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: 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: 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: The present paper summarizes and correlates broadly some of the research results applicable to fin-stabilized ammunition. The discussion and correlation are intended to be comprehensive, rather than detailed, in order to show general trends over the Mach number range up to 7.0. Some discussion of wings, bodies, and wing-body interference is presented, and a list of 179 papers containing further information is included. The present paper is intended to serve more as a bibliography and source of reference material than as a direct source of design information.

Description: Air-flow characteristics behind wings and wing-body combinations are described and are related to the downwash at specific tall locations for unseparated and separated flow conditions. The effects of various parameters and control devices on the air-flow characteristics and tail contribution are analyzed and demonstrated. An attempt has been made to summarize certain data by empirical correlation or theoretical means in a form useful for design. The experimental data herein were obtained mostly at Reynolds numbers greater than 4 x 10(exp 6) and at Mach numbers less than 0.25.

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: Measurements of the statistical properties of the fluctuating wall pressure produced by a subsonic turbulent boundary layer are described. The measurements provide additional information about the structure of the turbulent boundary layer; they are applicable to the problems of boundary-layer induced noise inside an airplane fuselage and to the generation of waves-on water. The spectrum of the wall pressure is presented in dimensionless form. The ratio of the root-mean-square wall pressure to the free-stream dynamic pressure is found to be a constant square root of bar P(sup 2)/q(sub infinity) = 0.006 independent of Mach number and Reynolds number. In addition, space- time correlation measurements in the stream direction show that pressure fluctuations whose scale is greater than or equal to 0.3 times the boundary-layer thickness are convected with the convection speed U(sub c) = 0.82U(sub infinity) where U(infinity) is the free-stream velocity and have lost their identity in a distance approximately equal to 10 boundary-layer thicknesses.

Description: Approximate analytical solutions are presented for two-dimensional and axisymmetric hypersonic flow over slender power law bodies. Both zero order (M approaches infinity) and first order (small but nonvanishing values of 1/(M(Delta)(sup 2) solutions are presented, where M is free-stream Mach number and Delta is a characteristic slope. These solutions are compared with exact numerical integration of the equations of motion and appear to be accurate particularly when the shock is relatively close to the body.

Description: An experimental investigation of the mixing of two coaxial gas streams was conducted over a range of subsonic jet Mach numbers and temperatures. Three configurations were investigated. One had no innerbody in the primary or inner pipe and was designed to give flat velocity profiles at the exit of the primary pipe. The other two configurations had innerbodies in the primary pipe. These were designed to give velocity profiles similar to those existing at the inlet of propulsive systems such as afterburners. Curves of axial velocity and temperature profiles across the radius are presented at various axial stations. For the two configurations with the innerbody, data are shown at stations out to approximately 8 primary-pipe diameters from the exit of the primary pipe. For the flat-velocity-profile configuration, data are shown at distances extending downstream at 22 primary-pipe diameters from the exit of the primary pipe.

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: 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: Techniques which have been used for finishing and quantitatively specifying surface roughness on boundary-layer-transition models are reviewed. The appearance of a surface as far as roughness is concerned can be misleading when viewed either by the eye or with the aid of a microscope. The multiple-beam interferometer and the wire shadow method provide the best simple means of obtaining quantitative measurements.

Description: Carrier landing-approach studies of a tailless delta-wing fighter airplane disclosed that approach speeds were limited by ability to control altitude and lateral-directional characteristics. More detailed flight studies of the handling-qualities characteristics of the airplane in the carrier-approach configuration documented a number of factors that contributed to the adverse comments on the lateral-directional characteristics. These were: (1) the tendency of the airplane to roll around the highly inclined longitudinal axis, so that significant sideslip angles developed in the roll as a result only of kinematic effects; (2) reduction of the rolling response to the ailerons because of the large dihedral effect in conjunction with the kinematically developed sideslip angles; and (3) the onset of rudder lock at moderate angles of sideslip at the lowest speeds with wing tanks installed. The first two of the factors listed are inseparably identified with this type of configuration which is being considered for many of the newer designs and may, therefore, represent a problem which will be encountered frequently in the future. The results are of added significance in the demonstration of a typical situation in which extraneous factors occupy so much of the pilot's attention that his capability of coping with the problems of precise flight-path control is reduced, and he accordingly demands a greater speed margin above the stall to allow for airspeed fluctuations.

Description: Wind-tunnel tests have been made to determine the location of the boundary-layer transition on three hemispheres having surface roughness (absolute) values of 50, 580, and 2760 microinches. After the initial test run of the smoothest (50 microinch) hemisphere, holes ranging in depth from 1500 to 2500 microinches were noticed in the meridian where transition was observed. The holes were believed to be caused by particles in the air stream. Shadowgraph pictures were obtained of all hemispheres and surface temperature measurements were made on one hemisphere (580 microinches). Tests at high Reynolds numbers (6.4 to 7.5 x 10(exp 6) and a Mach number of 2.48 did not indicate any transition on the 50-microinch surface hemisphere before the holes appeared. However, after the holes were noticed, transition locations as low as 50 deg(measured from the stagnation point) were observed at similar Reynolds numbers and Mach numbers. It is felt the transition resulted from the holes. Similar transition locations of approximately 500 were also observed in the tests of hemispheres with surface roughness values of 580 and 2760 microinches at high Reynolds numbers (6.4 x 10(exp 6) to 7.5 x 10(exp 6)) and at a Mach number of 2.48. The results at a Mach number of 2.48 indicate that an absolute surface roughness value of 50 microinches was not critical in causing boundary-layer tran sition at Reynolds numbers of 6.4 to 7.5 x 10(exp 6) whereas roughness values of 580 and 2760 microinches were greater than critical. Transition Reynolds numbers based on momentum thickness, R(sub phi T) varied over a range of approximately 480 to 300 for transition locations, alpha, on the hemisphere from 880 to 410 (measured from the stagnation point). A maximum value of R(phi) of 660 (based on alpha = 90 deg) was obtained with the 50-microinch surface hemisphere at a Mach number of 2.48.

