From Summary: "Flight measurements of the aerodynamic wing loads on the D-558-II airplane have been made in the Mach number range from 1.0 to 2.0. Nonlinear wing-panel characteristics occurred with variations in angle of attack. These nonlinear characteristics were apparent primarily at the lower supersonic speeds."

An investigation has been made at Mach numbers of 1.61 and 2.01 and Reynolds numbers from 1.7 X 10 to 7.6 X 10 to determine the pressure distributions over a 60 deg. delta wing having 20 different control configurations. Measurements were made at angles of attack from 0 deg to 15 deg for control deflections from -30 deg to 30 deg. This report presents the complete tabulated pressure data for the range of test conditions.

Memorandum presenting an investigation conducted to determine the feasibility of increasing the tip-section blade-element efficiency of a transonic rotor by reducing the blade loading for a given design inlet relative Mach number. This report primarily discusses the overall and blade-element performances of the modified rotor and compares them with the performance of the original rotor.

From Summary: "Data were obtained with five exhaust-nozzle areas and with the variable-area exhaust nozzle interlinked with the control system at conditions simulating flight at a Mach number of 0.8 and altitudes of 35,00 and 45,000 feet. Data simulating operation at zero flight Mach number at an altitude of 15,000 feet are also included. Engine component performance data are presented in addition to the overall engine performance."

"An investigation has been made at Mach numbers of 1.61 and 2.01 and Reynolds numbers from 1.7 x 10(exp 6) to 7.6 x 10(exp 6) to determine the pressure distributions over a 60 degree delta wing having 20 different control configurations. Measurements were made at angles of attack from 0 to 15 degrees for control deflections from -30 to 30 degrees. This report presents the complete tabulated pressure data for the range of test conditions" (p. 1).

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 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.

From Summary: "An analysis was made to permit the calculation of the effectiveness of oxide coatings in retarding the transient heat flow into turbine blades when the combustion gas temperature of a turbojet engine is suddenly changed. The analysis is checked with experimental data obtained from a turbojet engine whose blades were coated with two different coating materials (silicon dioxide and boric oxide) by adding silicone oil and tributyl borate to the engine fuel. The very thin coatings (approximately 0.001 in.) that formed on the blades produced a negligible effect on the turbine-blade transient temperature response. With the analysis discussed here, it was possible to predict the turbine rotor-blade temperature response with a maximum error of 40 F."

The design principle of injecting liquid fuel at more than one axial station in an annual turbojet combustor was investigated. Fuel was injected into the combustor as much as 5 inches downstream of the primary fuel injectors. Many fuel-injection configurations were examined and the performance results are presented for 11 configurations that best demonstrate the trends in performance obtained. The performance investigations were made at a constant combustor-inlet pressure of 15 inches of mercury absolute and at air flows up to 70 percent higher than values typical of current design practice. At these higher air flows, staging the fuel introduction improved the combustion efficiency considerably over that obtained in the combustor when no fuel staging was employed. At air flows currently encountered in turbojet engines, fuel staging was of minor value. Radial temperature distribution seemed relatively unaffected by the location of fuel-injection stations.

Memorandum presenting wing and tail-load data on a swept-wing airplane to determine the flight conditions and maneuvers in which maximum wing and tail loads were experienced, and, where pertinent, to relate these loads to the important stability and control changes that occurred. The results indicated that maximum wing loads and bending moments would be expected at relatively low Mach numbers. With increasing Mach number, a relieving effect on the wing-panel loading coefficients was noted, which appears to be due to premature flow separation on the outboard wing sections.

Memorandum presenting wing and tail-load data on a swept-wing fighter airplane analyzed to determine the flight conditions and maneuvers in which maximum wing and tail loads were experienced, and, where pertinent, to relate these loads to the important stability and control changes that occurred. The results indicated that maximum wing loads and bending moments would be expected at relatively low Mach numbers. Results are also presented on the use of controls in the various maneuvers for which loads data were obtained.

A wind-tunnel investigation to determine the downwash angles in a transverse plane behind the wings of six wing-body combinations at Mach number 0.25. A triangular wing of aspect ratio 2.0 and a trapezoidal wing of aspect ratio 3.0 were tested in combination with three geometrically similar slender bodies of revolution.

Report presenting a low-speed wind-tunnel investigation to determine the aerodynamic derivatives arising from acceleration in sideslip for a 60 degree triangular, a 45 degree sweptback, and an unswept wing. The magnitude of the sideslip acceleration derivatives encountered and the effect of wing plan form on the derivatives was small in the low angle-of-attack range. Results regarding acceleration derivatives and sideslip derivatives are provided.

From Introduction: "The general objective of this report is a study of the characteristics of the described form of peak holding control. The control has not yet been used in experimental turbojet-engine study.This study is concerned with only one method of extracting the maximum amount of information from a peak output for the purpose of maintaining that peak. The criteria sought in this control process are: maintenance of peak average output, minimum duration of time in undesirable operation, quickness in response to command to reach peak condition, least amount of excursion in the undesirable region, and type of stability."

Memorandum presenting a modified group method used to calculate the reactivity of the uranyl-fluoride-water experimental critical assemblies which are part of the Oak Ridge critical mass studies. The method relates the experimental slowing-down length for water to that in fluoride solutions and includes the effects of epithermal absorption and fission.

"The NACA is investigating a series of J-33 turbojet-engine compressors to determine the over-all and component performances and to improve theories of flow through large centrifugal compressors, The production model J-33-A-21 was operated over a range of inlet temperatures from 80 to -40 F and inlet pressures from 14 to 5 inches mercury absolute for equivalent impeller speeds from 6000 to 13,400 rpm. At the equivalent design speed of 11,500 rpm, the compressor had a peak pressure ratio of 3.98 at an equivalent weight flow of 73.4 pounds per second and an adiabatic temperature-rise , efficiency of 0.701" (p. 1).