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  • Inorganic Chemistry  (3,066)
  • Cell & Developmental Biology  (673)
  • Aerodynamics  (141)
  • 1955-1959  (3,880)
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Year
  • 1
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    Design Guide for Pitch-Up Evaluation and Investigation at High Subsonic Speeds of Possible Limitations Due to Wing-Aspect-Ratio Variations (1959)
    In:  CASI
    Details
    Publication Date: 2019-05-11
    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.
    Keywords: Aerodynamics
    Type: NASA-TM-X-26
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  • 2
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    An Investigation of Some Aspects of the Sonic Boom by Means of Wind-Tunnel Measurements of Pressures About Several Bodies at a Mach Number of 2.01 (1959)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    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.
    Keywords: Aerodynamics
    Type: NASA-TN-D-161
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  • 3
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    Ground Measurements of the Shock-Wave Noise from Airplanes in Level Flight at Mach Numbers to 1.4 and at Altitudes to 45000 Feet (1959)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    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.
    Keywords: Aerodynamics
    Type: NASA-TN-D-48
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  • 4
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    A Dynamic Model Investigation of the Effect of a Sharp-Edge Vertical Gust on Blade Periodic Flapping Angles and Bending Moments of a Two-Blade Rotor (1959)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    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.
    Keywords: Aerodynamics
    Type: NASA-TN-D-31
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  • 5
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    Longitudinal Aerodynamic Characteristics of a Wing-Body-Tail Model Having a Highly Tapered, Cambered 45 degree Swept Wing of Aspect Ratio 4 at Transonic Speeds (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    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.
    Keywords: Aerodynamics
    Type: NASA-TM-X-130
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  • 6
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    Investigation at Mach Numbers of 0.20 to 3.50 of a Blended Diamond Wing and Body Combination of Sonic Design but with Low Wave-Drag Increase with Increasing Mach Number (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    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.
    Keywords: Aerodynamics
    Type: NASA-TM-X-105
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  • 7
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    Transonic Aerodynamic Characteristics of a 45 deg Swept Wing Fuselage Model with a Finned and Unfinned Body Pylon Mounted Beneath the Fuselage or Wing, Including Measurements of Body Loads (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-4-20-59L , L-206
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  • 8
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    Analysis and Evaluation of Supersonic Underwing Heat Addition (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-3-17-59E
    Format: text
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  • 9
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    Performance Characteristics of Flush and Shielded Auxiliary Exits at Mach Numbers of 1.5 to 2.0 (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-5-18-59E , E-139
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  • 10
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    Effect on Inlet Performance of a Cowl Visor and an Internal-Contraction Cowl for Drag Reduction at Mach Numbers 3.07 and 1.89 (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-3-18-59E , E-173
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  • 11
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    Charts of the Induced Velocities Near a Lifting Rotor (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-4-15-59L
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  • 12
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    Low-Speed Tests of Semispan-Wing Models at Angles of Attack from 0 to 180 degrees (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-2-27-59A
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  • 13
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    The Effect of Leading-Edge Sweep and Surface Inclination on the Hypersonic Flow Field Over a Blunt Flat Plate (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-12-26-58A
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  • 14
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    Basic Pressure Measurements at Transonic Speeds on a Thin 45 deg Sweptback Highly Tapered Wing with Systematic Spanwise Twist Variations (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-5-12-59L
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  • 15
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    Surface Pressure Distribution at Hypersonic Speeds for Blunt Delta Wings at Angle of Attack (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-5-12-59A
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  • 16
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    Aerodynamic Heating and Fatigue (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-6-4-59W
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  • 17
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    Review of Aircraft Altitude Errors Due to Static-Pressure Source and Description of Nose-Boom Installations for Aerodynamic Compensation of Error (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-5-10-59L
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  • 18
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    The Effects of Streamwise-Deflected Wing Tips on the Aerodynamic Characteristics of an Aspect Ratio-2 Triangular Wing, Body, and Tail Combination (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-5-18-59A
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  • 19
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    An Investigation of the Pitching-Moment Contribution of a High Horizontal Tail on an Unswept-Wing and Body Combination at Mach Numbers from 0.80 to 1.40 (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    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.
    Keywords: Aerodynamics
    Type: NASA-TM-X-43
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  • 20
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    Free-Spinning-Tunnel Investigation of a 1/25-Scale Model of the Chance Vought F8U-1P Airplane (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    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.
    Keywords: Aerodynamics
    Type: NASA-TM-SX-196 , L-714 , NASA-AD-3137
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  • 21
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    Basic Pressure Measurements at Transonic Speeds on a Thin 45 deg Sweptback highly Tapered Wing With Systematic Spanwise Twist Variations (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-2-24-59L , L-207
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  • 22
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    Flow Characteristics About Two Thin Wings of Low Aspect Ratio Determined from Surface Pressure Measurements Obtained in Flight at Mach Numbers from 0.73 to 1.90 (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-5-1-59H
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  • 23
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    A Hydrogen Peroxide Hot-Jet Simulator for Wind-Tunnel Tests of Turbojet-Exit Models (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-1-10-59L
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  • 24
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    The Effect of Moment of Area Rule Modifications on the Drag, Lift and Pitching Moment Characteristics of an Unswept Aspect Ratio 6 Wing and Body Combination (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-2-24-59A , A-145
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  • 25
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    Low-Speed Wind-Tunnel Investigation of Blowing Boundary-Layer Control on Leading- and Trailing-Edge Flaps of a Large-Scale, Low-Aspect-Ratio, 45 Swept-wing Airplane Configuration (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-1-23-59A
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  • 26
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    Maximum Mean Lift Coefficient Characteristics at Low Tip Mach Numbers of a Hovering Helicopter Rotor Having an NACA 64(1)A012 Airfoil Section (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-1-23-59L
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  • 27
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    Transonic Aerodynamic Characteristics of Two Wedge Airfoil Sections Including Unsteady Flow Studies (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-4-30-59L
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  • 28
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    Free-Flight Investigation of a Rocket-Propelled Model to Determine the Aerodynamic Heating on a Thin, Unswept, Untapered, Multispar, Aluminum-Alloy Wing at Mach Numbers up to 2.22 (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-12-15-58L
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  • 29
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    Tables for the Rapid Estimation of Downwash and Sidewash Behind Wings Performing Various Motions at Supersonic Speeds (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-2-20-59L
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  • 30
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    Supersonic and Moment-of-Area Rules Combined for Rapid Zero-Lift Wave-Drag Calculations (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-4-19-59A , A-158
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  • 31
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    Free-Flight Test of a Technique for Inflating an NASA 12-Foot-Diameter Sphere at High Altitudes (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-2-5-59L , L-214
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  • 32
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    Boundary-Induced Downwash Due to Lift in a Two-Dimensional Slotted Wind Tunnel (1959)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    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.
    Keywords: Aerodynamics
    Type: NASA-TR-R-25
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  • 33
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    Investigation on the use of a Freely Rotating Rotor at the Cowl Face of a Supersonic Conical Inlet to Reduce Inlet Flow Distortion (1959)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-5-28-59L
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  • 34
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    Effect of Multiple-Jets Exits on the Base Pressure of a Simple Wing-Body Combination at Mach Numbers of 0.6 to 1.27 (1959)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    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.
    Keywords: Aerodynamics
    Type: NASA-TM-X-25
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  • 35
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    Effects of Outboard Thickened and Blunted Leading Edges on the Wave Drag of a 45 Degree Swept-Wing and Body Combination (1959)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    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).
    Keywords: Aerodynamics
    Type: NASA-TM-X-27
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  • 36
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    An Experimental Investigation of the Pressure Distribution on A 1/15-Scale Model of the Lockheed WS-117L Vehicle Plus Booster "B" at Mach Numbers from 0.70 to 1.45 (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    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.