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: Tests were made on a 10-foot-diameter hemispherical nose at Reynolds numbers up to 10 x 10(exp 6) and at a maximum Mach number of about 0.1 to determine the effects of a highly favorable pressure gradient on boundary-layer transition caused by roughness. Both two-dimensional and three-dimensional roughness particles were used, and the transition of the boundary layer was determined by hot-wire anemometers. The roughness Reynolds number for transition R(sub k,t) caused by three-dimensional particles such as Carborundum grains, spherical particles, and rimmed craters was found. The results show that for particles immersed in the boundary layer, R(sub k,t) is independent of the particle size or position on the hemispherical nose and depends mainly on the height-to-width ratio of the particle. The values of R(sub k,t) found on the hemispherical nose compare closely with those previously found on a flat plate and on airfoils with roughness. For two-dimensional roughness, the ratio of roughness height to boundary-layer displacement thickness necessary to cause transition was found to increase appreciably as the roughness was moved forward on the nose. Also included in the investigation were studies of the spread of turbulence behind a single particle of roughness and the effect of holes such as pressure orifices.

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 theory is derived for determining the loads and motions of a deeply immersed prismatic body. The method makes use of a two-dimensional water-mass variation and an aspect-ratio correction for three-dimensional flow. The equations of motion are generalized by using a mean value of the aspect-ratio correction and by assuming a variation of the two-dimensional water mass for the deeply immersed body. These equations lead to impact coefficients that depend on an approach parameter which, in turn, depends upon the initial trim and flight-path angles. Comparison of experiment with theory is shown at maximum load and maximum penetration for the flat-bottom (0 deg dead-rise angle) model with bean-loading coefficients from 36.5 to 133.7 over a wide range of initial conditions. A dead-rise angle correction is applied and maximum-load data are compared with theory for the case of a model with 300 dead-rise angle and beam-loading coefficients from 208 to 530.

Description: An investigation has been made to determine the effect of wing fences, fuselage contouring, varying wing sweepback angle from 40 deg. to 45 deg., mounting the horizontal tail on an outboard boom) and wing thickness distribution upon the buffeting response of typical airplane configurations employing sweptback wings of high aspect ratio. The tests were conducted through an angle-of-attack range at Mach numbers varying from 0.60 to 0.92 at a Reynolds number of 2 million. For the combinations with 40 deg. of sweepback, the addition of multiple wing fences usually decreased the buffeting at moderate and high lift coefficients and reduced the erratic variation of buffet intensities with increasing lift coefficient and Mach number. Fuselage contouring also reduced buffeting but was not as effective as the wing fences. At most Mach numbers, buffeting occurred at higher lift coefficients for the combination with the NACA 64A thickness distributions than for the combination with the NACA four-digit thickness distributions. At high subsonic speeds, heavy buffeting was usually indicated at lift coefficients which were lower than the lift coefficients for static-longitudinal instability. The addition of wing fences improved the pitching-moment characteristics but had little effect on the onset of buffeting. For most test conditions and model configurations, the root-mean- square and the maximum values measured for relative buffeting indicated similar effects and trends; however, the maximum buffeting loads were usually two to three times the root-mean-square intensities.

Description: An analytical heat transfer solution is derived and evaluated for the general case of a turbulently flowing liquid metal which suddenly encounters a step-function boundary temperature in a channel system. Local Nusselt moduli, dimensionless mixed-mean fluid temperatures, and arithmetic-mean Nusselt moduli are given as functions of Reynolds and Prandtl moduli and a dimensionless axial-distance modulus. These solutions are compared with known solutions of more specific systems as well as with a set of experimental liquid-metal heat transfer data for a thermal entrance region.

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 wind-tunnel investigation has been made to determine the aerodynamic characteristics of a 1/4-scale model of a tilt-wing vertical-take-off-and-landing aircraft. The model had two 3-blade single-rotation propellers with hinged (flapping) blades mounted on the wing, which could be tilted from an incidence of 4 deg for forward flight to 86 deg for hovering flight. The investigation included measurements of both the longitudinal and lateral stability and control characteristics in both the normal forward flight and the transition ranges. Tests in the forward-flight condition were made for several values of thrust coefficient, and tests in the transition condition were made at several values of wing incidence with the power varied to cover a range of flight conditions from forward-acceleration (or climb) conditions to deceleration (or descent) conditions The control effectiveness of the all-movable horizontal tail, the ailerons and the differential propeller pitch control was also determined. The data are presented without analysis.

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: Heat-transfer and pressure-drop data were obtained experimentally for the gas side of a liquid-metal to air, compact finned-tube heat exchanger. The heat exchanger was fabricated from 0.185-inch Inconel tubing in an inline array. The fins were made of 310 stainless-steel- clad copper with a total thickness of 0.010 inch, and the fin pitch was 15.3 fins per inch. The liquid used as the heating medium was sodium. The heat-exchanger inlet gas temperature was varied from 5100 to 1260 R by burning JP fuel for airflow rates of 0.4 to 10.5 pounds per second corresponding to an approximate Reynolds number range of 300 to 9000. The sodium inlet temperature was held at 1400 R with the exception of a few runs taken at 1700 and 1960 R. The maximum ratio of surface temperature to air bulk temperature was 1.45. Friction-factor data with heat transfer were best represented by a single line when the density and viscosity of Reynolds number were evaluated at the average film temperature. At the lower Reynolds numbers reported, the friction data with heat transfer plotted slightly above the friction data without heat transfer. The density of the friction factor was calculated at the average bulk temperature. Heat-transfer results of this investigation were correlated by evaluating the physical properties of air (specific heat, viscosity, and thermal conductivity) at the film temperature.

