ALBERT

Description: A design guide is suggested as a basis for indicating combinations of airplane design variables for which the possibilities of pitch-up are minimized for tail-behind-wing and tailless airplane configurations. The guide specifies wing plan forms that would be expected to show increased tail-off stability with increasing lift and plan forms that show decreased tail-off stability with increasing lift. Boundaries indicating tail-behind-wing positions that should be considered along with given tail-off characteristics also are suggested. An investigation of one possible limitation of the guide with respect to the effects of wing-aspect-ratio variations on the contribution to stability of a high tail has been made in the Langley high-speed 7- by 10-foot tunnel through a Mach number range from 0.60 to 0.92. The measured pitching-moment characteristics were found to be consistent with those of the design guide through the lift range for aspect ratios from 3.0 to 2.0. However, a configuration with an aspect ratio of 1.55 failed t o provide the predicted pitch-up warning characterized by sharply increasing stability at the high lifts following the initial stall before pitching up. Thus, it appears that the design guide presented herein might not be applicable when the wing aspect ratios lower than about 2.0.

Description: Wind tunnel data of X-15 and B-52 aircraft models carry loads and mutual interference

Description: Wind tunnel tests - effect of wind induced loads on dynamically scaled model of large missile in launching position

Description: An experimental investigation was conducted to determine the performance characteristics an underslung nose-scoop air-induction system for a supersonic airplane. Five different nose shapes, three lip shapes, and two internal diffusers were investigated. Tests were made at Mach numbers from 0 to 1.9, angles of attack from 0 deg to approximately l5 deg, and mass-flow ratios from 0 to maximum obtainable. It was found that the underslung nose-scoop inlet was able to operate at Mach numbers from 0.6 to 1.9 over a large positive angle-of-attack range without adverse effects on the pressure recovery. Although there was no one inlet configuration that was markedly superior over the entire range of operating variables, the arrangement having a nose designed to give increased supersonic compression at low angles of attack, and a sharp lip (configuration designated N3L3) showed the most favorable performance characteristics over the supersonic Mach number range. Inlets with sizable lip radii gave satisfactory performance up to a Mach number of 1.5; however, as a result of an increase in drag, the performance of such inlets was markedly inferior to the sharp-lip configuration above Mach numbers of 1.5. Throughout the range of test Mach numbers all inlet configurations evidenced stable air-flow characteristics over the mass-flow range for normal engine operation. Analysis of the inlet performance on the basis of a propulsive thrust parameter showed that a fixed inlet area could be used for Mach numbers up to 1.5 with only a small sacrifice in performance.

Description: High subsonic speed of static longitudinal aerodynamic characteristics of delta wing configuration for angle of attack from 0 deg to 90 deg

Description: Stability and control of variable sweep wing configuration with outboard wing panels swept back 75 degrees at Mach 2.01

Description: Zero angle of attack performance of isentropic spike inlet designed for maximum external compression at hypersonic speed

Description: An investigation of some aspects of the sonic boom has been made with the aid of wind-tunnel measurements of the pressure distributions about bodies of various shapes. The tests were made in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 2.01 and at a Reynolds number per foot of 2.5 x 10(exp 6). Measurements of the pressure field were made at orifices in the surface of a boundary-layer bypass plate. The models which represented both fuselage and wing types of thickness distributions were small enough to allow measurements as far away as 8 body lengths or 64 chords. The results are compared with estimates made using existing theory. To the first order, the boom-producing pressure rise across the bow shock is dependent on the longitudinal development of body area and not on local details. Nonaxisymmetrical shapes may be replaced by equivalent bodies of revolution to obtain satisfactory theoretical estimates of the far-field pressures.

Description: Time histories of noise pressures near ground level were measured during flight tests of fighter-type airplanes over fairly flat, partly wooded terrain in the e Mach number range between 1.13 and 1.4 and at altitudes from 25,000 to 45,000 feet. Atmospheric soundings and radar tracking studies were made for correlation with the measured noise data. The measured and calculated values of the pressure rise across the shock wave were generally in good agreement. There is a tendency for the theory to overestimate the pressure at locations remote from the track and to underestimate the pressures for conditions of high tailwind at altitude. The measured values of ground-reflection factor averaged about 1.8 f or the surface tested as compared to a theoretical value of 2.0. P o booms were measured in all cases. The observers also generally reported two booms; although, in some cases, only one boom was reported. The shock-wave noise associated with some of the flight tests was judged to be objectionable by ground observers, and in one case the cracking of a plate-glass store window was correlated in time with the passage of the airplane at an altitude of 25,000 feet.

Description: Transonic wind tunnel study of aerodynamic characteristics of blunt reentry vehicles at varying angles of attack

Description: Hypersonic flutter tests on rectangular flat-plate models and double-wedge airfoils in helium flow

Description: Wind tunnel studies at supersonic and transonic speeds to determine aerodynamic characteristics of variable sweep wing aircraft - configuration

Description: Mach number and air temperature effect on hypersonic flow over blunt bodies

Description: Overall stage and stator blade element performance with straight stator and tilted stator in transonic axial flow compressor stage

Description: Pressure measurements in flight over conically cambered delta wing of F-102A aircraft at transonic speeds

Description: A solution of the equations of the compressible laminar boundary layer including the effects of transpiration cooling is presented. The analysis applies to the flow over an isothermal porous plate with a velocity of fluid injection proportional to the reciprocal of the square root of the distance from the leading edge. The effect of several flow parameters on coolant-flow rates is discussed with the aid of representative examples. A stability analysis indicates that, although transpiration cooling requires a lower surface temperature for stable flow than does internal wall cooling, this lower temperature can be obtained with a smaller expenditure of coolant.

Description: The temperature distributions encountered in thin solid wings subjected to aerodynamic heating induce thermal stresses that may effectively reduce the stiffness of the wing. The effects of this reduction in stiffness were investigated experimentally by rapidly heating the edges of a cantilever plate. The midplane thermal stresses imposed by the nonuniform temperature distribution caused the plate to buckle torsionally, increased the deformations of the plate under a constant applied torque, and reduced the frequency of the first two natural modes of vibration. By using small-deflection theory and employing energy methods, the effect of nonuniform heating on the plate stiffness was calculated. The theory predicts the general effects of the thermal stresses, but becomes inadequate as the temperature difference increases and plate deflections become large.

Description: Low speed measurements of oscillatory lateral stability derivatives of 60 degree delta wing bomber model

Description: Pitch and control stiffness effects on flutter characteristics of all-moveable wing and vertical and horizontal tails on fighter aircraft at supersonic speeds

Description: Aerodynamic effects of airfoil thickness on transonic flutter characteristics of swept and unswept wings

Description: Aircraft body flare for pitch stability and body flap for pitch control in hypersonic flight

Description: Effect of forebody strakes on aerodynamic characteristics in sideslip and pitch of hypersonic aircraft configurations

Description: Determination of loads due to wing twist at transonic and low supersonic speeds

Description: Investigation of amplitude and phase shift of static pressure variations in supersonic diffuser for separate oscillation of spike and bypass

Description: Fighter aircraft external stores ejection at transonic and supersonic speeds

Description: Aerodynamic characteristics of variable sweep aircraft configurations - low altitude supersonic vehicle

Description: A two-blade rotor having a diameter of 4 feet and a solidity of 0.037 was subjected to sharp-edge vertical gusts while being operated at various forward speeds to study the effect of the gusts on the blade periodic bending moments and flapping angles. Variables studied included gust velocity, collective pitch angle, flapping hinge offset, and tip-speed ratio. Dimensionless coefficients are derived for the periodic components of the incremental changes in blade flapping angles and bending moments which arise when a rotor blade penetrates a sharp-edge gust. Mental changes in both the flapping angles and bending moments are essentially proportional to gust velocity, and the coefficients express the ratio of these increments to gust velccity. The results show that the flapping coefficient usually increases with an increase in collective pitch angle, is generally dependent on tip-speed ratio, and is essentially independent of the amount of flapping hinge offset. The bending-moment coefficient is also dependent on collective pitch angle and tip-speed ratio. Expected reductions in bending moments are realized by the use of flapping hinges, and further reductions in bending moments are achieved as the amount of flapping hinge offset is increased. Comparison of the experimental results of this investigation with limited available theoretical results shows substantial agreement but indicates that the assumption that the response of the rotor to a sharp-edge gust is independent of the collective pitch angle prior to gust entry is probably inadequate.

