Report presenting testing of a semispan model of a wing of triangular plan form and aspect ratio 2 in the 12-foot pressure tunnel to determine the aerodynamic characteristics of the wing as influenced by the independent effects of Reynolds number and Mach number up to Mach numbers approaching unity. Results regarding the effect of body and effect of wing-profile modification are also provided.

Report presenting an investigation in the free-flight tunnel to determine the dynamic longitudinal stability and control characteristics of tandem-coupled bomber-fighter airplane models. Results regarding bomber alone tests, freely coupled combination, and rigidly coupled combination are provided.

Report presenting longitudinal-damping data in the form of the pitching-moment derivatives and summarized from flight tests of rocket-propelled models and full-scale airplanes. 22 models and 4 airplanes are examined in the study and the damping derivative results are generalized.

Report presenting an investigation of a supersonic aircraft configuration and various combinations of its components at a range of Mach numbers. Longitudinal- and lateral-force characteristics of the configurations, as well as longitudinal- and lateral-stability derivatives, are provided.

Memorandum presenting a method for calculating the stick-fixed longitudinal stability of a wing-fuselage configuration at subcritical Mach numbers. The method applies to unswept- and swept-wing configurations. The stability parameters estimated by the method show reasonable agreement with the experimental values for the 23 configurations used in the comparison.

Report presenting an investigation at Mach number 1.61 to determine the hinge-moment characteristics of seven tip controls on a 60 degree sweptback delta wing. Testing occurred over a range of angles of attack and control deflection. The results indicated that the most important parameter in designing delta-wing tip control is the ratio of control-surface area ahead of the hinge line to total control-surface area.

Report presenting testing of a shielded total-pressure tube with a curved venturi entry in order to obtain the variation of total-pressure error with angle of attack from a range of 0 to 60 degrees and for a range of Mach numbers. The critical angle was found to vary from about 57 degrees at Mach number 0.90 to 56 degrees at Mach number 1.10.

A summary of some effective aerodynamic twisting-moment coefficients of various wing-control configurations for use at a range of Mach numbers. The values obtained were determined by the combined use of experimentally determined free-flight data and subsonic aerodynamic theory.

Report presenting the results of an investigation to determine the pressure ratios required to operate a slotted tunnel through a range of Mach numbers where the speed variation is affected by removal of air from the main stream by auxiliary pumping, use of a main-stream-operated ejector located downstream of the test section, and a combination of these methods. Testing was conducted in order to improve the accuracy of transonic wind-tunnel testing.

Report presenting two-group theory calculations to determine the effect of nonuniform uranium loading as compared to uniform loading on the reflector control effectiveness attainable in a large thermal reactor usable for aircraft power application. Results regarding reactivities and investments, uranium distributions, and power distributions are provided.

Report presenting a flight and preflight evaluation of an automatic thrust-coefficient control system in a twin-engine ram-jet missile. A flicker-type single-loop servocontrol system is shown to be a usable way of controlling ram-jet thrust coefficients.

"Flow-angle and pressure surveys behind five, thin, pointed-tip wings of varying plan form have been made at Mach numbers 1.62 and 2.41. Schlieren studies at a Mach number 1.93 for the same five plan-form wings were made to illustrate the behavior of the vortex sheet. The surveys were conducted at 1.5, 3, and 4 root chords behind three triangular wings of 50 degree, 63 degree, and 72 degree leading-edge sweep angle, and behind the 50 degree triangular wing reversed" (p. 1067).

"A number of noise-suppression nozzles were tested on full-scale engines. In general, these nozzles achieved noise reduction by the mixing interference of adjacent jets, that is, by using multiple-slot-nozzles. Several of the nozzles achieved reductions in sound power of approximately 5 decibels (nearly 70 percent) with small thrust losses (approx. 1 percent). The maximum sound-pressure level was reduced by as much as 18 decibels in particular frequency bands" (p. 1249).

From Summary: "As part of a general investigation of propellers at high forward speeds, tests of two 2-blade propellers having the NACA 4-(3)(8)-03 and NACA 4-(3)(8)-45 blade designs have been made in the Langley 8-foot high-speed tunnel through a range of blade angle from 20 degrees to 60 degrees for forward Mach numbers from 0.165 to 0.725 to establish in detail the changes in propeller characteristics due to compressibility effects. These propellers differed primarily only in blade solidity, one propeller having 50 percent and more solidity than the other. Serious losses in propeller efficiency were found as the propeller tip Mach number exceeded 0.91, irrespective of forward speed or blade angle."

