The third report in a series of five that present unanalyzed pressure data obtained in tests of five full-scale propellers with NACA 16-series blade sections. Pressure distributions on the blade sections were measured under operating conditions to determine the aerodynamic characteristics of each blade section. This particular report presents information for nine radial stations of the NACA 10-(3)(090)-03 propeller.

Report presenting exploratory testing obtained with leading-edge wing chord-extensions on two 52 degree sweptback wings. The wings had the same aspect ratio but different airfoil sections. Results regarding force characteristics and flow characteristics are provided.

"A complete stage of an axial-flow compressor was designed and built to investigate the possibility of obtaining a high pressure ratio with an acceptable efficiency through the use of the optimum combination of high blade loading and high relative inlet Mach number. Over-all stage performance was investigated over a range of flows at equivalent tip speeds of 418 to 836 feet per second. At design speed (836 ft/sec), a peak total-pressure ration of 1.445 was obtained with an adiabatic efficiency of 0.89. For design angle of attack at the mean radius, a total-pressure ratio of 1.392 was obtained" (p. 1).

Report presenting a wind-tunnel investigation to explore the possibility of employing boundary-layer suction slots as a means for delaying laminar separation at the leading edge and turbulent separation over the rear portions of the airfoil section at high lift coefficients. Results regarding lift and drag are explored.

Memorandum presenting an investigation of 4811C ceramic blades and the problems involved in the adaption of ceramic materials to stressed turbine components. Results regarding spin tests and rotating heat-shock runs are provided.

Report presenting an investigation of a submerged air scoop consisting of a conventional scoop located in a dimple in the fuselage surface at low speeds. Results regarding a study of the basic inlet without suction, a comparison of arrangements using boundary-layer control, the performances of the most desirable configurations with a certain boundary layer, and variation of boundary-layer thickness are provided.

Report presenting an investigation by the transonic-bump method through a range of Mach numbers in the high-speed 7- by 10-foot tunnel to determine the lateral-control characteristics of 30-percent-chord flap-type controls of various spans and locations. The wing of the semispan fuselage-wing combination had 60 degrees of sweepback of the quarter-chord line, a taper ratio of 0.6, an aspect ratio of 4.0, and an NACA 65A006 airfoil section parallel to the free air stream.

From Introduction: "This paper presents the results of an investigation of force and moment characteristics for a wing with an unswept quarter-chord line, aspect ratio 2, taper ratio 0.78, and an NACA 65A004 airfoil section parallel to the free stream."

Report presenting testing of a semispan model of a wing with the leading edge swept back 37.25 degrees, an aspect ratio of 6.04, and a taper ratio of 0.5 in order to ascertain the compressibility effects on the forces, moments, and surface pressures. Lift, drag, and pitching-moment data along with chordwise distribution of static pressure at five spanwise stations are presented for a range of Mach numbers and a constant Reynolds number. Results regarding force and moment characteristics, critical flow and drag increase at high subsonic Mach numbers, and pressure changes with increasing Mach number are provided.

Report presenting an analytical method for calculating the maximum wall temperature of a tail-pipe burner or a ramjet combustion chamber cooled by air or gas flowing through a surrounding annular passage.

Report presenting a theoretical study of the dynamic lateral stability characteristics of an airplane with a 60 degree triangular wing. Calculations included the determination of the neutral-lateral-oscillatory-stability boundary, the period and time to damp to one-half amplitude of the lateral oscillation, and the time to damp to one-half amplitude for the spiral mode. The airplane was generally found to be stable for all conditions.

Report presenting an investigation of the low-speed longitudinal stability and control characteristics of a scale model of a Bell X-5 airplane design with various leading-edge slat and trailing-edge flap arrangements. A marked increase in stability with increasing sweep angle at low lift coefficients was noted. Results regarding the basic longitudinal characteristics, leading-edge slats, trailing-edge flaps, and longitudinal control are provided.

Report presenting an investigation of the longitudinal aerodynamic characteristics of a 47.7 degree sweptback wing in the 19-foot pressure tunnel in order to provide large-scale data on relatively high-aspect-ratio sweptback wings. Results regarding pitching-moment characteristics, lift characteristics, lift-drag ratio, and effect of roughness are provided.