Description: Measurements of near- and far-field noise pressures are presented for a 1,500-pound-thrust engine and for several 5,000-pound-thrust engines for which the nozzle exit pressure was changed systematically in order to study its effects on the noise level and spectra. Near-field surveys indicated that the highest noise pressure occurred at about 20 exit diameters downstream if the nozzle near the transition from super-sonic to subsonic flow. The acoustical power radiated from all engines averaged about 0.5 percent of the mechanical power of the exhaust stream, the least noise being radiated by the nozzle having an exit pressure less than atmospheric. The rocket engines of these tests radiate more power per cycle at the lower frequencies than arte reported for subsonic jets in other related studies.

Description: Constant chord, trailing edge, control deflection effects on aerodynamic loading characteristics of 60 degree delta wing-body combination at transonic speeds

Description: Wind tunnel determination of supersonic flow properties in vicinity swept, delta, and trapezoidal wing-body combinations

Description: Full-scale wind tunnel determination of effect of corrugated canister surface on static aerodynamic characteristics of reentry capsule

Description: Pressure distribution on hypersonic glide configuration having 79.5 deg sweepback and 45 deg dihedral at Mach number of 4.95

Description: The longitudinal aerodynamic characteristics of a wing-body-horizontal-tail configuration designed for efficient performance at transonic speeds has been investigated at Mach numbers from 0.80 to 1.03 in the Langley 16-foot transonic tunnel. The effect of adding an outboard leading-edge chord-extension to the highly tapered 45 deg. swept wing was also obtained. The average Reynolds number for this investigation was 6.7 x 10(exp 6) based on the wing mean aerodynamic chord. The relatively low tail placement as well as the addition of a chord-extension achieved some alleviation of the pitchup tendencies of the wing-fuselage configuration. The maximum trimmed lift-drag ratio was 16.5 up to a Mach number of 0.9, with the moment center located at the quarter-chord point of the mean aerodynamic chord. For the untrimmed case, the maximum lift-drag ratio was approximately 19.5 up to a Mach number of 0.9.

Description: The importance of atomizing and mixing liquid oxygen and heptane was studied in a 200-pound-thrust rocket engine. Ten injector elements were used with both steel and transparent chambers. Characteristic velocity was measured over a range of mixture ratios. Combustion gas-flow and luminosity patterns within the chamber were obtained by photographic methods. The results show that, for efficient combustion, the propellants should be both atomized and mixed. Heptane atomization controlled the combustion rate to a much larger extent than oxygen atomization. Induced mixing, however, was required to complete combustion in the smallest volume. For stable, high-efficiency combustion and smooth engine starts, mixing after atomization was most promising.

Description: During an investigation of the J57-P-1 turbojet engine in the Lewis altitude wind tunnel, effects of inlet-flow distortion on engine stall characteristics and operating limits were determined. In addition to a uniform inlet-flow profile, the inlet-pressure distortions imposed included two radial, two circumferential, and one combined radial-circumferential profile. Data were obtained over a range of compressor speeds at an altitude of 50,000 and a flight Mach number of 0.8; in addition, the high- and low-speed engine operating limits were investigated up to the maximum operable altitude. The effect of changing the compressor bleed position on the stall and operating limits was determined for one of the inlet distortions. The circumferential distortions lowered the compressor stall pressure ratios; this resulted in less fuel-flow margin between steady-state operation and compressor stall. Consequently, the altitude operating Limits with circumferential distortions were reduced compared with the uniform inlet profile. Radial inlet-pressure distortions increased the pressure ratio required for compressor stall over that obtained with uniform inlet flow; this resulted in higher altitude operating limits. Likewise, the stall-limit fuel flows required with the radial inlet-pressure distortions were considerably higher than those obtained with the uniform inlet-pressure profile. A combined radial-circumferential inlet distortion had effects on the engine similar to the circumferential distortion. Bleeding air between the two compressors eliminated the low-speed stall limit and thus permitted higher altitude operation than was possible without compressor bleed.

Description: A linear stability analysis and flight-test investigation has been performed on a rolleron-type roll-rate stabilization system for a canard-type missile configuration through a Mach number range from 0.9 to 2.3. This type damper provides roll damping by the action of gyro-actuated uncoupled wing-tip ailerons. A dynamic roll instability predicted by the analysis was confirmed by flight testing and was subsequently eliminated by the introduction of control-surface damping about the rolleron hinge line. The control-surface damping was provided by an orifice-type damper contained within the control surface. Steady-state rolling velocities were at all times less than 1 radian per second between the Mach numbers of 0.9 to 2.3 on the configurations tested. No adverse longitudinal effects were experienced in flight because of the tendency of the free-floating rollerons to couple into the pitching motion at the low angles of attack and disturbance levels investigated herein after the introduction of control-surface damping.

Description: Two full-scale models of an inline, cruciform, canard missile configuration having a low-aspect-ratio wing equipped with flap-type controls were flight tested in order to determine the missile's longitudinal aerodynamic characteristics. Stability derivatives and control and drag characteristics are presented for a range of Mach number from 0.7 to 1.8. Nonlinear lift and moment curves were noted for the angle - of-attack range of this test (0 deg to 8 deg). The aerodynamic-center location for angles of attack near 50 remained nearly constant for supersonic speeds at 13.5 percent of the mean aerodynamic chord; whereas for angles of attack near 0 deg, there was a rapid forward movement of the aerodynamic center as the Mach number increased. At a control deflection of 0 deg, the missile's response to the longitudinal control was in an essentially fixed space plane which was not coincident with the pitch plane as a result of the missile rolling. As a consequence, stability characteristics were determined from the resultant of pitch and yaw motions. The damping-in-pitch derivatives for the two angle -of-attack ranges of the test are in close agreement and varied only slightly with Mach number. The horn-balanced trailing-edge flap was effective in producing angle of attack over the Mach number range.

