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  • Inorganic Chemistry  (677)
  • Industrial Chemistry  (233)
  • Cell & Developmental Biology  (107)
  • Aerodynamics  (53)
  • Aircraft Design, Testing and Performance  (30)
  • 1960-1964  (1,100)
  • 1950-1954
  • 1930-1934
  • 1920-1924
  • 1915-1919
  • 1960  (1,100)
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  • 1960-1964  (1,100)
  • 1950-1954
  • 1930-1934
  • 1920-1924
  • 1915-1919
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  • 1
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    A Method of Reducing Aeroelastic Effects of Highly Swept Wings (1960)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: The greatest efficiency for a lifting surface at supersonic speeds, according to the theoretical considerations of reference 1, can be attained if the leading edge is swept well behind the Mach cone and the highest aspect ratio which is structurally possible is employed. Such a wing, designed for a Mach number of 3.0, would have 80 deg. of sweepback. Aeroelastic effects have 〈 been shown 3 to be considerable for a wing with 60deg of sweepback and designed for a Mach number of 2.0. The wing shown was found theoretically to have considerable loss in maximum lift-drag ratio attributable to aeroelasticity. This wing has 12-per cent-thick Clark-Y airfoils normal to the wing leading edge. If it were of solid aluminum and flying at a dynamic pressure of 2,400 lbs./sq.ft. (flexibility parameter qb(exp. 4) /El(0) = 7.8), analysis indicates that the wing would deflect so as to reduce the maximum lift-drag ratio about 30 per cent.
    Keywords: Aerodynamics
    Type: Journal of the Aerospace Sciences; Volume 27; No. 8; 634-635
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  • 2
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    Measurements of Skin Friction of the Compressible Turbulent Boundary Layer on a Cone with Foreign Gas Injection (1960)
    In:  Other Sources
    Details
    Publication Date: 2018-06-05
    Description: Measurements of average skin friction of the turbulent boundary layer have been made on a 15deg total included angle cone with foreign gas injection. Measurements of total skin-friction drag were obtained at free-stream Mach numbers of 0.3, 0.7, 3.5, and 4.7 and within a Reynolds number range from 0.9 x 10(exp 6) to 5.9 x 10(exp 6) with injection of helium, air, and Freon-12 (CCl2F2) through the porous wall. Substantial reductions in skin friction are realized with gas injection within the range of Mach numbers of this test. The relative reduction in skin friction is in accordance with theory-that is, the light gases are most effective when compared on a mass flow basis. There is a marked effect of Mach number on the reduction of average skin friction; this effect is not shown by the available theories. Limited transition location measurements indicate that the boundary layer does not fully trip with gas injection but that the transition point approaches a forward limit with increasing injection. The variation of the skin-friction coefficient, for the lower injection rates with natural transition, is dependent on the flow Reynolds number and type of injected gas; and at the high injection rates the skin friction is in fair agreement with the turbulent boundary layer results.
    Keywords: Aerodynamics
    Type: Journal of Aerospace Sciences; Volume 27; No. 5; 321-333
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  • 3
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    Effects of Fixing Boundary-Layer Transition for a Swept- and a Triangular-Wing and Body Combination at Mach Numbers from 0.60 to 1.40 (1960)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The aerodynamic effects of fixing boundary-layer transition for a swept- and a triangular-wing configuration have been determined from tests of two small-scale wing-body models. The wings had an aspect ratio of 2.99 and 3-percent-thick biconvex sections. Lift, pitching-moment, and drag data were obtained at Mach numbers ranging from 0.60 to 1.40 for angles of attack between -2 deg and about 15 deg. The Reynolds number of the tests was generally 1.5 million; however, minimum drag measurements were made for both models over a range of Reynolds numbers from 1.0 million to about 3.0 or 4.0 million.
    Keywords: Aerodynamics
    Type: NASA-TN-D-312
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  • 4
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    Experimental Determination of the Effects of Frequency and Amplitude of Oscillation on the Roll-Stability Derivatives for a 60 Degree Delta-Wing Airplane Model (1960)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A 60' delta-wing airplane model was oscillated in roll for several frequencies and amplitudes of oscillation to determine the effects of the oscillatory motion on the roll-stability derivatives for the model. The derivatives were measured at a Reynolds number of 1,600,000 for the wing alone, the wing-fuselage combination, and the complete model which included a triangular-plan-form vertical tail. Both rolling and yawing moments due to rolling velocity exhibited large frequency effects for angles of attack higher than 16 degrees. Variations in these derivatives were measured for the lowest frequencies of oscillation; as the frequency increased, the derivatives because more nearly linear with angle of attack. Both velocity derivatives were considerably different at high angles of attack from the corresponding derivatives measured by the steady-state rolling-flow technique. Rolling and yawing moments due to rolling acceleration were measured and similarly found to be highly dependent on frequency at high angles of attack. Some period and time-to-damp computations, which were made to reveal the significance of the acceleration derivatives, indicated that inclusion of the measured derivatives in the equations of motion lengthened the period of the lateral oscillation by 10 percent for a typical delta-wing airplane and increased the time to damp to one-half amplitude by 50 percent.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-232
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  • 5
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    Ground Effect for Lifting Rotors in Forward Flight (1960)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A theoretical analysis indicates that, for rotors, ground effect decreases rapidly with increases in either height above the ground or forward speed. The decrease with height above the ground in forward flights is greater than that in hovering. The major part of the decrease in ground effect with forward speed occurs at speeds less than 1.5 times the hovering mean induced velocity. Consequently, the total induced velocity at the rotor center increases rather than decreases when a helicopter gathers speed at low height above the ground.
    Keywords: Aerodynamics
    Type: NASA-TN-D-234
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  • 6
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    Effect of Ground Proximity on Aerodynamic Characteristics of Two Horizontal-Attitude Jet Vertical-Take-Off-and-Landing Airplane Models (1960)
    In:  CASI
    Details
    Publication Date: 2019-06-25
    Description: An investigation has been made to study the effect of ground proximity on the aerodynamic characteristics of two jet vertical-take-off-and-landing airplane models in which the fuselage remains in a horizontal attitude for the take-off and landing. The first model (called the tilt-wing model) had a tilting wing-engine assembly which was set at 90 deg incidence for the take-off and landing. The second model, called the deflected-jet model) had a cascade of retractable turning vanes to deflect the exhaust of the horizontally mounted jet engines downward for vertical take-off and landing while the entire model remained in a horizontal attitude. With the models at various heights above the ground in the take-off and landing configuration, the lift, drag, and pitching moment were measured and tuft surveys were made to determine the flow field caused by the jet exhaust. The tilt-wing model experienced a loss of lift of less than 3 percent near the ground. The deflected-jet model, however, suffered losses in lift as high as 45 percent near the ground because of a low pressure region under the model caused by the entrainment of air by the jet exhaust as it spread out along the ground. This loss in lift for the deflected-jet configuration could probably be reduced to less than 5 percent by the use of a longer landing gear and a high wing location.
    Keywords: Aerodynamics
    Type: NASA-TN-D-419 , L-1059
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  • 7
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    Measured Base Pressures on Several Twin Rocket-Nozzle Configurations at Mach Numbers of 0.6 to 1.4 with Effects Due to Nozzle Canting and Stabilizing Fins (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: An experimental investigation has been conducted at Mach numbers of 0.6 to 1.4 to determine the base pressures on several cylindrical afterbody configurations having two propulsive nozzles and to determine the effect on base pressure of stabilizing fins and the canting outward of the propulsive nozzles. Nozzle design Mach numbers of 2.0 and 3.43 were employed in this investigation and cold air at total pressures up to 120 times the free-stream static pressure was used to simulate nozzle flow. The results show that canting the nozzles outward 11 degrees was effective in increasing base pressures at supersonic speeds and that stabilizing fins caused a decrease in base pressure. The magnitudes of base pressure coefficients obtained in this investigation were consistent with those obtained on similar configurations in previous jet-effect investigations.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA/TN-D-544 , L-861
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  • 8
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    Hovering Measurements for Twin-Rotor Configurations with and Without Overlap (1960)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: Results of an investigation in the Langley full-scale tunnel of the hovering performance of large-scale twin-rotor-helicopter models are presented. Measurements of thrust, torque, and rotor flapping are given for overlapped (approximately 76 percent of blade radius) and nonoverlapped configurations and for two different rotor solidities. The measured performance is compared with single-rotor measurements and with available rotor theory. These tests show that the hovering performance of a single rotor and of two rotors without overlap or vertical offset are the same and hence may be calculated by single-rotor theory. These tests in conjunction with results of previous coaxial-rotor tests show that the performance of highly overlapped rotors can be reasonably predicted by available rotor theory.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-534 , L-95399
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  • 9
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    Tests of Vortex-Ring Parachute at Supersonic Speed in the Langley Unitary Plan Wind Tunnel (1960)
    In:  Other Sources
    Details
    Publication Date: 2019-07-12
    Description: For the test, the 12-inch-diameter "Vortex-Ring" parachute was towed behind a conical-nosed cylindrical body 2.25 inches in diameter. The tow-cable length was 24 inches, and was attached to the cylindrical body through a large swivel and to the parachute through a smaller swivel. The attachment between the large swivel an the cylindrical body failed after about 1 minute's operation. Mach number was approximately 2.2, dynamic pressure was approximately 150 pounds per square foot, and camera speed was approximately 3000 frames per second.
    Keywords: Aerodynamics
    Type: L-560
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  • 10
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    Measurement of the Maximum Altitude Attained by the X-15 Airplane Powered with Interim Rocket Engines (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: On August 12, 1960, an X-15 flight was made to achieve essentially the maximum altitude expected to be possible with the interim rocket engines. N l y corrected altitude measurements showed that the maxhum geometric altitude was 136,500 feet k600 and the maximum pressure altitude, referred to the tables of the 0. S . Extension to the ICAO Standard Atmosphere, was indicated to be 133,900 feet.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-623 , H-206
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  • 11
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    Aerodynamic Characteristics at Low Speed of a Reentry Configuration Having Rigid Retractable Conical Lifting Surfaces (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: An exploratory investigation has been made in the Langley 300 MPH 7 by 10 foot tunnel to study the low-speed static longitudinal and lateral stability characteristics of a reentry configuration having rigid retractable conical lifting surfaces that unfolded from the surface of a conical fuselage. The model also had curved tail surfaces that unfolded from a cylindrical aft section attached to the cone. Longitudinal tests were made through an angle-of-attack range from -4 deg to 90 deg and limited lateral tests were made through an angle-of-sideslip range from -12 deg to 32 deg at an angle of attack of 0 deg. The tail surface provided longitudinal trim to maximum lift and beyond and up to an angle of attack of 51 deg for a center-of-moment location of 42.9 percent mean aerodynamic chord. For this center-of-moment position the model had a static margin of 12 percent mean aerodynamic chord at the lower lift coefficients and was longitudinally stable up to a lift coefficient between 1.0 and 1.2. Neutral stability occurred from lift coefficient of 1.0 up to near maximum lift coefficient. The maximum value of trimmed lift-drag ratio was 4.85 at a lift coefficient of approximately 0.3 and a trimmed angle of attack of approximately 10 deg. The configuration was directionally stable throughout the test angle of sideslip range for an angle of attack of 0 deg.
