"Tabulated pressure coefficients and aerodynamic characteristics obtained in flight from pressure distributions over six chordwise rows of orifices on the right wing of the D-558-I research airplane (BuAero No. 37972) are presented. The data were obtained through a Mach number range of 0.80 to 0.89 and throughout the normal-force-coefficient range at Mach numbers of 0.61, 0.70, 0.855, and 0.88. This paper supplements similar tabulated data which have been presented in NACA RM L50J10 and NACA RM L50L12a" (p. 1).

Report presenting an investigation conducted at low speed of an air-induction system with twin scoop-type inlets on the sides of the fuselage of a model of an airplane designed for flight at supersonic Mach numbers. The leading edge of the lip of the scoop was rounded and there were provisions for control of the boundary layer ahead of the inlet. Results regarding the characteristics of the lips investigated, characteristics of the main duct with lip E, characteristics of the boundary layer control duct, and pressure coefficient distributions and boundary-layer characteristics are provided.

Report presenting results from a tank investigation of a powered dynamic model of the Grumman JRF-5 airplane fitted with a single hydro-ski and extended afterbody. Results indicated that the afterbody extension in place of a tail ski reduced the maximum resistance by 10 percent. Removal of the wing-tip skids gave a further reduction in maximum resistance of 3 percent.

Report presenting an investigation in the 300 mph tunnel to determine the low-speed stability and control characteristics and the jet-engine duct-inlet pressure recovery characteristics of a model of the Edo 142 hydro-ski research airplane. Results regarding the longitudinal stability and control, stall-control devices, alighting gear, lateral stability, directional control, lateral control, and duct pressure recovery are provided.

"A rocket-propelled model of the X-3 configuration equipped with an all-movable tail of aspect ratio 3.0 has been flown to determine the longitudinal characteristics of this configuration at high lift coefficients. An analysis of the response of the model to rapid deflections of the horizontal tail gave information on lift, drag, longitudinal stability, and longitudinal trim change. The primary result of the tests was that the configuration was indicated to have very unstable tendencies at lift coefficients above the stall and at Mach number near 0.7" (p. 1).

An investigation was conducted to study the mechanisms of blade failure of forged Hastelloy B and cast Stellite 21. The blades were mounted in a 16-25-6 alloy rotor and subjected to 20-minute cycles consisting of 15 minutes at rated speed and approximately 5 minutes at idle. The first failures of the Hastelloy B and Stellite 21 blades were probably the result of excessive vibratory stresses and occurred after 14.25 and 16.75 hours, respectively. After 28.75 hours of operation, all but 3 of the original 25 Hastelloy B blades had either failed or contained stress-rupture-type cracks and four of the original 27 Stellite 21 blades contained stress-rupture-type cracks.

Report presenting testing of wing-body combinations incorporating several wing plan forms of theoretical interest at Mach numbers of 1.62, 1.93, and 2.41 in the supersonic tunnel. One triangular, one arrow, and tow diamond plan forms were tested. Results regarding minimum drag, lift-curve slope, center-of-pressure positions, drag due to lift, and lift-drag ratios are provided.

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.

Report presenting an investigation of a wing-body combination with a plane triangular wing of aspect ratio 4 and 3-percent-thick rounded nose 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.

Report presenting buffet boundaries, buffeting-load increments for the stabilizers and elevators, and buffeting bending-moment increments for the stabilizers and wings for a jet-powered bomber airplane equipped with reflexed flaps and ailerons and tail-tip-incidence changes.

Memorandum presenting a satisfactory correlation obtained between experimental sea-level transient data at constant flight speed and solutions from the analog representation. The representation is accomplished by transfer functions formed from a frequency-response analysis, which is then used to compute system response at altitude.

Report presenting an investigation by the wing-flow method to determine the snaking characteristics of a scale partial-model of a research airplane. Additionally, the snaking characteristics of 11 modified configurations of the research-airplane model were determined as well as a model that had a greater fineness ratio.

"An investigation of two semispan wings swept back 50 degrees 38 minutes was conducted in the Langley high-speed 7- by 10-foot tunnel over a Mach number range of 0.68 to 1.15 by use of the transonic-bump technique. This paper presents the results of the investigation of wing-alone and wing-fuselage configurations of the two wings; one was an untwisted uncambered wing and the other was the same wing but incorporated twist and camber designed to give uniform load at a lift coefficient of 0.25 at a Mach number of 1.10. The semispan wings had their quarter-chord lines swept back 50 degrees 38 minutes, aspect ratios 2.98, taper ratios 0.45, and modified NACA 64A-series airfoil sections tapered in thickness" (p. 1).