Description: A diamond wing and body combination was designed to have an area distribution which would result in near optimum zero-lift wave-drag coefficients at a Mach number of 1.00, and decreasing wave-drag coefficient with increasing Mach number up to near sonic leading-edge conditions for the wing. The airfoil section were computed by varying their shape along with the body radii (blending process) to match the selected area distribution and the given plan form. The exposed wing section had an average maximum thickness of about 3 percent of the local chords, and the maximum thickness of the center-line chord was 5.49 percent. The wing had an aspect ratio of 2 and a leading-edge sweep of 45 deg. Test data were obtained throughout the Mach number range from 0.20 to 3.50 at Reynolds numbers based on the mean aerodynamic chord of roughly 6,000,000 to 9,000,000. The zero-lift wave-drag coefficients of the diamond model satisfied the design objectives and were equal to the low values for the Mach number 1.00 equivalent body up to the limit of the transonic tests. From the peak drag coefficient near M = 1.00 there was a gradual decrease in wave-drag coefficient up to M = 1.20. Above sonic leading-edge conditions of the wing there was a rise in the wave-drag coefficient which was attributed in part to the body contouring as well as to the wing geometry. The diamond model had good lift characteristics, in spite of the prediction from low-aspect-ratio theory that the rear half of the diamond wing would carry little lift. The experimental lift-curve slope obtained at supersonic speeds were equal to or greater than the values predicted by linear theory. Similarly the other basic aerodynamic parameters, aerodynamic center position, and maximum lift-drag ratios were satisfactorily predicted at supersonic speeds.

Description: Two rocket configurations with turbopump drive were investigated analytically. In one configuration the inlet pressure to the turbine was fixed at the design value. The second configuration employed a "bootstrap" technique for supplying energy to the turbine. An injector was the chief resistance between the pump and the rocket combustion chamber. From the analysis two parameters were developed from which the speed response time of the turbopump, the flow response time, and the maximum dynamic line loss could be evaluated. These parameters were functions of turbopump moment of inertia, design performance of the turbine, and flow-system geometry. The moment of inertia of the turbopump and the ratio of turbine torque at zero speed to design torque had the most influence on the starting dynamics of the flow system. These parameters were also applicable to the bootstrap configuration as long as the inlet pressure to the turbine exceeded half the design value.

Description: An investigation of a model of a standard size body in combination with a representative 45 deg swept-wing-fuselage model has been conducted in the Langley 8-foot transonic pressure tunnel over a Mach number range from 0.80 to 1.43. The body, with a fineness ratio of 8.5, was tested with and without fins, and was pylon-mounted beneath the fuselage or wing. Force measurements were obtained on the wing-fuselage model with and without the body, for an angle-of-attack range from -2 deg to approximately 12 deg and an angle-of-sideslip range from -8 deg to 8 deg. In addition, body loads were measured over the same angle-of-attack and angle-of-sideslip range. The Reynolds number for the investigation, based on the wing mean aerodynamic chord, varied from 1.85 x 10(exp 6) to 2.85 x 10(exp 6). The addition of the body beneath the fuselage or the wing increased the drag coefficient of the complete model over the Mach number range tested. On the basis of the drag increase per body, the under-fuselage position was the more favorable. Furthermore, the bodies tended to increase the lateral stability of the complete model. The variation of body loads with angle of attack for the unfinned bodies was generally small and linear over the Mach number range tested with the addition of fins causing large increases in the rates of change of normal-force coefficient and nose-down pitching-moment coefficient. The variation of body side-force coefficient with sideslip for the unfinned body beneath the fuselage was at least twice as large as the variation of this load for the unfinned body beneath the wing. The addition of fins to the body beneath either the fuselage or the wing approximately doubled the rate of change of body side-force coefficient with sideslip. Furthermore, the variation of body side-force coefficient with sideslip for the body beneath the wing was at least twice as large as the variation of this load with angle of attack.

Description: The linearized theory for heat addition under a wing has been developed to optimize wing geometry, heat addition, and angle of attack. The optimum wing has all of the thickness on the underside of the airfoil, with maximum-thickness point well downstream, has a moderate thickness ratio, and operates at an optimum angle of attack. The heat addition is confined between the fore Mach waves from under the trailing surface of the wing. By linearized theory, a wing at optimum angle of attack may have a range efficiency about twice that of a wing at zero angle of attack. More rigorous calculations using the method of characteristics for particular flow models were made for heating under a flat-plate wing and for several wings with thickness, both with heat additions concentrated near the wing. The more rigorous calculations yield in practical cases efficiencies about half those estimated by linear theory. An analysis indicates that distributing the heat addition between the fore waves from the undertrailing portion of the wing is a way of improving the performance, and further calculations appear desirable. A comparison of the conventional ramjet-plus wing with underwing heat addition when the heat addition is concentrated near the wing shows the ramjet to be superior on a range basis up to Mach number of about B. The heat distribution under the wing and the assumed ramjet and airframe performance may have a marked effect on this conclusion. Underwing heat addition can be useful in providing high-altitude maneuver capability at high flight Mach numbers for an airplane powered by conventional ramjets during cruise.

Description: The performance characteristics of several flush and shielded auxiliary exits were investigated at Mach numbers of 1.5 to 2.0, and jet pressure ratios from jet off to 10. The results indicate that the shielded configurations produced better overall performance than the corresponding flush exits over the Mach-number and pressure-ratio ranges investigated. Furthermore, the full-length shielded exit was highest in performance of all the configurations. The flat-exit nozzle block provided considerably improved performance compared with the curved-exit nozzle block.

Description: A five-stage solid-fuel sounding-rocket system which can boost a payload of 25 pounds to an altitude of 525 nautical miles and that of 100 pounds to 300 nautical miles is described. Data obtained from a typical flight test of the system are discussed.

Description: Two methods for reducing the external cowl angle, and hence the cowl pressure drag, were investigated on a two-dimensional model. One method used at both on- and off-design Mach numbers was the addition of a cowl visor that had the inner surface parallel to the free stream at 0 deg angle of attack. The other method investigated consisted in replacing the original cowl by a flatter cowl that also provided internal contraction. Both the visor and the internal-contraction cowl reduced the cowl pressure drag 64 percent or more. The visor had little effect on inlet performance at the design Mach number except to reduce the stability range slightly. At off-design, the visor caused an increase in critical pressure recovery.

Description: A compilation of charts of the induced velocities near a lifting rotor is presented. The charts cover uniform as well as various non-uniform distributions of disk loading and should be applicable to many aerodynamic interference problems involving rotors.

Description: Semispan-wing models were tested at angles of attack from 0 to 180 deg at low subsonic speeds. Eight plan forms were considered, both swept and unswept with aspect ratios ranging from 2 to 6. Except for a delta-wing model of aspect ratio 2. all models had a taper ratio of 0.5 and an NACA 64AO10 airfoil section. The delta-wing model had an NACA 0005 (modified) airfoil section. With two exceptions, the models were tested both with and without a full-span trailing-edge flap deflected 25 deg. The Reynolds numbers based on the mean aerodynamic chord were between 1.5 and 2.2 million. Lift, drag, and pitching-moment coefficients are presented as functions of angle of attack. Approximate corrections for the effects of blockage were applied to the data.

Description: An investigation of the effects of variation of leading-edge sweep and surface inclination on the flow over blunt flat plates was conducted at Mach numbers of 4 and 5.7 at free-stream Reynolds numbers per inch of 6,600 and 20,000, respectively. Surface pressures were measured on a flat plate blunted by a semicylindrical leading edge over a range of sweep angles from 0 deg to 60 deg and a range of surface inclinations from -10 deg to +10 deg. The surface pressures were predicted within an average error of +/- 8 percent by a combination of blast-wave and boundary-layer theory extended herein to include effects of sweep and surface inclination. This combination applied equally well to similar data of other investigations. The local Reynolds number per inch was found to be lower than the free-stream Reynolds number per inch. The reduction in local Reynolds number was mitigated by increasing the sweep of the leading edge. Boundary-layer thickness and shock-wave shape were changed little by the sweep of the leading edge.