    Keywords: Aerodynamics
    Type: NASA-TN-D-622 , L-1180
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  • 12
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    Aerodynamic Characteristics of a Model of an Inflatable-Sphere Launching Vehicle under Simulated Conditions of Mach Number and Altitude (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: An investigation has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel to determine the aerodynamic characteristics in pitch of a two-stage-rocket model configuration which simulated the last two stages of the launching vehicle for an inflatable sphere. Tests were made through an angle-of-attack range from -6 deg to 18 deg at dynamic pressures of 102 and 255 pounds per square foot with corresponding Mach numbers of 1.89 and 1.98 for the model both with and without a bumper arrangement designed to protect the rocket casing from the outer shell of the vehicle.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-640 , L-911
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  • 13
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    Thrust Characteristics of Multiple Lifting Jets in Ground Proximity (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: An investigation has been made to determine the thrust characteristics within ground proximity of a series of models which might represent vertical take-off-and-landing (VTOL) aircraft with multiple exit jet engines exhausting vertically downward beneath a lifting surface. Variations in simulated engine configurations were provided by a series of nozzle insert plugs in which the number of jet exits, located symmetrically on a fixed circle, was varied, or the diameter of the circle was varied for a given number of jet exits. represent lifting surfaces, and high-pressure air was used to simulate jet-engine exhaust. Plywood plates were used to The results of the investigation showed that increasing the number of exits, such that an annular jet configuration was approached, provided more favorable thrust characteristics within ground proximity than any other variation in the geometry of these multiple jets. Tests of a configuration with two nozzles approximating a fan-in-wing VTOL aircraft with fans located at different spanwise locations indicated that the augmentation in thrust within ground proximity was greater for the arrangement with the more inboard location of the nozzles.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-513 , L-868
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  • 14
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    An Investigation of Landing-Contact Conditions for a Large Turbojet Transport During Routine Daylight Operations (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: An investigation has been made by the NASA to obtain statistical measurements of landing-contact conditions for a large turbojet transport in commercial airline operations. The investigation was conducted at the Los Angeles International Airport in Los Angeles, California. Measurements were taken photographically during routine daylight operations. The quantities determined were vertical velocity, horizontal velocity, rolling velocity, bank angle, and distance from runway threshold, just prior to ground contact. The results indicated that the mean vertical velocity for the turbojet-transport landings was 1.62 feet per second and that 1 landing out of 100 would be expected to equal or exceed about 4.0 feet per second. The mean airspeed at contact was 132.0 knots, with 1 landing in 100 likely to equal or exceed about 153.0 knots. The mean rolling velocity was about 1.6 deg per second. One lending in 100 would probably equal or exceed a rolling velocity of about 4.0 deg. per second in the direction of the first wheel to touch. The mean bank angle for the turbojet transports was 1.04 deg, and right and left angles of bank were about evenly divided. One lending in 100 would be likely to equal or exceed a bank angle of about 3.5 deg. The mean value of distance to touchdown from the runway threshold was 1,560 feet. One lending in 100 would be expected to touchdown at or beyond about 2,700 feet from the runway threshold. The mean values for vertical velocity, airspeed, and distance t o touch-down for the turbojet transports were somewhat higher than those found previously for piston-engine transports. No significant differences were found for values of rolling velocity and bank angle.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-527 , L-1009
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  • 15
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    An Analysis of the Flapwise Bending Frequencies and Mode Shapes of Rotor Blades Having Two Flapping Hinges to Reduce Vibration Levels (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: This paper presents the analysis of the flapwise natural bending frequencies and mode shapes of rotor blades with two flapping hinges located at arbitrary blade radii. The equations of motion are derived for a blade of variable mass and stiffness distribution. Solutions to the equations (natural frequencies and mode shapes) are presented for a typical blade of constant cross section having a wide range of hinge locations. The results show that the natural frequencies of the blades can be changed appreciably by varying the locations of the blade hinges, and that with two properly located flapping hinges, blade designs are possible which eliminate or greatly reduce conditions of resonance between the blade natural frequencies and the frequencies of the harmonic air loads. The results also show that ratios of natural frequency to rotor speed below a value of 6.0 are essentially constant for variations in rotor speed consistent with helicopter and VTOL applications.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-633
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  • 16
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    A Technique for Determining Relaxation Times by Free-Flight Tests of Low-Fineness-Ratio Cones; with Experimental Results for Air at Equilibrium Temperatures up to 3440 K (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: This report describes a technique which combines theory and experiments for determining relaxation times in gases. The technique is based on the measurement of shapes of the bow shock waves of low-fineness-ratio cones fired from high-velocity guns. The theory presented in the report provides a means by which shadowgraph data showing the bow waves can be analyzed so as to furnish effective relaxation times. Relaxation times in air were obtained by this technique and the results have been compared with values estimated from shock tube measurements in pure oxygen and nitrogen. The tests were made at velocities ranging from 4600 to 12,000 feet per second corresponding to equilibrium temperatures from 35900 R (19900 K) to 6200 R (34400 K), under which conditions, at all but the highest temperatures, the effective relaxation times were determined primarily by the relaxation time for oxygen and nitrogen vibrations.
    Keywords: Aerodynamics
    Type: NASA-TN-D-327
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  • 17
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    Aerodynamic and Landing Measurements Obtained During the First Powered Flight of the North American X-15 Research Airplane (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: During the first powered flight of the North American X-15 research airplane on September 17, 1959, a Mach number of 2.1 and an altitude of 52,000 feet were attained. Static and dynamic maneuvers were performed to evaluate the characteristics of the airplane at subsonic and supersonic speeds. Data from these maneuvers as well as from the launch and landing phases are presented, discussed, and compared with predicted values. The rate of separation of the X-15 from the B-52 carrier airplane at launch was less than that predicted by wind-tunnel studies and was less rapid than in the lightweight condition of the initial glide flight. In addition, the angular motions and bank angle attained following the launch were of lesser magnitude than in the glide flight. Stable longitudinal-stability trends were apparent during the acceleration to maximum speed, and the pilot reported experiencing little or no transonic trim excursions. An inexplicable high-frequency vibration, which occurred at Mach numbers above 1.4, is being investigated further. Essentially linear lift and stability characteristics were indicated within the limited ranges of angle of attack and angle of sideslip investigated. The dynamic longitudinal and lateral-directional stability and control-effectiveness characteristics appeared satisfactory to the pilot. Although the longitudinal- and lateral-directional-damping ratios showed no significant change from subsonic to supersonic speeds, on the basis of time to damp, the damping characteristics at supersonic speeds appeared to the pilot to be somewhat improved over those at subsonic speeds.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TM-X-269
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  • 18
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    Wind-Tunnel Investigation of a Small-Scale Model of an Aerial Vehicle Supported by Tilting Ducted Fans (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: A wind-tunnel investigation has been made to study the static longitudinal and lateral stability characteristics of a simplified aerial vehicle supported by ducted fans that tilt relative to the airframe. The ducts were in a triangular arrangement with one duct in front and two at the rear in order to minimize the influence of the downwash of the front duct on the rear ducts. The results of the investigation were compared with those of a similar investigation for a tandem two-duct arrangement in which the ducts were fixed (rather than tiltable) relative to the airframe, since the three-duct configuration had been devised in an attempt to avoid some of the deficiencies of the tandem fixed-duct configuration. The results of the investigation indicated that the tilting-duct arrangement had less noseup pitching moment for a given forward speed than the tandem fixed-duct arrangement. The model had less angle-of-attack instability than the tandem fixed-duct arrangement. The model was directionally unstable but had a positive dihedral effect throughout the test speed range.
    Keywords: Aerodynamics
    Type: NASA-TN-D-409 , L-961
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  • 19
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    The Longitudinal Aerodynamic Characteristics of a Sweptback Wing-Body Combination With and Without End Plates at Mach Numbers from 0.40 to 0.93 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: An investigation was made at high subsonic speeds in the Langley high-speed 7- by 10-foot tunnel to determine the effect of end plates on the longitudinal aerodynamic characteristics of a sweptback wing-body combination with and without drooped chord-extensions. The wing had 45 deg sweepback of the quarter-chord line, an aspect ratio of 4, a taper ratio of 0.3, and NACA 65AO06 airfoil sections parallel to the plane of symmetry, and was mounted near the rear of a body of revolution having a fineness ratio of approximately 8. The results indicated that the addition of the end plates to either the wing with drooped chord-extensions or to the wing without drooped chord-extensions slightly increased the lift in the low angle-of-attack range but slightly decreased the lift at moderate and high angles of attack. The addition of the end plates to the wing without the chord-extensions caused a small increase in the maximum lift-drag ratio at Mach numbers below 0.65 and a slight decrease at the higher Mach numbers; however, for the addition of the end plates to the wing with the chord- extensions the maximum lift-drag ratio was slightly decreased below a Mach number of 0.88, while a slight increase occurred for the higher Mach numbers. The addition of the end plates to the wings with and without the chord-extensions caused the static longitudinal stability to increase considerably for all Mach numbers; however, only a slight reduction in the aerodynamic-center variation with Mach number was observed.
    Keywords: Aerodynamics
    Type: NASA-TN-D-389 , L-834
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  • 20
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    Some Examples of the Applications of the Transonic and Supersonic Area Rules to the Prediction of Wave Drag (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: The experimental wave drags of bodies and wing-body combinations over a wide range of Mach numbers are compared with the computed drags utilizing a 24-term Fourier series application of the supersonic area rule and with the results of equivalent-body tests. The results indicate that the equivalent-body technique provides a good method for predicting the wave drag of certain wing-body combinations at and below a Mach number of 1. At Mach numbers greater than 1, the equivalent-body wave drags can be misleading. The wave drags computed using the supersonic area rule are shown to be in best agreement with the experimental results for configurations employing the thinnest wings. The wave drags for the bodies of revolution presented in this report are predicted to a greater degree of accuracy by using the frontal projections of oblique areas than by using normal areas. A rapid method of computing wing area distributions and area-distribution slopes is given in an appendix.
    Keywords: Aerodynamics
    Type: NASA-TN-D-446 , L-1000
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  • 21
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    Data from a Static-Thrust Investigation of Large-Scale General Research VTOL-STOL Model in Ground Effect (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: The model was tested at two different elevations with the wing pivot at 1.008 and 2.425 propeller diameters above the ground. The slipstream of the propellers was deflected by tilting the wing and propellers, by deflections of large-chord trailing-edge flaps, and by combinations of flap deflection and wing tilt. Tests were conducted over a range of propeller disk loadings from 7.41 to 29.70 pounds per square foot. Force data for the complete model and pressure distributions for the wing and flaps behind one propeller were recorded and are presented in tabular form without analysis.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-397 , L-987
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  • 22
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    Effect of Blade Cutout on Power Required by Helicopters Operating at High Tip-Speed Ratios (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: A numerical study was made of the effects of blade cutout on the power required by a sample helicopter rotor traveling at tip-speed ratios of 0.3, 0.4, and 0.5. The amount of cutout varied from 0 to 0.5 of the rotor radius and the calculations were carried out for a thrust coefficient-solidity ratio of 0.04. In these calculations the blade within the cutout radius was assumed to have zero chord. The effect of such cutout on profile-drag power ranged from almost no effect at a tip-speed ratio of 0.3 to as much as a 60 percent reduction at a tip-speed ratio of 0.5. Optimum cutout was about 0.3 of the rotor radius. Part of the large power reduction at a tip-speed ratio of 0.5 resulted from a reduction in tip-region stall, brought about by cutout. For tip-speed ratios greater than 0.3, cutout also effected a significant increase in the ability of the rotor to overcome helicopter parasite drag. It is thus seen that the adverse trends (at high tip-speed ratios) indicated by the uniform-chord theoretical charts are caused in large measure by the center portion of the rotor. The extent to which a modified-design rotor can actually be made more efficient at high speeds than a uniform-chord rotor will depend in practice on the degree of success in minimizing the blade plan form near the center and on special modifications in center-section profiles. A few suggestions and estimates in regard to such modifications are included herein.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-382 , L-696
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  • 23
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    Heat Transfer to 36.75 and 45 degree Swept Blunt Leading Edges in Free Flight at Mach Numbers from 1.70 to 2.99 and From 2.50 to 4.05 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: A flight investigation has been conducted to study the heat transfer to swept-wing leading edges. A rocket-powered model was used for the investigation and provided data for Mach number ranges of 1.78 to 2.99 and 2.50 to 4.05 with corresponding free-stream Reynolds number per foot ranges of 13.32 x 10(exp 6) to 19.90 x 10(exp 6) and 2.85 x 10(exp 6) to 4.55 x 10(exp 6). The leading edges employed were cylindrically blunted wedges ', three of which were swept 450 with leading-edge diameters of 1/4, 1/2, and 3/4 inch and one swept 36-750 with a leading-edge diameter of 1/2 inch. In the high Reynolds number range, measured values of heat transfer were found to be much higher than those predicted by laminar theory and at the larger values of leading-edge diameter were approaching the values predicted by turbulent theory. For the low Reynolds number range a comparison between measured and theoretical heat transfer showed that increasing the leading-edge diameter resulted in turbulent flow on the cylindrical portion of the leading edge.