Description: Three numerical solutions of the partial differential equations describing the compressible laminar boundary layer are obtained by the finite difference method described in reports by I. Flugge-Lotz, D.C. Baxter, and this author. The solutions apply to steady-state supersonic flow without pressure gradient, over a cold wall and over an adiabatic wall, both having transpiration cooling upstream, and over an adiabatic wall with upstream cooling but without upstream transpiration. It is shown that for a given upstream wall temperature, upstream transpiration cooling affords much better protection to the adiabatic solid wall than does upstream cooling without transpiration. The results of the numerical solutions are compared with those of approximate solutions. The thermal results of the finite difference solution lie between the results of Rubesin and Inouye, and those of Libby and Pallone. When the skin-friction results of one finite difference solution are used in the thermal analysis of Rubesin and Inouye, improved agreement between the thermal results of the two methods of solution is obtained.

Description: The measured static-pressure distributions at the model surface and in the surrounding flow field are presented for a basic parabolic-arc body having a fineness ratio of 14 and for three additional bodies obtained by modifying the basic parabolic-arc body along the middle portion of the body length by adding a bump, by indenting, or by quadripole shaping. The data were obtained with the various bodies at zero angle of attack. The Mach number varied from 0.80 to 1.20 with a corresponding Reynolds number (based on body length) variation of 27 x 10(exp 6) to 38 x 10(exp 6). The data are subject to tunnel-wall interference and do not represent free-air conditions.

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 investigation of expanded duct sections and the effect of their design parameters on flow distortion over a duct Mach number range of 0.19 to 0.67 was conducted in the small tunnel facility of the Lewis Research Center. The parameters investigated were: (1) entrance angle of expanded section, (2) length of expanded section, (3) area ratio of expanded section, (4) location of expanded section relative to the engine face, and (5) the use of screens of varying solidities and mesh. Expansion half-angles of deg, 15 deg, and 30 deg reduced the total-pressure distortions induced in the duct. The larger expansion angles reduced circumferential distortion more effectively than radial distortion. However, the half-angle of 15 deg appeared to be optimum for reducing both radial and circumferential distortions while still maintaining a high total-pressure recovery. Increasing the expanded-section area ratio and increasing the expanded-section lengths with-the 150 expansion half-angle led to less total-pressure distortion with no appreciable loss in pressure recovery. Screens incorporated in the expanded section indicated that 22.2-percent- solidity screens decreased distortion still further.while 37.3-percent- solidity screens generally increased distortion above that of a constant- area duct incorporating the same solidity screen.

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: Wind-tunnel measurements were made of the static and dynamic rotary stability derivatives of an airplane model having sweptback wing and tail surfaces. The Mach number range of the tests was from 0.23 to 0.94. The components of the model were tested in various combinations so that the separate contribution to the stability derivatives of the component parts and the interference effects could be determined. Estimates of the dynamic rotary derivatives based on some of the simpler existing procedures which utilize static force data were found to be in reasonable agreement with the experimental results at low angles of attack. The results of the static and dynamic measurements were used to compute the short-period oscillatory characteristics of an airplane geometrically similar to the test model. The results of these calculations are compared with military flying qualities requirements.

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: Hot-wire anemometer measurements were made of several statistical properties of approximately homogeneous and isotropic fields of turbulence and temperature fluctuations generated by a warm grid in a uniform airstream sent through a 4-to-1 contraction. These measurements were made both in the contraction and in the axisymmetric domain farther downstream. In addition to confirming the well-known turbulence anisotropy induced by strain, the data show effects on the skewnesses of both longitudinal velocity fluctuation (which has zero skewness in isotropic turbulence) and its derivative. The concomitant anisotropy in the temperature field accelerates the decay of temperature fluctuations.

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: A procedure based on the method of similar solutions is presented by which the skin friction, heat transfer, and boundary-layer thickness in a laminar hypersonic flow with pressure gradient may be rapidly evaluated if the pressure distribution is known. This solution, which at present is. restricted to power-law variations of pressure with surface distance, is presented for a wide range of exponents in the power law corresponding to both favorable and adverse pressure gradients. This theory has been compared to results from heat-transfer experiments on blunt-nose flat plates and a hemisphere cylinder at free-stream Mach numbers of 4 and 6.8. The flat-plate experiments included tests made at a Mach number of 6.8 over a range of angle of attack of +/- 10 deg. Reasonable agreement of the experimental and theoretical heat-transfer coefficients has been obtained as well as good correlation of the experimental results over the entire range of angle of attack studied. A similar comparison of theory with experiment was not feasible for boundary-layer-thickness data; however, the hypersonic similarity theory was found to account satisfactorily for the variation in boundary-layer thickness due to local pressure distribution for several sets of measurements.

Description: A simplified analysis is made of ablation cooling near the stagnation point of a two-dimensional or axisymmetric body which occurs as the body vaporizes directly from the solid state. The automatic shielding mechanism Is discussed and the important thermal properties required by a good ablation material are given. The results of the analysis are given in terms of dimensionless parameters.

Description: An investigation of the use of ballast at the leading edge of a sweptback wing as a flutter fix has been made. The investigation was conducted in the Langley transonic blowdown tunnel with wing models which had an aspect ratio of 4, sweepback of the quarter-chord line of 450, and a taper ratio of 0.2. Four ballast configurations, which included different amounts of ballast distributed at two different span-wise locations, were investigated. Full-span sting-mounted models were employed. Data were obtained over a Mach number range from 0.65 to 1.32. Comparison of the data for the ballasted wings with data for a similar wing without ballast shows that in the often critical Mach number range between 0.85 and 1.05, the dynamic pressure required for flutter is increased by as much as 100 percent due to the addition of about 6 percent of the wing mass as ballast at the leading edge of the outboard sections. Furthermore, there are indications that similar benefits of leading-edge ballast can be obtained at Mach numbers above M = 1.1. Changing the spanwise location of the ballast and increasing the amount of the ballast by a factor of about 2 had very little additional effect on the dynamic pressure required for flutter. The possibility, therefore, exists that the beneficial effects obtained may be accomplished by using less than the minimum of about 6 percent of the wing mass as ballast as investigated in this paper.