Description: Pressure distributions obtained in the Langley 8-foot transonic pressure tunnel on a thin, highly tapered, twisted, 450 sweptback wing in combination with a body are presented. The wing has a cubic spanwise twist variation from 0 deg. at 10 percent of the semispan to 60 at the tip. The tip is at a lower angle of attack than the root. Tests were made at stagnation pressures of 1.0 and 0.5 atmosphere, at Mach numbers from 0 0.800 to 1.200, and at angles of attack from -4 deg. to 20 deg.

Description: Surface pressures were measured over a blunt 60 deg delta wing with extended trailing edge at a Mach number of 5.7, a free-stream Reynolds number of 20,000 per inch, and angles of attack from -10 to +10 deg. Aft of four leading-edge thicknesses the pressure distributions evidenced no appreciable three-dimensional effects and were predicted qualitatively by a method described herein for calculation of pressure distribution in two-dimensional flow. Results of tests performed elsewhere on blunt triangular wings were found to substantiate the near two-dimensionality of the flow and were used to extend the range of applicability of the method of surface pressure predictions to Mach numbers of 11.5 in air and 13.3 in helium.

Description: An experimental study shows that 2 percent by weight ozone in oxygen has little effect on overall reactivity for a range of oxidant-fuel weight ratios from 1 to 6. This conclusion is based on characteristic-velocity measurements in 200-pound-thrust chambers at a pressure of 300 pounds per square inch absolute with low-efficiency injectors. The presence of 9 percent ozone in oxygen also did not affect performance in an efficient chamber. Explosions were encountered when equipment or procedure permitted ozone to concentrate locally. These experiments indicate that even small amounts of ozone in oxygen can cause operational problems.

Description: A review of the physical condition's under which future airplanes will operate has been made and the necessity for considering fatigue in the design has been established. A survey of the literature shows what phases of elevated-temperature fatigue have been investigated. Other studies that would yield data of particular interest to the designer of aircraft structures are indicated.

Description: A brief review of airplane altitude errors due to typical pressure installations at the fuselage nose, the wing tip, and the vertical fins is presented. A static-pressure tube designed to compensate for the position errors of fuselage-nose installations in the subsonic speed range is described. This type of tube has an ogival nose shape with the static-pressure orifices located in the low-pressure region near the tip. The results of wind-tunnel tests of these compensated tubes at two distances ahead of a model of an aircraft showed the position errors to be compensated to within 1/2 percent of the static pressure through a Mach number range up to about 1.0. This accuracy of sensing free-stream static pressure was extended up to a Mach number of about 1.15 by use of an orifice arrangement for producing approximate free-stream pressures at supersonic speeds and induced pressures for compensation of error at subsonic speeds.

Description: An investigation has been conducted on a triangular wing and body combination to determine the effects on the aerodynamic characteristics resulting from deflecting portions of the wing near the tips 900 to the wing surface about streamwise hinge lines. Experimental data were obtained for Mach numbers of 0.70, 1.30, 1.70, and 2.22 and for angles of attack ranging from -5 deg to +18 deg at sideslip angles of 0 deg and 5 deg. The results showed that the aerodynamic center shift experienced by the triangular wing and body combination as the Mach number was increased from subsonic to supersonic could be reduced by about 40 percent by deflecting the outboard 4 percent of the total area of each wing panel. Deflection about the same hinge line of additional inboard surfaces consisting of 2 percent of the total area of each wing panel resulted in a further reduction of the aerodynamic center travel of 10 percent. The resulting reductions in the stability were accompanied by increases in the drag due to lift and, for the case of the configuration with all surfaces deflected, in the minimum drag. The combined effects of reduced stability and increased drag of the untrimmed configuration on the trimmed lift-drag ratios were estimated from an analysis of the cases in which the wing-body combination with or without tips deflected was assumed to be controlled by a canard. The configurations with deflected surfaces had higher trimmed lift-drag ratios than the model with undeflected surfaces at Mach numbers up to about 1.70. Deflecting either the outboard surfaces or all of the surfaces caused the directional stability to be increased by increments that were approximately constant with increasing angle of attack at each Mach number. The effective dihedral was decreased at all angles of attack and Mach numbers when the surfaces were deflected.

Description: With the advent of the space age, new adjustments in technical thinking and engineering experience are necessary. There is an increasing and extensive interest in the utilization of materials for components to be used at temperatures ranging from -423 to over 3500 deg F. This paper presents a description of the materials problems associated with the various components of chemical liquid rocket systems. These components include cooled and uncooled thrust chambers, injectors, turbine drive systems, propellant tanks, and cryogenic propellant containers. In addition to materials limitations associated with these components, suggested research approaches for improving materials properties are made. Materials such as high-temperature alloys, cermets, carbides, nonferrous alloys, plastics, refractory metals, and porous materials are considered.

Description: An investigation has been conducted to determine the effects of a high positioned horizontal tail on a wing-body configuration having a thin unswept wing of aspect ratio 3.09. Lift and pitching-moment coefficients were obtained for Mach numbers from 0.80 to 1.40 at Reynolds numbers of 1.0 and 1.5 million and for angles of attack to 20 deg. An experimental study of the pitching-moment contribution of the horizontal tail indicated that the marked destabilizing effect of the horizontal tail at high angles of attack for Mach numbers of 0.80 to 1.00 was associated with the formation of completely separated flow on the upper surface of the wing. Computations of the interference effects of the wing-body combination on the tail for Mach numbers of 0.80 and 0.94 and high angles of attack confirmed this conclusion. For a Mach number of 1.40, and high angles of attack, computations disclosed that the destabilizing effect primarily resulted from the trailing vortices of the wing. Two modifications to the basic wing plan form, which consisted of chord extensions, were generally unsuccessful in reducing the destabilizing contributions of the horizontal tail at high angles of attack.

Description: An investigation has been made in the Langley 20-foot free-spinning tunnel on a 1/25-scale dynamic model to determine the spin and recovery characteristics of the Chance Vought F8U-1P airplane. Results indicated that the F8U-IP airplane would have spin-recovery characteristics similar to the XF8U-1 design, a model of which was tested and the results of the tests reported in NACA Research Memorandum SL56L31b. The results indicate that some modification in the design, or some special technique for recovery, is required in order to insure satisfactory recovery from fully developed erect spins. The recommended recovery technique for the F8U-lP will be full rudder reversal and movement of ailerons full with the spin (stick right in a right spin) with full deflection of the wing leading- edge flap. Inverted spins will be difficult to obtain and any inverted spin obtained should be readily terminated by full rudder reversal to oppose the yawing rotation and neutralization of the longitudinal and lateral controls. In an emergency, the same size parachute recommended for the XFBU-1 airplane will be adequate for termination of the spin: a stable parachute 17.7 feet in diameter (projected) with a drag coefficient of 1.14 (based on projected diameter) and a towline length of 36.5 feet.

Description: Pressure distributions obtained in the Langley 8-foot transonic pressure tunnel on a thin highly tapered twisted 45 deg sweptback wing-body combination are presented. The wing has a quadratic spanwise twist variation from 0 deg at 10 percent of the semispan to 6 deg at the tip. The tip is at a lower angle of attack than the root. Tests were made at stagnation pressures of both 0.5 and 1.0 atmosphere at Mach numbers from 0.800 to 1.200 through an angle-of-attack range from -4 deg to 20 deg.