    Keywords: Aerodynamics
    Type: NASA-TM-X-208
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  • 24
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    The Shock-Wave Patterns on a Cranked-Wing Configuration (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: The shock-wave patterns of a complex configuration with cranked cruciform wings and a cone-cylinder body were examined to determine the interaction of the body bow wave with the flow field about the wing. Also of interest, was the interaction of the forward (760 sweptback) wing leading-edge wave with the rear (600 sweptback) wing leading-edge wave. The shadowgraph pictures of the model in free flight at a Mach number of 4.9, although not definitive, appear to indicate that the body bow wave crosses the outer wing panel after first being refracted either by the leading-edge wave of the 600 sweptback wing or by pressure fields in the flow crossing the wing.
    Keywords: Aerodynamics
    Type: NASA-TN-D-346 , A-433
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  • 25
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    Axial-Force Reduction by Interference Between Jet and Neighboring Afterbody (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: Experimental results are presented for an exploratory investigation of the effectiveness of interference between jet and afterbody in reducing the axial force on an afterbody with a neighboring jet. In addition to the interference axial force., measurements are presented of the interference normal force and the center of pressure of the interference normal force. The free-stream Mach number was 2.94, the jet-exit Mach number was 2.71, and the Reynolds number was 0.25 x 10, based on body diameter. The variables investigated include static-pressure ratio of the jet (up to 9), nacelle position relative to afterbody, angle of attack (-5 deg to 10 deg), and afterbody shape. Two families of afterbody shapes were tested. One family consisted of tangent-ogive bodies of revolution with varying length and base areas. The other family was formed by taking a planar slice off a circular cylinder with varying angle between the plane and cylinder. The trends with these variables are shown for conditions near maximum jet-afterbody interference. The interference axial forces are large and favorable. For several configurations the total afterbody axial force is reduced to zero by the interference.
    Keywords: Aerodynamics
    Type: NASA-TN-D-332
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  • 26
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    Aerodynamic Performance and Static Stability at Mach Number 3.3 of an Aircraft Configuration Employing Three Triangular Wing Panels and a Body Equal Length (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: An aircraft configuration, previously conceived as a means to achieve favorable aerodynamic stability characteristics., high lift-drag ratio, and low heating rates at high supersonic speeds., was modified in an attempt to increase further the lift-drag ratio without adversely affecting the other desirable characteristics. The original configuration consisted of three identical triangular wing panels symmetrically disposed about an ogive-cylinder body equal in length to the root chord of the panels. This configuration was modified by altering the angular disposition of the wing panels, by reducing the area of the panel forming the vertical fin, and by reshaping the body to produce interference lift. Six-component force and moment tests of the modified configuration at combined angles of attack and sideslip were made at a Mach number of 3.3 and a Reynolds number of 5.46 million. A maximum lift-drag ratio of 6.65 (excluding base drag) was measured at a lift coefficient of 0.100 and an angle of attack of 3.60. The lift-drag ratio remained greater than 3 up to lift coefficient of 0.35. Performance estimates, which predicted a maximum lift-drag ratio for the modified configuration 27 percent greater than that of the original configuration, agreed well with experiment. The modified configuration exhibited favorable static stability characteristics within the test range. Longitudinal and directional centers of pressure were slightly aft of the respective centroids of projected plan-form and side area.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-330
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  • 27
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    Large-Scale Wind-Tunnel Tests and Evaluation of the Low-Speed Performance of a 35 deg Sweptback Wing Jet Transport Model Equipped with a Blowing Boundary-Layer-Control Flap and Leading-Edge Slat (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: A wind-tunnel investigation was conducted to determine the effect of trailing-edge flaps with blowing-type boundary-layer control and leading-edge slats on the low-speed performance of a large-scale jet transport model with four engines and a 35 deg. sweptback wing of aspect ratio 7. Two spanwise extents and several deflections of the trailing-edge flap were tested. Results were obtained with a normal leading-edge and with full-span leading-edge slats. Three-component longitudinal force and moment data and boundary-layer-control flow requirements are presented. The test results are analyzed in terms of possible improvements in low-speed performance. The effect on performance of the source of boundary-layer-control air flow is considered in the analysis.
    Keywords: Aerodynamics
    Type: NASA-TN-D-333 , A-340
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  • 28
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    Evaluation of Blended Wing-Body Combinations with Curved Plan Forms at Mach Numbers Up to 3.50 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: This investigation is a continuation of the experimental and theoretical evaluation of the effects of wing plan-form variations on the aerodynamic performance characteristics of blended wing-body combinations. The present report compares previously tested straight-edged delta and arrow models which have leading-edge sweeps of 59.04 and 70-82 deg., respectively, with related models which have plan forms with curved leading and trailing edges designed to result in the same average sweeps in each case. All the models were symmetrical, without camber, and were generally similar having the same span, length, and aspect ratios. The wing sections had an average value of maximum thickness ratio of about 4 percent of the local wing chords in a streamwise direction. The wing sections were computed by varying their shapes along with the body radii (blending process) to match the selected area distribution and the given plan form. The models were tested with transition fixed at Reynolds numbers of roughly 4,000,000 to 9,000,000, based on the mean aerodynamic chord of the wing. The characteristic effect of the wing curvature of the delta and arrow models was an increase at subsonic and transonic speeds in the lift-curve slopes which was partially reflected in increased maximum lift-drag ratios. Curved edges were not evaluated on a diamond plan form because a preliminary investigation indicated that the curvature considered would increase the supersonic zero-lift wave drag. However, after the test program was completed, a suitable modification for the diamond plan form was discovered. The analysis presented in the appendix indicates that large reductions in the zero-lift wave drag would be obtained at supersonic Mach numbers if the leading- and trailing-edge sweeps are made to differ by indenting the trailing edge and extending the root of the leading edge.
    Keywords: Aerodynamics
    Type: NASA-TM-X-379
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  • 29
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    A Wind-Tunnel Investigation of the Development of Lift on Wings in Accelerated Longitudinal Motion (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: An investigation was made in the Langley 300 MPH 7- by 10-foot tunnel to determine the development of lift on a wing during a simulated constant-acceleration catapult take-off. The investigation included models of a two-dimensional wing, an unswept wing having an aspect ratio of 6, a 35 deg. swept wing having an aspect ratio of 3.05, and a 60 deg. delta wing having an aspect ratio of 2.31. All the wings investigated developed at least 90 percent of their steady-state lift in the first 7 chord lengths of travel. The development of lift was essentially independent of the acceleration when based on chord lengths traveled, and was in qualitative agreement with theory.
    Keywords: Aerodynamics
    Type: NASA-TN-D-422 , L-1027
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  • 30
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    Transition Reynolds Numbers of Separated Flows at Supersonic Speeds (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: Experimental research has been conducted on the effects of wall cooling, Mach number, and unit Reynolds number on the transition Reynolds number of cylindrical separated boundary layers on an ogive-cylinder model. Results were obtained from pressure and temperature measurements and shadowgraph observations. The maximum scope of measurements encompassed Mach numbers between 2.06 and 4.24, Reynolds numbers (based on length of separation) between 60,000 and 400,000, and ratios of wall temperature to adiabatic wall temperature between 0.35 and 1.0. Within the range of tile present tests, the transition Reynolds number was observed to decrease with increasing wall cooling, increase with increasing Mach number, and increase with increasing unit Reynolds number. The wall cooling effect was found to be four times as great when the attached boundary layer upstream of separation was cooled in conjunction with cooling of the separated boundary layer as when only the separated boundary layer was cooled. Wall cooling of both the attached and separated flow regions also caused, in some cases, reattachment in the otherwise separated region. Cavity resonance present in the separated region for some model configurations was accompanied by a large decrease in transition Reynolds number at the lower test Mach numbers.
    Keywords: Aerodynamics
    Type: NASA-TN-D-349 , A-178
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  • 31
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    Large-Scale Wind-Tunnel Tests of a Wingless Vertical Take-Off and Landing Aircraft: Preliminary Results (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: Large-scale wind-tunnel tests were made of a wingless vertical take-off and landing aircraft at zero sideslip to determine performance and longitudinal stability and control characteristics at airspeeds from 0 to 70 knots. Roll control and rudder effectiveness were also obtained. Limitations in the propulsion system restricted the lift for which level flight could be simulated to approximately 1500 pounds. Test variables with roll control and rudder undeflected were airspeed, vane setting, angle of attack, elevator deflection, and power. In most of the tests angle of attack, elevator, and power were varied individually while the other four parameters were held constant at previously determined values required for simulating trimmed level flight. The majority of the tests were made with power on and tail on at airspeeds between 20 and 70 knots. However, a limited number of data were obtained for the following conditions: (1) at zero velocity, horizontal tail on, power on; (2) at forward velocity, tail off and power on; and (3) at forward velocity, tail on, but with power off.
    Keywords: Aerodynamics
    Type: NASA-TN-D-326
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  • 32
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    Transonic Aerodynamic Characteristics of a Model of a Proposed Six-Engine Hull-Type Seaplane Designed for Supersonic Flight (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: Force tests of a model of a proposed six-engine hull-type seaplane were performed in the Langley 8-foot transonic pressure tunnel. The results of these tests have indicated that the model had a subsonic zero-lift drag coefficient of 0.0240 with the highest zero-lift drag coefficient slightly greater than twice the subsonic drag level. Pitchup tendencies were noted for subsonic Mach numbers at relatively high lift coefficients. Wing leading-edge droop increased the maximum lift-drag ratio approximately 8 percent at a Mach number of 0.80 but this effect was negligible at a Mach number of 0.90 and above. The configuration exhibited stable lateral characteristics over the test Mach number range.
    Keywords: Aerodynamics
    Type: NASA-TM-X-246
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  • 33
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    Characteristics at Mach Number of 2.03 of a Series of Wings Having Various Spanwise Distributions of Thickness Ratio and Chord (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: A series of semispan wing models having various spanwise distributions of both thickness ratio and chord but having the same effective thickness ratio was tested in the Langley 4-by 4-foot supersonic pressure tunnel at Mach number 2.03 and Reynolds numbers from 1.9 x 10(exp 6) to 6.5 x 10(exp 6) complex wing forms with thickened roots, extended root chords, and higher volumes show appreciably lower zero-lift wave drag coefficients than the plain swept wings. A calculative technique for the determination of wave drag has been applied to one of the complex wings of the series and good agreement is shown with experimental results. The complex wing forms showed higher drags due to lift than the plain swept wings.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-631
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  • 34
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    Aerodynamic Characteristics of a Large-Scale Unswept Wing-Body-Tail Configuration with Blowing Applied Over the Flap and Wind Leading Edge (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: An investigation has been conducted in the Langley full-scale tunnel to determine the effects of a blowing boundary-layer-control lift-augmentation system on the aerodynamic characteristics of a large-scale model of a fighter-type airplane. The wing was unswept at the 70-percent- chord station, had an aspect ratio of 2.86, a taper ratio of 0.40, and 4-percent-thick biconvex airfoil sections parallel to the plane of symmetry. The tests were conducted over a range of angles of attack from approximately -4 deg to 23 deg for a Reynolds number of approximately 5.2 x 10(exp 6) which corresponds to a Mach number of 0.08. Blowing rates were normally restricted to values just sufficient to control air-flow separation. The results of this investigation showed that wing leading-edge blowing in combination with large values of wing leading-edge-flap deflection was a very effective leading-edge flow-control device for wings having highly loaded trailing-edge flaps. With leading-edge blowing there was no hysteresis of the lift, drag, and pitching-moment characteristics upon recovery from stall. End plates were found to improve the lift and drag characteristics of the test configuration in the moderate angle-of-attack range, and blockage to one-quarter of the blowing-slot area was not detrimental to the aerodynamic characteristics. Blowing boundary-layer control resulted in a considerably reduced landing speed and reduced landing and take-off distances. The ailerons were very effective lateral-control devices when used with blowing flaps.