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: Free-oscillation tests were made in the Langley high-speed 7- by 10-foot tunnel to determine the effects of wing thickness and wing sweep on the hinge-moment and flutter characteristics of a trailing-edge flap-type control. The untapered semispan wings had full-span aspect ratios of 5 and NACA 65A-series airfoil sections. Unswept wings having ratios of wing thickness to chord of 0.04, 0.06, 0.08, and 0.10 were investigated. The swept wings were 6 percent thick and had sweep angles of 30 deg and 45 deg. The full-span flap-type controls had a total chord of 50 percent of the wing chord and were hinged at the 0.765-wing-chord line. Tests were made at zero angle of attack over a Mach number range from 0.60 to 1.02, control oscillation amplitudes up to about 12 deg, and a range of control-reduced frequencies. Static hinge-moment data were also obtained. Results indicate that the control aerodynamic damping for the 4-percent-thick wing-control model was unstable in the Mach number range from 0.92 to 1.02 (maximum for these tests). Increasing the ratio of wing thickness to chord to 0.06, 0.08, and then to 0.10 had a stabilizing effect on the aerodynamic damping in this speed range so that the aerodynamic damping was stable for the 10-percent-thick model at all Mach numbers. The 6-percent-thick unswept-wing-control model generally had unstable aerodynamic damping in the Mach number range from 0.96 to 1.02. Increasing the wing sweep resulted in a general decrease in the stable aerodynamic damping at the lower Mach numbers and in the unstable aerodynamic damping at the higher Mach numbers. The one-degree-of-freedom control-surface flutter which occurred in the transonic Mach number range (0.92 to 1.02) for the 4-, 6-, and 8-percent-thick unswept-wing-control models could be eliminated by further increasing the ratio of thickness to chord to 0.10. Flutter could also be eliminated by increasing the wing sweep angle to either 30 deg or 45 deg. The magnitude of variation in spring moment derivative with Mach number at transonic speeds was decreased by either increasing the ratio of wing thickness to chord or increasing the wing sweep angle.

Description: The momentum integral equations are derived for the boundary layer on an arbitrary curved surface, using a streamline coordinate system. Computations of the turbulent boundary layer on a slightly yawed cone are made for a Prandtl number of 0.729, wall to free-stream temperature ratios of 1/2, 1, and 2, and Mach numbers from 1 to 4. Deflection of the fluid in the boundary layer from outer stream direction, local friction coefficient, displacement surface, lift coefficient, and pitching-moment coefficient are presented.

Description: An investigation was conducted to obtain the aerodynamic characteristics of a model of a fighter-type airplane embodying partial body indentation. The wing had an aspect ratio of 4, taper ratio of 0.5, 35 deg sweepback of the 0.25-chord line, and a modified NACA 65A006 airfoil section at the root and a modified NACA 65A004 airfoil section at the tip. The fuselage has been indented in the region of the wing in order to obtain a favorable area distribution. The results reported herein consist of the performance and of the static longitudinal and lateral stability and control characteristics of the complete model. The Mach number range extended from 0.60 to 1.13, and the corresponding Reynolds number based on the wing mean aerodynamic chord varied from 1.77 x 10(exp 6) to 2.15 x 10(exp 6). The drag rise for both the cambered leading edge and symmetrical wing sections occurred at a Mach number of 0.95. Certain local modifications to the body which further improved the distribution of cross-sectional area gave additional reductions in drag at a Mach number of 1.00. The basic configuration indicated a mild pitch-up tendency at lift coefficients near 0.70 for the Mach number range from 0.80 to 0.90; however, the pitch-up instability may not be too objectionable on the basis of dynamic-stability considerations. The basic configuration indicated positive directional stability and positive effective dihedral through the angle-of-attack range and Mach number range with the exception of a region of negative effective dihedral at low lifts at Mach numbers of 1.00 and slightly above.

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: An analytical investigation has been carried out to determine the responses of a flicker-type roll control incorporated in a missile which traverses a range of Mach number of 6.3 at an altitude of 82,000 feet to 5.26 at an altitude of 282,000 feet. The missile has 80 deg delta wings in a cruciform arrangement with aerodynamic controls attached to the fuselage near the wing trailing edge and indexed 450 to the wings. Most of the investigation was carried out on an analog computer. Results showed that roll stabilization that may be adequate for many cases can be obtained over the altitude range considered, providing that the rate factor can be changed with altitude. The response would be improved if the control deflection were made larger at the higher altitudes. lag times less than 0.04 second improve the response appreciably. Asymmetries that produce steady rolling moments can be very detrimental to the response in some cases. The wing damping made a negligible contribution to the response.

Description: Measurements of the heat transfer from a horizontal cylinder rotating about its axis have been made with oil as the surrounding fluid to provide an addition to the heat-transfer results for this system heretofore available only for air. The results embrace a Prandtl number range from about 130 to 660, with Reynolds numbers up to 3 x 10(exp 4), and show an increasing dependence of free-convection heat transfer on rotation as the Prandtl number is increased by reducing the oil temperature. Some correlation of this effect, which agrees with the prior results for air, has been achieved. At higher rotative speeds the flow becomes turbulent, the free- convection effect vanishes, and the results with oil can be correlated generally with those for air and with mass-transfer results for even higher Prandtl numbers. For this system, however, the analogy calculations which have successfully related the heat transfer to the friction for pipe flows at high Prandtl numbers fail.

Description: The effects of Mach number and surface-roughness variation on boundary-layer transition were studied using fin-stabilized hollow-tube models in free flight. The tests were conducted over the Mach number range from 2.8 to 7 at a nominally constant unit Reynolds number of 3 million per inch, and with heat transfer to the model surface. A screwthread type of distributed two-dimensional roughness was used. Nominal thread heights varied from 100 microinches to 2100 microinches. Transition Reynolds number was found to increase with increasing Mach number at a rate depending simultaneously on Mach number and roughness height. The laminar boundary layer was found to tolerate increasing amounts of roughness as Mach number increased. For a given Mach number an optimum roughness height was found which gave a maximum laminar run greater than was obtained with a smooth surface.

