oscillating wing, both for a free stream Mach number equal to one. Momentum relations for the evaluation of wave and vortex drag are also discussed. (author).

"In order to compare the relative maneuverability of two fighting airplanes and to accumulate additional data to assist in establishing a satisfactory criterion for the maneuverability of any airplane, the National Advisory Committee for Aeronautics has conducted maneuverability investigations on the F6C-3 (water-cooled engine) and the F6C-4 (air-cooled engine) airplanes. The investigation made on the F6C-3 airplane was reported in NACA-TR-369. This report contains the results of the investigation made on the F6C-4 airplane" (p. 475).

From Summary: "Pressures and temperatures throughout the X24C-4B turbojet engine are presented in both tabular and graphical forms to show the effect of altitude, flight Mach number, and engine speed on the internal operation of the engine. These data were obtained in the NACA Cleveland altitude wind tunnel at simulated altitudes from 5000 to 45,000 feet, simulated flight Mach numbers from 0.25 to 1.08, and engine speeds from 4000 to 12,500 rpm. Location and detail drawings of the instrumentation installed at seven survey stations in the engine are shown."