    Keywords: Aerodynamics
    Type: NASA-TN-D-407 , L-927
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  • 35
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    Some Divergence Characteristics of Low-Aspect-Ratio Wings at Transonic and Supersonic Speeds (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: The problem of chordwise, or camber, divergence at transonic and supersonic speeds is treated with primary emphasis on slender delta wings having a cantilever support at the trailing edge. Experimental and analytical results are presented for four wing models having apex half-angles of 5 deg, 10 deg, 15 deg, and 20 deg. A Mach number range from 0.8 to 7.3 is covered. The analytical results include calculations based on small-aspect-ratio theory, lifting-surface theory, and strip theory. A closed-form solution of the equilibrium equation is given, which is based on low-aspect-ratio theory but which applies only to certain stiffness distributions. Also presented is an iterative procedure for use with other aerodynamic theories and with arbitrary stiffness distribution.
    Keywords: Aerodynamics
    Type: NASA-TN-D-461 , L-582
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  • 36
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    Free-Spinning-Tunnel Investigation of a 1/20-Scale Model of an Unswept-Wing Jet-Propelled Trainer Airplane (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: A flutter analysis employing the kernel function for three- dimensional, subsonic, compressible flow is applied to a flutter-tested tail surface which has an aspect ratio of 3.5, a taper ratio of 0.15, and a leading-edge sweep of 30 deg. Theoretical and experimental results are compared at Mach numbers from 0.75 to 0.98. Good agreement between theoretical and experimental flutter dynamic pressures and frequencies is achieved at Mach numbers to 0.92. At Mach numbers from 0.92 to 0.98, however, a second solution to the flutter determinant results in a spurious theoretical flutter boundary which is at a much lower dynamic pressure and at a much higher frequency than the experimental boundary.
    Keywords: Aerodynamics
    Type: NASA-TN-D-381 , L-872
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  • 37
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    Flight Investigation of the Lift and Drag Characteristics of a Swept-Wing, Multijet, Transport-Type Airplane (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: The lift and drag characteristics of a Boeing KC-135 airplane were determined during maneuvering flight over the Mach number range from 0.70 to 0.85 for the airplane in the clean configuration at an altitude of 26,000 feet. Data were also obtained over the speed range of 130 knots to 160 knots at 9,000 feet for various flap deflections with gear down.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-30 , H-119
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  • 38
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    Preliminary Full-Scale Power-Off Drag of the X-15 Airplane for Mach Numbers from 0.7 to 3.1 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: Drag characteristics have been obtained for the X-15 airplane during unpowered flight. These data represent a Mach number range from about 0.7 to 3.1 and a Reynolds number range from 13.9 x 10(exp 6) to 28 x 10(exp 8), based on the mean aerodynamic chord. The full-scale data are compared with estimates compiled from several wind-tunnel facilities. The agreement between wind-tunnel and full-scale supersonic drag, uncorrected for Reynolds number effects, is reasonably close except at low supersonic Mach numbers where the flight values are significantly higher.
    Keywords: Aerodynamics
    Type: NASA-TM-X-430
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  • 39
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    Investigation at Mach Numbers of 0.20 to 3.50 of Blended Wing-Body Combinations of Sonic Design with Diamond, Delta, and Arrow Plan Forms (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The models had aspect-ratio-2 diamond, delta, and arrow wings with the leading edges swept 45.00 deg, 59.04 deg, and 70.82 deg, respectively. The wing sections were computed by varying the section shape along with the body radii (blending process) to match the prescribed area distribution and wing plan form. The wing sections had an average value of maximum thickness ratio of about 4 percent of the local chords in a streamwise direction. The models were tested with transition fixed at Reynolds numbers of about 4,000,000 to 9,000,0000, based on the mean aerodynamic chord of the wings. The effect of varying Reynolds number was checked at both subsonic and supersonic speeds. The diamond model was superior to the other plan forms at transonic speeds ((L/D)max = 11.00 to 9.52) because of its higher lift-curve slope and near optimum wave drag due to the blending process. For the wing thickness tested with the diamond model, the marked body and wing contouring required for transonic conditions resulted in a large wave-drag penalty at the higher supersonic Mach numbers where the leading and trailing edges of the wing were supersonic. Because of the low sweep of the trailing edge of the delta model, this configuration was less adaptable to the blending process. Removing a body bump prescribed by the Mach number 1.00 design resulted in a good supersonic design. This delta model with 10 percent less volume was superior to the other plan forms at Mach numbers of 1.55 to 2.35 ((L/D)max = 8.65 to 7.24), but it and the arrow model were equally good at Mach numbers of 2.50 to 3.50 ((L/D)max - 6.85 to O.39). At transonic speeds the arrow model was inferior because of the reduced lift-curve slope associated with its increased sweep and also because of the wing base drag. The wing base-drag coefficients of the arrow model based on the wing planform area decreased from a peak value of 0.0029 at Mach number 1.55 to 0.0003 at Mach number 3.50. Linear supersonic theory was satisfactory for predicting the aerodynamic trends at Mach numbers from 1.55 to 3.50 of lift-curve slope, wave drag, drag due to lift, aerodynamic-center location, and maximum lift-drag ratios for each of the models.
    Keywords: Aerodynamics
    Type: NASA-TM-X-372
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  • 40
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    Subsonic Flight Tests of a 1/7-Scale Radio-Controlled Model of the North American X-15 Airplane with Particular Reference to High Angel-of-Attack Conditions (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An investigation of the subsonic stability and control characteristics of an unpowered 1/7-scale model based on the North American X-15 airplane was conducted by using a radio-controlled model launched from a helicopter and flown in free-gliding flight. At angles of attack below about 20 deg. where the model motions represent those of the X-15 airplane, the model was found to be both longitudinally and laterally stable, and the all-movable tail surfaces were found to be very effective. The model could also be flown at much higher angles of attack where the model motions did not necessarily represent those of the airplane because of slight geometrical differences and Reynolds number effects, but these test results are useful in evaluating the effectiveness at these angles of the type of lateral control system used in the X-15 airplane. In some cases, the model was flown to angles of attack as high as 60 or 70 deg. without encountering divergent or uncontrollable conditions. For some flights in which the model was subjected to rapid maneuvers, spinning motions were generated by application of corrective controls to oppose the direction of rotation. Rapid recoveries from this type of motion were achieved by applying roll control in the direction of rotation.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TM-X-283
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  • 41
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    Friction and Pressure Drag of Boundary-Layer Diverter Systems at Mach Number of 3.0 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An experimental investigation was performed at a Mach number of 3.0 to determine the friction and pressure drags of a pylon and a 20 deg- and a 40 deg-included-angle wedge diverter over a range of Reynolds number. The results indicated that the measured friction drag coefficients agreed reasonably with that predicted by flat-plate theory. The pressure drag coefficients of the 20 and 40 deg wedges agreed with those presented in the literature. The total drag coefficient of the pylon and the 20 deg wedge diverter was about 0.36, based on diverter frontal area, while the drag coefficient of the 40 deg wedge was about 0.47.
    Keywords: Aerodynamics
    Type: NASA-TM-X-147
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  • 42
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    Free-spinning-tunnel Investigation of a 1/30 Scale Model of a Twin-jet-swept-wing Fighter Airplane (1960)
    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 1/30-scale dynamic model of a twin-jet swept-wing fighter airplane. The model results indicate that the optimum erect spin recovery technique determined (simultaneous rudder reversal to full against the spin and aileron deflection to full with the spin) will provide satisfactory recovery from steep-type spins obtained on the airplane. It is considered that the air-plane will not readily enter flat-type spins, also indicated as possible by the model tests, but developed-spin conditions should be avoided in as much as the optimum recovery procedure may not provide satisfactory recovery if the airplane encounters a flat-type developed spin. Satisfactory recovery from inverted spins will be obtained on the airplane by neutralization of all controls. A 30-foot- diameter (laid-out-flat) stable tail parachute having a drag coefficient of 0.67 and a towline length of 27.5 feet will be satisfactory for emergency spin recovery.
    Keywords: Aerodynamics
    Type: NASA-TM-SX-446 , L-1191 , N5154
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  • 43
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    Ground Measurements of Airplane Shock-Wave Noise at Mach Numbers to 2.0 and at Altitudes to 60,000 Feet (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-14
    Description: The intensity of shock-wave noise at the ground resulting from flights at Mach numbers to 2.0 and altitudes to 60,000 feet was measured. Meagurements near the ground track for flights of a supersonic fighter and one flight of a supersonic bomber are presented. Level cruising flight at an altitude of 60,000 feet and a Mach number of 2.0 produced sonic booms which were considered to be tolerable, and it is reasonable t o expect that cruising flight at higher altitudes will produce booms of tolerable intensity for airplanes of the size and weight of the test airplanes. The measured variation of sonic-boom intensity with altitude was in good agreement with the variation calculated by an equation given in NASA Technical Note D-48. The effect of Mach number on the ground overpressure is small between Mach numbers of 1.4 and 2.0, a result in agreement with the theory. No amplification of the shock-wave overpressures due to refraction effects was apparent near the cutoff Mach number. A method for estimating the effect of fligh-path angle on cutoff Mach number is shown. Experimental results indicate agreement with the method, since a climb maneuver produced booms of a much decreased intensity as compared with the intensity of those measured in level flight at about the same altitude and Mach number. Comparison of sound pressure levels for the fighter and bomber airp lanes indicated little effect of either airplane size or weight at an altitude of 40,000 feet.
    Keywords: Aerodynamics
    Type: NASA-TN-D-235
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  • 44
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    Shock Losses in Transonic Compressor Blade Rows (1960)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The rather extensive study of the shock losses in transonic compressors can he summarized by the following remarks: 1. A simple flow model can be used to estimate shock losses at the design point for transonic compressor blade rows and results iii reasonable correlation of loss data. It is indicated that shock losses can constitute a sizable portion of the total losses in it transonic compressor rotor. This includes all blade elements at which sonic or higher relative velocities are obtained. 2. Shock losses can he shown to exist across the blade passage (free-stream loss) and by the method of superposition with the blade profile losses result in an estimated design total loss coefficient. 3. The shock configuration was experimentally determined by the rapid pressure rise between the blades as measured by the use of barium titanate crystals. At the minimum loss operating conditions the shock is very similar to that assumed in the simple How model. 4. Shock losses obtained from a more detailed flow model were compared with the losses obtained by the simple flow model. Measured loss distribution from blade to blade closely approaches the analytical shock loss distribution. The measured distribution shows the effect of a shock boundary layer interaction. 5. The analytical method (from the detailed flow model) of determining the shock location ahead of the blade seems to apply reasonably well over a range of incidence angles. The analytical shock losses do not vary a great deal with blade element incidence angles.
    Keywords: Aerodynamics
    Type: ASME Paper No. 60-WA-77 , ASME Winter Annual Meeting; Nov 27, 1960 - Dec 02, 1960; New York, NY; United States
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  • 45
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    High Speed Schlieren Studies of Flow Over Mercury Atlas Vehicle in the Langley 2-Foot Transonic Aeroplasticity Tunnel (1960)
    In:  Other Sources
    Details
    Publication Date: 2019-07-12
    Description: Test conditions for the studies are: Mach number varying continuously from approximately 0.8 to 1.1 and Reynolds number (based on maximum diameter of Atlas) approximately 0.451 x 10(exp 6). Camera speed is 2000 frames per second.
    Keywords: Aerodynamics
    Type: L-583
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  • 46
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    Flow Over Blunt Body at M equals 20 in 2-Inch Helium Tunnel (1960)
    In:  Other Sources
    Details
    Publication Date: 2019-07-12
    Description: The film shows flow over blunt body alone, with internal spike, and with external spikes.
    Keywords: Aerodynamics
    Type: L-562
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  • 47
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    High-Speed Schlieren Movies of Decelerators at Supersonic Speeds (1960)
    In:  Other Sources
    Details
    Publication Date: 2019-07-12
    Description: Tests were conducted on several types of porous parachutes, a paraglider, and a simulated retrorocket. Mach numbers ranged from 1.8-3.0, porosity from 20-80 percent, and camera speeds from 1680-3000 feet per second (fps) in trials with porous parachutes. Trials of reefed parachutes were conducted at Mach number 2.0 and reefing of 12-33 percent at camera speeds of 600 fps. A flexible parachute with an inflatable ring in the periphery of the canopy was tested at Reynolds number 750,000 per foot, Mach number 2.85, porosity of 28 percent, and camera speed of 36oo fps. A vortex-ring parachute was tested at Mach number 2.2 and camera speed of 3000 fps. The paraglider, with a sweepback of 45 degrees at an angle of attack of 45 degrees was tested at Mach number 2.65, drag coefficient of 0.200, and lift coefficient of 0.278 at a camera speed of 600 fps. A cold air jet exhausting upstream from the center of a bluff body was used to simulate a retrorocket. The free-stream Mach number was 2.0, free-stream dynamic pressure was 620 lb/sq ft, jet-exit static pressure ratio was 10.9, and camera speed was 600 fps.