Description: Surface pressure measurements were obtained at three chordwise stations on the wings of the X-3 and X-lE airplanes at Mach numbers from 0.73 to 1.13 for the X-3, and from 0.82 to 1.90 for the X-IE. Leading-edge separation is present on the X-3 wing at a Mach number of about 0.73 and an angle of attack of about 6 deg. However., when the Mach number is increased to 0.88, the trailing-edge separation dominates the pressure distribution and no leading-edge separation is visible although it is anticipated at the higher angles of attack shown. Conversely, the X-lE wing shows no indication of leading-edge separation within the scope of this investigation, but an overexpansion immediately behind the leading edge is present at a Mach number of approximately 0.82. Two separate normal shocks are present on the X-3 wing at a Mach number of about 0.88 and at a low angle of attack as an effect of wing geometry. These shocks merge to form a single shock when the angle of attack is increased to about 6 deg. At supersonic speeds the upper-surface expansion on the X-lE wing is limited by the approach of the pressure coefficients to the pressure coefficient for a vacuum.

Description: A turbojet-engine-exhaust simulator which utilizes a hydrogen peroxide gas generator has been developed for powered-model testing in wind tunnels with air exchange. Catalytic decomposition of concentrated hydrogen peroxide provides a convenient and easily controlled method of providing a hot jet with characteristics that correspond closely to the jet of a gas turbine engine. The problems associated with simulation of jet exhausts in a transonic wind tunnel which led to the selection of a liquid monopropellant are discussed. The operation of the jet simulator consisting of a thrust balance, gas generator, exit nozzle, and auxiliary control system is described. Static-test data obtained with convergent nozzles are presented and shown to be in good agreement with ideal calculated values.

Description: Two full-scale models of an inline, cruciform, canard missile configuration having a low-aspect-ratio wing equipped with flap-type controls were flight tested in order to determine the missile's longitudinal aerodynamic characteristics. Stability derivatives and control and drag characteristics are presented for a range of Mach number from 0.7 to 1.8. Nonlinear lift and moment curves were noted for the angle-of-attack range of this test (0 deg to 8 deg ). The aerodynamic-center location for angles of attack near 5 deg remained nearly constant for supersonic speeds at 13.5 percent of the mean aerodynamic chord; whereas for angles of attack near O deg, there was a rapid forward movement of the aerodynamic center as the Mach number increased. At a control deflection of O deg, the missile's response to the longitudinal control was in an essentially fixed space plane which was not coincident with the pitch plane as a result of the missile rolling. As a consequence, stability characteristics were determined from the resultant of pitch and yaw motions. The damping-in-pitch derivatives for the two angle-of-attack ranges of the test are in close agreement and varied only slightly with Mach number. The horn-balanced trailing-edge flap was effective in producing angle of attack over the Mach number range.

Description: The effect of turbine-inlet temperature on rocket gross weight was investigated for three high-energy long-range rockets in order to explore the desirability of turbine cooling in rocket turbodrive applications. Temperatures above and below the maximum that is permissible in uncooled turbines were included. Turbine bleed rate and stage number were considered as independent variables. The gross weight of the hydrogen-reactor system was more sensitive to changes in turbine-inlet temperature than either the hydrogen-oxygen or the hydrogen-fluorine systems. Gross weight of the hydrogen-reactor system could be reduced by 2.6 percent by the use of cooling and a turbine-inlet temperature of 3000 R. The reductions in the first stages of the hydrogen-oxygen and hydrogen-fluorine systems were 0.7 and 0.2 percent, respectively. The effect of turbine-inlet temperature on rocket gross weight was small because the resulting turbine weight and bleed rate variations were small. Since these small gains must be balanced against considerations of greater cost, weight, and complexity as well as lessened reliability with a system utilizing a cooled turbine, none of the systems investigated showed gains warranting the use of turbine cooling.

Description: An experimental investigation was conducted to determine the effect of moment-of-area-rule modifications on the drag, lift, and pitching-moment characteristics of a wing-body combination with a relatively high aspect-ratio unswept wing. The basic configuration consisted of an aspect-ratio-6 wing with a sharp leading edge and a thickness ratio of 0.06 mounted on a cut-off Sears-Haack body. The model with full moment-of-area-rule modifications had four contoured pods mounted on the wing and indentations in the body to improve the longitudinal distributions of area and moments of area. Also investigated were modifications employing pods and indentations that were only half the size of the full modifications and modifications with partial body indentations. The models were tested at angles of attack from -2 deg to +12 deg at Mach numbers from 0.6 to 1.4. In general, the moment-of-area-rule modifications had a large effect on the drag characteristics of the models but only a small effect on their lift and pitching-moment characteristics. The modifications provided substantial reductions in the zero-lift drag at transonic and low supersonic speeds, but at subsonic speeds the drag was increased. Near Mach number 1.0, the model with full modification provided the greatest reduction in drag, but at the highest test Mach numbers the half modification gave the largest drag reduction. In general, the percent reductions of zero- lift drag obtained with the aspect-ratio-6 wing were as great or greater than those previously obtained with aspect-ratio-3 wings. The effect of the modifications on the drag due to lift was small except at Mach num- bers below 0.9 where the modified models had higher drag-rise factors. Above Mach number 0.9, the modified models had higher lift-drag ratios than the basic model. The modified models also had higher lift curve slopes and generally were slightly more stable than the basic configuration.

Description: Blowing boundary-layer control was applied to the leading- and trailing-edge flaps of a 45 deg sweptback-wing complete model in a full-scale low-speed wind-tunnel study. The principal purpose of the study was to determine the effects of leading-edge flap deflection and boundary-layer control on maximum lift and longitudinal stability. Leading-edge flap deflection alone was sufficient to maintain static longitudinal stability without trailing-edge flaps. However, leading-edge flap blowing was required to maintain longitudinal stability by delaying leading-edge flow separation when trailing-edge flaps were deflected either with or without blowing. Partial-span leading-edge flaps deflected 60 deg with moderate blowing gave the major increase in maximum lift, although higher deflection and additional blowing gave some further increase. Inboard of 0.4 semispan leading-edge flap deflection could be reduced to 40 deg and/or blowing could be omitted with only small loss in maximum lift. Trailing-edge flap lift increments were increased by boundary-layer control for deflections greater than 45 deg. Maximum lift was not increased with deflected trailing-edge flaps with blowing.

Description: An investigation has been conducted on the Langley helicopter test tower to determine experimentally the maximum mean lift-coefficient characteristics at low tip Mach number and a limited amount of drag- divergence data at high tip Mach number of a helicopter rotor having an NACA 64(1)AO12 airfoil section and 8 deg of linear washout. Data are presented for blade tip Mach numbers M(t) of 0.29 to 0.74 with corresponding values 6 6 of tip Reynolds number of 2.59 x 10(exp 6) and 6.58 x 10(exp 6). Comparisons are made between the data from the present rotor with results previously obtained from two other rotors: one having NACA 0012 airfoil sections and the other having an NACA 0009 airfoil tip section. At low tip Mach numbers, the maximum mean lift coefficient for the blade having the NACA 64(1)AO12 section was about 0.08 less than that obtained with the blade having the NACA 0009 tip section and 0.21 less than the value obtained with the blade having the NACA 0012 tip section. Blade maximum mean lift coefficient values were not obtained for Mach number values greater than 0.47 because of a blade failure encountered during the tests. The effective mean lift-curve slope required for predicting rotor thrust varied from 5.8 for the tip Mach nuniber range of 0.29 to 0.55 to a value of 6.65 for a tip Mach number of 0.71. The blade pitching-moment coefficients were small and relatively unaffected by changes in thrust coefficient and Mach number. In the instances in which stall was reached, the break in the blade pitching-moment curve was in a stable direction. The efficiency of the rotor decreased with an increase in tip speed. Expressed as figure of merit, at a tip Mach number of 0.29 the maximum value was about 0.74. Similar measurements made on another rotor having an NACA 0012 airfoil and with a rotor having an NACA 0009 tip section, showed a value of 0.75. Synthesized section lift and profile-drag characteristics for the rotor-blade airfoil section are presented as an aid in predicting the high-tip-speed performance of rotors having similar airfoils.

