Report presenting the average heat-transfer coefficients and equilibrium temperatures for the front half of an isothermal cylinder with a laminar boundary layer as determined by wind tunnel testing at Mach number 6.9 and a range of Reynolds numbers and sweep angles.

Report presenting a transonic compressor rotor with double-circular-arc blade sections, which were designed and tested to investigate the aerodynamic effects of using a short blade chord and a high aspect ratio. The report presents the blade-element performance and performance analysis.

Report presenting static-pressure changes resulting from steady combustion of aluminum borohydride in a supersonic wind tunnel. Static pressures were measured along the top wall of a tunnel adjacent to the flame that filled the upper portion of the test section. Results regarding the static-pressure increases, effect of change in fuel-injection rate, and effect of tunnel boundary layer on the character of the heat-addition region are provided.

Report presenting heat-transfer data obtained in free flight at supersonic Mach numbers up to 4.90 and local Reynolds numbers per foot up to 27.7 million. Two 20 degree-included-angle cone-cylinder models were launched from a carrier airplane at 36,000 feet of altitude. Results regarding primary data and heat-transfer results are provided.

Report presenting an investigation to determine the effects of inlet flow distortion on the compressor stall and engine acceleration characteristics of a J65-B-3 turbojet engine. A sinusoidal circumferential distortion and two abrupt radial tip flow distortions were introduced at the engine inlet at several altitudes and Mach number 0.8. Results regarding the definition of inlet flow distortions, acceleration limits, compressor pressure ratio at stall, fuel flow for stall, and maximum acceleration rate are provided.

From Summary: "The performance of the XRJ43-MA-3 flight-weight ram-jet engine has been investigated in a free-jet facility as part of the development program for the "Bomarc," ram-jet powered, interceptor-type missile. The engine performance was obtained at Mach numbers of 2.35, 2.50, and 2.70, altitudes from 50,000 to 75,000 feet, and Miami hot day and cold day inlet temperatures at an angle of attack of +4 deg. At each flight condition investigated, the engine control set an inner-ring fuel flow which was held constant as outer-ring fuel flow was varied."

Report presenting an investigation with sting-mounted rocket-propelled models in free flight at approximately zero lift to determine the effect of wing flexibility on the damping-in-roll characteristics of a wing-body combination over a range of Mach numbers.

From Introduction: "Results from the tests in the Langley 4- by 4-foot supersonic pressure tunnel at Mach numbers of 1.61 and 2.01 are presented in reference 1. Reported herein are results obtained from the tests in the Langley 8-foot transonic tunnel of the model with no control deflections at Mach numbers 0.6 to 1.12 for angles of attack up to 34^o

Memorandum presenting an investigation conducted in the 4- by 4-foot supersonic pressure tunnel to determine the aerodynamic characteristics of a supersonic aircraft configuration equipped with a canard control surface at a Mach number of 1.89. The model had a 40 degree sweptback tapered wing with an aspect ratio of 4 and 10-percent-thick circular-arc sections normal to the quarter-chord line. Results regarding the pitching-moment characteristics, lift and longitudinal-force characteristics, longitudinal stability and control, and lateral stability characteristics are provided.

Report presenting a free-flight investigation in a range of Mach numbers to determine the maximum deflection, zero-lift rolling effectiveness, and general operating characteristics of a bellows-actuated aileron control system energized by the pitot (impact) pressure. Testing of a system with a partial-span spit-flap aileron on a 60 degree delta wing showed that it is promising for obtaining lateral control at supersonic speeds. Results regarding aileron rolling effectiveness, aileron hinge moments, system time lag, estimation of aileron deflection, and aileron deflection control are provided.

Memorandum presenting an investigation made to determine the low-subsonic-speed static longitudinal stability characteristics of a composite-plan-form wing tested as a wing-fuselage combination and as a complete model. Also included is a comparison of the transonic-speed characteristics of the composite-plan-form wing alone and the 45 degree sweptback wing from which the composite wing was derived.

Memorandum presenting wind-tunnel tests of a semispan model of an unswept wing of aspect ratio 4 and a taper ratio of 0.50, equipped with leading- and trailing-edge flaps. The purpose of the tests was to determine the low-speed aerodynamic characteristics of the wing as affected by the separate or combined deflections of a full-span, constant-chord, leading-edge, plain flap and a partial-plan, constant-chord, trailing-edge flap of either the plain or split type. Results regarding the maximum lift characteristics, pitching-moment characteristics, drag coefficients, lift-drag ratio, and effect of Reynolds number and profile modification are provided.