    Keywords: Aerodynamics
    Type: L-569
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  • 48
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    High-Speed Schlieren Movies of the Flow About Reefed Parachute Models Towed at Supersonic Speeds Behind a Conical Body (4.875 Inches in Diameter). Drag Values Based on the Unreefed Diameter of 1.73 F. Porosity of Unreefed Parachute is 28 Percent. (1960)
    In:  Other Sources
    Details
    Publication Date: 2019-07-12
    Description: Flexible parachute models reefed to one-eighth, one-fourth, one-third, and four tenths of its diameter were towed at speeds of Mach 1.80, 2.00, 2.20 and 2.87. Towline lengths tested were 23.40, 24.38, 26.81, and 29.25 inches. High-speed Schlieren movies of the flow are shown.
    Keywords: Aerodynamics
    Type: L-556
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  • 49
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    Calculation of Aerodynamic Loading and Twist Characteristics of a Flexible Wing at Mach Numbers Approaching 1.0 and Comparison with Experiment (1960)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: An iteration method is presented by which the detailed aerodynamic loading and twist characteristics of a flexible wing with known elastic properties may be calculated. The method is applicable at Mach numbers approaching 1.0 as well as at subsonic Mach numbers. Calculations were made for a wing-body combination; the wing was swept back 45 deg and had an aspect ratio of 4. Comparisons were made with experimental results at Mach numbers from.0.80 to 0.98.
    Keywords: Aerodynamics
    Type: NASA-TR-R-58
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  • 50
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    Recent Hypersonic Studies of Wings and Bodies (1960)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: Even though a great deal of theoretical and experimental information has been obtained in recent years on the flow over simple shapes in hypersonic flow a great deal of confusion still exists on how to interpret and extrapolate the results obtained. This paper offers information recently obtained at Langley at Mach numbers ranging from 7 to 21 encompassing both work in air and helium on shapes ranging from rods to delta wings. The results indicate that in most cases methods for making useful estimates of pressure are in hand for simple shapes. However, three-dimensional effects and the interaction between the components considerably complicates the flow fields over delta wings at low angles of attack.
    Keywords: Aerodynamics
    Type: ARS Semi-Annual Meeting; May 09, 1960 - May 12, 1960; Los Angeles, CA; United States
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  • 51
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    Landing-Impact Characteristics of Load-Alleviating Struts on a Model of a Winged Space Vehicle (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A brief experimental investigation was made of the landing-impact characteristics of a 1/9-scale dynamic model of a winged space vehicle. The landing tests were made by catapulting a free model onto a hard; surface runway and onto water. The model had a conical fuselage and a flat - plate wing with a basic delta planform and 75 deg sweepback of the leading edge. The use of yielding-metal shock absorbers and various landing-gear arrangements was investigated during landing impact. The basic landing gear consisted of a dual rubber-tired nose wheel and twin main skids aft of the center of gravity near the wing tips. landing motion and acceleration data were obtained over a range of landing attitudes, gross weights, and initial sinking speeds. Brief tests were made with an alternate nose-wheel location. An all-skid configuration also was briefly evaluated for hard-surface and water landings. The landing gear employing yielding struts for impact-energy absorption during hard-surface landings resulted in accelerations of approximately 5 1/2 g near the nose gear over a range of landing parameters. Replacing the nose wheel and tire with a skid did not significantly change the accelerations. Landings in smooth water with rigid struts and adequate planing area at the nose skid resulted in a maximum landing acceleration of approximately 4g.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-541 , L-958
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  • 52
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    Effects of Fineness Ratio and Reynolds Number on the Low-Speed Crosswind Drag Characteristics of Circular and Modified-Square Cylinders (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A wind-tunnel investigation has been made on modified-square and circular cylinders to determine the effects of fineness ratio and Reynolds numbers on the crosswind drag characteristics. Fineness ratios from 2 to 14 were investigated over a Reynolds number range from approximately 300,000 to 1,650,000 which corresponded to Mach numbers from 0.057 to 0.377.The result of the investigation show that at supercraft Reynolds numbers the drag coefficient of the circular cylinder increases with increasing Reynolds number for all fineness ratios but at low fineness ratios this effect is considerably less than at higher fineness ratios. For circular cylinders in the high fineness-ratio range there is a reduction in drag as the fineness ratio is decreased except for Reynolds numbers of 900,000 and 1,000,000, whereas at low fineness ratios the opposite trend generally occurs. The addition of hemispherical ends to the circular cylinder gave a substantial decrease in drag at a fineness ratio of 3.27 but the effect was negligible at fineness ratios of 5.27 and 10. The finite-length modified-square cylinder gave the reduction in drag over the two-dimensional modified-square cylinder for the complete range of test Reynolds numbers with the lowest fineness ratio giving the lowest drag at Reynolds numbers above 3O0,OOO.
    Keywords: Aerodynamics
    Type: NASA-TN-D-540 , L-1020
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  • 53
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    Flight Evaluation of Several Spring Force Gradients and a Bobweight in the Cyclic-Power-Control System of a Light Helicopter (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Several feel springs of different rates were evaluated in the power-control system of a light helicopter. In addition, a bobweight and viscous damper for providing maneuvering forces were evaluated. The evaluation was qualitative, based upon the combined opinions of eight research pilots and four non-pilot engineers from NASA. The evaluation revealed that desirable all-around forces for the helicopter were obtained with a 1/2-lb/in. feel spring for both longitudinal and lateral control combined with a 14-lb/g bobweight. Further investigation proved the necessity of the viscous damper in the bobweight system.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA/TN-D-537 , L-643
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  • 54
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    Effect of Base Bleed and Terminal Fairings on the Performance of Exhaust-Nozzle-Afterbody Combinations at Mach Numbers of 1.93, 2.55, and 3.05 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An investigation of four exhaust-nozzle-afterbody combinations has been conducted in the Langley 9- by 12-inch blowdown tunnel at Mach numbers of 1.93, 2.55, and 3.05. The models were tested on a pylon-mounted nacelle and the jet exhaust was simulated with cold air. Base bleed w a s varied from 0 to about 12 percent of the primary jet weight flow and was discharged in to the base region through either a sonic or supersonic bleed nozzle. The models were tested at zero degree angle of attack and the Reynolds number range was from 8 x 10(exp 6) to 9 x 10(exp 6) per foot. The results indicate that the base pressure and the performance of the exhaust-nozzle-afterbody combinations were little affected gy the high-velocity base bleed. The efficiency of the terminal-fairing model was only slightly less than that of the convergent-divergent nozzle-afterbody combinations; this difference indicates the loss associated with improved transonic efficiency at higher Mach numbers.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA/TN-D-539 , L-977
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  • 55
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    Measurement of the Maximum Speed Attained by the X-15 Airplane Powered with Interim Rocket Engines (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: On August 4, 1960, a flight was made with the X-15 airplane to the maximum speed expected with the interim rocket engines. Fully corrected airspeed measurements showed that the maximum Mach number of 3.1 +/- 0.04 and maximum true airspeed of 2,196 mph +/- 35 were attained at an altitude of 69,600 feet. At Mach numbers greater than 2.0 the pitot-static tube exhibited a negative static-pressure error which resulted in a Mach number correction of -0.18 at the maximum speed.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-615 , H-187
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  • 56
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    Experimental and Calculated Flow Fields Produced by Airplanes Flying at Supersonic Speeds (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Results are presented of a flight investigation conducted to survey the flow field generated by airplanes flying a t supersonic speeds. The pressure signatures of an F-100, an F-104, and a B-58 airplane, representing widely varying configurations, a t distances from 120 t o 425 f e e from the generating aircraft and at Mach numbers from 1.2 t o 1.8 are shown. Calculations were made by using Whitham's method and were compared with the experimental results.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-621 , H-190
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  • 57
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    A Supersonic Area Rule and an Application to the Design of a Wing-Body Combination with High Lift-Drag Ratios (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A concept for interrelating the wave drags of wing-body combinations at supersonic speeds with axial developments of cross-sectional area is presented. A swept-wing-indented-body combination designed on the basis of this concept to have significantly improved maximum lift-drag ratios over a range of transonic and moderate supersonic speeds is described. Experimental results have been obtained for this configuration at Mach numbers from 0.80 to 2.01. Maximum lift-drag ratios of approximately 14 and 9 were measured at Mach numbers of 1.15 and 1.41, respectively.
    Keywords: Aerodynamics
    Type: NASA-TR-R-72
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  • 58
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    Response of Rotor Lift to an Increase in Collective Pitch in the Case of Descending Flight, the Regime of the Rotor Being Near Autorotation (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An elementary calculation inspired by the classic treatment for the steady state permits the determination of the induced velocity and the overall lift of the rotor as a function of the collective pitch for all values of the advance per turn. The nature of the lift response is shown to be essentially a function of the rate of pitch change.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TT-F-18 , L-455 , Comptes Rendus; 247; 9; 738-741
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  • 59
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    The Numerical Calculation of Flow Past Conical Bodies Supporting Elliptic Conical Shock Waves at Finite Angles of Incidence (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The inverse method, with the shock wave prescribed to be an elliptic cone at a finite angle of incidence, is applied to calculate numerically the supersonic perfect-gas flow past conical bodies not having axial symmetry. Two formulations of the problem are employed, one using a pair of stream functions and the other involving entropy and components of velocity. A number of solutions are presented, illustrating the numerical methods employed, and showing the effects of moderate variation of the initial parameters.
    Keywords: Aerodynamics
    Type: NASA-TN-D-340 , A-385
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  • 60
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    A Flight Study of the Conversion Maneuver of a Tilt-Duct VTOL Aircraft (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Flight records are presented from an early flight test of a wing-tip mounted tilting-ducted-fan, vertical-take-off and landing (VTOL) aircraft configuration. Time histories of the aircraft motions, control positions, and duct pitching-moment variation are presented to illustrate the characteristics of the aircraft in hovering, in conversion from hovering to forward flight, and in conversion from forward flight to hovering. The results indicate that during essentially continuous slow level- flight conversions, this aircraft experiences excessive longitudinal trim changes. Studies have shown that the large trim changes are caused primarily by the variation of aerodynamic moments acting on the duct units. Action of the duct-induced downwash on the horizontal stabilizer during the conversion also contributes to the longitudinal trim variations. Time histories of hovering and slow vertical descent in the final stages of landing in calm air show angular motions of the aircraft as great as +/- 10 deg. about all axes. Stick and pedal displacements required to control the aircraft during the landing maneuver were on the order of 50 to 60 percent of the total travel available.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-372 , L-891
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  • 61
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    Equations for the Induced Velocities Near a Lifting Rotor with Nonuniform Azimuthwise Vorticity Distribution (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Equations, which can be integrated on high-speed computing machines, are developed for all three components of induced velocity at an arbitrary point near the rotor and for an arbitrary harmonic variation of vorticity. Sample calculations for vorticity which varies as the sine of the azimuth angle indicate that the normal component of induced velocity is, in this case, uniform along either side of the lateral axis.