Description: Tests have been conducted to determine the starting characteristics of a 50,000-pound-thrust rocket engine with the conditions of a quantity of fuel lying dormant in the simulated main thrust chamber. Ignition was provided by a smaller rocket firing rearwardly along the center line. Both alcohol-water and anhydrous ammonia were used as the residual fuel. The igniter successfully expelled the maximum amount of residual fuel (3 1/2 gal) in 2.9 seconds when the igniter.was equipped with a sonic discharge nozzle operating at propellant flow rates of 3 pounds per second. Lesser amounts of residual fuel required correspondingly lower expulsion times. When the igniter was equipped with a supersonic exhaust nozzle operating at a flow of 4 pounds per second, a slightly less effective expulsion rate was encountered.

Description: A two-dimensional wind-tunnel investigation has been conducted on a 20-percent-thick single-wedge airfoil section. Steady-state forces and moments were determined from pressure measurements at Mach numbers from 0.70 to about 1.25. Additional information on the flows about the single wedge is provided by means of instantaneous pressure measurements at Mach numbers up to unity. Pressure distributions were also obtained on a symmetrical double-wedge or diamond-shaped profile which had the same leading-edge included angle as the single-wedge airfoil. A comparison of the data on the two profiles to provide information on the effects of the afterbody showed that with the exception of drag, the single-wedge profile proved to be aerodynamically superior to the diamond profile in all respects. The lift effectiveness of the single-wedge airfoil section far exceeded that of conventional thin airfoil sections over the speed range of the investigation. Pitching-moment irregularities, caused by negative loadings near the trailing edge, generally associated with conventional airfoils of equivalent thicknesses were not exhibited by the single-wedge profile. Moderately high pulsating pressures existing over the base of the single-wedge airfoil section were significantly reduced as the Mach number was increased beyond 0.92 and the boundaries of the dead airspace at the base of the model converged to eliminate the vortex street in the wake. Increasing the leading-edge radius from 0 to 1 percent of the chord had a minor effect on the steady-state forces and generally raised the level of pressure pulsations over the forward part of the single-wedge profile.

Description: Two 10-inch-diameter spherical rocket motors have been flight tested at the NASA Wallops Station. These tests were conducted to measure "spin-up" or amplification of the spinning velocity of the motor during the thrusting process due to internal swirling of the exhaust gases. Model 1, a heavy-wall motor, experienced an increase in spin rate during thrusting of about 10 percent, whereas model 2, a flight-type motor with a lightweight motor case, experienced an increase of about 19 percent. The propellant weight and geometry were the same for both motors. A simple relationship for "spin-up" which satisfies these measured results is reported herein. Both models were spin stabilized throughout their flights. A theoretical method of predicting spin-up was derived and used to extend the measured 10-inch-motor results to spherical rocket motors of other sizes having a similar propellant geometry. This method is presented and its predictions are shown to compare favorably with the measured flight results.

Description: A free-flight investigation has been made to determine some effects of aerodynamic heating on the structural behavior of a wing at supersonic speeds. The test wing was a thin, unswept, untapered, multispar, aluminum-alloy wing having a 20-inch chord, a 20-inch exposed semispan, and a circular-arc airfoil section with a thickness ratio of 5 percent. The wing was tested on a model propelled by a two-stage rocket-propulsion system to a Mach number of 2.22 and a corresponding Reynolds number per foot of 13.2 x 10(6) Reasonably good agreement was obtained between Stanton numbers obtained from measured temperature-time data and values obtained by the theory of Van Driest for flat plates having turbulent boundary layers. Temperature measurements made in the skin of the wing and in the internal structures agreed well with calculated values. The wing was instrumented to detect any apparent fluttering motion in the wing, but no evidence of flutter was observed throughout the flight.

Description: Equations for the downwash and sidewash due to supersonic yawed and unswept horseshoe vortices have been utilized in formulating tables and charts to permit a rapid estimation of the flow velocities behind wings performing various steady motions. Tabulations are presented of the downwash and sidewash in the wing vertical plane of symmetry due to a unit-strength yawed horseshoe vortex located at 20 equally spaced spanwise positions along lifting lines of various sweeps. (The bound portion of the yawed vortex is coincident with the lifting line.) Charts are presented for the purpose of estimating the spanwise variations of the flow-field velocities and give longitudinal variations of the downwash and sidewash at a nuMber of vertical and spanwise locations due to a unit-strength unswept horseshoe vortex. Use of the tables and charts to calculate wing downwash or sidewash requires a knowledge of the wing spanwise distribution of circulation. Sample computations for the rolling sidewash and angle-of-attack downwash behind a typical swept wing are presented to demonstrate the use of the tables and charts.

Description: The concepts of the supersonic area rule and the moment-of-area rule are combined to develop a new method for calculating zero-lift wave drag which is amenable to the use of ordinary desk calculators. The total zero-lift wave drag of a configuration is calculated by the new method as the sum of the wave drag of each component alone plus the interference between components. In calculating the separate contributions each component or pair of components is analyzed over the smallest allowable length in order to improve the convergence of the series expression for the wave drag. The accuracy of the present method is evaluated by comparing the total zero-lift wave-drag solutions for several simplified configurations obtained by the present method with solutions given by slender-body and linearized theory. The accuracy and computational time required by the present method are also evaluated relative to the supersonic area rule and the moment-of-area rule. The results of the evaluation indicate that total zero-lift wave-drag solutions for simplified configurations can be obtained by the present method which differ from solutions given by slender-body and linearized theory by less than 6 percent. This accuracy for simplified configurations was obtained from only nine terms of the series expression for the wave drag as a result of calculating the total zero-lift wave drag by parts. For the same number of terms these results represent an accuracy greater than that for solutions obtained by either of the two methods upon which the present method is based, except in a few isolated cases. For the excepted cases, solutions by the present method and the supersonic area rule are identical. Solutions by the present method are obtained in one fifth the computing time required by the supersonic area rule. This difference in computing time of course would be substantially reduced if the complete procedures for both methods were programmed on electronic computing machines.

Description: A conceptual design of a nuclear turboelectric powerplant, producing 20,000 kilowatts of power suitable for manned space vehicles is presented. The study indicates that the radiator necessary for rejecting cycle waste heat is the dominant weight, and emphasis is placed on the selection of cycle operating conditions in order to reduce this weight. A thermodynamic cycle using sodium vapor as the working fluid and operating at a turbine-inlet temperature of 2500 R was selected. The total powerplant weight was calculated to be approximately 6 pounds per kilowatt. The radiator contributes approximately 2.1 pounds per kilowatt to the total weight and the reactor and reactor shield contribute approximately 0.24 and 1.2 pounds per kilowatt, respectively. The generator, turbine, and piping add significantly to the total weight (between 0.5 and 0.6 lb/kw), but the heat exchanger, pumps, and so on are less important. Several important research areas associated with the development of a reliable nuclear turboelectric powerplant of the type analyzed are discussed.