    Keywords: Aerodynamics
    Type: NASA-MEMO-3-12-59A , A-217 , AF-AM-163
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  • 37
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    An Investigation of the Pressure Distribution on a 1/15-Scale Model of the Lockheed WS-117L Vehicle Plus Booster "B" at Mach Numbers from 1.55 to 2.35 (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Pressure coefficients were measured over the vehicle and over the forward part of the booster at Reynolds numbers of 3.0 x 10(exp 6) per foot. Tabular results are presented for two nose shapes at Mach numbers of 1.55, 1.75, 2.00, and 2.35, at angles of attack from -4 deg to +10 deg, and at 0 deg sideslip.
    Keywords: Aerodynamics
    Type: NASA-MEMO-3-13-59A , AF-AM-163
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  • 38
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    Full-Scale Wind-Tunnel Investigation of a Jet Flap in Conjunction with a Plain Flap with Blowing Boundary-Layer Control on a 35 deg Sweptback-Wing Airplane (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Previous investigations have shown that increased blowing at the hinge-line radius of a plain flap will give flap lift increases above that realized with boundary-layer control. Other experiments and theory have shown that blowing from a wing trailing edge, through the jet flap effect, produced lift increases. The present investigation was made to determine whether blowing simultaneously at the hinge-line radius and trailing edge would be more effective than blowing separately at either location. The tests were made at a Reynolds number of 4.5 x 10(exp 6) with a 35 deg sweptback-wing airplane. For this report, only the lift data are presented. Of the three flap blowing arrangements tested, blowing distributed between the trailing edge and the hinge-line radius of a plain flap was found to be superior to blowing at either location separately at the plain flap deflections of interest. Comparison of estimated and experimental jet flap effectiveness was fair.
    Keywords: Aerodynamics
    Type: NASA-MEMO-2-20-59A
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  • 39
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    Influence of Large Positive Dihedral on Heat Transfer to Leading Edges of Highly Swept Wings at Very High Mach Numbers (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A geometric study has been made of some of the effects of dihedral on the heat transfer to swept delta wings. The results of this study show that the incorporation of large positive dihedral on highly swept wings can shift, even at moderately low angles of attack, the stagnation-line heat-transfer problem from the leading edges to the axis of symmetry (ridge line). An order-of-magnitude analysis (assuming laminar flow) indicates conditions for which it may be possible to reduce the heating at the ridge line (except in the vicinity of the wing apex) to a small fraction of the leading-edge heat transfer of a flat wing at the same lift. Furthermore, conditions are indicated where dihedral reduces the leading-edge heat transfer for angles of attack less than those required to shift the stagnation line from the leading edge to the ridge line.
    Keywords: Aerodynamics
    Type: NASA-MEMO-3-7-59L
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  • 40
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    Wind-Tunnel Measurements of Effect of Dive-Recovery Flaps at Transonic Speeds on Models of a Seaplane and a Transport (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The effects of wing-lower-surface dive-recovery flaps on the aero- dynamic characteristics of a transonic seaplane model and a transonic transport model having 40 deg swept wings have been investigated in the Langley 16-foot transonic tunnel. The seaplane model had a wing with an aspect ratio of 5.26, a taper ratio of 0.333, and NACA 63A series airfoil sections streamwise. The transport model had a wing with an aspect ratio of 8, a taper ratio of 0.3, and NACA 65A series airfoil sections perpendicular to the quarter-chord line. The effects of flap deflection, flap longitudinal location, and flap sweep were generally investigated for both horizontal-tail-on and horizontal-tail-off configurations. Model force and moment measurements were made for model angles of attack from -5 deg to 14 deg in the Mach number range from 0.70 to 1.075 at Reynolds numbers of 2.95 x 10(exp 6) to 4.35 x 10(exp 6). With proper longitudinal location, wing-lower-surface dive-recovery flaps produced lift and pitching-moment increments that increased with flap deflection. For the transport model a flap located aft on the wing proved to be more effective than one located more forward., both flaps having the same span and approximately the same deflection. For the seaplane model a high horizontal tail provided added effectiveness for the deflected-flap configuration.
    Keywords: Aerodynamics
    Type: NASA-MEMO-6-9-59L , L-292
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  • 41
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    Wind-Tunnel Investigation of a Small-Scale Sweptback-Wing Jet-Transport Model Equipped with an External-Flow Jet-Augmented Double Slotted Flap (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A wind-tunnel investigation at low speeds has been made to study the aerodynamic characteristics of a small-scale sweptback-wing Jet-transport model equipped with an external-flow jet-augmented double slotted flap. Included in the investigation were tests of the wing alone to study the effects of varying the spanwise extent of blowing on the full-span flap. The results indicated that the double-slotted-flap arrangement of the present investigation was more efficient in terms of lift and drag than were the external-flow single-slotted-flap arrangements previously tested and gave a substantial reduction In the thrust-weight ratio required for a given lift coefficient under trimmed drag conditions. An increase in the spanwise extent of blowing on the full-span flap was also found to increase the efficiency of the model in terms of the lift and drag but, as would be expected on a sweptback-wing configuration, was accompanied by significant increases in negative pitching moment.
    Keywords: Aerodynamics
    Type: NASA-MEMO-3-8-59L
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  • 42
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    A Numerical Method for Calculating the Wave Drag of a Configuration from the Second Derivative of the Area Distribution of a Series of Equivalent Bodies of Revolution (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A method based on linearized and slender-body theories, which is easily adapted to electronic-machine computing equipment, is developed for calculating the zero-lift wave drag of single- and multiple-component configurations from a knowledge of the second derivative of the area distribution of a series of equivalent bodies of revolution. The accuracy and computational time required of the method to calculate zero-lift wave drag is evaluated relative to another numerical method which employs the Tchebichef form of harmonic analysis of the area distribution of a series of equivalent bodies of revolution. The results of the evaluation indicate that the total zero-lift wave drag of a multiple-component configuration can generally be calculated most accurately as the sum of the zero-lift wave drag of each component alone plus the zero-lift interference wave drag between all pairs of components. The accuracy and computational time required of both methods to calculate total zero-lift wave drag at supersonic Mach numbers is comparable for airplane-type configurations. For systems of bodies of revolution both methods yield similar results with comparable accuracy; however, the present method only requires up to 60 percent of the computing time required of the harmonic-analysis method for two bodies of revolution and less time for a larger number of bodies.
    Keywords: Aerodynamics
    Type: NASA-MEMO-1-16-59A
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  • 43
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    Wind-Tunnel Investigation of Subsonic Longitudinal Aerodynamic Characteristics of a Tiltable-Wing Vertical-Take-Off-and-Landing Supersonic Bomber Configuration Including Turbojet Power Effects (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Jet-powered model tests were made to determine the low-speed longitudinal aerodynamic characteristics of a vertical-take-off and-landing supersonic bomber configuration. The configuration has an unique engine-wing arrangement wherein six large turbojet engines (three on each side of the fuselage) are buried in a low-aspect-ratio wing which is tilted into the vertical plane for take-off. An essentially two-dimensional variable inlet, spanning the leading edge of each wing semispan, provides air for the engines. Jet flow conditions were simulated for a range of military (nonafterburner) and afterburner turbojet-powered flight at subsonic speeds. Three horizontal tails were tested at a station down-stream of the jet exit and at three heights above the jet axes. A semi-span model was used and test parameters covered wing-fuselage incidence angles from 0 deg to 15 deg, wing angles of attack from -4 deg to 36 deg, a variable range of horizontal-tail incidence angles, and some variations in power simulation conditions. Results show that, with all horizontal tails tested, there were large variations in static stability throughout the lift range. When the wing and fuselage were alined, the model was statically stable throughout the test range only with the largest tail tested (tail span of 1.25 wing span) and only when the tail was located in the low test position which placed the tail nearest to the undeflected jet. For transition flight conditions, none of the tail configurations provided satisfactory longitudinal stability or trim throughout the lift range. Jet flow was destabilizing for most of the test conditions, and varying the jet-exit flow conditions at a constant thrust coefficient had little effect on the stability of this model. Wing leading-edge simulation had some important effects on the longitudinal aerodynamic characteristics.
