Report presenting a consideration of the problem of shaping an adjoining fuselage for a given wing and supersonic Mach number so that the combination will have a low wave drag. Only fuselages that can be simulated by singularities distributed along the body axis are studied.

Report presenting a description of the Ames 2- by 2-foot transonic wind tunnel, its uses in testing, and a preliminary evaluation of the wall interference in the tunnel.

"Aerodynamic data have been obtained for a fuselage forebody alone and for canopy-body configurations consisting of four different canopies mounted on a fuselage forebody. Two of the canopies had the same shape and size rearward of the windshield but one had a "flat" and the other a "vee" windshield. The remaining two canopies were located at different body stations and were geometrically similar. The data indicated that the drag of the flat-windshield model was consistently lower than that of the vee-windshield model" (p. 1).

Memorandum presenting experimental flutter Mach numbers for several solid, thin, rectangular cantilever wings with uniform section properties, low aspect ratio, and high relative density estimated from the results of previous tests at zero angle of attack. The experimental values are considered estimates, rather than determinations, in the high subsonic speed range because in that range the amplitude criterion was necessarily arbitrary. Results regarding conservatism of standard analysis relative to the experiment, closeness of analytical and experimental results, and interpretation of the analytical and experimental uncertainties are provided.

Memorandum presenting a systematic transonic zero-lift drag investigation utilizing a swept-wing configuration with three different ratios of wing to fuselage size utilizing a 6-inch helium gun. The experiments and comparisons with other results indicated that the reduction of pressure drag obtainable from partial indentation is approximately proportional to the amount of indentation employed up to a Mach number of 1.3. Results regarding the basic data, total drag, pressure drag, equivalent bodies, and special projections are provided.

From Introduction: "The investigation reported herein was initiated to define further the effects on the aerodynamic characteristics and load distribution of a thin, sweptback wing of a low-pressure blowing system and also to provide information on which to base a more thorough study of a complete airplane configuration."

Report presenting shock-position controls designed to actuate the translating spike and variable bypass of a variable-geometry inlet in the 8- by 6-foot supersonic tunnel. The operation of the inlet was observed in combination with a J34 turbojet engine at a range of Mach numbers. Results regarding terminal-shock-positioning control systems and combination control are provided.

Report presenting an investigation of a turbojet engine with variable inlet guide vanes in an altitude test facility to determine the steady-state compressor performance and surge characteristics. Compressor surge pressure ratios were found to be unaffected by changes in flight condition for either position of the inlet guide vanes. Decreasing the inlet Reynolds number was found to reduce the corrected fuel flow required for compressor surge.

Report presenting chordwise pressure distributions measured over the left wing of a Convair XF-92A delta-wing airplane at low lift to determine the effect of Mach number on the wing characteristics at transonic speeds. The critical Mach number of the basic wing was 0.83 for an airplane normal-force coefficient at 0.09. The spanwise load distributions were roughly triangular in shape.

Report presenting a program to investigate the aerodynamic characteristics of a family of delta wings with a blunt double-wedge section in the hypersonic tunnel at Mach number 6.9. The wings varied at semiapex angle and were tested over a range of angles of attack and Reynolds numbers. Results regarding lift and drag characteristics, center of pressure and moment coefficient, Schlieren photographs, and surface film flow studies are presented.

Memorandum presenting a correlation of analyses of the effect of distributed surface injection on the heat transfer and skin friction associated with the turbulent boundary layer at high speeds to eliminate the effects of Mach and Reynolds number. Data for heat transfer and skin friction at three Mach numbers are compared with the analyses and the agreement is good.

Report presenting an experimental investigation in the transonic blowdown tunnel to determine effects of center-of-gravity location on the transonic flutter characteristics of a 45 degree sweptback-wing plan form of aspect ratio 4.0 and taper ratio 0.6. Solid-construction models with streamwise NACA 65A004 airfoil sections and center-of-gravity locations at three different chords were fluttered at a range of Mach numbers.

From Summary: "Hydrodynamic characteristics have been determined for a 1/13.33-scale jet-powered dynamic model of the Martin XP6M-1 flying boat with the forebody modified so as to increase the depth of step and the angle between the forebody and afterbody keels. Longitudinal stability during takeoff and landing in smooth water and resistance of the complete model in smooth water and in waves are presented."