Description: An investigation was made to determine the characteristics of a nonlinear linkage installed in a power control system incorporated in a ground simulator. The nonlinear linkage provided for increased control-stick motion for relatively small simulator response at control motions near neutral. The quality of the control system was rated on the ease and precision with which various tracking tasks were performed by the pilots who operated the simulator. The results obtained with the nonlinear linkage installed in the control system were compared with those obtained by using the normal linear control system. Several combinations of nonlinearity of the linkage were tested for various dynamic characteristics of the simulator. It was found that the pilots were able to track almost as well with the nonlinear linkage installed as with the normal system. All of the pilots were of the opinion, however, that the nonlinearity was an undesirable feature in the control system because of the apparent lack of simulator response through the neutral range of the linkage where relatively large stick deflections could be made with very little simulator motion. The results showed that increased lag between the target and chair position, higher stick-force levels, and uneven stick forces due to the dynamics of the linkage were general characteristics of all the nonlinear linkage conditions tested. It was also found that for cases of low simulator damping, rapid control motions caused considerably higher overshoots when the nonlinear linkage was installed than were obtained for the normal linear control system. These characteristics were considered to be sufficiently undesirable to out-weigh the advantages to be gained from the use of a nonlinear linkage in the control system of an airplane.

Description: The possibility of obtaining useful estimates of the static longitudinal stability of aircraft flying at high supersonic Mach numbers at angles of attack between 0 and +/-180 deg is explored. Existing theories, empirical formulas, and graphical procedures are employed to estimate the normal-force and pitching-moment characteristics of an example airplane configuration consisting of an ogive-cylinder body, trapezoidal wing, and cruciform trapezoidal tail. Existing wind-tunnel data for this configuration at a Mach number of 6.86 provide an evaluation of the estimates up to an angle of attack of 35 deg. Evaluation at higher angles of attack is afforded by data obtained from wind-tunnel tests made with the same configuration at angles of attack between 30 and 150 deg at five Mach numbers between 2.5 and 3.55. Over the ranges of Mach numbers and angles of attack investigated, predictions of normal force and center-of-pressure locations for the configuration considered agree well with those obtained experimentally, particularly at the higher Mach numbers.

Description: A wind-tunnel investigation was made at low speed in the Langley stability tunnel in order to determine the effects of fuselage nose length and a canopy on the oscillatory yawing derivatives of a complete swept-wing model configuration. The changes in nose length caused the fuselage fineness ratio to vary from 6.67 to 9.18. Data were obtained at various frequencies and amplitudes for angles of attack from 0 deg. to about 32 deg. Static lateral and longitudinal stability data are also presented.

Description: Results of an investigation of the static longitudinal stability and control characteristics of an aspect-ratio-3.1, unswept wing configuration equipped with an aspect-ratio-4, unswept horizontal tail are presented without analysis for the Mach number range from 0.70 to 2.22. The hinge line of the all-movable horizontal tail was in the extended wing chord plane, 1.66 wing mean aerodynamic chords behind the reference center of moments. The ratio of the area of the exposed horizontal-tail panels to the total area of the wing was 13.3 percent and the ratio of the total areas was 19.9 percent. Data are presented at angles of attack ranging"from -6 deg to +18 deg for the horizontal tail set at angles ranging from +5 deg to -20 deg and for the tail removed.

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: The fluid-dynamic characteristics of flat plates, 5 deg and 10 deg wedges, and 5 deg and 10 deg cones have been investigated at Mach numbers from 16.3 to 23.9 in helium flow. The flat-plate results are for a leading-edge Reynolds number range of 584 to 19,500 and show that the induced pressure distribution is essentially linear with the hypersonic viscous interaction parameter bar X within the scope of this investigation. It is also shown that the rate at which the induced pressure varies with bar X is a linear function of the leading-edge Reynolds number. The wedge and cone results show that as the flow-deflection angle increases, the induced-pressure effects decrease and the measured pressures approach those predicted by inviscid shock theory.

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: An experimental investigation has been made to determine the static stability characteristics of three thick wing models with parabolic plan forms at a Mach number of 3.11 for angles of attack from about -6 to 16 deg. The primary variable was aspect ratio, with the plan-form area and the ratio of base height to span kept the same for all three models. All models had stable, linear pitching-moment curves about the quarter chord of the wing mean aerodynamic chord. The model with the lowest aspect ratio attained a maximum untrimmed lift-drag ratio of about 5.0 at an angle of attack of about 8 deg. Increasing the aspect ratio (which was accompanied by an increase in base area because the ratio of the base height to span was kept constant) caused a decrease in maximum lift-drag ratio. All models were directionally stable for the range of angle of attack of the tests. Addition of a vertical tail to the models caused an increase in the directional stability over the angle-of-attack range. In general, the lateral aerodynamic characteristics of the models were not linear functions of angle of attack over any appreciable angle-of-attack range.

Description: An investigation of the static stability characteristics of several hypersonic boost-glide configurations has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel at Mach numbers of 1.41 and 2.01 (with Reynolds numbers per foot of 2.90 x 10(exp 6) and 2.41 x 10(exp 6) respectively). This series of configurations consisted of a cone, with and without cruciform fins, a trihedron, two low-aspect-ratio delta wings that differed primarily in cross-sectional shape, and two wing-body configurations. All configurations indicated reasonably linear pitching-, yawing-, and rolling-moment characteristics for angles of attack to at least 12 deg. The maximum lift-drag ratio for the zero-thrust condition (base drag included) was about 3 for the delta-wing configurations and about 4 for the wing-body configurations.

Description: Hypersonic-slender-body theory, in the limit as the free-stream Mach number becomes infinite, is used to find the effect of slightly perturbing the surface of slender two-dimensional and axisymmetric power law bodies, The body perturbations are assumed to have a power law variation (with streamwise distance downstream of the nose of the body). Numerical results are presented for (1) the effect of boundary-layer development on two dimensional and axisymmetric bodies, (2) the effect of very small angles of attack (on tow[dimensional bodies), and (3) the effect of blunting the nose of very slender wedges and cones.

