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  • Other Sources  (382)
  • Aerodynamics  (382)
  • 1960-1964  (150)
  • 1955-1959  (141)
  • 1950-1954  (76)
  • 1935-1939  (15)
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  • Other Sources  (382)
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  • 1
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    Effect of Rapid Pressure Decay on Solid Propellant Combustion (1961)
    In:  Other Sources
    Details
    Publication Date: 2015-04-02
    Description: Effect of rapid pressure decay on solid propellant combustion
    Keywords: Aerodynamics
    Type: ARS Journal; Volume 31; No. 11; 1584-1586
    Format: text
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  • 2
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    Effects of Mass Addition on the Stability of Slender Cones at Hypersonic Speeds (1963)
    In:  Other Sources
    Details
    Publication Date: 2017-03-07
    Description: The effect of mass addition on the flow over bodies moving at hypersonic speeds has been studied by several investigators (e.g., Cresci and Libby). In most of this work, primary attention logically has been directed toward the effects of foreign-gas injection on heat transfer and pressure distributions, and, principally for this reason, most of the work ha been done at zero angle of attack. The foreign gas can be provided either by some active injection system or by the action of an ablation heat shield. With increasing rates of injection, the basic flow about the body can be affected significantly. One such effect was observed in the paper by Cresci and Libby, where it was shown that the shockwave standoff distance can be increased by gas injection at the nose of a body.
    Keywords: Aerodynamics
    Type: AIAA Journal; Volume 1; No. 4; 939-940
    Format: text
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  • 3
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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
    Format: application/pdf
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  • 4
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    Preliminary Wind-Tunnel Investigation of the Performance of Republic F-105 Wing-Root Inlet Configurations at Various Angles of Attack and a Mach Number of 2.01 (1957)
    In:  CASI
    Details
    Publication Date: 2019-05-31
    Description: A 1/13-scale model of the forebody of the Republic F-105 with twin-duct wing-root inlets was tested in the Langley 4- by 4-foot supersonic pressure tunnel through a range of angle of attack from -4 deg to 15 deg at a Mach number of 2.01 and a Reynolds number of approximately 3.4 x 10(exp 6) per foot. The tests were made with four configurations which incorporated varying amounts of sweep and stagger of the inlet leading edges, modifications to the areas of the boundary-layer diverter floor plate, and modifications to the area of the boundary-layer diverter bleed slots. The highest overall pressure recovery at an angle of attack of 0 deg (average total-pressure recovery, 0.84 mass-flow ratio, 0.98) was achieved with configuration having an inlet leading-edge sweep angle of 58 deg with no leading-edge stagger. Stagger was found to improve the angle-of- attack performance, but at a sacrifice in inlet efficiency for an angle of attack of 0 deg. The boundary-layer diverter floor height, of the order of one boundary-layer thickness, was satisfactory for bypassing the fuselage boundary layer. The boundary-layer diverter-plate bleed slots were effective in increasing the total-pressure recovery of the inlet. The total-pressure-recovery contour plots, taken at the compressor-face station, indicate the existence of high-velocity "cores" throughout the inlet operating range.
    Keywords: Aerodynamics
    Type: NACA-RM-SL56L12
    Format: application/pdf
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  • 5
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    Design Guide for Pitch-Up Evaluation and Investigation at High Subsonic Speeds of Possible Limitations Due to Wing-Aspect-Ratio Variations (1959)
    In:  CASI
    Details
    Publication Date: 2019-05-11
    Description: A design guide is suggested as a basis for indicating combinations of airplane design variables for which the possibilities of pitch-up are minimized for tail-behind-wing and tailless airplane configurations. The guide specifies wing plan forms that would be expected to show increased tail-off stability with increasing lift and plan forms that show decreased tail-off stability with increasing lift. Boundaries indicating tail-behind-wing positions that should be considered along with given tail-off characteristics also are suggested. An investigation of one possible limitation of the guide with respect to the effects of wing-aspect-ratio variations on the contribution to stability of a high tail has been made in the Langley high-speed 7- by 10-foot tunnel through a Mach number range from 0.60 to 0.92. The measured pitching-moment characteristics were found to be consistent with those of the design guide through the lift range for aspect ratios from 3.0 to 2.0. However, a configuration with an aspect ratio of 1.55 failed t o provide the predicted pitch-up warning characterized by sharply increasing stability at the high lifts following the initial stall before pitching up. Thus, it appears that the design guide presented herein might not be applicable when the wing aspect ratios lower than about 2.0.
