The performance of NACA 65-series compressor blade section in cascade has been investigated systematically in a low-speed cascade tunnel. Porous test-section side walls and for high-pressure-rise conditions, porous flexible end walls were employed to establish conditions closely simulating two-dimensional flow. Blade sections of design lift coefficients from 0 to 2.7 were tested over the usable angle-of-attack range for various combinations of inlet-flow angle. A sufficient number of combinations were tested to permit interpolation and extrapolation of the data to all conditions within the usual range of application. The results of this investigation indicate a continuous variation of blade-section performance as the major cascade parameters, blade camber, inlet angle, and solidity were varied over the test range. Summary curves of the results have been prepared to enable compressor designers to select the proper blade camber and angle of attack when the compressor velocity diagram and desired solidity have been determined.
Forces and moments of store-pylon combination mounting on swept wing-fuselage configuration in supersonic pressure tunnel. The results of the investigation indicate that the most important source of store-pylon side forces is the pylon itself.
Report presenting an investigation of a semispan model of a possible vertical-take-off-and-landing jet bomber configuration at a range of Mach numbers. The primary objective was to determine to the effects of jet interference, horizontal-tail location, and canard controls on the longitudinal stability characteristics of the model.
Report presenting the performance of a two-dimensional side inlet with a technique of varying compression-surface angle while retaining a fixed-geometry diffuser at several Mach numbers and zero angle of attack. A 12 degree compression ramp was faired into the diffuser contour in this conventional manner. Results regarding the inlet flow field, application to reduced engine speeds, and a inlet performance with a sudden expansion in the diffuser are provided.
Report discussing heat-transfer and pressure measurements obtained from flight testing of a model of the Titan intercontinental ballistic missile up to a specified Mach and Reynolds number. The heat-transfer coefficients were compared to theoretical results and a discrepancy was found during the accelerating portions of flight. Drag coefficients were obtained for a range of Mach numbers.
Report presenting heat-transfer and pressure measurements obtained from flight tests of a model of the Titan intercontinental ballistic missile. Turbulent flow was observed over the model throughout the flight with the exception of one station on the nose. Results regarding pressure measurements, heat transfer, and drag measurements are required.
Report presenting an investigation to determine the effectiveness of a leading-edge slat and a trailing-edge split flap in improving the high subsonic speed aerodynamic characteristics of a model of a wing-fuselage combination with a nearly triangular wing. Results regarding the effects of Reynolds number, effects of slats, and effects of flaps are provided.
Report presenting a free-flight investigation of the zero-lift control effectiveness of a steady-flow jet-spoiler roll control on a cruciform 80 degree delta-wing missile over a range of Mach numbers. Measurements were made of the rolling-moment, damping-in-roll derivative, drag, and pressures at the inlet and on the wing near the jet exit. Results indicated that the wing and jet combination magnified the thrust force of the isolated jet alone by factors of 11 at subsonic speeds to 3 at supersonic speeds.
Memorandum presenting a transonic flutter investigation of models of the all-movable horizontal tail of a fighter airplane in the transonic blowdown tunnel. The models were dynamically and elastically scaled by criteria which provide a flutter safety margin. Results regarding some general comments, simulated airplane tests, effects of pitch stiffness with rearward pitch axis, and effect of pitch-axis location are provided.
Report presenting an investigation of the transonic flutter characteristics of models of an all-movable horizontal tail of a fighter airplane. Results regarding simulated airplane tests, the effects of pitch stiffness with rearward pitch axis, and the effect of pitch-axis location are provided.
Report presenting a wind-tunnel investigation to determine the effectiveness of area suction in increasing the lift of a moderately thick straight wing encountering trailing-edge air flow separation. Results regarding a model with undeflected leading-edge flap and tip tanks on and a model with the tip tanks removed are provided.
Report presenting some measurements of the vortex movements with time about an airfoil undergoing a blast of sufficient strength to exceed the stall angle by a large amount. Measurements were also obtained without flight simulation but with blast orientation and strength and compared to the first test. Results regarding Schileren photographs, presentation of results, and discussion of results are provided.
Report presenting measurements of aerodynamic heat transfer at a number of stations along a cone-cylinder-flare model with 15 degree total-angle conical nose and a 10 degree half-angle flare skirt. Results regarding temperature measurements and heating rates, local flow parameters, heat transfer with theoretical recovery factors, experimental recovery factors, prediction of skin temperatures, and transition are provided.