    Keywords: Aerodynamics
    Type: NASA-TN-D-394 , L-797
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  • 62
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    Some Landing Studies Pertinent to Glider-Reentry Vehicles (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Results are presented of some landing studies that may serve as guidelines in the consideration of landing problems of glider-reentry configurations. The effect of the initial conditions of sinking velocity, angle of attack, and pitch rate on impact severity and the effect of locating the rear gear in various positions are discussed. Some information is included regarding the influence of landing-gear location on effective masses. Preliminary experimental results on the slideout phase of landing include sliding and rolling friction coefficients that have been determined from tests of various skids and all-metal wheels.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-448 , L-1066
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  • 63
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    Flight Investigation of the Low-Speed Characteristics of a 45 deg Swept-Wing Fighter-Type Airplane with Blowing Boundary-Layer Control Applied to the Leading- and Trailing-Edge Flaps (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A flight investigation has been conducted to study how pilots use the high lift available with blowing-type boundary-layer control applied to the leading- and trailing-edge flaps of a 45 deg. swept-wing airplane. The study includes documentation of the low-speed handling qualities as well as the pilots' evaluations of the landing-approach characteristics. All the pilots who flew the airplane considered it more comfortable to fly at low speeds than any other F-100 configuration they had flown. The major improvements noted were the reduced stall speed, the improved longitudinal stability at high lift, and the reduction in low-speed buffet. The study has shown the minimum comfortable landing-approach speeds are between 120.5 and 126.5 knots compared to 134 for the airplane with a slatted leading edge and the same trailing-edge flap. The limiting factors in the pilots' choices of landing-approach speeds were the limits of ability to control flight-path angle, lack of visibility, trim change with thrust, low static directional stability, and sluggish longitudinal control. Several of these factors were found to be associated with the high angles of attack, between 13 deg. and 15 deg., required for the low approach speeds. The angle of attack for maximum lift coefficient was 28 deg.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-321
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  • 64
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    Low-Speed Aerodynamic Characteristics of a Model of a Hypersonic Research Airplane at Angles of Attack up to 90 deg for a Range of Reynolds Numbers (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Static force tests have been made at low subsonic speeds for a model of a hypersonic research airplane in the Langley high-speed 7- by 10-foot tunnel to determine the aerodynamic forces and moments up to an angle of attack of 90 deg for a range of Reynolds numbers. The Reynolds numbers, based on the mean aerodynamic chord, ranged from 740,000 to 1,900,000, which correspond to dynamic pressures from 15 to 100 lb/sq ft (Mach numbers from 0.10 to 0.27). The model was tested in the clean configuration with various horizontal-tail settings, horizontal tail off, lower rudder off, fuselage alone, and with various size strakes and slats on the nose of the model. Representative results of the present investigation are presented in plotted form, and a tabulation of all the data obtained is presented in a table. Appreciable effects on side force, yawing moment, and pitching moment are indicated by changes in Reynolds number for angles of attack of 40 to 90 deg.
    Keywords: Aerodynamics
    Type: NASA-TN-D-403 , L-905
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  • 65
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    Supersonic Aerodynamic Characteristics of a Low-Drag Aircraft Configuration having an Arrow Wing of Aspect Ratio 1.86 and a Body of Fineness Ratio 20 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: A free-flight rocket-propelled-model investigation was conducted at Mach numbers of 1.2 to 1.9 to determine the longitudinal and lateral aero-dynamic characteristics of a low-drag aircraft configuration. The model consisted of an aspect-ratio -1.86 arrow wing with 67.5 deg. leading-edge sweep and NACA 65A004 airfoil section and a triangular vertical tail with 60 deg. sweep and NACA 65A003 section in combination with a body of fineness ratio 20. Aerodynamic data in pitch, yaw, and roll were obtained from transient motions induced by small pulse rockets firing at intervals in the pitch and yaw directions. From the results of this brief aerodynamic investigation, it is observed that very slender body shapes can provide increased volumetric capacity with little or no increase in zero-lift drag and that body fineness ratios of the order of 20 should be considered in the design of long-range supersonic aircraft. The zero-lift drag and the drag-due-to-lift parameter of the test configuration varied linearly with Mach number. The maximum lift-drag ratio was 7.0 at a Mach number of 1.25 and decreased slightly to a value of 6.6 at a Mach number of 1.81. The optimum lift coefficient, normal-force-curve slope, lateral-force-curve slope, static stability in pitch and yaw, time to damp to one-half amplitude in pitch and yaw, the sum of the rotary damping derivatives in pitch and also in yaw, and the static rolling derivatives all decreased with an increase in Mach number. Values of certain rolling derivatives were obtained by application of the least-squares method to the differential equation of rolling motion. A comparison of the experimental and calculated total rolling-moment-coefficient variation during transient oscillations of the model indicated good agreement when the damping-in-roll contribution was included with the static rolling-moment terms.
    Keywords: Aerodynamics
    Type: NASA-TN-D-509 , L-894
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  • 66
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    Effect of Reynolds Number on the Force and Pressure Distribution Characteristics of a Two-Dimensional Lifting Circular Cylinder (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A two-dimensional lifting circular cylinder has been tested over a Mach number range from 0.011 to 0.32 and a Reynolds number range from 135,000 to 1,580,000 to determine the force and pressure distribution characteristics. Two flaps having chords of 0.37 and 6 percent of the cylinder diameter, respectively, and attached normal to the surface were used to generate lift. A third configuration which had 6-percent flaps 1800 apart was also investigated. All flaps were tested through a range of angular positions. The investigation also included tests of a plain cylinder without flaps. The lift coefficient showed a wide variation with Reynolds number for the 6-percent flap mounted on the bottom surface at the 50-percent-diameter station, varying from a low of about 0.2 at a Reynolds number of 165,000 to a high of 1.54 at a Reynolds number of 350,000 and then decreasing almost linearly to a value of 1.0 at a Reynolds number of 1,580,000. The pressure distribution showed that the loss of lift with Reynolds number above the critical was the result of the separation point moving forward on the upper surface. Pressure distributions on a plain cylinder also showed similar trends with respect to the separation point. The variation of drag coefficient with Reynolds number was in direct contrast to the lift coefficient with the minimum drag coefficient of 0.6 occurring at a Reynolds number of 360,000. At this point the lift-drag ratios were a maximum at a value of 2.54. Tests of a flap with a chord of 0.0037 diameter gave a lift coefficient of 0.85 at a Reynolds number of 520,000 with the same lift-drag ratio as the larger flap but the position of the flap for maximum lift was considerably farther forward than on the larger flap. Tests of two 6-percent flaps spaced 180 deg apart showed a change in the sign of the lift developed for angular positions of the flap greater than 132 deg at subcriti- cal Reynolds numbers. These results may find use in application to air- craft using forebody strakes. The drag coefficient developed by the flaps when normal to the relative airstream was approximately equal to that developed by a flat plate in a similar attitude.
    Keywords: Aerodynamics
    Type: NASA-TN-D-455 , L-936
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  • 67
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    Theoretical Calculations of the Pressures, Forces, and Moments Due to Various Lateral Motions Acting on Tapered Sweptback Vertical Tails with Supersonic Leading and Trailing Edges (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Based on expressions for the linearized velocity potentials and pressure distributions given in NACA Technical Report 1268, formulas for the span load distribution, forces, and moments are derived for families of thin isolated vertical tails with arbitrary aspect ratio, taper ratio, and sweepback performing the motions constant sideslip, steady rolling, steady yawing, and constant lateral acceleration. The range of Mach number considered corresponds, in general, to the condition that the tail leading and trailing edges are supersonic. To supplement the analytical results, design-type charts are presented which enable rapid estimation of the forces and moments (expressed as stability derivatives) for given combinations of geometry parameters and Mach number.
    Keywords: Aerodynamics
    Type: NASA-TN-D-383 , L-780
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  • 68
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    A Summary of Operating Conditions Experienced by Three Military Helicopters and a Mountain-Based Commercial Helicopter (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The results of a survey of the flight conditions experienced by three military helicopters engaged in simulated and actual military missions, and a commercial helicopter operated in the mountainous terrain surrounding Denver, CO, are presented. The data, obtained with NASA helicopter VGHN recorders, represent 813 flights or 359 flying hours, and are compared where applicable to previous survey results. The current survey results show that none of the helicopters exceeded the maximum design airspeed. One military helicopter, used for instrument flight training, never exceeded 70 percent of its maximum design airspeed. The rates of climb and descent utilized by the IFR training helicopter and of the mountain-based helicopter were generally narrowly distributed within all the airspeed ranges. The number of landings per hour for all four of the helicopters ranged from 1.6 to 3.3. The turbine-engine helicopter experienced more frequent normal-acceleration increments above a threshold of +/-0.4g (where g is acceleration due to gravity) than the mountain-based helicopter, but the mountain-based helicopter experienced acceleration increments of greater magnitude. Limited rotor rotational speed time histories showed that all the helicopters were operated at normal rotor speeds during all flight conditions.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-432 , L-1157
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  • 69
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    Effect of Body-Mounted Lateral Controls and Speed Brakes on the Aerodynamic Load Distribution over a 45 deg Swept Wing at Mach Numbers from 0.80 to 0.98 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An investigation has been made in the Langley 16-foot transonic tunnel to determine the effect of body-mounted lateral controls and speed brakes on the aerodynamic load distribution over a swept wing. The lateral controls and speed brakes consisted of flat plates which rotated out of the side of the fuselage, were approximately perpendicular to the wing chord plane, and extended either above or below the chord plane. The wing had 45 deg sweep of the quarter-chord line, an aspect ratio of 3, a taper ratio of0.2, and 4-percent-thick airfoil section. Data were obtained at Mach numbers of 0.80, 0.94, and 0.98 fir angels of attack that usually ranged from about 0 deg to 21 deg. The results show that at the higher angles of attack a lower-surface body-mounted lateral control located along the wing trailing edge had higher effectiveness than a similar upper-surface control. Reduction in span from 0.3 to 0.2 of the wing semispan of an upper-surface body-mounted lateral control located along the wing trailing edge resulted in a less than proportiona1,change in control effectiveness.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA/TN-D-522 , L-789
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  • 70
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    Transonic and Supersonic Wind-Tunnel Tests of Wing-Body Combinations Designed for High Efficiency at a Mach Number of 1.41 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Wind-tunnel force tests of a number of wing-body combinations designed for high lift-drag ratio at a Mach number of 1.41 are reported. Five wings and six bodies were used in making up the various wing-body combinations investigated. All the wings had the same highly swept dis- continuously tapered plan form with NACA 65A-series airfoil sections 4 percent thick at the root tapering linearly to 3 percent thick at the tip. The bodies were based on the area distribution of a Sears-Haack body of revolution for minimum drag with a given length and volume. These wings and bodies were used to determine the effects of wing twist., wing twist and camber, wing leading-edge droop, a change from circular to elliptical body cross-sectional shape, and body indentation by the area-rule and streamline methods. The supersonic test Mach numbers were 1.41 and 2.01. The transonic test Mach number range was from 0.6 to 1.2. For the transition-fixed condition and at a Reynolds number of 2.7 x 10(exp 6) based on the mean aerodynamic chord, the maximum value of lift- drag ratio at a Mach number of 1.41 was 9.6 for a combination with a twisted wing and an indented body of elliptical cross section. The tests indicated that the transonic rise in minimum drag was low and did not change appreciably up to the highest test Mach number of 2.01. The lower values of lift-drag ratio obtained at a Mach number of 2.01 can be attributed to the increase of drag due to lift with Mach number.
