Report presenting steady straight-and-level and steady turning tracking runs against an aerial target using an F-51H airplane equipped with a fixed optical sight and with various combinations of maneuvering stick-force and stick-deflection gradients. Results regarding aim wander, elevator movement, and stick-force variation for various test conditions are provided.

The possibility of overshooting the anticipated normal acceleration as a result of the artificial-feel characteristics of the F-89C airplane at a condition of minimum static stability was investigated analytically by means of an electronic simulator. Several methods of improving the stick-force characteristics were studied. It is shown that, due to the lag in build-up of the portion of the stick force introduced by the bobweight, it would be possible for excessive overshoots of normal acceleration to occur in abrupt maneuvers with reasonable assumed control movements. The addition of a transient stick force proportional to pitching acceleration (which leads the normal acceleration) to prevent this occurring would not be practical due to the introduction of an oscillatory mode to the stick-position response. A device to introduce a viscous damping force would Improve the stick-force characteristics so that normal acceleration overshoots would not be likely, and the variation of the maximum stick force in rapid pulse-type maneuvers with duration of the maneuver then would have a favorable trend.

Report presenting an investigation of the feasibility of modifying the transient portion of the stick force in abrupt maneuvers in order to eliminate inadvertent normal-acceleration overshoots. The modification consists of additional stick force proportional to a quantity which leads the normal acceleration. The characteristics introduced by the inclusion of the force were considered to be very desirable by pilots.

Report presenting an investigation of internal film cooling of the exhaust nozzle of a 1000-pound-thrust liquid ammonia liquid-oxygen rocket engine. With water as a coolant, approximately 16 percent of the total propellant and coolant flow was required to film-cool the entire nozzle and with anhydrous liquid ammonia, approximately 19 percent of the total flow was required. Results regarding the coolant results and performance results are provided.

Note describing the variation in the number of revolutions per minute and provides simple formulas that allow for advanced calculation to be made of the variation of propeller speed with variation of one or more dimensions of the propeller and allow the necessary corrections to be applied.

Memorandum presenting an investigation in natural icing conditions to determine the effect of ice formations on the performance of an axial-flow turbojet engine. A description of the change over time in tail-pipe temperature, engine jet thrust, fuel flow, ice collection, and acceleration qualities are provided.

Memorandum presenting a flight investigation in natural icing conditions to determine the effect of ice formations on the performance of an axial-flow turbojet engine. Tail-pipe temperature increased from 761 to 1065 degrees Fahrenheit and the jet thrust decreased from 1234 to 910 pounds during a period of 45 minutes in icing. No general conclusions can be reached from the data because the icing condition was relatively light.

Report presenting a flight investigation in natural icing conditions to determine the effect of ice formations on the performance of an axial-flow turbojet engine. Results regarding the tail-pipe temperature, engine jet thrust, and characteristics of ice formation are provided. No general conclusions can be reached from the data because the icing condition was relatively light.

Report presenting a flight investigation to determine the comparative performance of AN-F-58 and AN-F-32 fuels in a 4000-pound-thrust turbojet engine. The fuels were equivalent over the range of conditions investigated. Results regarding corrected net thrust, corrected jet-fuel consumption, variation of corrected tail-pipe temperature, combustor blow-out speeds, and visual observations of the jet exhaust are provided.

A flight investigation in natural icing conditions was conducted to determine the effect of inlet ice formations on the performance of axial-flow turbojet engines. The results are presented for icing conditions ranging from a liquid-water content of 0.1 to 0.9 gram per cubic meter and water-droplet size from 10 to 27 microns at ambient-air temperature from 13 to 26 degrees F. The data show time histories of jet thrust, air flow, tail-pipe temperature, compressor efficiency, and icing parameters for each icing encounter. The effect of inlet-guide-vane icing was isolated and shown to account for approximately one-half the total reduction in performance caused by inlet icing.

Flight and ground investigations have been made to compare an exhaust-ejector installation with a standard exhaust-collector-ring installation on air-cooled aircraft engines in a twin-engine airplane. The ground investigation allowed that, whereas the standard engine would have overheated above 600 horsepower, the engine with exhaust ejectors cooled at take-off operating conditions at zero ram. The exhaust ejectors provided as much cooling with cowl flaps closed as the conventional cowl flaps induced when full open at low airspeeds. The propulsive thrust of the exhaust-ejector installation was calculated to be slightly less than the thrust of the collector-ring-installation.