    Keywords: Aerodynamics
    Type: NASA-MEMO-1-8-59L
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  • 44
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    Chordwise Pressure Distributions over Several NACA 16-Series Airfoils at Transonic Mach Numbers up to 1.25 (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A two-dimensional wind-tunnel investigation of the pressure distributions over several NACA 16-series airfoils with thicknesses of 4, 6, 9, and 12 percent of the chord and design lift coefficients of 0, 0.2, 1 and 0.5 has been conducted in the Langley airfoil test apparatus at transonic Mach numbers from 0.7 to 1.25. The tests ranged in Reynolds number from 2.4 x 10 (exp 6) to 2.8 x 10 (exp 6) and in angle of attack from -10 to 12 degrees. Chordwise pressure distributions and schlieren flow photographs are presented without analysis.
    Keywords: Aerodynamics
    Type: NASA-MEMO-6-1-59L
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  • 45
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    Modified Matrix Method for Calculating Steady-State Span Loading on Flexible Wings in Subsonic Flight (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A method is presented for shortening the computations required to determine the steady-state span loading on flexible wings in subsonic flight. The method makes use of tables of downwash factors to find the necessary aerodynamic-influence coefficients for the application of lifting-line theory. Explicit matrix equations of equilibrium are converted into a matrix power series with a finite number of terms by utilizing certain characteristic properties of matrices. The number of terms in the series is determined by a trial-and-error process dependent upon the required accuracy of the solution. Spanwise distributions of angle of attack, airload, shear, bending moment, and pitching moment are readily obtained as functions of qm(sub R) where q denotes the dynamic pressure and mR denotes the lift-curve slope of a rigid wing. This method is intended primarily to make it practical to solve steady-state aeroelastic problems on the ordinary manually operated desk calculators, but the method is also readily adaptable to automatic computing equipment.
    Keywords: Aerodynamics
    Type: NASA-MEMO-5-26-59L
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  • 46
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    Inlet Performance Characteristics from Wind-Tunnel Tests of a 0.10-Scale Air-Induction System Model of the YF-108A Airplane at Mach Numbers of 2.50, 2.76, and 3.00 (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Inlet-performance and external-drag-coefficient characteristics are presented without analysis. Effects are shown of variations of fuselage boundary-layer diverter profile, bleed-surface porosity, bleed-exit area, and inlet ramp, and lip angle.
    Keywords: Aerodynamics
    Type: NASA-MEMO-7-18-59A , AF-AM-157
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  • 47
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    Pressure Drag of Axisymmetric Cowls Having Large Initial Lip Angles at Mach Numbers from 1.90 to 4.90 (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The results of experimental and theoretical data on nine cowls are presented to determine the effect of initial lip angle and projected frontal area on the cowl pressure drag coefficient at Mach numbers from 1.90 to 4.90. The experimental drag coefficients were approximated well with two-dimensional shock-expansion theory at the lower cowl-projected areas, but the difference between theory and experiment increased as the cowl area ratio was increased or as shock detachment at the cowl lips was approached. An empirical chart is presented, which can be used to estimate the cowl pressure drag coefficient of cowls approaching an elliptic contour.
    Keywords: Aerodynamics
    Type: NASA-MEMO-1-10-59E
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  • 48
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    Aerodynamic Loading Characteristics Including Effects of Aeroelasticity of a Thin-Trapezoidal-Wing-Body Combination at Mach Number of 1.43 (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Results have been obtained in the Langley 8-foot transonic pressure tunnel at a Mach number of 1.43 and at angles of attack from 0 deg to about 24 deg which indicate the static-aerodynamic-loads characteristics for a 2-percent-thick trapezoidal wing in combination with a body. Included are the effects of changing Reynolds number and of fixing boundary-layer transition. The results show that aerodynamic loading characteristics at a Mach number of 1.43 are similar to those reported in NACA RM L56Jl2a for the same configuration at a Mach number of 1.115. Reducing the Reynolds number resulted in reductions in the deflection of the wing and caused a slight increase in the relative loading over the outboard wing sections since the deflections were in a direction to unload the tip sections. Little or no effects were seen to result from fixing boundary-layer transition at a tunnel stagnation pressure of 1,950 pounds per square foot.
    Keywords: Aerodynamics
    Type: NASA-TM-X-119
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  • 49
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    An Experimental Study at a Mach Number of 3 of the Effect of Turbulence Level and Sandpaper Type Roughness on Transition on a Flat Plate (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An investigation has been conducted at a Mach number of 3 of the effect of turbulence level and sandpaper-type roughness on transition for a flat plate. The Reynolds number varied from 0.8 x 10(exp 6) to 1.8 x 10(exp 6) per inch; the settling-chamber turbulence level varied from 0.7 percent to 35 percent; and the heat transfer between the plate and the stream was negligible. Transition locations were determined by an optical method. This method was indicative of a permanent change in the boundary-layer density distribution rather than the onset of turbulent bursts. Results showed that, when transition was influenced by roughness, it moved in a way similar to its movement on a smooth plate. That is, it gradually approached the roughness location with either an increase in unit Reynolds number or an increase in turbulence level. For roughness submerged in the linear portion of the boundary-layer velocity profile, the square root of the roughness Reynolds number and the ratio of roughness height to boundary-layer displacement thickness gave similar results as parameters for predicting the effects of roughness. A range of each of these parameters which moved transition less than 10 percent was found and this range was a function of turbulence level.
    Keywords: Aerodynamics
    Type: NASA-MEMO-2-9-59L
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  • 50
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    A Systematic Kernel Function Procedure for Determining Aerodynamic Forces on Oscillating or Steady Finite Wings at Subsonic Speeds (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Details are given of a numerical solution of the integral equation which relates oscillatory or steady lift and downwash distributions in subsonic flow. The procedure has been programmed for the IBM 704 electronic data processing machine and yields the pressure distribution and some of its integrated properties for a given Mach number and frequency and for several modes of oscillation in from 3 to 4 minutes, results of several applications are presented.
    Keywords: Aerodynamics
    Type: NASA-TR-R-48
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  • 51
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    Pressure and Force Characteristics of Noncircular Cylinders as Affected by Reynolds Number with a Method Included for Determining the Potential Flow About Arbitrary Shapes (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The low-speed pressure-distribution and force characteristics of several noncircular two-dimensional cylinders were measured in wind tunnel through a range of Reynolds numbers and flow incidences. A method of determining the potential-flow pressure distribution for arbitrary cross sections is described. Application of the data in predicting the spin characteristics of fuselages is briefly discussed.
    Keywords: Aerodynamics
    Type: NASA-TR-R-46
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  • 52
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    Wind-Tunnel Investigation at Mach Numbers from 0.40 to 1.14 of the Static Aerodynamic Characteristics of a Nonlifting Vehicle Suitable for Reentry (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An investigation was conducted in the Langley 8-foot transonic pressure tunnel to investigate the static pitching-moment, normal-force, and axial-force characteristics on a model of a nonlifting vehicle suit- able for reentry. The vehicle was designed to use a heat sink and blunt shape to alleviate the effects of the heating encountered during reentry of the earth's atmosphere. The effects of modifying the intersection of the face of the model with the afterbody from a sharp corner to a rounded edge were also investigated. Tests were conducted at Mach numbers from 0.40 to 1.14 and at angles of attack from approximately -3 deg to 20 deg. The Reynolds number varied from about 2.0 x 10(exp 6) to 3.6 x 10(exp 6). The results show that the model had a low positive static-stability level, low normal-force coefficients, and large axial-force coefficients. The model trimmed, for the angle-of-attack range investigated, at angles of attack near zero. The effects on the stability as a result of rounding the corner were negligible.