    Keywords: Aerodynamics
    Type: NASA-TN-D-435 , L-260
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  • 71
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    Hydrodynamic and Aerodynamic Characteristics of a Model of a Supersonic Multijet Water-Based Aircraft Equipped with Supercavitating Hydrofoils (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The hydrodynamic and aerodynamic characteristics of a model of a multijet water-based Mach 2.0 aircraft equipped with hydrofoils have been determined. Takeoff stability and spray characteristics were very good, and sufficient excess thrust was available for takeoff in approximately 32 seconds and 4,700 feet at a gross weight of 225,000 pounds. Longitudinal and lateral stability during smooth-water landings were good. Lateral stability was good during rough-water landings, but forward location of the hydrofoils or added pitch damping was required to prevent diving. Hydrofoils were found to increase the aerodynamic lift-curve slope and to increase the aerodynamic drag coefficient in the transonic speed range, and the maximum lift-drag ratio decreased from 7.6 to 7.2 at the cruise Mach number of 0.9. The hydrofoils provided an increment of positive pitching moment over the Mach number range of the tests (0.6 to 1.42) and reduced the effective dihedral and directional stability.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TM-X-191
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  • 72
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    Experimental and Theoretical Study of a Rectangular Wing in a Vortical Wake at Low Speed (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A systematic study has been made, experimentally and theoretically, of the effects of a vortical wake on the aerodynamic characteristics of a rectangular wing at subsonic speed. The vortex generator and wing were mounted on a reflection plane to avoid body-wing interference. Vortex position, relative to the wing, was varied both in the spanwise direction and normal to the wing. Angle of attack of the wing was varied from -40 to +60. Both chordwise and spanwise pressure distributions were obtained with the wing in uniform and vortical flow fields. Stream surveys were made to determine the flow characteristics in the vortical wake. The vortex-induced lift was calculated by several theoretical methods including strip theory, reverse-flow theory, and reverse-flow theory including a finite vortex core. In addition, the Prandtl lifting-line theory and the Weissinger theory were used to calculate the spanwise distribution of vortex-induced loads. With reverse-flow theory, predictions of the interference lift were generally good, and with Weissinger's theory the agreement between the theoretical spanwise variation of induced load and the experimental variation was good. Results of the stream survey show that the vortex generated by a lifting surface of rectangular plan form tends to trail back streamwise from the tip and does not approach the theoretical location, or centroid of circulation, given by theory. This discrepancy introduced errors in the prediction of vortex interference, especially when the vortex core passed immediately outboard of the wing tip. The wake produced by the vortex generator in these tests was not fully rolled up into a circular vortex, and so lacked symmetry in the vertical direction of the transverse plane. It was found that the direction of circulation affected the induced loads on the wing either when the wing was at angle of attack or when the vortex was some distance away from the plane of the wing.
    Keywords: Aerodynamics
    Type: NASA-TN-D-339
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  • 73
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    Aerodynamic Characteristics at Mach Number 2.05 of a Series of Highly Swept Arrow Wings Employing Various Degrees of Twist and Camber (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A series of arrow wings employing various degrees of twist and camber were tested in the Langley 4- by 4-foot supersonic pressure tunnel. Aerodynamic forces and moments in pitch were measured at a Mach number of 2.05 and at a Reynolds number of 4.4 x 10(exp 6) based on the mean aerodynamic chord. Three of the wings, having a leading-edge sweep angle of 70 deg. and an aspect ratio of 2.24, were designed to produce a minimum drag (in comparison with that produced for other wings in the family) at lift coefficients of 0. 0.08, and 0.16. A fourth and a fifth wing, having a 75 deg. swept leading edge and an aspect ratio of 1.65, were designed for lift coefficients of 0 and 0.16, respectively. A 70 deg. swept arrow wing with twist and camber designed for an optimum loading at a lift coefficient considerably less than that for maximum lift-drag ratio gave the highest lift-drag ratio of all the wings tested a value of 8.8 compared with a value of 8.1 for the corresponding wing without twist and camber. Two twisted and cambered wings designed for optimum loading at the lift coefficient for maximum lift-drag ratio gave only small increases in maximum lift-drag ratios over that obtained for the corresponding flat wings. However, in all cases, the lift-drag ratios obtained were far below the theoretical estimates.
    Keywords: Aerodynamics
    Type: NASA-TM-X-332-1 , L-876
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  • 74
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    Basic Pressure Measurements at a Mach Number of 1.43 on a Thin 45 deg Sweptback Highly Tapered Wing with Systematic Spanwise Twist Variations (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: Pressure distributions are presented on four wings: an untwisted wing to serve as a reference, and wings with linear, quadratic and cubic twist variations along the span. All the twisted wings had 0deg twist at the 10-percent-semispan station and 6deg twist at the tip. The tests were made at a Mach number of 1.43 and covered an angle-of-attack range from -4deg to 20deg. The average Reynolds number based on the wing mean aerodynamic chord was 2.9 x 10(exp 6) during tests at a stagnation pressure of 1.0 atmosphere and 1.5 x 10(exp 6) during tests at a stagnation pressure of 0.5 atmosphere.
    Keywords: Aerodynamics
    Type: NASA-TN-D-528 , L-854
    Format: text
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  • 75
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    Aerodynamic Characteristics at Mach Numbers of 1.41 and 2.01 of a Series of Cranked Wings Ranging in Aspect Ratio from 4.00 to 1.74 in Combination with a Body (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: A program has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel to determine the effects of certain wing plan-form variations on the aerodynamic characteristics of wing-body combinations at supersonic speeds. The present report deals with the results of tests of a family of cranked wing plan forms in combination with an ogive-cylinder body of revolution. Tests were made at Mach numbers of 1.41 and 2.01 at corresponding values of Reynolds number per foot of 3.0 x 10(exp 6) and 2.5 x 10(exp 6). Results of the tests indicate that the best overall characteristics were obtained with the low-aspect-ratio wings. Plan-form changes which involved decreasing the aspect ratio resulted in higher values of maximum lift-drag ratio, in addition to large increases in wing volume. Indications are that this trend would have continued to exist at aspect ratios even lower than the lowest considered in the present tests. Increases in the maximum lift-drag ratio of about 15 percent over the basic wing were achieved with practically no increase in drag. The severe longitudinal stability associated with the basic cranked wing was no longer present (within the limits of the present tests) on the wings of lower aspect ratio formed by sweeping forward the inboard portion of the trailing edge.
    Keywords: Aerodynamics
    Type: NASA-TM-X-172 , L-261
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  • 76
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    Effect of Afterbody Terminal Fairings on the Performance of a Pylon-Mounted Turbojet-Nacelle Model (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: An investigation of the effect of afterbody terminal fairings on the performance of a pylon-mounted turbojet-nacelle model has been conducted in the Langley 16-foot transonic tunnel. A basic afterbody having a boattail angle of 16 deg was investigated with and without terminal fairings. The equivalent boattail angle, based on the cross-sectional area of the afterbody and terminal fairings, was 8 deg. Therefore, a simple body of revolution with a boattail angle of 8 deg was included for comparison. The tests were made at an angle of attack of 0 deg, Mach numbers of 0.80 to 1.05, jet total-pressure ratio of 1 to approximately 5, and an average Reynolds number per foot of 4.1 x 10(exp 6). A hydrogen peroxide jet simulator was used to supply the hot-jet exhaust. The results indicate that addition of terminal fairings to a 16 deg boattail afterbody increased the thrust-minus-drag coefficients and provided the lowest effective drag of the three configurations tested.
    Keywords: Aerodynamics
    Type: NASA-TM-X-215
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  • 77
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    Longitudinal Aerodynamic Characteristics of a Four-Propeller Deflected Slipstream VTOL Model Including the Effects of Ground Proximity (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: Results are presented of a wind-tunnel investigation of the longitudinal stability, control, and performance characteristics of a model of a four-propeller deflected-slipstream VTOL airplane in the transition speed range. These results indicate that steady level-flight transition and descending flight-path angles up to 7 or 8 deg. out of the region of ground effect can be accomplished without wing stall being encountered. In general, the pitching moments out of ground proximity can be adequately trimmed by programming the stabilizer incidence to increase with increasing flap deflection, except for a relatively large diving moment in the hovering condition. The deflection of the slipstream onto the horizontal tail in proximity of the ground substantially increases the diving moment in hovering, unless the tail is set at a large nosedown incidence.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-248 , L-735
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  • 78
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    Preliminary Investigation of a Paraglider (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: A preliminary investigation of the aerodynamic and control characteristics of a flexible glider similar to a parachute in construction has been made at the Langley Research Center to evaluate its capabilities as a reentry glider. Preliminary weight estimates of the proposed vehicle indicate that such a structure can be made with extremely low wing loading. Maximum temperatures during the reentry maneuver might be held as low as about 1,500 F. The results of wind-tunnel and free-glide tests show that the glider when constructed of nonporous material performed extremely well at subsonic speeds and could be flown at angles of attack from about 200 to 900. At supersonic speeds the wing showed none of the unfavorable tendencies exhibited by conventional parachutes at these speeds, such as squidding and breathing. Several methods of packing and deploying the glider have been successfully demonstrated. The results of this study indicate that this flexible-lifting-surface concept may provide a lightweight controllable paraglider for manned space vehicles.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-443 , L-827
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  • 79
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    Investigation at Mach Numbers of 0.60 to 3.50 of Blended Wing-Body Combinations with Cambered and Twisted Wings with Diamond, Delta and Arrow Plan Forms (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: This investigation is a continuation of the experimental and theoretical evaluation of blended wing-body combinations. The basic diamond, delta, and arrow plan forms which had an aspect ratio of 2 with leading-edge sweeps of 45.00 deg., 59.04 deg., and 70.82 deg. and trailing edge of -45.00 deg., -18.43 deg., and 41.19 deg., respectively, are used herein as standards for evaluating the effects of camber and warp. The wing thickness distributions were computed by varying the section shape along with the body radii (blending process) to match the prescribed area distribution and wing plan form. The wing camber and warp were computed to try to obtain nearly elliptical spanwise and chordwise load distributions for each plan form and thus to obtain low drag due to lift for a range of Mach numbers for which the velocities normal to the wing leading edge are subsonic. Elliptical chordwise load distributions were not possible for the plan forms and design conditions selected, so these distributions were somewhat different for each plan form. The models were tested with transition fixed at Mach numbers from 0.60 to 3.50 and at Reynolds numbers, based on the mean aerodynamic chord of the wing, of roughly 4,000,000 to 9,000,000. At speeds where the velocities normal to the wing leading edges were supersonic, an increase in the experimental wave-drag coefficients due to camber and twist was evident, but this penalty decreased with increased sweep. Thus the minimum wave-drag coefficients for the cambered arrow model were almost identical with the zero-lift wave- drag coefficients for the uncambered arrow model at all test Mach numbers.
    Keywords: Aerodynamics
    Type: NASA-TM-X-390
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  • 80
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    Calculated Effects of Body Shape on the Bow-Shock Overpressures in the Far Field of Bodies in Supersonic Flow (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: A theory for the supersonic flow about bodies in uniform flight in a homogeneous medium is reviewed and an integral which expresses the effect of body shape upon the flow parameters in the far field is reduced to a form which may be readily evaluated for arbitrary body shapes. This expression is then used to investigate the effect of nose angle, fineness ratio, and location of maximum body cross section upon the far-field pressure jump across the bow-shock of slender bodies. Curves are presented showing the variation of the shock strength with each of these parameters. It is found that, for a wide variety of shapes having equal fineness ratios, the integral has nearly a constant value.
    Keywords: Aerodynamics
    Type: NASA-TR-R-76
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  • 81
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    Low-Speed Investigation of a Full-Span Internal-Flow Jet-Augmented Flap on a High-Wing Model with a 35 deg Swept Wing of Aspect Ratio 7.0 (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An investigation of a full-span 17-percent-chord internal-flow jet-augmented flap on an aspect-ratio-7.0 wing with 35 deg of sweepback has been made in the Langley 300-MPH 7- by 10-foot tunnel. Blowing over the conventional elevator and blowing down from a nose jet were investigated as a means of trimming the large diving moments at the high momentum and high lift coefficients. The results of the investigation showed that the model with the horizontal tail 0.928 mean aerodynamic chord above the wing-chord plane was stable to the maximum lift coefficient. The large diving-moment coefficients could be trimmed either with a downward blowing nose jet or by blowing over the elevator. Neither the downward blowing nose jet nor blowing over the elevator greatly affected the static longitudinal stability of the model. Trimmed lift coefficients up to 8.8 with blowing over the elevator and up to 11.4 with blowing down at the nose were obtained when the flap was deflected 70 deg and the total momentum coefficients were 3.26 and 4.69.
    Keywords: Aerodynamics
    Type: NASA-TN-D-434 , L-931
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  • 82
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    Tests of Aerodynamically Heated Multiweb Wing Structures in a Free Jet at Mach Number 2: Five Aluminum-Alloy Models of 20-Inch Chord with 0.064-Inch-Thick Skin, 0.025-Inch-Thick Webs, and Various Chordwise Stiffening at 2 deg Angle of Attack (1960)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An experimental study was made on five 2024-T3 aluminum-alloy multiweb wing structures (MW-2-(4), MW-4-(3), mw-16, MW-17, and MW-18), at a Mach number of 2 and an angle of attack of 2 deg under simulated supersonic flight conditions. These models, of 20-inch chord and semi-span and 5-percent-thick circular-arc airfoil section, were identical except for the type and amount of chordwise stiffening. One model with no chordwise ribs between root and tip bulkhead fluttered and failed dynamically partway through its test. Another model with no chordwise ribs (and a thinner tip bulkhead) experienced a static bending type of failure while undergoing flutter. The three remaining models with one, two, or three chordwise ribs survived their tests. The test results indicate that the chordwise shear rigidity imparted to the models by the addition of even one chordwise rib precludes flutter and subsequent failure under the imposed test conditions. This paper presents temperature and strain data obtained from the tests and discusses the behavior of the models.