From Introduction: "This report describes some measurements of these quantities obtained by crashing fighter aircraft under circumstances approximating those observed in service."

We are undertaking the task of computing the air forces on a slightly cambered airfoil in the absence of friction and with an infinite aspect ratio. We also assume in advance that the leading edge is very sharp and that its tangent lies in the direction of motion.

It is stated that the index value 6000, as found in normal tests of wing sections with a 20 cm chord, falls in the same region where the transition of laminar to turbulent flow takes place on thin flat plates. It is to be expected that slightly cambered, thin wing sections will behave similarly. The following test of two such wing sections were made for the purpose of verifying this supposition.

"The cavitation in nozzles on airfoils of various shape and on a sphere are experimentally investigated. The limits of cavitation and the extension of the zone of the bubbles in different stages of cavitation are photographically established. The pressure in the bubble area is constant and very low, jumping to high values at the end of the area. The analogy with the gas compression shock is adduced and discussed" (p. 1).

"Airfoils with their trailing edge cut away are often found on aircraft, as the fins on the hulls of flying boats and the central section of the wings for affording better visibility. It was therefore of some interest to discover the effect of such cutaways on the lift and drag and on the position of the center of pressure. For this purpose, systematic experiments were performed on two different airfoils, a symmetrical airfoil and an airfoil of medium thickness, with successive shortenings of their chords" (p. 1).

Wind tunnel construction and design is discussed especially in relation to subsonic and supersonic speeds. Reynolds Numbers and the theory of compressible flows are also taken into consideration in designing new tunnels.

Some of the problems considered in this report include: thermodynamics of surface friction, application of thick wing sections, special applications of controllable propellers, and gas turbines for aircraft.

This report presents the results of various experiments carried out at the Gottingen Aerodynamic Institute. These include: experiments with Joukowski wing profiles; experiments on an airplane model with a built-in motor and functioning propeller; and the rotating cylinder (Magnus Effect).

"Through the utilization of the "Magnus effect" on the Flettner rotor ship, the attention of the public has been directed to the underlying physical principle. It has been found that the Prandtl boundary-layer theory furnishes a satisfactory explanation of the observed phenomena. The present article deals with the prevention of this separation or detachment of the flow by drawing the boundary layer into the inside of a body through a slot or slots in its surface" (p. 1).

"Hot-air engines with dynamic compressors and turbines offer new prospects of success through utilization of units of high efficiencies and through the employment of modern materials of great strength at high temperature. Particular consideration is given to an aerodynamic prime mover operating on a closed circuit and heated externally. Increase of the pressure level of the circulating air permits a great increase of limit load of the unit. This also affords a possibility of regulation for which the internal efficiency of the unit changes but slightly. The effect of pressure and temperature losses is investigated" (p. 1).

"Our attempts to improve the properties of airfoils by removing the boundary layer by suction, go back to 1922. The object of the suction is chiefly to prevent the detachment of the boundary layer from the surface of the airfoil. At large angles of attack, such detachment prevents the attainment of the great lift promised by the theory, besides greatly increasing the drag, especially of thick airfoils. This report gives results of those experiments" (p. 1).

The mutual influences of compression shocks and friction boundary layers were investigated by means of high speed wind tunnels.Schlieren optics provided a clear picture of the flow phenomena and were used for determining the location of the compression shocks, measurement of shock angles, and also for Mach angles. Pressure measurement and humidity measurements were also taken into consideration.Results along with a mathematical model are described.

Drag tests at zero lift have been made at Mach numbers from 0.7 to approximately 0.95 in the high speed wind tunnel of the Institute of Aerodynamics, ETH, Zurich, on a group of untapered wings of aspect ratio 3.25, having sweep angles of 0 degree and 35 degrees. For each sweep angle, a series of geometrically similar models was tested at a constant Reynolds number to provide a verification of computed tunnel blocking corrections. Tests were also made for wings having thickness ratios of 0.09 and 0.12 and the results compared with results predicted by von Karman's similarity law.