Description: Heat-transfer coefficients and pressure distributions were obtained on a 4-inch-diameter flat-face cylinder in the Langley Unitary Plan wind tunnel. The measured stagnation heat-transfer coefficient agrees well with 55 percent of the theoretical value predicted by the modified Sibulkin method for a hemisphere. Pressure measurements indicated the dimensionless velocity gradient parameter r du\ a(sub t) dx, where x=0 at the stagnation point was approximately 0.3 and invariant throughout the Mach number range from 2.49 to 4.44 and the Reynolds number range from 0.77 x 10(exp 6) to 1.46 x 10(exp 6). The heat-transfer coefficients on the cylindrical afterbody could be predicted with reasonable accuracy by flat-plate theory at an angle of attack of 0 deg. At angles of attack the cylindrical afterbody stagnation-line heat transfer could be computed from swept-cylinder theory for large distances back of the nose when the Reynolds number is based on the distance from the flow reattachment points.

Description: Results of tests at Mach numbers of 3.0 and 7.3 for possible wing flutter of a series of models of a boost-glide-vehicle wing are presented herein. All of the models were tested at conditions which exceeded the proposed nominal design requirements for the full-scale vehicle; namely, dynamic pressure of 1,000 pounds per square foot at the test Mach numbers. None of the models experienced flutter; therefore, large margins of safety from wing flutter are indicated. However, the effects of body freedoms on the flutter characteristics and local types of flutter were not investigated.

Description: Heat-transfer data were evaluated from temperature time histories measured on a cooled cylindrical model with a cone-shaped nose and with turbulent flow at Mach numbers 3.00, 3.44, 4.08, 4.56, and 5.04. The experimental data were compared with calculated values using a modified Reynold's analogy between skin-friction and heat-transfer. Theoretical skin- friction coefficients were calculated using the method of Van Driest the method of Sommer and Short. The heat-transfer data obtained from the model were found to correlate when the 'T' method of Sommer and Short was used. The increase in turbulent heat-transfer rate with a reduction in wall to freestream temperature ratio was of the same order of magnitude as has been found for the turbulent skin-friction coefficient.

Description: The results of several flutter investigations to determine the effects of plan-form variations on the flutter characteristics of thin cantilevered wings at transonic Mach numbers have been reported previously. In the present investigation the data are extended to include a wing having an aspect ratio of 4, 45 of sweepback, and a taper ratio of 0.2. The data were obtained in the Langley transonic blowdown tunnel over a Mach number range from 0.6 to 1.4. The experimental results indicate an abrupt and rather large increase in both a flutter-speed parameter and a flutter-frequency parameter as the Mach number is increased from 1.05 to 1.10. The foregoing is interpreted as indicating a marked change in the flutter mode. Calculated flutter speeds, based on incompressible-flow aerodynamic coefficients, were too high by 20 percent or more throughout the subsonic Mach number range of the investigation. Calculated flutter frequencies were about 7 percent too high at a Mach number of 0.65 and were about 20 percent too high at a Mach number of 0.9. No significant independent effects of thickness were indicated for the plan form investigated as the thickness was changed from 3 to 4 percent chord.

Description: Thrust, air-handling, and base-pressure characteristics of five ejector configurations were investigated in the Lewis 8-by 6-foot wind tunnel at free-stream Mach numbers from 0 to 2.0 over ranges of primary-jet pressure ratio up to 24 and corrected secondary weight-flow ratio up to 13 percent. The ejector-shroud geometries varied from convergent to divergent. Base pressure ratio and ejector performance were interrelated by means of an exit-momentum parameter. Correlations, to at least a first approximation, with base pressure ratio, of both internal-ejector-flow separation and external-flow separation over the model boattail were shown. Furthermore, it was shown that magnitudes and exact trends in base pressure ratio depended largely, and in a complicated fashion, on ejector geometry and amount of secondary flow. External-stream effects on ejector jet thrust were determined for a typical schedule of jet-engine pressure ratios. With the exception of the ejector having the largest (1.81) shroud-exit-to primary-diameter ratio, there were no stream effects at Mach numbers from 1.5 to 2.0 and variations from quiescent-air thrust data were less than 2.5 percent at the subsonic speed investigated.

Description: A general relation, empirical in origin, for the mean velocity distribution of both laminar and turbulent boundary layers is proposed. The equation, in general, accurately describes the profiles in both laminar and turbulent flows. The calculation of profiles is based on a prior knowledge of momentum, displacement, and boundary-layer thickness together with free-stream conditions. The form for turbulent layers agrees with the present concepts of similarity of the outer layer. For the inner region or turbulent boundary layers the present relation agrees very closely with experimental measurements even in cases where the logarithmic law of the wall is inadequate. A unique relation between profile form factors and the ratio of displacement thickness to boundary-layer thickness is obtained for turbulent separation. A similar criterion is also obtained for laminar separation. These relations are demonstrated to serve as an accurate criterion for identifying separation in known profiles.

Description: Tests have been conducted in the Langley high-speed 7- by 10-foot tunnel to determine the effect of tail dihedral on lateral control effectiveness of a complete-model configuration having differentially deflected horizontal-tail surfaces. Limited tests were made to determine the lateral characteristics as well as the longitudinal characteristics in sideslip. The wing had an aspect ratio of 3, a taper ratio of 0.14, 28.80 deg sweep of the quarter-chord line with zero sweep at the 80-percent-chord line, and NACA 65A004 airfoil sections. The test Mach number range extended from 0.60 to 0.92. There are only small variations in the roll effectiveness parameter C(sub iota delta) with negative tail dihedral angle. The tail size used on the test model, however, is perhaps inadequate for providing the roll rates specified by current military requirements at subsonic speeds. The lateral aerodynamic characteristics were essentially constant throughout the range of sideslip angle from 12 deg to -12 deg. A general increase in yawing moment was noted with increased negative dihedral throughout the Mach number range.

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: The results of some experimental and theoretical studies of the interaction of oblique shock waves with laminar boundary layers are presented. Detailed measurements of pressure distribution, shear distribution, and velocity profiles were made during the interaction of oblique shock waves with laminar boundary layers on a flat plate. From these measurements a model was derived to predict the pressure levels characteristic of separation and the length of the separated region.