Description: A free-flight test has been conducted to check a technique for inflating an NASA 12-foot-diameter inflatable sphere at high altitudes. Flight records indicated that the nose section was successfully separated from the booster rocket, that the sphere was ejected, and that the nose section was jettisoned from the fully inflated sphere. On the basis of preflight and flight records, it is believed that the sphere was fully inflated by the time of peak altitude (239,000 feet). Calculations showed that during descent, jettison of the nose section occurred above an altitude of 150,000 feet. The inflatable sphere was estimated to start to deform during descent at an altitude of about 120,000 feet.

Description: An investigation was made of a 1/10-scale dynamically similar model of the Grumman FgF-2 airplane to study its behavior when ditched. The model was landed in calm water at the Langley Tank No. 2 monorail. Various landing attitudes, speeds, and configurations were investigated. The behavior of the model was determined from visual observations, acceleration records, and motion-picture records of the ditchings. Data are presented in tabular form, sequence photographs, time-history acceleration curves, and plots of attitude and speed against distance after contact.

Description: No abstract available

Description: Low-lift drag data are presented herein for one 1/7.5-scale rocket-boosted model and three 1/45.85-scale equivalent-body models of the Grumman F9F-9 airplane, The data were obtained over a Reynolds number range of about 5 x 10(exp 6) to 10 x 10(exp 6) based on wing mean aerodynamic chord for the rocket model and total body length for the equivalent-body models. The rocket-boosted model showed a drag rise of about 0,037 (based on included wing area) between the subsonic level and the peak supersonic drag coefficient at the maximum Mach number of this test. The base drag coefficient measured on this model varied from a value of -0,0015 in the subsonic range to a maximum of about 0.0020 at a Mach number of 1.28, Drag coefficients for the equivalent-body models varied from about 0.125 (based on body maximum area) in the subsonic range to about 0.300 at a Mach number of 1.25. Increasing the total fineness ratio by a small amount raised the drag-rise Mach number slightly.

Description: A comparison of the zero-lift drag coefficients at Mach numbers from 0.81 to 1.41 of a fin-stabilized parabolic body of revolution as measured in the Langley transonic blowdown tunnel has been made with measurements obtained in free-flight on a larger but geometrically similar model. The absolute values of drag coefficient obtained in the slotted wind tunnel were equivalent to the free-flight drag-coefficient values up to a Mach number of 1.4 when adjustments were made for the effect on viscous drag of differences in Reynolds number between the two test conditions. Excellent agreement was obtained between the two tests for the pressure-drag variation with Mach number, regardless of whether the scale effect on skin friction was considered. Favorable agreement was also obtained between the pressure-drag increments due t o the presence of the stabilizing fins as determined in the wine tunnel from fins-on and fins-off tests and as obtained by a different method in free flight. Tests of a specific airplane configuration to obtain an indication of the problems involved in the construction and tests of small-scale (approximately 7-inch span) complete airplane configuration with internal air flow indicated that reliable zero-lift drag-coefficient measurements at Mach numbers up to 1.4 can be attained with such models, provided the model is constructed with a high but not an unreasonable degree of accuracy.

Description: A free-flight investigation over a Mach number range from 0.6 to 2.0 has been conducted to determine the longitudinal aerodynamic characteristics and effect of rocket jet on zero-lift drag of 1/5-scale models of two ballistic-type missiles, the Hermes A-3A and A-3B. Models of both types of missiles exhibited very nearly linear normal forces and pitching moments over the angle-of-attack range of 8 deg to -4 deg and Mach number range tested. The centers of pressure for both missiles were not appreciably affected by Mach number over the subsonic range; however, between a Mach number of 1.02 and 1.50 the center of pressure for the A-3A model moved forward 0.34 caliber with increasing Mach number. At a trim angle-of-attack of approximately 30 deg, the A-3A model indicated a total drag coefficient 30% higher than the power-off zero-lift drag over the subsonic Mach number range and 10% higher over the supersonic range. Under the conditions of the present test, and excluding the effect of the jet on base drag, there was no indicated effect of the propulsive jet on the total drag of the A-3A model. The propulsive jet operating at a jet pressure ratio p(sub j)/p(sub o) of 0.8 caused approximately 100% increase in base drag over the Mach number range M = 0.6 to 1.0. This increase in base drag amounts to 15% of the total drag. An underexpanded jet operating at jet pressure ratios corresponding approximately to those of the full-scale missile caused a 22% reduction in base drag at M = 1.55 (p(sub j)/p(sub o) = 1.76) but indicated no change at M = 1.30 (p(sub j)/p(sub o) = 1.43). At M = 1.1 and p(sub j)/p(sub o) = 1.55, the jet caused a 50% increase in base drag.

Description: A solution has been obtained for the complete tunnel-interference flow for a lifting vortex in a two-dimensional slotted tunnel. Curves are presented for the longitudinal distribution of tunnel-induced downwash angle for various values of the boundary openness parameter and for various heights of the vortex above the tunnel center line. Some quantitative discussion is given of the use of these results in calculating the tunnel interference for three-dimensional wings in rectangular tunnels with closed side walls and slotted top and bottom.

Description: An investigation is being conducted in the Langley 20-foot free-spinning tunnel on a 1/24-scale model of the Grumman F11F-1 airplane to determine spin and recovery characteristics and the minimum-size parachute required to satisfactorily terminate the spin in an emergency. Results obtained to date are presented herein. Test results indicate that it may be difficult to obtain an erect or inverted spin on the airplane, but, if a spin is obtained, the spin will be very oscillatory and recovery from the developed erect spin by rudder reversal may not be possible. The lateral controls will have no appreciable effect on recoveries from erect.spins. Recovery from the inverted spin by merely neutralizing the rudder will be satisfactory. After recoveries by rudder reversal and after recoveries from spins without control movement (no spins), the model oftentimes rolled very rapidly about the X-axis. Based on limited preliminary tests made in this investigation to make the model recover satisfactorily, it appears that canards near the nose of the airplane or differentially operated horizontal tails may be utilized to provide rapid recoveries. The parachute test results indicate that an 11-foot-diameter (laid-out-flat) parachute with a drag coefficient of 0.650 (based on the laid- out-flat diameter) and with a towline length equal to the wing span is the minimum-size parachute required to satisfactorily terminate an erect or inverted spin in an emergency.

Description: An investigation is being conducted in the Langley 20-foot free-spinning tunnel on a l/18 scale model of the Ryan X-13 airplane to determine its spin and recovery characteristics. The spin and recovery characteristics determined to date are presented in this report.

Description: Tests have been conducted in the Langley 8-foot transonic tunnel on a 0.04956-scale model of the Convair F-102A airplane which employed an indented and extended fuselage, cambered wing leading edges, and deflected wing tips. Force and moment characteristics were obtained for Mach numbers from 0.60 to 1.135 at angles of attack up to 20 . In addition, tests were made over a limited angle-of-attack range to determine the effects of the cambered leading edges, deflected tips, and a nose section with a smooth area distribution. Fuselage modifications employed on the F-102A were responsible for a 25.percent reduction in the minimum drag-coefficient rise between the Mach numbers of 0.85 and 1.075 when compared with that for the earlier versions of the F-102. Although the wing modifications increased the F-102A subsonic minimum drag-coefficient level approximately 0.0020, they produced large decreases in drag at lifting conditions over that for the original (plane-wing) F-102. The F-102A had 15 to 25 percent higher maximum lift-drag ratios than did the original F-102. The F-102A had about 15 percent lower maximum lift-drag ratios at Mach numbers below 0.95 and slightly higher maximum lift-drag ratios at supersonic speeds when compared with those ratios for sn earlier modified-wing version of the F-102. Chordwise wing fences which provided suitable longitudinal stability for the original F-102 were not adequate for the cambered-wing F-102A The pitching-moment curves indicated a region of near neutral stability with possible pitch-up tendencies for the F-102A at high subsonic Mach numbers for lift coefficients between about 0.4 and 0.5.

