From Summary: "Icing and drag investigations were conducted in the NACA Lewis icing research tunnel employing a combination thermal-pneumatic de-icer mounted on a 42-inch-chord NACA 0018 airfoil. The de-icer consisted of a 3-inch-wide electrically heated strip symmetrically located about the leading edge with inflatable tubes on the upper and lower airfoil surfaces aft of the heated area. The entire de-icer extended to approximately 25 percent of chord. A maximum power density of 9.25 watts per square inch was required for marginal ice protection on the airfoil leading edge at an air temperature of 00 F and an airspeed of 300 miles per hour."

Report presenting a wing-body combination with a plane triangular wing of aspect ratio 3 and NACA 0003-63 sections in streamwise planes at subsonic and supersonic Mach numbers. Lift, drag, and pitching moment are presented for a range of Mach and Reynolds numbers.

Memorandum presenting the observed gust experience of an airplane and information on the associated meteorological conditions to obtain a simple empirical relation for estimating the intensity of turbulence in the earth's friction layer. The relation does not discriminate between differences in turbulence intensity resulting from variations of flight altitude or diurnal variations of turbulence.

Effects of air-cooling turbine rotor blades on performance of a turbojet engine were calculated for a range of altitudes from sea level to 40,000 feet and a range of coolant flows up to 3 percent of compressor air flow, for two conditions of coolant bleed from the compressor. Bleeding at required coolant pressure resulted in a sea-level thrust reduction approximately twice the percentage coolant flow and in an increase in specific fuel consumption approximately equal to percentage coolant flow. For any fixed value of coolant flow ratio the percentage thrust reduction and percentage increase in specific fuel consumption decreased with altitude. Bleeding coolant at the compressor discharge resulted in an additional 1 percent loss in performance at sea level and in smaller increase in loss of performance at higher altitudes.

"Surface temperatures and heat transfer to the air stream have been measured for turbulent flow over a flat plate at a Mach number of 1.5 and at a Reynolds number, based on the momentum thickness of the boundary layer, of approximately 5000. Preliminary data are presented and the surface heat-transfer coefficients calculated from these data are considered to be accurate to plus or minus 2.6 percent at a temperature potential of 50 degrees Fahrenheit. These data are in good agreement with the results produced by applying modifications obtained from published information to existing subsonic theories" (p. 1).

Report presenting testing of a series of three wings over a range of Mach numbers by the use of the transonic-bump technique. The lift, drag, pitching-moment, and root-bending-moment data of wings of a variety of aspect ratios but with the same quarter-chord lines and airfoil sections were obtained.

The permeability characteristics and tensile strength of a porous material developed from stainless-steel corduroy wire cloth for use in transpiration-cooled walls where the primary stresses are in one direction were investigated. The results of this investigation are presented and compared with similar results obtained with porous sintered metal compacts. A much wider range of permeabilities is obtainable with the wire cloth than with the porous metal compacts considered and the ultimate tensile strength in the direction of the primary stresses for porous materials produced from three mesh sizes of wire cloth are from two to three times the ultimate tensile strengths of the porous metal compacts.

Memorandum presenting an investigation to determine the low-speed aerodynamic characteristics of a 45 degree swept wing with leading-edge inlets. The wing had a constant chord and completely spanned the wind tunnel. Results regarding surface-pressure characteristics, lift and pitching-moment characteristics, wake-drag characteristics, and internal-flow characteristics are provided.

Report presenting tests of NACA 65-series compressor blades cambered to an isolated airfoil lift coefficient of 1.2 at several maximum section thicknesses to obtain the effect of maximum section thickness on section operating characteristics. Information about surface pressure trends, design angle-of-attack selection, drag, operating range, and critical Mach number is also provided. Changing the section thickness was not found to significantly affect the design angle of attack selected.

An experimental investigation of the aerodynamic characteristics of a slender, square-based body of revolution was conducted at a Mach number of 3.12 for angles of attack from 0 degree to 10 degrees and for Reynolds numbers from 2 x 10(exp 6) to 15 x 10(exp 6). Boundary-layer measurements at zero angle of attack are compared with several compressible flow formulating for predicting boundary-layer characteristics. Comparison of experimental pressure and force values with theoretical values showed good agreement for low angles of attack. The measured mean skin-friction coefficients agreed well with those predicted by Mangler's transformation for laminar flow over cones.

Performance of four experimentally investigated axially symmetric spike-type nose inlets is compared on basis of ram-jet-engine aircraft range and operational problems. At design conditions, calculated peak engine efficiencies varied 25 percent from the highest value which indicates importance of inlet design. Calculations for a typical supersonic aircraft indicate possible increase in range if engine is flown at moderate angle of attack and result in engine lift utilized. For engines with fixed exhaust nozzle, propulsive thrust increases with increasing heat addition in subcritical flow region in spite of increasing additive drag. For the perforated inlet there is a range of increasing total-temperature ratios in subcritical flow region that does not yield an increase in propulsive thrust. Effects of inlet characteristics on speed stability of a typical aircraft for three types of fuel control is discussed.

Report presenting testing of a series of seven 10-percent-thick unswept and untapered airfoils of aspect ratio 8.0 using the wing-flow method at a range of Mach numbers. The most significant characteristic noted was the occurrence of an unstable pitching moment variation with angle of attack for small angles of attack at Mach numbers of 0.85 to 0.90 for all airfoils with a conventional thickness distribution. Results regarding lift characteristics, moment characteristics, and drag characteristics are provided.