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  • Chemistry  (2,763)
  • Inorganic Chemistry  (433)
  • Aerodynamics  (67)
  • Aircraft Stability and Control  (39)
  • Fluid Mechanics and Thermodynamics  (20)
  • SPACE SCIENCES
  • 1950-1954  (1,919)
  • 1945-1949  (970)
  • 1951  (1,919)
  • 1947  (970)
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Keywords
Publisher
Years
  • 1950-1954  (1,919)
  • 1945-1949  (970)
Year
  • 1
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    Calculations of Laminar Heat Transfer Around Cylinders of Arbitrary Cross Section and Transpiration-Cooled Walls with Application to Turbine Blade Cooling (1951)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An approximate method for development of flow and thermal boundary layers in laminar regime on cylinders with arbitrary cross section and transpiration-cooled walls is obtained by use of Karman's integrated momentum equation and an analogous heat-flow equation. Incompressible flow with constant property values throughout boundary layer is assumed. Shape parameters for approximated velocity and temperature profiles and functions necessary for solution of boundary-layer equations are presented as charts, reducing calculations to a minimum. The method is applied to determine local heat-transfer coefficients and surface temperature-cooled turbine blades for a given flow rate. Coolant flow distributions necessary for maintaining uniform blade temperatures are also determined.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NACA-RM-E51F22
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Investigation of Pressure Distribution over an Extended Leading-Edge Flap on a 42 Degrees Sweptback Wing (1947)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Pressure distribution over an extended leading-edge flap on a 42 degree swept-back wing was investigated. Results indicate that the flap normal-force coefficient increased almost linearly with the angle of attack to a maximum value of 3.25. The maximum section normal-force coefficient was located about 30 percent of the flap span outboard of the inboard end and had a value of 3.75. Peak negative pressures built up at the flap leading edge as the angle of attack was increased and caused the chordwise location of the flap center of pressure to be move forward.
    Keywords: Aerodynamics
    Type: NACA-RM-L7J03
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    Investigation of Thrust Augmentation of a 1600-pound Thrust Centrifugal-flow-type Turbojet Engine by Injection of Refrigerants at Compressor Inlets (1947)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Investigations were conducted to determine effectiveness of refrigerants in increasing thrust of turbojet engines. Mixtures of water an alcohol were injected for a range of total flows up to 2.2 lb/sec. Kerosene was injected into inlets covering a range of injected flows up to approximately 30% of normal engine fuel flow. Injection of 2.0 lb/sec of water alone produced an increase in thrust of 35.8% of rate engine conditions and kerosene produced a negligible increase in thrust. Carbon dioxide increased thrust 23.5 percent.
    Keywords: Aerodynamics
    Type: NACA-RM-E7G23
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    Experimental Investigation of a Preloaded Spring-tab Flutter Model (1947)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental investigation was made of a preloaded spring-tab flutter model to determine the effects on flutter speed of aspect ratio, tab frequency, and preloaded spring constant. The rudder was mass-balanced, and the flutter mode studied was essentially one of three degrees of freedom (fin bending coupled with rudder and tab oscillations). Inasmuch as the spring was preloaded, the tab-spring system was a nonlinear one. Frequency of the tab was the most significant parameter in this study, and an increase in flutter speed with increasing frequency is indicated. At a given frequency, the tab of high aspect ratio is shown to have a slightly lower flutter speed than the one of low aspect ratio. Because the frequency of the preloaded spring tab was found to vary radically with amplitude, the flutter speed decreased with increase in initial displacement of the tab.
    Keywords: Aircraft Stability and Control
    Type: NACA-RM-L7G18
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Observations on an Aileron-Flutter Instability Encountered on a 45 Degree Swept-Back Wing in Transonic and Supersonic Flight (1947)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: In the course of a flight test of a supersonic research pilotless aircraft (the NACA RM-1), large-amplitude aileron oscillations, probably aileron compressibility flutter, were encountered in the transonic and supersonic speed ranges. The wing was oscillating at the same frequency as the aileron. The aircraft was equipped with 45 degree swept-back wings of symmetrical NASA 65-010 airfoil section. Completely mass-balanced ailerons with 20 degree beveled trailing edges were installed on the wings. The ailerons were free floating with no mechanical restraining force other than the friction of the aileron hinges and servomechanism bearings throughout the high-speed interval of flight.