    Keywords: Aerodynamics
    Type: NASA-MEMO-4-13-59L , L-437
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  • 53
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    Effect of Afterbody-Ejector Configurations on the Performance at Transonic Speeds of a Pylon-Supported Nacelle Model having a Hot-Jet Exhaust (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An investigation of several afterbody-ejector configurations on a pylon-supported nacelle model has been completed in the Langley 16-foot transonic tunnel at Mach numbers from 0.80 to 1.05. The propulsive performance of two nacelle afterbodies with low boattailing and long ejector spacing was compared with a configuration corresponding to a turbojet-engine installation having a highly boattailed afterbody with a short ejector. The jet exhaust was simulated with a hydrogen peroxide turbojet simulator. The angle of attack was maintained at 0 deg, and the average Reynolds number based on body length was 20 x 10(exp 6). The results of the investigation indicated that the configuration with a conical afterbody with smooth transition to a 15 deg boattail angle had large beneficial jet effects on afterbody pressure-drag coefficient and had the best thrust-minus-drag performance of the afterbody-ejector configurations investigated.
    Keywords: Aerodynamics
    Type: NASA-MEMO-1-4-59L , L-133
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  • 54
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    Free-Spinning-Tunnel Investigation of a 1/40-Scale Model of the McConnell F-101A Airplane (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An investigation has been made in the Langley 20-foot free-spinning tunnel of a 1/40-scale model of the McDonnell F-101A airplane to alleviate the unfavorable spinning characteristics encountered with the airplane. The model results indicate that a suitable strake extended on the inboard side of the nose of the airplane (right side in a right spin) in conjunction with the use of optimum control recovery technique will terminate spin rotation of the airplane. It may be difficult to recover from subsequent high angle-of-attack trimmed flight attitudes even by forward stick movement. The optimum spin-recovery control technique for the McDonnell F-101A is simultaneous full rudder reversal to against the spin and aileron movement to full with the spin (stick full right in a right erect spin) and forward movement of the stick immediately after rotation stops.
    Keywords: Aerodynamics
    Type: NASA-MEMO-3-14-59L , AF-AM-87
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  • 55
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    Low-Speed Wind-Tunnel Investigation to Determine the Aerodynamic Characteristics of a Rectangular Wing Equipped with a Full-Span and an Inboard Half-Span Jet-Augmented Flap Deflected 55 deg (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An investigation to determine the aerodynamic characteristics of a rectangular wing equipped with a full-span and an inboard half-span jet-augmented flap has been made in the Langley 300 MPH 7- by 10-foot tunnel. The wing had an aspect ratio of 8.3 and a thickness-chord ratio of 0.167. A jet of air was blown backward through a small gap, tangentially to the upper surface of a round trailing edge, and was separated from the trailing edge by a very small flap at an angle of 55 deg with respect to the wing-chord plane. The results of the investigation showed that the ratio of total lift to jet-reaction lift for the wing was about 35 percent less for the half-span jet-augmented flap than for the full-span jet-augmented flap. The reduction of the span of the jet-augmented flap from full to half span reduced the maximum value of jet-circulation lift coefficient that could be produced from about 6.8 to a value of about 2.2. The half-span jet- augmented flap gave thrust recoveries considerably poorer than those obtained with the full-span jet-augmented flap. Large nose-down pitching- moment coefficients were produced by the half-span flap, with the greater part of these being the result of the larger jet reactions required to produce a given lift for the half-spin flap compared with that required for the full-span flap.
    Keywords: Aerodynamics
    Type: NASA-MEMO-1-27-59L , L-156
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  • 56
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    Free-Spinning-Tunnel Investigation of a 1/20-Scale Model of the North American T2J-1 Airplane (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An investigation has been made in the Langley 20-foot free-spinning tunnel to determine the erect and inverted spin and recovery characteristics of a 1/20-scale dynamic model of the North American T2J-1 airplane. The model results indicate that the optimum technique for recovery from erect spins of the airplane will be dependent on the distribution of the disposable load. The recommended recovery procedure for spins encountered at the flight design gross weight is simultaneous rudder reversal to against the spin and aileron movement to with the spin. With full wingtip tanks plus rocket installation and full internal fuel load, rudder reversal should be followed by a downward movement of the elevator. For the flight design gross weight plus partially full wingtip tanks, recovery should be attempted by simultaneous rudder reversal to against the spin, movement of ailerons to with the spin, and ejection of the wing-tip tanks. The optimum recovery technique for airplane-inverted spins is rudder reversal to against the spin with the stick maintained longitudinally and laterally neutral.
    Keywords: Aerodynamics
    Type: NASA-TM-SX-245 , L-872 , NASA-AD-3136
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  • 57
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    Lift, Drag, and Pitching Moments of an Arrow Wing Having 80 Degree of Sweepback at Mach Numbers from 2.48 to 3.51 and Reynolds Numbers up to 11.0 Million (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Measurements were made of the lift, drag, and pitching moments on an arrow wing (taper ratio of zero) having an aspect ratio of 1.4 and a leading-edge sweepback of 80 (degrees). The wing was designed to have a subsonic leading-edge and a Clark-Y airfoil with a thickness ratio of 12 percent of the chord perpendicular to the wing leading edge. The wing was tested both with and without the wing tips bent upward in an attempt to alleviate possible flow separation in the vicinity of the wing tips. Small jets of air were used to fix transition near the wing leading edge. Force results are presented for Mach numbers of 2.48, 2.75, 3.04, 3.28, and 3.51 at Reynolds numbers of 3.5 and 9.0 million and for a Mach number of 3.04 at a Reynolds number of 11.0 million. The measured aerodynamic characteristics are compared with those estimated by linear theory. The maximum lift-drag ratio measured was much less than that predicted. This difference is attributed to lack of full leading-edge thrust and to the experimental lift-curve slope being about 20 percent below the theoretical value.
    Keywords: Aerodynamics
    Type: NASA-TM-X-22
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  • 58
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    Effect of Flow Incidence and Reynolds Number on Low-Speed Aerodynamic Characteristics of Several Noncircular Cylinders with Applications to Directional Stability and Spinning (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-26
    Description: The aerodynamic characteristics of several noncircular two-dimensional cylinders with axes normal to the stream at various flow incidences (analogous to angles of attack of a two-dimensional airfoil and obtained by rotating the cylinders about their axes) for a range of Reynolds numbers have been determined from low-speed wind-tunnel tests. The results indicate that these parameters have rather large effects on the drag and side force developed on these cylinders. The side force is especially critical and very often undergoes a change in sign with a change in Reynolds number. Since the flow incidences correspond to combined angles of attack and sideslip in the crossflow plane of three-dimensional bodies, these two-dimensional results appear to have strong implications with regard to directional stability of fuselages at high angles of attack. These implications, along with those associated with the spin-recovery characteristics of aircraft, are briefly discussed.
    Keywords: Aerodynamics
    Type: NASA-TR-R-29
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  • 59
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    Semiempirical Procedure for Estimating Lift and Drag Characteristics of Propeller-Wing-Flap Configurations for Vertical-and Short-Take-Off-and-Landing Airplanes (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: The analysis presented uses the momentum theory as a starting point in developing semiempirical expressions for calculating the effect of propeller thrust and slipstream on the lift and drag characteristics of wing-flap configurations that would be suitable for vertical-take-off-and-landing (VTOL) and short-take-off-and-landing (STOL) airplanes. The method uses power-off forward-speed information and measured slipstream deflection data at zero forward speed to provide a basis for estimating the lift and drag at combined forward speed and power-on conditions. A correlation of slipstream deflection data is also included. The procedure is applicable only in the unstalled flight regime; nevertheless, it should be useful in preliminary design estimates of the performance that may be expected of VTOL and STOL airplanes.