"An investigation was conducted in a 3.4- by 3.4-inch duct to determine the characteristics of the supersonic flow downstream of four wire-mesh screen nozzles with nominal design Mach numbers in the range between 1.97 and 2.58. Two types of disturbances were observed in the flow field: a fine network of interacting expansion and compression waves which were formed immediately downstream of the screens and appeared to dissipate within 25 to 40 wave intersections; and relatively strong oblique shock waves that originated at the junctions of the screens and the walls and were reflected throughout the length of the duct. Regions of fairly uniform flow were found to exist" (p. 1).

Report presenting the results of an experimental investigation of the downwash and longitudinal stability characteristics of a semispan, triangular-wing airplane model with a horizontal tail. Lift, drag, and pitching-moment data re presented for the wing-fuselage combination. Results indicated that the model with the tail on or below the wing-chord plane possessed satisfactory stability characteristics throughout the test range of lift coefficient.

The nature of the flow at the exit of a row of turbine blades for the range of conditions represented by four different blade configurations was evaluated by the conservation-of-momentum principle using static-pressure surveys and by analysis of Schlieren photographs of the flow. It was found that for blades of the type investigated, the maximum exit tangential-velocity component is a function of the blade geometry only and can be accurately predicted by the method of characteristics. A maximum value of exit velocity coefficient is obtained at a pressure ratio immediately below that required for maximum blade loading followed by a sharp drop after maximum blade loading occurs.

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.

Investigation to determine lift, drag, and pitching-moment characteristics of several engine-strut-body combinations was conducted over range of angles of attack from 0 degrees to 10 degrees at Mach numbers of 1.8 and 2.0. The average Reynolds number based on body length was 28x106. Data are presented without analysis and indicate decreases in minimum drag and lift curve slope with decreasing in minimum drag and lift curve slope with decreasing strut length. Decreases in minimum drag also noted with rear-ward movement of engines.

Report presenting the results of theoretical incompressible-flow calculations of the spanwise lift distributions, lift-curve slopes, spanwise centers of pressure, aerodynamic centers, coefficients of damping in roll, rolling-moment coefficients due to aileron deflection, and induced-drag coefficients of twenty M, W, and delta wings.

Report presenting the results of nine experimental flutter tests involving two flat-plate models and two rib and spar models of NACA 0012 airfoil sections in order to give a comparison of the flutter characteristics of wings with M and W type plan forms. The W plan form generally gave higher flutter speeds than the M plan forms.

Report presenting tests using the NACA wing-flow method at a range of Mach numbers to determine the aerodynamic characteristics of four triangular-wing-fuselage models. Measurements of the normal force, chord force, and pitching moment for various angles of attack. Two models had wings of aspect ratio 2.31, one with an NACA 65-009 airfoil section and one with a 9-percent-thick biconvex section, and two with wings of aspect ratio 4, one with an NACA 65-006 airfoil and one with a 6-percent-thick double-wedge airfoil.

A preliminary analytical evaluation of the air and fuel specific-impulse characteristics of hydrogen, a-methylnapthalene, and graphite carbon has been made. Adiabatic constant-pressure combustion flame temperatures for each fuel at several equivalence ratios were calculated for an initial air temperature of 560 degrees R and a pressure of 2 atmospheres.

Memorandum presenting an investigation in the 20-foot free-spinning tunnel on a 1/20-scale model to determine the effect of two alternate nose designs on the spin and recovery characterstics of the McDonnel XF3H-1 airplane. The results indicated that, in order to insure satisfactory recovery from any developed spin obtained in the airplane with either of the two alternate nose designs, it may be necessary to use ailerons deflected with the spin in conjunction with rudder reversal, as was the case for the original nose design.

Report presenting an investigation in the free-flight tunnel to determine the dynamic lateral stability and control characteristics of a high-aspect-ratio bomber model with self-supporting, free-floating fuel tanks attached to the wing tips. The results showed that the flight behavior of the configuration could be made satisfactory when sufficient restoring moment was supplied. Results regarding the flight behavior, rolling motions, motions of the bomber and attached tanks, and flight records for different gearing ratios are provided.