Memorandum presenting the results of a simulator stud of the factors affecting the design of a device, a stick pusher, for preventing a representative supersonic airplane from entering the pitch-up region. The effects of varying the stick-pusher-activation boundaries, sensing parameters, and magnitude of stick-pusher force on the controllability of the airplane pitch-up were investigated.
Report presenting the results of an investigation of the static longitudinal stability and control characteristics of a canard airplane configuration without analysis for a range of Mach numbers. Data are presented for a variety of angles of attack and canard angles.
Report presenting a boundary-layer-control investigation to determine the longitudinal aerodynamic characteristics and chordwise load distribution on a thin, untapered, semispan wing with an aspect ratio of 3.33 and NACA 65A004 airfoil sections. The results are presented without discussion.
Memorandum presenting flight tests conducted with an F-86D airplane equipped with a director-type radar fire-control system with scope presentation of the attack display. The effects of two attack-computer parameters and one attack-display parameter on the tracking performance in the manual mode of operation were investigated. Results regarding fixed-sight tracking and tracking with scope presentation are provided.
Report presenting flight testis using an F-86D airplane equipped with a director-type radar fire-control system with scope presentation of the attack display. The effects of two attack-computer parameters and one attack-display parameter on the tracing performance in manual mode were investigated.
Report presenting an investigation of the changes in the internal flow conditions resulting from adjusting the first-stator throat area of the J71 experimental three-stage turbine. The results of interstage flow measurements obtained from a radial survey investigation conducted on the turbine when equipped with a first-stage stator with the throat area increased to 132 percent of the design area and operated at a single predetermined turbine match point are also provided.
Memorandum presenting the overall performance results obtained for a 7-inch transonic turbine and a comparison with the results of a 14-inch turbine of geometrically similar design that had been previously investigated. The peak efficiency obtained with the turbine was 0.85, which was 2 points lower than that obtained previously with the 14-inch turbine.
Report presenting performance results for a 7-inch transonic turbine and a comparison of said turbine with a 14-inch turbine of geometrically similar design. The size effect on performance was found to generally be due to the inability to fabricate the smaller blades to the same percentage of dimensional deviation as that of larger blades. Results regarding overall performance and performance under specific inlet conditions are provided.
Memorandum presenting an experimental turbine designed for a high weight flow per unit frontal area, a high specific work output, a relative critical velocity ratio of 0.82 at the rotor hub inlet, and zero rotor blade suction-surface diffusion. Results regarding a comparison of overall performance of subject turbine with that of configurations I and II and comparison of rotor blade momentum thickness of subject turbine with that of configurations I and II are provided.
Report presenting an experimental turbine designed for high weight flow per unit frontal area, a high specific work output, a relative critical velocity ratio of 0.82 at the rotor hub inlet, and zero rotor blade suction-surface diffusion. At the equivalent design blade speed and work out put, the brake internal efficiency based on the actual overall total-pressure ratio was 0.889. Results regarding the comparison of overall performance of subject turbine with that of configurations I and II and comparison of rotor blade momentum thickness are provided.
Report presenting an investigation to determine the effectiveness of experimental control signals applied to a theoretical inlet throat Mach number control system and normal-shock-position control system for varying the inlet geometry of a twin-duct prototype aircraft. Results regarding inlet instability, control requirements, individual throat Mach number control signals, average throat Mach number control signals, theoretical thraot Mach number control analysis, effect of angle of attack, and theoretical normal-shock-position control are provided.
Report presenting an investigation in the supersonic wind tunnel to determine the effectiveness of experimental control signals applied to a theoretical inlet throat Mach number control system and normal-shock-position control system for varying the inlet geometry of a twin-duct, side-inlet, fuselage forebody model of a prototype aircraft. These types of control systems have been found to be particularly helpful in supersonic flight. Results regarding inlet stability, control requirements, individual throat Mach number control signals, average throat Mach number control signals, effect of angle of attack, and theoretical normal-shock-position control are provided.
Report presenting an investigation of a full-scale, four-shock, side-inlet diffuser in a free-jet facility at a Mach number of 3.16. The diffuser was a heavy-duty version of the diffuser to be used with the XRJ47-W-7 engine. Results regarding the supercritical capture-area ratio, critical diffuser total-pressure recovery, and diffuser-exit flow profiles are provided.