A blade design study is presented for a two-stage air-cooled turbine suitable for flight at a Mach number of 2.5 for which velocity diagrams have been previously obtained. The detailed procedure used in the design of the blades is given. In addition, the design blade shapes, surface velocity distributions, inner and outer wall contours, and other design data are presented. Of all the blade rows, the first-stage rotor has the highest solidity, with a value of 2.289 at the mean section. The second-stage stator also had a high mean-section solidity of 1.927, mainly because of its high inlet whirl. The second-stage rotor has the highest value of the suction-surface diffusion parameter, with a value of 0.151. All other blade rows have values for this parameter under 0.100.

Report presenting the results of an investigation of the static longitudinal stability and control characteristics of a canard airplane configuration without analysis for a range of Mach numbers. Data are presented for a variety of angles of attack and canard angles.

"Drag characteristics of a series of related conical and circular-arc afterbodies are presented for Mach numbers from 0.6 to 1.3. Drag was obtained from pressure measurements on the boattail and solid base. The boattail angles tested ranged from 0 degrees to 45 degrees for ratios of base diameter to maximum body diameter ranging from 0 to 1.0" (p. 1).

Report presenting an investigation at transonic speeds to determine the effects of a sonic propulsive jet on the aerodynamic characteristics of the body from which it issues. A variety of afterbody shapes and their pressure distributions are described. Correlations are also drawn between certain characteristics of the afterbody and the effect of temperatures.

Report presenting measurements of aerodynamic heat transfer at a number of stations along a cone-cylinder-flare model with 15 degree total-angle conical nose and a 10 degree half-angle flare skirt. Results regarding temperature measurements and heating rates, local flow parameters, heat transfer with theoretical recovery factors, experimental recovery factors, prediction of skin temperatures, and transition are provided.

From Summary: "Drag measurements at transonic speeds on rectangular airfoils and on airfoils swept back 450 are reported. These airfoils, which were mounted on cylindrical test bodies, are part of a series being tested in free drops from high altitude to determine the effect of variation of basic airfoil parameters on airfoil drag characteristics at transonic speeds. These rectangular and swept-back airfoils had the same span, airfoil section (NACA 65-009), and chord perpendicular to the leading edge. The tests were made to compare the drag of rectangular and sweptback airfoils at a higher aspect ratio than had been used in a similar comparison reported previously."

Memorandum presenting the results of a simulator stud of the factors affecting the design of a device, a stick pusher, for preventing a representative supersonic airplane from entering the pitch-up region. The effects of varying the stick-pusher-activation boundaries, sensing parameters, and magnitude of stick-pusher force on the controllability of the airplane pitch-up were investigated.

Report presenting some measurements of the vortex movements with time about an airfoil undergoing a blast of sufficient strength to exceed the stall angle by a large amount. Measurements were also obtained without flight simulation but with blast orientation and strength and compared to the first test. Results regarding Schileren photographs, presentation of results, and discussion of results are provided.

"An investigation of the XJ-41-V turbojet-engine compressor was conducted to determine the performance of the compressor and to obtain fundamental information on the aerodynamic problems associated with large centrifugal-type compressors. The results of the research conducted on the original compressor indicated the compressor would not meet the desired engine-design air-flow requirements because of an air-flow restriction in the vaned collector. The compressor air-flow choking point occurred near the entrance to the vaned-collector passage and was instigated by a poor mass-flow distribution at the vane entrance and from relatively large negative angles of attack of the air stream along the entrance edges of the vanes at the outer passage wall and large positive angles of attack at the inner passage wall" (p. 1).

Report presenting an investigation of three different nacelles in an underwing position at the 0.46 semispan station of a 45 degree sweptback wing at three chordwise positions for a range of Mach numbers. The nacelle profiles were an ogive cylinder, an NACA 65A-series airfoil, and an NACA 0-series airfoil (reversed). Results regarding drag characteristics, lift-drag ratios, lift characteristics, pitch characteristics, and lateral center of pressure are provided.