Description: Theoretical analysis of the longitudinal behavior of an automatically controlled supersonic interceptor during the attack phase against a nonmaneuvering target is presented. Control of the interceptor's flight path is obtained by use of a pitch rate command system. Topics lift, and pitching moment, effects of initial tracking errors, discussion of normal acceleration limited, limitations of control surface rate and deflection, and effects of neglecting forward velocity changes of interceptor during attack phase.

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: Reported herein are the results of observations and measurements made in connection with a study of the phenomenon of the development of atmosphere-connected cavities about surface-piercing struts. Conditions for the existence of such ventilated flows which have been derived from the experimental data are presented. In addition, certain broad conclusions pertinent to model testing and full-scale design are reached. Further experimentation to define the inception of ventilation as a function of boundary-layer state or Reynolds number is required.

Description: Normal forces, axial forces, pitching moments, and rolling moments on the model and hinge moments on each of the four control surfaces were measured. Control surfaces were deflected from -35 deg to 15 deg in various combinations to produce pitching, yawing, and rolling moments on the model over a range of angles of attack from -5 deg to 25 deg at roll angles from -135 deg to 45 deg.

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: The effect of an external boundary layer on the performance of an axisymmetric external-internal-compression inlet was evaluated at Mach numbers of 3.0 and 2.5 and Reynolds numbers from 2.2 to 0.5 x 10(exp 6) per foot. The inlet was tested at locations up to two-thirds of the way into the 1.7- and 9.0-inch boundary layers generated by a flat plate and the tunnel floor, respectively. The inlet could be readily started at all conditions tested, including those where the boundary layer was separated upstream of the inlet by the various shock systems during the restart cycle. Although the inlet performance decreased with increasing immersion into the boundary layer at both Mach numbers, the inlet was more sensitive to boundary-layer ingestion at the design Mach number of 3.0.

Description: Slender-body theory for subsonic and supersonic flow past bodies of revolution is extended to a second approximation, Methods are developed for handling the difficulties that arise at round ends, Comparison is made with experiment and with other theories for several simple shapes.

Description: Some 100 numerical computations have been carried out for unyawed bodies of revolution with detached bow waves. The gas is assumed perfect with gamma = 5/3, 7/5, or 1. Free-stream Mach numbers are taken as 1.2, 1.5, 2, 3, 4, 6, 10, and infinity. The results are summarized with emphasis on the sphere and paraboloid.

Description: A previous analysis of turbulent heat transfer and flow with variable fluid properties in smooth passages is extended to flow over a flat plate at high Mach numbers, and the results are compared with experimental data. Velocity and temperature distributions are calculated for a boundary layer with appreciative effects of frictional heating and external heat transfer. Viscosity and thermal conductivity are assumed to vary as a power or the temperature, while Prandtl number and specific heat are taken as constant. Skin-friction and heat-transfer coefficients are calculated and compared with the incompressible values. The rate of boundary-layer growth is obtained for various Mach numbers.

Description: Exploratory tests of a circular internal-contraction inlet were made at Mach numbers of 2.00 and 2.35 to determine the effect of a cowl-type boundary-layer control located downstream of the inlet throat. The inlet was designed for a Mach number of 2.5. Tests were also made of the inlet modified to correspond to design Mach numbers of 2.35 and 2.25. Surveys near the minimum area section of the inlet without boundary-layer control indicated maximum averaged pressure recoveries between 0.90 and 0.92 at a free-stream Mach number, M(sub infinity), of 2.35 for the inlets. Farther downstream, after partial subsonic diffusion, a maximum pressure recovery of 0.842 was obtained with the inlet at M(sub infinity) = 2.35. The pressure recovery of the inlet was increased by 0.03 at a Mach number of 2.35 and decreased by 0.02 at a Mach number of 2.00 by the application of cowl-type boundary-layer control. Further investigation with the inlet without bleed demonstrated that an increase of angle of attack from 0 deg to 3 deg reduced the pressure recovery 0.04. The effect of Reynolds number was to increase pressure recovery 0.07 (from 0.785 to 0.855) with an increase in Reynolds number (based on inlet diameter) from 0.79 x 10(exp 6) to 3.19 x 10(exp 6).

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: A detailed report is given of exact (numerical) solutions of the laminar-boundary-layer equations for the Prandtl number range appropriate to liquid metals (0.003 to 0.03). Consideration is given to the following situations: (1) forced convection over a flat plate for the conditions of uniform wall temperature and uniform wall heat flux, and (2) free convection over an isothermal vertical plate. Tabulations of the new solutions are given in detail. Results are presented for the heat-transfer and shear-stress characteristics; temperature and velocity distributions are also shown. The heat-transfer results are correlated in terms of dimensionless parameters that vary only slightly over the entire liquid-metal range. Previous analytical and experimental work on low Prandtl number boundary layers is surveyed and compared with the new exact solutions.

Description: A hydrodynamic investigation was made in Langley tank no. 1 of a planing surface which was curved longitudinally in the shape of a circular arc with the center of curvature above the model and had a beam of inches and a radius of curvature of 20 beams. The planing surface had length-beam ratio of 9 and an angle of dead rise of 0 deg. Wetted length, resistance, and trimming moment were determined for values of load coefficient C(sub Delta) from -4.2 to 63.9 and values of speed coefficient C(sub V) from 6 to 25. The effects of convexity were to increase the wetted length-beam ratio (for a given lift), to decrease the lift-drag ratio, to move the center of pressure forward, and ta increase the trim for maximum lift-drag ratio as compared with values for a flat surface. The effects were greatest at low trims and large drafts. The maximum negative lift coefficient C(sub L,b) obtainable with a ratio of the radius of curvature to the beam of 20 was -0.02. The effects of camber were greater in magnitude for convexity than for the same amount of concavity.