    Keywords: Aerodynamics
    Type: NASA-TM-X-26
    Format: application/pdf
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  • 6
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    An Investigation of Some Aspects of the Sonic Boom by Means of Wind-Tunnel Measurements of Pressures About Several Bodies at a Mach Number of 2.01 (1959)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation of some aspects of the sonic boom has been made with the aid of wind-tunnel measurements of the pressure distributions about bodies of various shapes. The tests were made in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 2.01 and at a Reynolds number per foot of 2.5 x 10(exp 6). Measurements of the pressure field were made at orifices in the surface of a boundary-layer bypass plate. The models which represented both fuselage and wing types of thickness distributions were small enough to allow measurements as far away as 8 body lengths or 64 chords. The results are compared with estimates made using existing theory. To the first order, the boom-producing pressure rise across the bow shock is dependent on the longitudinal development of body area and not on local details. Nonaxisymmetrical shapes may be replaced by equivalent bodies of revolution to obtain satisfactory theoretical estimates of the far-field pressures.
    Keywords: Aerodynamics
    Type: NASA-TN-D-161
    Format: application/pdf
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  • 7
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    Ground Measurements of the Shock-Wave Noise from Airplanes in Level Flight at Mach Numbers to 1.4 and at Altitudes to 45000 Feet (1959)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Time histories of noise pressures near ground level were measured during flight tests of fighter-type airplanes over fairly flat, partly wooded terrain in the e Mach number range between 1.13 and 1.4 and at altitudes from 25,000 to 45,000 feet. Atmospheric soundings and radar tracking studies were made for correlation with the measured noise data. The measured and calculated values of the pressure rise across the shock wave were generally in good agreement. There is a tendency for the theory to overestimate the pressure at locations remote from the track and to underestimate the pressures for conditions of high tailwind at altitude. The measured values of ground-reflection factor averaged about 1.8 f or the surface tested as compared to a theoretical value of 2.0. P o booms were measured in all cases. The observers also generally reported two booms; although, in some cases, only one boom was reported. The shock-wave noise associated with some of the flight tests was judged to be objectionable by ground observers, and in one case the cracking of a plate-glass store window was correlated in time with the passage of the airplane at an altitude of 25,000 feet.
    Keywords: Aerodynamics
    Type: NASA-TN-D-48
    Format: application/pdf
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  • 8
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    A Simplified Method for Estimating Subsonic Lift-Curve Slope at Low Angles of Attack for Irregular Planform Wings (1961)
    In:  CASI
    Details
    Publication Date: 2019-05-25
    Description: A simplified method is presented for estimating the lift-curve slope of irregular planform wings at subsonic speeds and low angles of attack. The present process is an extension of the method derived in NACA Technical Note 3911 and enables quick estimates of subsonic liftcurve slope, to be made whereas more refined procedures require considerable time and computation. Comparison of experimental and estimated values for a wide range of wing planforms having discontinuous spanwise sweep variation indicates good agreement. A comparison of the present procedure with a 20-step vortex method (NACA Research Memorandum L50L13) indicated good agreement for a variable-sweep configuration.
    Keywords: Aerodynamics
    Type: NASA-TM-X-525
    Format: application/pdf
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  • 9
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    Investigation of a 1/4-Scale Model of the Republic F-105 Airplane in the Langley 19-Foot Pressure Tunnel: Longitudinal Stability and Control of the Model Equipped with a Supersonic-Type Elliptical Wing-Root Inlet (1954)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: No abstract available
    Keywords: Aerodynamics
    Type: NACA-RM-SL54F28
    Format: application/pdf
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  • 10
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    A Correlation by Means of the Transonic Similarity Rules of the Experimentally Determined Characteristics of 22 Rectangular Wings of Symmetrical Profile (1952)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The transonic similarity rules have been applied to the correlation of experimental data for a series of 22 rectangular wings having symmetrical NACA 63A-series sections, aspect ratios from 1/2 to 6, and thicknesses from 2 to 10 percent. The data were obtained by use of the transonic bump technique over a Mach number range from 0.40 to 1.10, corresponding to a Reynolds number range from 1.25 to 2.05 million. The results show that it is possible to correlate experimental data throughout the subsonic, transonic, and moderate supersonic regimes by using the transonic similarity parameters in forms which are consistent with the Prandtl-Glauert rule of linearized theory. The multiple families of basic data curves for the various aspect ratios and thickness ratios have been summarized in single presentations involving only one geometric variable - the product of the aspect ratio and the l/3 power of the thickness ratio.
    Keywords: Aerodynamics
    Type: NACA-RM-A51L17b
    Format: application/pdf
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