    Keywords: Aerodynamics
    Type: NASA-MEMO-1-16-59L , L-144
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  • 60
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    The Total-Pressure Recovery and Drag Characteristics of Several Auxiliary Inlets at Transonic Speeds (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: Several flush and scoop-type auxiliary inlets have been tested for a range of Mach numbers from 0.55 to 1.3 to determine their transonic total-pressure recovery and drag characteristics. The inlet dimensions were comparable with the thickness of the boundary layer in which they were tested. Results indicate that flush inlets should be inclined at very shallow angles with respect to the surface for optimum total-pressure recovery and drag characteristics. Deep, narrow inlets have lower drag than wide shallow ones at Mach numbers greater than 0.9 but at lower Mach numbers the wider inlets proved superior. Inlets with a shallow approach ramp, 7 deg, and diverging ramp walls which incorporated boundary-layer bypass had lower drag than any other inlet tested for Mach numbers up to 1.2 and had the highest pressure recovery of all of the flush inlets. The scoop inlets, which operated in a higher velocity flow than the flush inlets, had higher drag coefficients. Several of these auxiliary inlets projected multiple, periodic shock waves into the stream when they were operated at low mass-flow ratios.
    Keywords: Aerodynamics
    Type: NASA-MEMO-12-21-58L
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  • 61
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    The Stabilizing Effectiveness of Conical Flares on Bodies with Conical Noses (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: An analysis is presented of published results of force tests on 80 cone-cylinder-flare configurations at Mach numbers of 2.18, 2.81, and 4.04. The contributions, excluding interference effects, of the cone-cylinder bodies to the over-all normal force derivatives have been removed by means of the second-order shock-expansion method, and the normal force derivatives at zero angle of attack due to the flares alone are shown. The results from a wide variety of configurations are correlated by plotting ratios of the normal force derivatives of the flares to the normal force derivatives of cones having the same included angle. Comparisons are made of the experimental normal force results with the normal force derivatives obtained by assuming conical flow over the flares and with those obtained by use of the second-order shock-expansion method. The comparisons show that use of the second-order shock-expansion method is generally the superior of the two, and in most cases gives values of the normal force derivatives of the flares which agree very well with the experimental results. Centers of pressure of the flares are presented and comparisons are made with results obtained from the theories mentioned. In general, the comparisons show that the assumption of conical flow over the flares is comparable to use of the second-order shock-expansion method in determining the centers of pressure, and in many cases both methods give values which agree closely with the experimental results.
    Keywords: Aerodynamics
    Type: NASA-TM-X-30
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  • 62
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    Parasite-Drag Measurements of Five Helicopter Rotor Hubs (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: An investigation has been conducted in the Langley full-scale tunnel to determine the parasite drag of five production-type helicopter rotor hubs. Some simple fairing arrangements were attempted in an effort to reduce the hub drag. The results indicate that, within the range of the tests, changes in angle of attack, hub rotational speed, and forward speed generally had only a small effect on the equivalent flat-plate area representing parasite drag. The drag coefficients of the basic hubs, based on projected hub frontal area, increased with hub area and varied from 0.5 to 0.76 for the hubs tested.
    Keywords: Aerodynamics
    Type: NASA-MEMO-1-31-59L
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  • 63
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    Subsonic Aerodynamic Characteristics of an Airplane Configuration with a 63 deg Sweptback Wing and Twin-Boom Tails (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: A wind-tunnel investigation has been conducted to determine the effects of an unconventional tail arrangement on the subsonic static longitudinal and lateral stability characteristics of a model having a 63 deg sweptback wing of aspect ratio 3.5 and a fuselage. Tail booms, extending rearward from approximately the midsemispan of each wing panel, supported independent tail assemblies well outboard of the usual position at the rear of the fuselage. The horizontal-tail surfaces had the leading edge swept back 45 deg and an aspect ratio of 2.4. The vertical tail surfaces were geometrically similar to one panel of the horizontal tail. For comparative purposes, the wing-body combination was also tested with conventional fuselage-mounted tail surfaces. The wind-tunnel tests were conducted at Mach numbers from 0.25 to 0.95 with a Reynolds number of 2,000,000, at a Mach number of 0.46 with a Reynolds number of 3,500,000, and at a Mach number of 0.20 with a Reynolds number of 7,000,000. The results of the investigation indicate that longitudinal stability existed to considerably higher lift coefficients for the outboard tail configuration than for the configuration with conventional tail. Wing fences were necessary with both configurations for the elimination of sudden changes in longitudinal stability at lift coefficients between 0.3 and 0.5. Sideslip angles up to 15 deg had only small effects upon the pitching-moment characteristics of the outboard tail configuration. There was an increase in the directional stability for the outboard tail configuration at the higher angles of attack as opposed to a decrease for the conventional tail configuration at most of the Mach numbers and Reynolds numbers of this investigation. The dihedral effect increased rapidly with increasing angle of attack for both the outboard and the conventional tail configurations but the increase was greater for the outboard tail configuration. The data indicate that the outboard tail is an effective roll control.
    Keywords: Aerodynamics
    Type: NASA-MEMO-3-3-59A
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  • 64
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    Predicted Static Aeroelastic Effects on Wings with Supersonic Leading Edges and Streamwise Tips (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: A method is presented for calculation of static aeroelastic effects on wings with supersonic leading edges and streamwise tips. Both chord-wise and spanwise deflections are taken into account. Aerodynamic and structural forces are introduced in influence coefficient form; the former are developed from linearized supersonic wing theory and the latter are assumed to be known from load-deflection tests or theory. The predicted effects of flexibility on lateral-control effectiveness, damping in roll, and lift-curve slope are shown for a low-aspect-ratio wing at Mach numbers of 1.25 and 2.60. The control effectiveness is shown for a trailing-edge aileron, a tip aileron, and a slot-deflector spoiler located along the 0.70 chord line. The calculations indicate that the tip aileron is particularly attractive from an aeroelastic standpoint, because the changes in effectiveness with dynamic pressure are small compared to the changes in effectiveness of the trailing-edge aileron and slot-deflector spoiler. The effects of making several simplifying assumptions in the example calculations are shown. The use of a modified strip theory to determine the aerodynamic influence coefficients gave adequate results only for the high Mach number case. Elimination of chordwise bending in the structural influence coefficients exaggerated the aeroelastic effects on rolling-moment and lift coefficients for both Mach numbers.
    Keywords: Aerodynamics
    Type: NASA-MEMO-4-18-59A , A-159
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  • 65
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    Basic Pressure Measurements at Transonic Speeds on a Thin 45 deg Sweptback Highly Tapered Wing With Systematic Spanwise Twist Variations (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: Pressure distributions obtained in the Langley 8-foot transonic pressure tunnel on a thin, highly tapered, twisted, 45 deg sweptback wing in combination with a body are presented. The wing has a linear span-wise 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 1.0 and 0.5 atmosphere, at Mach numbers from 0.800 to 1.200, and at angles of attack from -4 to 12 deg.
    Keywords: Aerodynamics
    Type: NASA-MEMO-12-28-58L
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  • 66
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    Wind-Tunnel Investigation of the Effect of Angle of Attack and Flapping-Hinge Offset on Periodic Bending Moments and Flapping of a Small Rotor (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A two-blade rotor having a diameter of 4 feet and a solidity of 0.037 was tested in the Langley 300-MPH 7- by 10-foot tunnel to obtain information on the effect of certain rotor variables on the blade periodic bending moments and flapping angles during the various stages of transformation between the helicopter and autogiro configuration. Variables studied included collective pitch angle, flapping-hinge offset, rotor angle of attack, and tip-speed ratio. The results show that the blade periodic bending moments generally increase with tip-speed ratio up into the transition region, diminish over a certain range of tip-speed ratio, and increase again at higher tip-speed ratios. Above the transition region, the bending moments increase with collective pitch angle and rotor angle of attack. The absence of a flapping hinge results in a significant amplification of the periodic bending moments, the magnitudes of which increase with tip-speed ratio. When the flapping hinge is used, an increase in flapping-hinge offset results in reduced period bending moments. The aforementioned trends exhibited by the bending moments for changes in the variables are essentially duplicated by the periodic flapping motions. The existence of substantial amounts of blade stall increased both the periodic bending moments and the flapping angles. Harmonic analysis of the bending moments shows significant contributions of the higher harmonics, particularly in the transition region.