Report presenting testing of an NACA octagonal transonic slotted test section has been modified to include slots with point origins located at the tunnel aerodynamic minimum and which diverged linearly over a length of 1.5 jet diameters. Pressure distributions along the axis and one wall of the slotted configuration are presented and indicate the attainment of a test region with satisfactory uniform Mach number distributions.

Report presenting an analysis of the flow conditions downstream of a cascade of turbine rotor blades at critical and supercritical pressure ratios. Five theoretical methods are compared with an experimental method using the conservation-of-momentum principle and static-pressure surveys as well as schileren photographs of the flow downstream of the blades. Results regarding static-pressure surveys and schileren photographs are provided.

Report presenting an investigation using the transonic-bump method through a range of Mach numbers in the high speed 7- by 10-foot tunnel to determine the lateral-control characteristics of 30-percent-chord flap-type controls with various spans and spanwise locations. Rolling moments, pitching moments, and lift were obtained through a small range of control deflections.

Memorandum presenting pressure-distribution and force data at various angles of attack of a triangular wing of aspect ratio 2 and having an NACA 0005 modified section. The wing had a plain, constant-chord, trailing-edge flap which was deflected 0 and 10 degrees. Results regarding the gross force characteristics of the wing, including lift, drag, pitching-moment, and flap hinge-moment coefficients are provided.

Report presenting an investigation of a rectangular wing with a subsonic-type round-leading-edge airfoil section (NACA 65-009) at various Mach numbers. The results generally agreed with calculated values with the exception of the laminar-separated region on the low-pressure surface of the model. Results regarding the calculation of pressure distribution, experimental pressure distributions, aerodynamic characteristics, and shadowgraph studies are provided.

From Summary: "By application of a hyperbolic approximation to the form of the bow waves caused by blunt leading edges on an infinite cascade of supersonic blades, the approximate losses in relative total pressure due to the external bow-wave system arising from blunt edges and subsonic axial entrance velocities were computed. The losses increase linearly with leading-edge radius for any given relative Mach number. For a relative Mach number of 1.60, leading-edge radii may be approximately 1.5 percent of the normal blade gap with a 1-percent loss of relative total pressure."

Memorandum presenting wind tunnel tests to determine the aerodynamic characteristics of a semispan wing with a modified NACA 0012-64 section and a 26.6-percent-chord, plain, trailing-edge aileron. Results are shown for the wing with the 0.229-chord line unswept and swept back 45 degrees and for a Mach number range from 0.40 to 0.925.

Report presenting a free-flight investigation of two rocket-powered model configurations to determine the damping in roll. The model had rectangular wings of 4.5 aspect ratio and were the same except for the airfoil sections. Results regarding rolling velocity and total-drag coefficient are provided.

An investigation was conducted to determine the effect of the inlet Mach number and entrance-boundary-layer thickness on the performance of a 23 degree 21-inch conical-diffuser - tail-pipe combination with a 2:1 area ratio. The air flows used in this investigation covered an inlet Mach number range from 0.17 to 0.89 and corresponding Reynolds numbers of 1,700,000 to 7,070,000. Results are reported for two inlet-boundary-layer thicknesses.

"The power plant from a Mark 25 aerial torpedo was investigated both as a two-stage turbine and as a single-stage modified turbine to determine the effect on overall performance of nozzle size and shape, first-stage rotor-blade configuration, and axial nozzle-rotor running clearance. Performance was evaluated in terms of brake, rotor, and blade efficiencies. All the performance data were obtained for inlet total to outlet static pressure ratios of 8, 15 (design), and 20 with inlet conditions maintained constant at 95 pounds per square inch gage and 1000 F for rotor speeds from approximately 6000 to 18,000 rpm" (p. 1).

"At the request of the Air Materiel Command an investigation was made in the Langley free-flight tunnel to determine the static longitudinal stability and control characteristics of models coupled together in a tandem configuration proposed by All American Airways, Inc. Force tests were made using 1/20-scale models of B-29 end F-80 airplanes to determine the effects of coupling the fighter to the tail of the bomber. The results of the investigation showed that for the bomber alone the aerodynamic center was 0.21 mean aerodynamic chord behind the center of gravity (stable) but that for the tandem configuration the aerodynamic center was 0.09 mean aerodynamic chord forward of the center of gravity, of the combination (unstable)" (p. 1).