Report discussing findings on net thrust, fuel flow, and related performance indices for hydrogen-rich operation of nacelle-type and submerged ramjet engines at various Mach numbers. Testing concluded that the system could be helpful in a flight path involving acceleration from low to high flight speeds. Some trends of the effect of the operation conditions are noted, but no conclusions are drawn because of the preliminary nature of the report.
Report presenting an experimental investigation of a rectangular swept scoop inlet designed for and tested at a free-stream Mach number of 3.1 to determine the effect of inlet aspect ratio on the minimum inlet starting concentration ratio and the resulting inlet total-pressure recovery. Results regarding the determination of the starting contraction ratio, effect of inlet aspect ratio on the inlet starting characteristics, total-pressure recovery and mass-flow variations, and effects of inlet aspect ratio on the maximum total-pressure recovery are provided.
A brief investigation of a target-type thrust reverser on a single-engine fighter model has been conducted in the Langley 16-foot transonic tunnel at Mach numbers from 0.20 to 1.05. At Mach numbers of 0.80, 0.92, and 1.05, a hydrogen peroxide turbojet-engine simulator was operated with the thrust reverser extended. The angle of attack was varied from 0 degrees to 5 degrees at these Mach numbers.
Memorandum presenting a five-stage rocket-propelled model flown up to a Mach number of 10.9 and a corresponding free-stream Reynolds number of 6.57 x 10(exp 6) based on nose diameter. Temperatures were measured at 16 stations on the inside of a flat-faced cylinder made of copper. Results regarding the flight test, wall temperatures, heat-transfer data, measurement repeatibility and effects of angle of attack, comparison of measured and theoretical stagnation-point heating rates, and measured and theoretical heating rates over the front and side of the nose are provided.
Report presenting pressures measured during a free-flight test at zero angle of attack over a rocket model with a modified Von Karman nose in combination with a cylindrical center section and a 10 degree half-angle flare.
Report on an investigation to determine the erect and inverted spin and recovery characteristics of a model of the North American FJ-4 airplane. The testing found that either a flat-type of steep-type spin may be obtained when the airplane is spinning erect. The optimum recovery technique from inverted spins was full rudder reversal with simultaneous movement of the ailerons to full with the stick maintained full forward.
Report presenting eight thin magnesium fins, seven with the leading edges swept back 17 degrees and one with the leading edge swept back 45 degrees, in the preflight high-temperature jet. The investigation was made to determine the effectiveness of various protective coverings designed to alleviate aerodynamic-heating effects and intended for application on the first stage of rocket-propelled multistage hypersonic models.
"An investigation of tantalum was made in still air and in high-velocity air to determine its resistance to oxidation at high temperatures. Experiments show that, except for a narrow range of temperature, unprotected tantalum is not suitable for airframe parts exposed to air at high stagnation temperatures, because above this temperature range tantalum oxidizes or burns rapidly" (p. 1).
Report presenting a laboratory-scale vapor-deposition coating facility used in the field of high-temperature coatings. The vapor-deposition coatings were produced by the hydrogen reduction of halides and not by vacuum plating methods. The coated model was found to reach equilibrium temperature of about 3100 degrees Fahrenheit and was undamaged after 470 seconds at equilibrium temperature, while an uncoated model was destroyed in less than 6 seconds.
Memorandum presenting testing of several materials including graphite, silicon carbide, and a number of polymer-glass-cloth laminated constructions at temperatures of 3,000 and 4,000 degrees Fahrenheit in a laboratory-scale ceramic-heated air jet. The tests were made to utilize four possible mechanisms for the alleviation of the aerodynamic-heating problem of hypersonic aircraft, which included radiative heat transfer, pyrolysis, fusion, and mass-transfer cooling by ablation.
Report presenting testing of three uninstrumented tapered magnesium fins with the leading edges swept back 17 degrees in an ethylene-heated high-temperature jet. The testing was carried out to investigate some effects of leading-edge diameter and leading-edge shape on the aerodynamic heating by noting the time for melting to begin on the fins. Results indicated that increasing the diameter of cylindrical leading edge increased the time required for melting to start.