    Keywords: Aerodynamics
    Type: NASA-MEMO-3-3-59L
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  • 67
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    An Investigation of a 0.05-Scale Model of the XSM-64A Navaho Missile and Booster (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An investigation has been conducted in the Langley Unitary Plan wind tunnel to determine the aerodynamic loads and the static longitudinal and lateral stability of a 0.05-scale model of the XSM-64A Navaho missile and booster and its various components. Tests were conducted through a Mach number range of 1.77 to 3.51 with a corresponding Reynolds number range of 2.4 x 10(exp 6) to 2.9 x 10(exp 6). Results are presented for an angle-of-attack range of -8 deg to 4 deg for the missile-booster combination and -10 deg to 10 deg for the missile-alone configuration. Tests for both configurations were conducted through an angle-of-sideslip range of -8 deg to 8 deg. Also presented are some effects on the model characteristics of the deflection of various components including canard, tip aileron, vertical stabilizer, speed brakes, and booster pitch and yaw thrust chambers. The various components on which loads were measured include the wing, tip aileron, rudder, booster, booster separating surface, booster fin, and booster yaw and pitch thrust chambers. These data are presented without analysis.
    Keywords: Aerodynamics
    Type: NASA-MEMO-5-30-59L , L-348
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    Lift-Drag Ratios for an Arrow Wing With Bodies at Mach Number 3 (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Force and moment characteristics, including lift-drag ratios, have been measured for bodies of circular and elliptic cross section alone and combined with a warped arrow wing. The test Mach number was 2.94, and the Reynolds number was 3.5 x 10(exp 6) (based on wing mean aerodynamic chord). The experimental results show that for equal volume the use of an elliptical body can result in a noticeably higher maximum lift-drag ratio than that obtained through use of a circular body. Methods for estimating the aerodynamic characteristics have been assessed by comparing computed with experimental results. Because of good agreement of the predictions with experiment, maximum lift-drag ratios have been computed for the arrow wing in combination with bodies of various sizes. These calculations have shown that, for an efficient wing-body combination, little loss in maximum lift-drag ratio results from considerable extension of afterbody length. For example, for a wing-body configuration having a maximum lift-drag ratio of about 7.1, a loss in maximum lift-drag ratio of less than 0.2 results from a 40-percent increase in body volume by extension of afterbody length. It also appears that with body length fixed, maximum lift-drag ratio decreases almost linearly with increase in body diameter. For a wing- body combination employing a body of circular cross section, a decrease in maximum lift-drag ratio from about 9.1 for zero body diameter to about 4.6 for a body diameter of 13.5 percent of the body length was computed.
    Keywords: Aerodynamics
    Type: NASA-MEMO-4-27-59A
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  • 69
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    A Summary of Flight-Determined Transonic Lift and Drag Characteristics of Several Research Airplane Configurations (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Flight-determined lift and drag data from transonic flights of seven research airplane configurations of widely varying characteristics are presented and compared with wind-tunnel and rocket-model data. The airplanes are the X-5 (590 wing sweep), XF-92A, YF-102 with cambered wing, YF-102 with symmetrical wing, D-558-ii, X-3, and X-LE. The effects of some of the basic configuration differences on the lift and drag characteristics are demonstrated. As indicated by transonic similarity laws, most of the configurations demonstrate a relationship between the transonic increase in zero-lift drag and the maximum cross-sectional area. No such relationship was found between the drag-rise Mach number and its normally related parameters. A comparison of flight and wind-tunnel data shows a generally reasonable agreement, but Reynolds number differences can cause considerable variations in the drag levels of the flight and wind-tunnel tests. Maximum lift-drag ratios vary widely in the subsonic region as would be expected from differences in aspect ratio and wing thickness ratio; however, the variations diminish as the Mach number is increased through the transonic region. The attainment of maximum lift-drag ratio in level flight by several of the airplanes was limited by engine performance, stability characteristics, and buffet boundaries.
    Keywords: Aerodynamics
    Type: NASA-MEMO-3-3-59H
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  • 70
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    Several Methods for Reducing the Drag of Transport Configurations at High Subsonic Speeds (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Results of investigations of several promising methods for alleviating the drag rise of transport configurations at high subsonic speeds are reviewed briefly. The methods include a wing leading-edge extension, a fuselage addition, and additions on the wing. Also, results are presented for a complete, improved transport configuration which incorporates the fuselage and wing additions and show that the improved configuration could have considerably higher cruise speeds than do current designs.
    Keywords: Aerodynamics
    Type: NASA-MEMO-2-25-59L
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  • 71
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    A Wind-Tunnel Investigation of Rotor Behavior Under Extreme Operating Conditions with a Description of Blade Oscillations Attributed to Pitch-Lag Coupling (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A wind-tunnel investigation was made to study the behavior of a model helicopter rotor under extreme operating conditions. A 1/8-scale model of the front rotor of a tandem helicopter was built and tested to obtaining blade motion and rotor aerodynamic characteristics for conditions that could be encountered in high-speed pullout maneuvers. The data are presented without analysis. A description is given in an appendix of blade oscillations that were experienced during the course of the investigation and of the part that blade pitch-lag coupling played in contributing to the oscillatory condition.
    Keywords: Aerodynamics
    Type: NASA-MEMO-1-7-59L
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  • 72
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    Aluminiumchlorid-Katalysen, XXV. Eine neue Umlagerung der Phenylalkane (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 10-17 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Bei gelinder Erwärmung der Phenylalkane (R · C6H5; R ≧ C3) mit durch Wasser aktiviertem Aluminiumchlorid findet eine Umlagerung der Seitenkette statt. Es entstehen: aus Isopropylbenzol n-Propylbenzol, aus sek.-Butylbenzol Isobutylbenzol, aus tert.-Amylbenzol 2-Methyl-3-phenyl-butan und aus letzterem, bei längerer Einwirkung des Katalysators, Neopentylbenzol.  -  Die stabilen Produkte dieser Umlagerungen enthalten an primäre Kohlenstoffatome gebundene Phenylgruppen, im Gegensatz zu den Produkten der Friedel-Crafts-Alkylierungsreaktion, in welchen eine tertiäre bzw. eine sekundäre Stellung der Phenylgruppen bevorzugt wird.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920103
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  • 73
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    Über das Naphthyl-(2)-tellurjodid (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 40-46 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Die Darstellung des Naphthyl-(2)-tellurjodids wird beschrieben und der Beweis für seine Konstitution erbracht.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920107
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  • 74
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    Bathochromie bei Anilen durch Komplexbildung und Adsorption (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 36-39 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Lewis-Säuren sowie Silicagel bewirken bei Anilen bestimmter Konstitution eine Polarisierung in Richtung der dipolaren Grenzformel. Dieser Effekt ist mit einer bedeutenden Farbvertiefung verbunden. Es werden die konstitutionellen Voraussetzungen der Erscheinung erörtert.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920106
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  • 75
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    Nucleophile aromatische Substitutionen, IX. Ringschlußreaktionen zu N-Methyl-indolin und N-Methyl-tetrahydrochinolin und ihr Chemismus (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 203-213 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Die optimalen Reaktionsbedingungen für die Cyclisierung der N-Lithium-Derivate der sekundären Amine I, II, IX und X mit m- bzw. o-ständigem Chlor zu den in der Überschrift genannten Heterocyclen werden ermittelt. Im Gegensatz zur stets intramolekularen Basen-Addition an die intermediären Arine sind für die Freisetzung der Arine, also für den geschwindigkeitsbestimmenden Teilschritt, inter- und intramolekulare Chemismen zu diskutieren. Die experimentellen Daten weisen für N-Lithium-I und -II auf eine Hilfestellung zugesetzten, komplex-gebundenen Lithiumdiäthylamids bei der Arin-Bildung.