Report presenting testing of a low-acceleration transonic flutter test vehicle to obtain flutter data on two similar sweptback wings which indicated that wing flutter was symmetrical in mode. Results regarding flight and flutter characteristics for the FR-1-B are provided.

Report presenting an investigation of the effect of the rate of change of angle of attack on the maximum lift coefficient of a pursuit airplane equipped with a low-drag-type wing in stalls of varying abruptness over a range of Mach and Reynolds numbers. Maximum lift coefficients were found to increase linearly with increasing rate of change of angle of attack per chord length of travel up to the maximum rate attained in the tests.

"The frequency of blade temperature to exhaust gas temperature is presented for two locations in the blade and at several operating conditions. The frequency response was determined by Fourier analysis of transient data. Two analytical methods are presented, and results are compared with experimental data. Dynamic response of turbine-blade temperature to exhaust-gas temperature exhibited the form of an approximate first-order lag" (p. 1).

Report presenting an investigation at Mach number 1.6 and a range of Reynolds number of the skin-friction drag and boundary-layer transition of a body of revolution. The body had a parabolic-arc profile, a blunt base, and a fineness ratio of 12.2 (NACA RM-10). Results regarding the effect of surface condition, drag breakdown, and results of boundary-layer surveys are provided.

Report presenting an investigation to increase the compressor surge-limit pressure ratio of the XJ40-WE-6 turbojet engine at high equivalent speeds. The two compressor modifications investigated included twisting rotor blades to change blade section angles and inserting new stator diaphragms with different blade angles. Results regarding the effect of surge limit on engine performance, performance map of original compressor with mixer, effect of blade modifications on surge line, effect of mixer on surge limit, and effect of compressor modifications on performance are provided.

Report presenting a flight investigation using rocket-powered models to determine some of the effects of wing-to-tail distance on the stability and control characteristics of a canard-missile configuration with 60 degree delta wings and control surfaces. Both models were tested over a range of Mach numbers from 0.8 to 2.0, and canard hinge moment and model drag data are presented. Testing indicated that additional fuselage length increases the lift-curve slope slightly and has very little effect on minimum drag.

Report presenting a flight investigation using four rocket-propelled models of low-aspect-ratio tailless configurations between Mach numbers 0.7 and 1.35. The configurations investigated were a 45 degree delta model with NACA 63A010 airfoil sections parallel to free stream, a 45 degree delta-wing model with NACA 63A006 airfoil sections parallel to free stream, a 37.5 degree swept-tapered wing model with NACA 64(sub 1)A012 sections perpendicular to quarter chord, and a 39 degree swept-tapered-wing model with NACA 64A006 airfoil sections perpendicular to quarter chord. Results regarding time histories, static longitudinal stability, dynamic longitudinal stability, order of oscillation frequency, rotational damping in pitch, and drag are provided.

Report presenting an investigation of the influence of high-altitude supersonic flight on the operation and effectiveness of turbine-blade liquid-cooling systems for application in turbojet engines in guided missiles and in supersonic aircraft. The problems encountered in liquid-cooling systems were investigated with reference to several specific designs for alternate heat-rejection mediums.

Memorandum presenting tests of a triangular-wing-airplane model equipped with high-lift devices and lateral and directional controls. The model consisted of an aspect ratio 4 triangular wing in combination with a fuselage of fineness ratio 12.5, a thin, triangular, vertical tail with a constant-chord rudder, and a thin, unswept, all-movable horizontal tail. Results regarding the effects of flap deflection and tail incidence on the pitching moment characteristics, lift and drag characteristics, and rudder effects are provided.

Report presenting an investigation to determine the effects of turbine-blade coatings, supplied by fuel additives, on heat transfer with combustor exhaust gases to an S-816 alloy blade. Two fuel additives were used. The results indicated that the fuel additives provided an oxide coating on the combustor-liner wall and other metal parts, including the blade, exposed to the exhaust gas.

Report presenting wind-tunnel studies of the wake behind eight types of canard control surfaces of equal span mounted on a missile body at Mach number 1.93. The effect of simulated external rocket boosters on the pressure field was also determined for three control-surface plan forms. The investigation also included the effect of end plates attached to the tips of one of the controls surfaces.