Memorandum presenting an investigation of the distortion removal performance and associated total-pressure-loss characteristics of several freely rotating fans, three single-stage and two double-stage freely rotating fans over a range of radial and circumferential distortions of from 0 to 20 percent at inlet annulus Mach numbers from 0.30 to about 0.60. Results regarding individual fan performance, comparison of fan distortion removal characteristics, and fan total-pressure loss and operating speed characteristics are provided.
Memorandum presenting an investigation of the performance of a semielliptical scoop inlet with a two-dimensional flow field at the design Mach number of 2.0. The investigation included a study of the effects of inlet-leading-edge shape and boundary-layer bleed on the pressure recovery and total-pressure distribution. Results regarding performance of inlets without boundary-layer bleed, effect of diverter boundary-layer bleed, effect of leading-edge sweep angle, effect of compression-surface bleed, performance of inlet III-B, tests of buzz suppressors, and tests of rearward-facing control tubes are provided.
"Two short turbojet combustors designed for use with vaporized hydrocarbon fuels were tested in a one-quarter annular duct. The experimental combustors consisted of many small "swirl-can" combustor elements manifolded together. This design approach allowed the secondary mixing zone to be considerably reduced over that of conventional combustors" (p. 1).
"A program was conducted in an altitude facility at the NACA Lewis laboratory to investigate the effects of rapid inlet pressure oscillations on the operation of a current turbo jet engine. These pressure oscillations were approximately sinusoidal in form and were generated to cover a frequency range of 2 to 75 cycles per second and an amplitude range of 10 to 70 percent of the free-stream total pressure. As the oscillation progressed through the compressor, the amplitude was attenuated considerably and a relatively large phase shift (lag) occurred" (p. 1).
Report presenting an investigation of some effects of Reynolds number on the stability of a series of flared-body and blunted-cone models at a range of Mach numbers in three wind tunnels. The Reynolds number had a pronounced effect on the static stability of the flared-body models at the lower Mach numbers. Results regarding the static stability and damping in pitch are provided.
Report discussing an investigation of the stability of a series of flared-body and blunted-cone models at a range of Mach numbers. Increasing the flare length was found to increase static stability and the damping in pitch. The Reynolds number was found to have a large effect on static stability at lower Mach numbers, but little effect on the higher end of the range covered in testing.
Report presenting the aerodynamic characteristics of a new type of lateral control through a range of Mach numbers. The control consisted of airfoils mounted vertically at the tips of the wing and could be rotated to induce rolling moments or lift on the wing surface. Results regarding lateral-control characteristics and lift-control characteristics are provided.
Report presenting an investigation to determine the effects of magnitude and circumferential extent of inlet total-pressure distortions on the overall an component performance of a current turbojet engine. Results regarding the pressure and temperature profiles, component performance, engine pumping characteristics, net thrust and net-thrust specific fuel consumption are provided.
"Additional wind-tunnel tests were made of a 1/8-scale model of the Republic XP-91 airplane to determine its characteristics with various modifications. The modifications included a revised conventional tail, revised rocket arrangement, drooped wing tips, and revised landing gear and doors. Tests were also made to determine the effectiveness of the control surfaces of the model with the conventional tail and the effect of changing wing incidence and tail length. The revised rocket arrangement provided a considerable increase in the static directional stability contributed by the vee tail at small angles of yaw" (p. 1).
From Introduction: "In the present report, consideration is given to certain features intended to improve the accleration-limiting characteristics of a normal-accleration control system."
Report presenting a wind-tunnel investigation at Mach number 1.96 to determine the normal forces, pitching moments, and rolling moments contributed by each wing panel of a cruciform-wing and body combination over a wide range of combined angles of pitch and roll. The wings were triangular of aspect ratio 2 and the body was an ogive-cylinder combination. Results regarding individual panels and panel combinations are provided.
Note presenting a study of the corrosion resistance of 11 nickel-base compositions to molten sodium hydroxide at 1500 and 1700 degrees Fahrenheit in order to find a container material for the caustic at these temperatures. Results are provided for the solid-solution alloys, nickel with mechanically dispersed second phase materials, and the precipitation-hardened alloy.
Note presenting a method for solving problems associated with Laplace and Poisson equations which, in general, requires considerably fewer equations than the usual methods and which gives a convergent solution by the method of successive approximations.
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