Report presenting an investigation to determine the performance of a Bumblebee 18-inch ram jet with a rake-type flame holder was conducted using kerosene and propylene oxide as fuels. Experiments occurred at a range of pressure altitudes and ram-pressure ratios over the operable range of fuel-air ratios. Use of propylene oxide was found to improve the stability of combustion, increased the operable range of fuel-air ratios, and allowed ignition at high values of combustion-chamber-inlet velocities.

Report presenting an investigation of large-chord flap-type controls with sweptback hinge lines and with various sizes and shapes of horn balances on a triangular wing of aspect ratio 2. Changes in the following elements were investigated: shielding of the horn balance, percentage of horn balance, trailing-edge thickness, flap nose seal, contour of the horn balance, and Reynolds number.

Report presenting an investigation in the high-speed wind tunnel to determine the aerodynamic characteristics of three triangular wings and a trapezoidal wing through the transonic speed range by use of the bump technique. Data was obtained over a range of Mach and Reynolds numbers. Results regarding the lift, drag, pitching moment, and lift-drag ratio of the various wings are provided.

Report presenting testing on the NACA 16-009 airfoil and on eight airfoils of the NACA 6A series, which shows that the chordwise loading due to lift at a Mach number of 1.0 is insignificantly affected by thickness distribution within the range of airfoil shapes presented and only slightly affected by thickness. Results regarding the experimental loading on symmetrical airfoils, method of estimating pressure distribution, experimental loadings on cambered airfoils, and scope of method are provided.

Fundamental flame velocities are reported for pentane air, ethylene-air, and propylene-air mixtures for the concentration range 60 to 130 percent of stoichiometric. A form of the Tanford and Pease equation, which includes a small constant velocity term independent of diffusion, will predict the observed changes in flame velocity.

Memorandum presenting an investigation conducted to determine the effects of scale and compressibility on the forces, moments, and pressure distribution on a cambered and twisted wing with an aspect ratio of 3.0 and a taper ratio of 0.5. Lift, drag, and pitching-moment data and the chordwise distribution of static pressure at seven spanwise stations are presented for specified Mach and Reynolds numbers. In order to determine the effects of camber and twist, the force, moment, and pressure data are compared with the data for an uncambered and untwisted wing of the same plan form.

Report presenting a method based on geometric factors of the coolant passage was evolved for the rapid selection and evaluation of coolant passage configurations for air-cooled turbine rotor blades. The most promising coolant passage configurations are analyzed to obtain absolute values of the required cooling-air weight flow, pressure loss, and strength characteristics.

Memorandum presenting a measurement of the mean skin-friction coefficients on a flat-plate model, with and without initial roughness, and on a wind tunnel wall were measured at a nominal Mach number of 5.5 over a range of Reynolds numbers.

Report presenting a vibrational survey conducted during normal operation of a production eleven-stage axial-flow compressor powered as part of a complete turbojet engine. After serious third-stage vibration was detected, an investigation was conducted to study and eliminate the vibration with several component configurations. The vibrational source was determined to be an aerodynamic phenomenon related to a stall, but a new inlet-guide-vane assembly with a 4-percent increase in inlet area eliminated the vibration and increased the engine thrust 4 percent.

Report presenting an investigation of a series of wings of various plan forms in the high-velocity field of the side-wall reflection plate of the 7- by 10-foot tunnel at a range of angles of attack and Mach numbers. The results indicated that the maximum lift coefficients obtainable increased with increase in sweep angle and decreased with Mach number at the lower subsonic Mach numbers.

Report presenting an investigation of the stage work distribution and losses through the turbine at design conditions of the XJ35-A-23 two-stage turbine. The investigation showed that the poor performance was primarily due to large whirl velocities at the turbine exit, choking in the second-stage rotor, and a choking condition upstream of the physical blade throat. If these conditions are eliminated, satisfactory two-stage turbine performance can be obtained.

Report presenting testing of a parabolic body of revolution in the 4- by 4-foot supersonic tunnel at two Mach numbers. The investigation included the effects of Reynolds number, fins, internal contour of body base, and two support systems on body pressure and force drag at an angle of attack of 0 degrees.