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    URL: http://dx.doi.org/10.1002/cber.19590920123
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  • 76
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    Formamid-Reaktionen, XIII. Imidazole aus α-Diketonen (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 338-343 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Benzil und substituierte Benzile lassen sich mit Formamid und Aldehyden, ebenso mit Formamid in Gegenwart von Ameisensäure und reduzierenden Schwefelverbindungen, in substituierte Imidazole überführen.
    Additional Material: 3 Tab.
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    URL: http://dx.doi.org/10.1002/cber.19590920214
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  • 77
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    Über die Abhängigkeit der Reaktionsfähigkeit funktioneller Gruppen in Paraffinkohlenwasserstoffen von ihrer Stellung in der Molekel, X. Über die Zusammensetzung der Produkte der direkten Nitrierung von n-Hexatriacontan (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 347-362 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Die bei der direkten Nitrierung von n-Hexatriacontan nach drei verschiedenen Verfahren entstehenden Mononitrohexatriacontane sind identische Gemische sämtlicher theoretisch möglicher Stellungsisomeren. Ausschließliche Bildung von 2-Nitro-hexatriacontan findet entgegen den Befunden von NAMETKIN und Mitarbb. nicht statt.
    Additional Material: 15 Tab.
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    URL: http://dx.doi.org/10.1002/cber.19590920216
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  • 78
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    Über Triterpene, V. Zur Kenntnis der Crataegolsäure und über zwei neue Triterpencarbonsäuren aus Crataegus oxyacantha L. (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 320-328 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Aus den Blättern von Crataegus oxyacantha L. wurden zwei neue Triterpencarbonsäuren, Acantholsäure, C30H48O5, und Neotaegolsäure, C30H48O4, isoliert. Crataegolsäure wurde als ein Derivat der β-Amyrin-Reihe erkannt. Ein papierchromatographisches Verfahren zur Trennung von Triterpensäuregemischen wird beschrieben; das angegebene Lösungsmittelgemisch ließ sich erfolgreich auch für eine Verteilungschromatographie an der Cellulosesäure verwenden.
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    URL: http://dx.doi.org/10.1002/cber.19590920212
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  • 79
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    Über Sulfenyl-Verbindungen organischer Basen, II. Zur Darstellung von Sulfenamiden, Sulfenylamidinen und Sulfinsäureamid-imidenHerrn Dr. F. G. Mann in dankbarer Erinnerung an einen Aufenthalt in Cambridge. (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 1-9 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Chloramin T läßt sich mit Sulfenamiden und besonders glatt mit Sulfenyl-amidinen und verwandten Verbindungen kondensieren. Die Reaktionsprodukte werden als Derivate von Sulfinsäureamid-imiden aufgefaßt.
    Additional Material: 6 Tab.
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    URL: http://dx.doi.org/10.1002/cber.19590920102
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  • 80
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    Beispiele zur King-Reaktion (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 22-36 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: CH-Acide Verbindungen wie Chinaldin, Lepidin, 9-Methyl-acridin, die drei Acetylpyridine, Aceton, 2.4-Dinitro-toluol und ein Phenacyl-thioäther werden nach KING mit Jod/Pyridin, teilweise auch mit Jod/Thioharnstoff umgesetzt; mit einigen der so erhaltenen Salze werden Ringschluß-Reaktionen vollzogen. Mit Brom/Pyridin gelingt die King-Reaktion nur in wenigen Fällen. Mittels der Umsetzung mit p-Nitroso-dimethylanilin sowie der nach King wird gezeigt und begründet, daß die Reaktivität der Methylgruppe in der Reihe α- und γ-Picolin → Chinolin/Lepidin → 9-Methyl-acridin zu-, daß sie aber in deren Halogenalkylaten in der gleichen Reihenfolge abnimmt. Es werden Benzologe eines von BERSON und COHEN bei ihrer Pyridonsynthese angenommenen Zwischenproduktes in kristallisierter Form gefaßt.
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    URL: http://dx.doi.org/10.1002/cber.19590920105
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  • 81
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    Benzazole, IX. Über das chemische Verhalten der Nitro-benzoxazolone (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 407-414 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Die Nitro-benzoxazolone reagieren mit Diazomethan, Carbonsäure- und Sulfonsäurechloriden nur in der Lactamform. Durch die Nitrogruppe in verschiedenen Stellungen des Benzolringes werden die NH-Acidität und die Spaltbarkeit des Heteroringes beeinflußt. Im 4-Nitro-benzoxazolon ist die 3-Stellung durch die Nitrogruppe sterisch behindert.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920224
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  • 82
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    Nucleophile aromatische Substitutionen, XI. Weitere Ringschlußreaktionen üer Arine (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 429-441 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Das am N-Methyl-indolin und N-Methyl-tetrahydrochinolin ausgearbeitete Syntheseprinzip läßt sich auf andere Heterocyclen übertragen. Chlor- oder Brombenzol-Derivate mit o- oder m-ständiger basisch-substituierter Seitenkette werden inter- oder intramolekular in das Arin übergeführt; die anschließende intramolekulare Basen-Addition liefert unter Ringschluß tertiäre Amine: N-Phenyl-tetrahydrochinolin, N-Methyl-phenmorpholin, N.N′-Dimethyl-tetrahydrochinoxalin und N.N′-Dimethyl-indazolin. Auch die erfolgreiche Schließung eines 8 gliedrigen Ringes wird beschrieben.
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    URL: http://dx.doi.org/10.1002/cber.19590920227
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  • 83
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    Ringschlußreaktion des 4-Oxo-1-propionyl-pentancarbonsäure-(1)-äthylesters (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 468-471 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Die unter verschiedenen Bedingungen durchgeführte Cyclisierung des im Titel genannten Esters wird beschrieben.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920230
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  • 84
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    Über Sauerstoffradikale, XI. Das 4-Phenyl-2.6-di-tert.-butyl-phenoxyl-(1) (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 474-482 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Darstellung sowie physikalische und chemische Eigenschaften eines neuen stabilen Aroxyls. des 4-Phenyl-2.6-di-tert.-butyl-phenoxyls-(1) werden beschrieben. Die 4-ständige Phenylgruppe dieses Aroxyls bedingt, verglichen mit 2.4.6-Tri-tert.-butyl-phenoxyl-(1) und ähnlichen Verbindungen, einige Besonderheiten im reaktiven Verhalten dieses freien Radikals.
    Additional Material: 2 Ill.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920232
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  • 85
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    all-cis-3.4.5-Trihydroxy-cyclohexan-carbonsäure-(1) (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 213-218 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Die durch Hydrierung von Dihydrogallussäure erhältliche 3.4.5-Trihydroxy-cyclohexan-carbonsäure-(1) vom Schmp. 203〉 erweist sich durch Lactonbildung und Abbau zum γ-Pyrogallit als all-cis-Isomeres.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920124
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  • 86
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    Synthese der L-Apiose aus (+)-Weinsäure (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 535-540 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Die Kohlenstoffkette der (+)-Weinsäure wird an dem einen Ende über das Diazoketon verlängert. Die neue Ketogruppe liefert durch nochmalige Reaktion mit Diazomethan eine Kettenverzweigung, worauf durch Abbau nach RUFF an dem anderen Ende der Weinsäure L-Apiose entsteht.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920305
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  • 87
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    Über bi- und polycyclische Azulene, XXXVIII. Synthesen von Azulencarbonsäure-nitrilen (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 606-615 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: 1- bzw. 3-Cyan-azulene lassen sich aus Azulenaldoximen durch Wasserabspaltung mit Acetanhydrid darstellen. Verschiedentlich tritt daneben eine Umlagerung des Oxims zur Acetaminoverbindung ein. Bei Umsetzung von Azulenen mit Bromcyan entstehen Nitrile durch direkte Substitution. Die Nitrile sind sehr reaktionsträge Verbindungen. Ihre Inaktivität wird durch eine polare Struktur erklärt. Die spektralen Verschiebungen gegenüber den zugrundeliegenden Azulenen und die Halbstufenpotentiale werden angegeben.