    Keywords: Aerodynamics
    Type: NASA-TM-X-186
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  • 83
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    The Supersonic Transport: A Technical Summary (1960)
    In:  CASI
    Details
    Publication Date: 2019-12-11
    Description: No abstract available
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-TN-D-423
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  • 84
    Electronic Resource
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    Darstellung einiger 8-Alkyl- und 8-Alkenyl-theophylline (1960)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 93 (1960), S. 99-102 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: 8-Alkyl- und 8-Alkenyl-theophylline werden aus 1.3-Dimethyl-4.5-diaminouracil und gesättigten Carbonsäuren bzw. β.γ- und γ.δ-ungesättigten Carbonsäuren durch Kondensation nach TRAUBE dargestellt; α.β-ungesättigte Säuren reagieren nicht.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19600930117
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  • 85
    Electronic Resource
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    Notiz über 7-Mercapto-cholesterin (1960)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 93 (1960), S. 280-282 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: 7-Mercapto-cholesterin wurde durch Hydrogenolyse des aus 7-Brom-cholesterylbenzoat hergestellten 7-Acetylmercapto-cholesteryl-benzoats erhalten. Seine Konfiguration am Kohlenstoff 7 wird diskutiert.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19600930142
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  • 86
    Electronic Resource
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    Notiz über 5-Methyl-pyridin-aldehyd-(3) (1960)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 93 (1960), S. 286-288 
    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.19600930145
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  • 87
    Electronic Resource
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    Masthead (1960)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 93 (1960) 
    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.19600930101
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  • 88
    Electronic Resource
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    Darstellung der 1.3.5-Triacetyl-2-desoxy-β-D-ribofuranose (1960)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 93 (1960), S. 137-140 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Es wird die Synthese der 1.3.5-Triacetyl-2-desoxy-β-D-ribofuranose beschrieben. Die Struktur wird durch Messung der kernmagnetischen Resonanz unter Vergleich mit den beiden β-D-Ribose-tetraacetaten (1.3.4 bzw. 1.3.5) bewiesen.
    Additional Material: 2 Ill.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19600930122
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  • 89
    Electronic Resource
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    Silylierung von Aminosäuren (1960)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 93 (1960), S. 424-427 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Durch Einwirkung von Silazanen auf Aminosäuren entstehen N-Trialkylsilyl-aminosäure-trialkylsilylester und Aminosäure-trialkylsilylester.
    Additional Material: 2 Tab.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19600930223
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  • 90
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    Zur Kenntnis der Reaktionen des Kobalttetracarbonyls mit verschiedenartigen Basen, VII. Reaktionen des Kobaltcarbonyls mit Verbindungen des 3-wertigen Phosphors und seiner Homologen (1960)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 93 (1960), S. 462-467 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: In Fortführung der Versuche über phosphinhaltige Kobaltcarbonyle hat sich ergeben, daß der Reaktionsverlauf von der speziellen Art des Phosphins, d. h. den am Phosphor gebundenen organischen Resten, abhängt. So entstehen mit Triäthyl- und Tricyclohexyl-phosphin bevorzugt (ca. 10°) die 2-ionigen Verbindungen [Co(CO)3L2][Co(CO)4] (L = P(C2H5)3, P(C6H11)3); nur bei höherer Temperatur (ca. 35°) und in indifferenten Mitteln bildet sich das betr. echte Derivat des Kobaltcarbonyls [Co(CO)3L]2, das mit Triphenylphosphin auch schon bei tieferer Temperatur entsteht. Häufig liegen Gemische beider Verbindungstypen vor.  -  Mit Triphenyl-arsin und -stibin sind die gleichfalls bei tieferer Temperatur (0°) gebildeten 2-ionigen Verbindungen nur labil und wandeln sich spontan unter CO-Abgabe in die echten Derivate um.  -  Ergänzende weitere Versuche wurden mit Mangancarbonyl und der 2-zähligen Base Äthylen-bis-[diphenylphosphin] und mit Rheniumcarbonyl und Triphenylphosphin durchgeführt. Im letzteren Fall erhält man das diamagnetische Monosubstitutionsprodukt [Re(CO)4P(C6H5)3]2.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19600930228
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  • 91
    Electronic Resource
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    Über die Kupplungsaktivität einiger heterocyclischer Diazoverbindungen (1960)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 93 (1960), S. 397-405 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Eine Anzahl diazotierter heterocyclischer Amine wurde mit Phenoläthern und aromatischen Kohlenwasserstoffen auf ihr Kupplungsvermögen geprüft. Die Versuche zeigten, daß insbesondere Verbindungen aus der 1.2.4- und 1.3.4-Thiodiazol-Reihe eine sehr hohe Reaktivität besitzen, die diejenige des diazotierten 2.4-Dinitro-anilins deutlich übertrifft.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19600930220
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  • 92
    Electronic Resource
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    Über Thio-cyclopentanon-(2)-carbonsäure-(1)-ester und einige Derivate (1960)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 93 (1960), S. 428-434 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Während bei den Cyclopentanon-o-carbonsäureestern die C-Alkylierung und -Acylierung vorherrschen, sind bei den entsprechenden Thio-cyclopentanon-o-carbonsäureestern nur am Schwefel substituierte Abkömmlinge zu erhalten. -  Die Säurespaltung führt zur Thio-cyclopentanon-o-carbonsäure und nicht zu ringaufgespaltenen Produkten.  -  Bei der Oxydation entstehen äußerst stabile Disulfide.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19600930224
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  • 93
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    Synthese der den natürlichen entsprechenden 2-Desoxy-Nucleoside des Adenins, Guanins und Hypoxanthins (1960)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 93 (1960), S. 140-149 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Die Synthese der 2-Desoxy-Nucleoside des Adenins, Guanins und Hypoxanthins wird beschrieben. Auf Grund des Syntheseweges aus dem entsprechenden 2-Desoxy-riboseacetat sowie den optischen und papierchromatographischen Eigenschaften ergibt sich Identität mit den durch Abbau aus natürlichem Material gewonnenen 2-Desoxy-Nucleosiden.
    Additional Material: 6 Ill.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19600930123
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    Notiz über die Einwirkung von N-Brom-succinimid auf Disulfide (1960)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 93 (1960), S. 284-285 
    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.19600930144
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  • 95
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    Zur Konstitution des Xanthohumols (1960)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 93 (1960), S. 312-317 
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    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Es wird gezeigt, daß der aus Xanthohumol durch Alkali-Abbau entstehende 2-[γ.γ-Dimethyl-allyl]-phloroglucin-x-methyläther mit dem 1-Methyläther identisch sein dürfte. Die Synthese des Xanthohumol-dimethyläthers (Kondensation von 3-[γ.γ-Dimethyl-allyl]-phloracetophenon-4.6-dimethyläther mit Anisaldehyd) steht in Einklang mit der Konstitution des Xanthohumols als 4.2′.6′-Trihydroxy-4′-methoxy-3′-[γ.γ-dimethyl-allyl]-chalkon.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19600930206
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  • 96
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    Zur Kenntnis der Chemie der Silazane, II. Über die Synthese von Verbindungen mit Silicium-Stickstoff-Phosphor-Bindungen (1960)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 93 (1960), S. 326-332 
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    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Bei der Umsetzung von Hexamethyldisilazan und Trimethyl-methylaminosilan mit Amiden und Hydraziden von Phosphorsäureestern tritt Heterolyse der Si-N-Bindung ein, und es entstehen Trimethylsilylamide bzw. -hydrazide der entsprechenden Phosphorsäure.  -  Monochloride von Phosphorsäureestern reagieren mit Hexamethyldisilazan ebenfalls unter Spaltung der Si-N-Bindung, während bei der Umsetzung mit Trimethyl-methylamino-silan Substitution des am N-Atom sitzenden H-Atoms ohne Spaltung der Si-N-Bindung beobachtet wird.  -  Derivate der Dichlorphosphorsäure lassen sich aus Silazan und OPCl3 gewinnen.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19600930209
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  • 97
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    Untersuchungen über Stilbene, XXXI. Additionsreaktionen an α-Nitro-stilbene (1960)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 93 (1960), S. 500-505 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Die Bromwasserstoff-, Hydroxylamin- und Ammoniak-Addition an α-Nitrostilbene wird beschrieben und durch sterische Zuordnung der Reaktionsprodukte der Additionsverlauf aufgeklärt.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19600930235
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  • 98
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    Zur Reaktion der N.N-Dimethyl-arylamine mit Diacylperoxyden; der Mechanismus der Entmethylierung (1960)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 93 (1960), S. 363-374 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Der Vergleich der Systeme N.N-Dimethyl-arylamin + Diacylperoxyd und Aminoxyd + Carbonsäureanhydrid trägt zur mechanistischen Klärung bei. Im letzteren System sind N-Methyl-acet-arylide und o-Acyloxy-N.N-dimethyl-arylamine die Hauptprodukte, wobei das N-Acyloxy-anilinium-Ion als Zwischenstufe angenommen wurde19). Bei der Umsetzung des N.N-Dimethyl-p-toluidins mit Dibenzoylperoxyd in Chloroform wird die „Phenolumlagerung“ lediglich durch die basenkatalysierte Entmethylierung unterdrückt; im sauren Milieu tritt sie auf.  -  Der quantitative Produktvergleich in den Systemen N.N-Dimethyl-p-chloranilin + Diacetylperoxyd und N.N-Dimethyl-p-chloranilin-oxyd + Acetanhydrid lehrt, daß die Äquivalenz nicht vollständig ist. Im ersten System existiert noch ein unabhängiger Entmethylierungsweg, an welchem Radikale beteiligt zu sein scheinen; in Styrol oder Cumol als Lösungsmittel ging die Ausbeute am Entmethylierungsprodukt zurück.
    Additional Material: 2 Ill.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19600930215
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  • 99
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    Zur Kenntnis des Imins, III. Phosphin-iminium-chlorideDie früher von uns als Aminophosphonium-salze bezeichneten Verbindungen sind nach den Richtsätzen für die Nomenklatur der anorg. Chemie, Abschnitt 3, 153, jetzt als Phosphin-iminium-salze zu benennen. und Triphenylphosphin-imin (1960)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 93 (1960), S. 405-411 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Durch Einwirkung von Chloramin auf tertiäre Phosphine werden Phosphin-iminium-chloride, [R3PNH2]Cl, erhalten. Die Deprotonierung des Triphenylphosphin-iminium-chlorids mit Natriumamid in flüssigem Ammoniak liefert Triphenylphosphin-imin; es ist mit dem isosteren Triphenylphosphin-methylen zu vergleichen. Seine Reaktion mit polarisierbaren Sauerstoffverbindungen führt zu einem Austausch des Sauerstoffs gegen den NH-Rest; hierbei entstehen Triphenylphosphin-oxyd und die entsprechenden Iminverbindungen.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19600930221
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  • 100
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    Reaktionen mit L-Cystinderivaten, I. Chlorierende Spaltung (1960)
    Weinheim : Wiley-Blackwell
    Berichte der deutschen chemischen Gesellschaft 93 (1960), S. 782-784 
    Details
    ISSN: 0009-2940
    Keywords: Chemistry ; Inorganic Chemistry
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Chemistry and Pharmacology
    Notes: Durch chlorierende Spaltung des N.N′-Dibenzoyl-cystin-diäthylesters wird das entsprechende Sulfenylchlorid erhalten, daß sich durch Anlagerung von Äthylen und anschließende Oxydation als β-Carbäthoxy-β-benzoylamino-β′-chlor-diäthylsulfon charakterisieren ließ. Wurde dagegen Cystin-diäthylester-hydrochlorid chlorierend gespalten, so entstand in guter Ausbeute L( - )-α-Amino-β-chlor-propionsäureester.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/cber.19600930405
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