Description: Results of an investigation to determine the static longitudinal stability and control characteristics of an aspect-ratio-2 triangular wing and body configuration equipped with either a canard control, a trailing-edge-flap control, or a cambered forebody are presented without analysis for Mach numbers from 0.70 to 2.22. The canard surface had a triangular plan form and a ratio of exposed area to total wing area of 7.8 percent. The hinge line of the canard was in the extended wing chord plane, 0.83 wing mean aerodynamic chord ahead of the reference center of moments. The trailing-edge controls were constant-chord full-span flaps with exposed area equal to 10.7 percent of the total wing area. The cambered body was a modified Sears-Haack body with camber only ahead of the wing apex. Data are presented for various canard and flap deflections at angles of attack ranging from -6 deg to +18 deg.

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 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: Measurements of peak overpressure and Mach stem height were made at four burst heights. Data were obtained with instrumentation capable of directly observing the variation of shock wave movement with time. Good similarity of free air shock peak overpressure with larger scale data was found to exist. The net effect of surface roughness on shock peak overpressures slightly. Surface roughness delayed the Mach stem formation at the greatest charge height and lowered the growth at all burst heights. A similarity parameter was found which approximately correlates the triple point path at different burst heights.

Description: The profiles and thicknesses of normal shock waves in argon at Mach numbers of 1.335, 1.454, 1.576, and 1-713 were determined experimentally by means of a free-molecule probe whose equilibrium temperature is related by kinetic theory to the local flow properties and their gradients. Comparisons were made between the experimental shock profiles and the theoretical profiles calculated from the Navier-Stokes equations, the Grad 13-moment equations, and the Burnett equations. New, very accurate numerical integrations of the Burnett equations were obtained for this purpose with results quite different from those found by Zoller, to whom the solution of this problem is frequently attributed. The experimental shock profiles were predicted with approximately equal success by the Navier-Stokes and Burnett theories, while the 13-moment method was definitely less satisfactory. A surprising feature of the theoretical results is the relatively small difference in predictions between the Navier-Stokes and Burnett theories in the present range of shock strengths and the contrastingly large difference between predictions of Burnett and the 13-moment theories. It is concluded that the Navier-Stokes equations are correct for weak shocks and that within the present shock strength range the Burnett equations make no improvement which merits the trouble of solving them. For shocks of noticeably greater strength, say with a shock Mach number of more than 2.5, it remains fundamentally doubtful that any of these theories can be correct.

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: The results are presented in the form of preliminary design charts which give a comparison between the dynamic-response factors of the semi-rigid case and the airplane longitudinal short-period case and between the dynamic-response factors of the semi-rigid case and the steady-state value of the airplane longitudinal short-period response. These charts can be used to estimate the first-order effects of the addition of a wing-bending degree of freedom on the short-period dynamic-response factor and on the maximum dynamic-response factor when compared with the steady-state response of the system.

Description: An examination of the effects of compressibility, variable properties, and body forces on fully developed laminar flow has indicated several limitations on such streams. In the absence of a pressure gradient, but presence of a body force (e.g., gravity), an exact fully developed gas flow results. For a liquid this follows also for the case of a constant streamwise pressure gradient. These motions are exact in the sense of a Couette flow. In the liquid case two solutions (not a new result) can occur for the same boundary conditions. An approximate analytic solution was found which agrees closely with machine calculations.In the case of approximately exact flows, it turns out that for large temperature variations across the channel the effects of convection (due to, say, a wall temperature gradient) and frictional heating must be negligible. In such a case the energy and momentum equations are separated, and the solutions are readily obtained. If the temperature variations are small, then both convection effects and frictional heating can consistently be considered. This case becomes the constant-property incompressible case (or quasi-incompressible case for free-convection flows) considered by many authors. Finally there is a brief discussion of cases wherein streamwise variations of all quantities are allowed but only a such form that independent variables are separable. For the case where the streamwise velocity varies inversely as the square root distance along the channel a solution is given.

Description: A method is presented for the calculation of lift coefficients for rectangular lifting surfaces of aspect ratios from 0.125 to 10 operating at finite depths beneath the water surface, including the zero depth or planing condition. Theoretical values are compared with experimental values obtained at various depths of submergence with lifting surfaces of aspect ratios from 0.125 to 10. The method can also be applied to hydrofoils with dihedral. Lift coefficients computed by this method are in good agreement with existing experimental data for aspect ratios from 0.125 to 10 and dihedral angles up to 30 deg.

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: Some experimental and theoretical studies have been made of axisymmetric free jets exhausting from sonic and supersonic nozzles into still air and into supersonic streams with a view toward problems associated with propulsive jets and the investigation of these problems. For jets exhausting into still air, consideration is given to the effects of jet Mach number, nozzle divergence angle, and jet static pressure ratio upon jet structure, jet wavelength, and the shape and curvature of the jet boundary. Studies of the effects of the ratio of specific heats of the jets are included are observations pertaining to jet noise and jet simulation. For jets exhausting into supersonic streams, an attempt has been made to present primarily theoretical certain jet interference effects and in formulating experimental studies. The primary variables considered are jet Mach number, free stream Mach number, jet static pressure ratio, ratio of specific heats of the jet, nozzle exit angle, and boattail angle. The simulation problem and the case of a hypothetical hypersonic vehicle are examined, A few experimental observations are included.

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: An approximate theoretical analysis was made of the shielding mechanism whereby the rate of heat transfer to the forward stagnation point of blunt bodies is reduced by melting and evaporation. General qualitative results are given and a numerical example, the melting and evaporation of ice, is presented and discussed in detail.

Description: A 0.10-scale model of a swept-wing fighter airplane was tested in the Langley high-speed 7- by 10-foot tunnel at Mach numbers from 0.60 to 0.92 to determine the effects of adding underfuselage speed brakes. The results of brief spoiler-aileron lateral control tests also are included. The tests show acceptable trim and drag increments when the speed brakes are installed at the 32-71-inch fuselage station.

Description: Force tests of the static and dynamic lateral stability characteristics of a VTOL airplane having a triangular wing mounted high on the fuselage with a triangular vertical tail on top of the wing and no horizontal tail have been made in the Langley free-flight tunnel. The static lateral stability parameters and the rolling, yawing, and sideslipping dynamic stability derivatives are presented without analysis.

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.