    Keywords: Aerodynamics
    Type: NACA-RM-L6L09
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 6
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    High-Speed Wind-Tunnel Investigation of the Lateral Control Characteristics of Plain Ailerons on a Wing with Various Amounts of Sweep (1947)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A three-dimensional investigation of straight-sided-profile plain ailerons on a wing with 30 degrees and 45 degrees of sweepback and sweepforward was made in a high-speed wind tunnel for aileron deflections from -10 degrees to 10 degrees and at Mach numbers from 0.60 to 0.96. Wing configurations of 30 degrees generally reduced the severity of the large changes in rolling-moment and aileron hinge-moment coefficients experienced by the upswept wing configurations as the result of compression shock and extended to higher Mach numbers the speeds at which such changes occurred.
    Keywords: Aerodynamics
    Type: NACA-RM-L7I15
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 7
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    Low-Speed Wind-Tunnel Tests of a Pilotless Aircraft Having Horizontal and Vertical Wings and Cruciform Tail (1947)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Low-speed tests of a pilotless aircraft were conducted in the Langley propeller-research tunnel to provide information for the estimation of the longitudinal stability and. control, to measure the aileron effectiveness, and to calibrate the radome and the Machmeter pitot-static orifices. It was found that the model possessed a stEb.le variation of elevator angle required for trim throughout the speed range at the design angle of attack. A comparison of the airplane with and without JATO units and with an alternate rocket booster showed that a large loss in longitudinal stability and control resulting from the addition of the rocket booster to the aircraft was sufficient to make the rocket-booster assembly unsatisfactory as an alternate for the JATO units. Reversal of the aileron effectiveness was evident at positive deflections of the vertical wing flap indicating that the roll-stabilization system would produce roiling moments in a tight right turn contrary to its design purpose. Vertical-wing-flap deflections caused large errors in the static-pressure reading obtained by the original static-tube installation. A practical installation point on the fuselage was located which should yield reliable measurement of the free-stream static pressure.
    Keywords: Aircraft Stability and Control
    Type: NACA-RM-L6J18a
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 8
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    Calculations of the Supersonic Wave Drag of Nonlifting Wings with Arbitrary Sweepback and Aspect Ratio: Wings Swept Behind the Mach Lines (1947)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: On the basis of a recently developed theory for finite sweptback wings at supersonic speeds, calculations of the supersonic wave drag at zero lift were made for a series of wings having thin symmetrical biconvex sections with untapered plan forms and various angles of sweepback and aspect ratios. The results are presented in a unified form so that a single chart permits the direct determination of the wave drag for this family of airfoils for an extensive range of aspect ratio and sweepback angle for stream Mach numbers up to a value corresponding to that at which the Mach line coincides with the wing leading edge. The calculations showed that in general the wave-drag coefficient decreased with increasing sweepback. At Mach numbers for which the Mach lines are appreciably ahead of the wing leading edge, the 'wave-drag coefficient decreased to an important extent with increases in aspect ratio or slenderness ratio. At Mach numbers for which the Mach lines approach the wing leading edge (Mach numbers approaching a value equal to the secant of the angle of sweepback), the wave-drag coefficient decreased with reductions in aspect ratio or slenderness ratio. In order to check the results obtained by the theory, a comparison was made with the results of tests at the Langley Memorial Aeronautical Laboratory of sweptback wing attached to a freely falling body. The variation of the drag with Mach number and aspect ratio as given by the theory appeared to be in reasonable
    Keywords: Aerodynamics
    Type: NACA-RM-L6K29
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 9
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    Measurements of Static and Total Pressure Throughout the Transonic Speed Range as Obtained from an Airspeed Head Mounted on a Freely Falling Body (1947)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: No abstract available
    Keywords: Aerodynamics
    Type: NACA-RM-L7C04a
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 10
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    An Investigation of Convergent-Divergent Diffusers at Mach Number 1.85 (1947)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation has been conducted in the Cleveland 18- by 18-inch supersonic tunnel at a Mach number of 1.85 and angles of attack from 0 deg to 5 deg to determine optimum design configurations for a convergent-divergent type of supersonic diffuser with a subsonic diffuser of 5 deg included divergence angle. Total pressure recoveries in excess of theoretical recovery across a normal shock at a free-stream Mach number of 1.85 wore obtained with several configurations. The highest recovery for configurations without a cylindrical throat section was obtained with an inlet having an included convergence angle of 20 deg. Insertion of a 2-inch throat section between a 10 deg included angle inlet and the subsonic diffuser stabilized the shock inside the diffuser and resulted in recoveries as high as 0.838 free-stream total pressure at an angle of attack of 0 deg, corresponding to recovery of 92.4 percent of the kinetic energy of the free air stream. Use of the throat section also lessened the reduction in recovery of all configurations due to angle of attack.
    Keywords: Aerodynamics
    Type: NACA-RM-E6K21
    Format: application/pdf
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