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    URL: http://dx.doi.org/10.1002/cber.19590920313
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  • 88
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    Zur Kenntnis hydrierter Pyrido[1.2-c]pyrimidine (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 637-644 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Die Darstellung des 3-Aza-chinolizidins und des von diesem sich ableitenden tricyclischen Systems mit zwei Brückenstickstoffatomen, des Perhydro-dipyrido-[1.2-a, 1′.2′-c]pyrimidins (4a.8a-Diaza-perhydrophenanthrens) wird beschrieben.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920316
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  • 89
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    Reaktionen mit Nitrosodisulfonat, XVII. Zum Reaktionsverlauf der Bildung von Chinonen aus einwertigen Phenolen (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 667-673 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Am Beispiel der Bildung von 3-Carbäthoxy-2-methyl-6.7-benzo-indolchinon-(4.5) (VI) aus 5-Hydroxy-3-carbäthoxy-2-methyl-6.7-benzo-indol (II) wird gezeigt, daß die Zwischenstufe eines Substitutionsprodukts (III) durchlaufen wird, das als solches isoliert und zum erwarteten Chinon zerlegt werden kann.
    Additional Material: 3 Ill.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920320
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  • 90
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    Die Bestimmung von Δ1.9- und Δ9.10-Octalin nebeneinander (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 693-699 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Die UV-Absorption der Jodkomplexe von Cycloolefinen wird gemessen und nach der von LONG und NEUZIL für Olefine entwickelten Methode zur Analyse von Δ1.9- und Δ9.10-Octalingemischen verwendet.
    Additional Material: 3 Ill.
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    URL: http://dx.doi.org/10.1002/cber.19590920322
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    Alkylhydroperoxyde aus Alkylhalogeniden, II (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 699-704 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Der Ersatz von organisch gebundenem Chlor und Brom durch den Hydroperoxy-Rest des Wasserstoffperoxyds ist immer dann möglich, wenn das Alkylhalogenid genügend reaktionsfähig im Sinne des SNl-Schemas ist. Komplexbildner können die Reaktion erleichtern.
    Additional Material: 1 Tab.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920323
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  • 92
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    Über Verbindungen mit Urotropin-Struktur, XI. Zur Kenntnis der Bildung des 2.4.9-Trioxa-adamantan-Ringsystems aus Triketonen (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 732-735 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Ausgehend von Methantriessigsäure wurde ein Tris diazoketon hergestellt. Die hieraus mit Halogenwasserstoffsäuren primär gebildeten Halogenketone erleiden eine „intramolekulare Trimerisierung“ der Carbonylgruppen unter Bildung von 1.3.5-Tris-chlormethyl- und 1.3.5-Tris-brommethyl-2.4.9-trioxa-adamantan. Dagegen zeigt das durch Friedel-Crafts-Acylierung des Säurechlorides der Methantriessigsäure mit Benzol erhaltene Triketon der gleichen Struktur keine Tendenz zur Bildung des 2.4.9-Trioxa-adamantan-Ringsystems. Die Trimerisierung der Carbonylgruppen unterbleibt hier wegen der Mesomeriebeteiligung der Carbonylgruppen am aromatischen System.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920327
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  • 93
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    Konstitution von Reticulin, einem neuen Alkaloid aus Anona reticulata Linn. (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 776-779 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Aus Anona reticulata Linn. wurde ein neues Alkaloid, Reticulin, isoliert. Reticulin ist ein 2-Methyl-7-hydroxy-6-methoxy-1-[3-hydroxy-4-methoxy-benzyl]-1.2.3.4-tetrahydro-isochinolin, wie durch oxydativen Abbau gezeigt werden konnte.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920404
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  • 94
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    Über in 2-Stellung äthinylsubstituierte N-Heterocyclen. 2-Äthinyl-benzoxazol (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 791-793 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: α.β-Dibrom-β-[benzthiazolyl-(2)]-propionsäure spaltet unter der Einwirkung von Alkalien keinen Bromwasserstoff, sondern Brom ab. N-Methyl-chinolon, N-Methyl-chinoliniumjodid und N.N′-Dimethyl-benzimidazolon reagieren mit den Monometallverbindungen des Acetylens nicht. Dagegen gelang die Umsetzung von 2-Chlor-benzoxazol mit Acetylen-natrium zum 2-Äthinylbenzoxazol.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920406
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  • 95
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    Untersuchungen zur Komplexchemie der Brenztraubensäure, III. Struktur und Eigenschaften von Brenztraubensäure-Komplexen (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 793-806 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Brenztraubensäure besitzt in komplexer Bindung eine Reihe von Eigenschaften, die von der freien Säure nicht bekannt sind. So nimmt das Substitutionsvermögen gegenüber Halogen namentlich bei Gegenwart von guten Komplexbildnern beträchtlich zu. Im Falle der Jodsubstitution konnte die Abhängigkeit der Substitutionsgeschwindigkeit von der Komplexstabilität bewiesen und damit ein Einblick in die Struktur der als Liganden fixierten Brenztraubensäuremoleküle gewonnen werden.
    Additional Material: 5 Ill.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920407
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  • 96
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    Über die Schwefelnitride S15N2 und S16N2 (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 855-857 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Durch Umsetzung von SCl2 bzw. S2Cl2 mit S7NH konnten die Schwefelnitride S15N2 und S16N2 als kristallisierte Substanzen gewonnen werden. Von diesen ist besonders S15N2 resonanzstabilisiert.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920413
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  • 97
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    Thiazolochinoline, IV. Über die Chlorierung der 2-Mercaptoderivate des Thiazolo[4.5-b]-und -[5.4-b]chinolins (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 869-871 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Die Chlorierung des 2-Mercapto-thiazolo[4.5-b]- und -[5.4-b]chinolins in Essigsäure führt zu den entsprechenden 2-Chlorderivaten. Die Chloratome erwiesen sich als sehr labil und lassen sich leicht durch verschiedene nucleophile Substituenten ersetzen.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920416
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  • 98
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    Zur Synthese 3-substituierter 4-Hydroxy-pyrone-(2), II. Substitutionsreaktionen am Triacetsäurelacton (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 982-986 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Am Triacetsäurelacton werden Substitutionsreaktionen in 3-Stellung durchgeführt. Bei der Einwirkung von Benzoldiazoniumchlorid entsteht das 3-Phenyl-Derivat; Aryl-bzw. Aryl-alkyl-carbinole und Aralkylhalogenide reagieren beim Zusammenschmelzen mit Triacetsäurelacton zu den entsprechenden 3-Substitutionsprodukten. Die C-3-Acylierung gelingt in Gegenwart von wasser-freiem Aluminiumtrichlorid.
    Additional Material: 3 Tab.
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    URL: http://dx.doi.org/10.1002/cber.19590920433
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  • 99
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    Synthese von Furan-Derivaten, XVIII. Notiz über die Synthese von 4-Oxo-octahydro-dibenzofuran (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 998-999 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920436
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  • 100
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    2-[α-Amino-alkyl]-benzimidazole (1959)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 92 (1959), S. 977-981 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Die 2-[α-Amino-alkyl]-benzimidazole Ia bis IVa lassen sich mit guten Ausbeuten durch Kondensation von o-Phenylendiamin mit den N-p-Tosyl- oder N-Carbobenzoxy-Derivaten von Amino-iminoäther-hydrochloriden und nach-folgende Abspaltung der Tosyl- bzw. der Carbobenzoxy-Gruppe darstellen. Bei Verbindungen mit verzweigtem Alkyl-Rest tritt als Konkurrenz-Reaktion zum Benzimidazol-Ringschluß die Zersetzung des Iminoäther-hydrochlorids zum Säureamid ein.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19590920432
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