Description: An investigation has been made on the use of a freely rotating rotor at the cowl face of a supersonic conical diffuser to determine its effectiveness in reducing inlet flow distortion and the penalty in terms of total-pressure loss imposed by such a device when distortions are negligible. Tests were made with a rotor having an inlet tip diameter of 2.18 inches and a ratio of hub radius to tip radius of 0.52, in conjunction with a conical inlet having a 25 deg semi-vertex cone angle, at a Mach number of 2.1 over an angle-of-attack range of 0 deg to 8 deg. A simplified analysis showing that a supersonic, freely rotating rotor with maximum solidity for noninterference between blades will operate in an undistorted flow with a total-pressure defect of 1 percent or less was experimentally verified. Overall total-pressure distortions of 0.1 to 0.4 and Mach number distortions of 0.4 to 1.4, obtained at 4 deg to 8 deg angle of attack, were reduced about 30 percent and 23 percent, respectively, because of the presence of the rotor, with no measurable total-pressure loss. The rotor increased the peak total-pressure recovery at the simulated combustion chamber 1 1/2 and 3 1/2 percent at 6 deg and 8 deg angles of attack, respectively. This increase is attributed to lower diffusion duct losses as a consequence of a more uniform flow created by the rotor.

Description: A flight test of a rocket-propelled model of the Convair XFY-1 airplane was conducted to determine the lateral stability and control characteristics, The 0.133-scale model had windmilling propellers for this test, which covered a Mach number range of O.70 to 1.12. The center of gravity was located at 13.9 percent of the mean aerodynamic chord. The methods of analysis included both a solution by vector diagrams and simple one- and two-degree-of-freedom methods. The model was both statically and dynamically stable throughout the speed range of the testa The roll damping was good, and the slope of the side-force curve varied little with speed. The rudder was effective throughout the test speed range, although it was reduced to about 43 percent of its subsonic value at supersonic speeds.

Description: An investigation has been conducted at Mach numbers of 0.6 to 1.27 to determine the effect of multiple-jet exits on the base pressure of a simple wing-body combination. The design Mach number of the nozzles ranged from 1 to 3 at jet exit diameters equal to 36.4 to 75 percent of the model thickness. Jet total-pressure to free-stream static-pressure ratios ranged from 1 (no flow) to 34.2. The results show that the variation of base pressure coefficient with jet pressure ratio for the model tested was similar to that obtained for single nozzles in bodies of revolution in other investigations. As in the case for single jets the base pressure coefficient for the present model became less negative as the jet exit diameter increased. For a constant throat diameter and an assumed schedule of jet pressure ratio over the speed range of these tests, nozzle Mach number had only a small effect on base pressure coefficient.

Description: An investigation to evaluate the effects of thickened and blunted leading-edge modifications on the wave drag of a swept wing has been made at Mach numbers from 0.65 to 2.20 and at a Reynolds number of 2,580,000 based on the mean aerodynamic chord of the basic wing. Two leading-edge designs were investigated and they are referred to as the thickened and the blunted modifications although both sections had equally large leading-edge radii. The thickened leading edge was formed by increasing the thickness over the forward 40 percent of the basic wing section. The blunted modification was formed by reducing the wing chords about 1 percent and by increasing the section thickness slightly over the forward 6 percent of the basic section in a manner to keep the wing sweep and volume essentially equal to the respective values for the basic wing. The basic wing had an aspect ratio of 3, a leading-edge sweep of 45 deg., a taper ratio of 0.4, and NACA 64AO06 sections perpendicular to a line swept back 39.45 deg., the quarter-chord line of these sections. Test results indicated that the thickened modification resulted in an increase in zero-lift drag coefficient of from 0.0040 to 0.0060 over values for the basic model at Mach numbers at which the wing leading edge was sonic or supersonic. Although drag coefficients of both the basic and thickened models were reduced at all test Mach numbers by body indentations designed for the range of Mach numbers from 1.00 to 2.00, the greater drag of the thickened model relative to that of the basic model was not reduced. The blunted model, however, had less than one quarter of the drag penalty of the thickened model relative to the basic model at supersonic leading-edge conditions (M greater or equal to root-2).

Description: Limitations on fully developed laminar flows due to compressibility and property variations are examined. The cases, for liquids and for gases, wherein such motions are "exact" are determined and solutions are given. For more general conditions, not permitting an exact fully developed flow, limitations are set. Two cases arise depending on the size of the temperature variation across the channel. Both the forced and free flow are solved for the case of large temperature variation. Finally, there are described briefly some circumstances under which streamwise variations of velocity occur. The case where the velocity varies inversely with the square root of the distance is solved.

Description: A concept of combustion time lag that includes dependency on injection velocity is introduced. The concept is used in the formulation of chamber transfer functions and in an analysis of low-frequency combustion instability. Theoretical frequency responses and stability boundaries are compared with those obtained when the injection-velocity effect on the time lag to be an important consideration, in the theory of chamber dynamics and combustion instability

Description: Results obtained with two nose shapes tested at a Reynolds number per foot of 5 x 10(exp 6) at angles of attack from -4 deg to +10 deg at 0 deg angle of sideslip are presented in tabulated pressure coefficient form without analysis.

Description: Pressure coefficients were measured over the vehicle and over the forward part of the booster at Reynolds numbers of 3.0 x 10(exp 6) per foot. Tabular results are presented for two nose shapes at Mach numbers of 1.55, 1.75, 2.00, and 2.35, at angles of attack from -4 deg to +10 deg, and at 0 deg sideslip.

Description: Previous investigations have shown that increased blowing at the hinge-line radius of a plain flap will give flap lift increases above that realized with boundary-layer control. Other experiments and theory have shown that blowing from a wing trailing edge, through the jet flap effect, produced lift increases. The present investigation was made to determine whether blowing simultaneously at the hinge-line radius and trailing edge would be more effective than blowing separately at either location. The tests were made at a Reynolds number of 4.5 x 10(exp 6) with a 35 deg sweptback-wing airplane. For this report, only the lift data are presented. Of the three flap blowing arrangements tested, blowing distributed between the trailing edge and the hinge-line radius of a plain flap was found to be superior to blowing at either location separately at the plain flap deflections of interest. Comparison of estimated and experimental jet flap effectiveness was fair.

Description: A geometric study has been made of some of the effects of dihedral on the heat transfer to swept delta wings. The results of this study show that the incorporation of large positive dihedral on highly swept wings can shift, even at moderately low angles of attack, the stagnation-line heat-transfer problem from the leading edges to the axis of symmetry (ridge line). An order-of-magnitude analysis (assuming laminar flow) indicates conditions for which it may be possible to reduce the heating at the ridge line (except in the vicinity of the wing apex) to a small fraction of the leading-edge heat transfer of a flat wing at the same lift. Furthermore, conditions are indicated where dihedral reduces the leading-edge heat transfer for angles of attack less than those required to shift the stagnation line from the leading edge to the ridge line.