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  • Other Sources  (23)
  • Spacecraft Propulsion and Power  (20)
  • INSTRUMENTATION AND PHOTOGRAPHY
  • 1955-1959  (23)
  • 1
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    Experimental Study of Ballistic-Missile Base Heating with Operating Rocket (1958)
    In:  CASI
    Details
    Publication Date: 2019-05-11
    Description: A rocket of the 1000-pound-thrust class using liquid oxygen and JP-4 fuel as propellant was installed in the Lewis 8- by 6-foot tunnel to permit a controlled study of some of the factors affecting the heating of a rocket-missile base. Temperatures measured in the base region are presented from findings of three motor extension lengths relative to the base. Data are also presented for two combustion efficiency levels in the rocket motor. Temperature as high as 1200 F was measured in the base region because of the ignition of burnable rocket gases. combustibles that are dumped into the base by accessories seriously aggravate the base-burning temperature rise.
    Keywords: Spacecraft Propulsion and Power
    Type: NACA-RM-E58G17
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Performance Comparison at Mach Numbers 1.8 and 2.0 of Full Scale and Quarter Scale Translating-Spike Inlets (1957)
    In:  CASI
    Details
    Publication Date: 2019-05-22
    Description: The performance of a full-scale translating-spike inlet was obtained at Mach numbers of 1.8 and 2.0 and at angles of attach from 0 deg to 6 deg. Comparisons were made between the full-scale production inlet configuration and a geometrically similar quarter-scale model. The inlet pressure-recovery, cowl pressure-distribution, and compressor-face distortion characteristics of the full-scale inlet agreed fairly well with the quarter-scale results. In addition, the results indicated that bleeding around the periphery ahead of the compressor-face station improved pressure recovery and compressor-face distortion, especially at angle of attack.
    Keywords: Spacecraft Propulsion and Power
    Type: NACA-RM-E57D16
    Format: application/pdf
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  • 3
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    Theoretical Rocket Performance of JP-4 Fuel with Several Fluorine-Oxygen Mixtures Assuming Equilibrium Composition (1958)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Theoretical rocket performance for equilibrium composition during expansion was calculated for JP-4 fuel with several fluorine-oxygen mixtures for a range of pressure ratios and oxidant-fuel ratios. The parameters included are specific impulse, combustion-chamber temperature, nozzle-exit temperature, molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, thermal conductivity, and equilibrium gas compositions. A correlation is given for the effect of chamber pressure on several of the parameters. The maximum value of specific impulse for a chamber pressure of 600 pounds per square inch absolute (40.827 atm) and an exit pressure of 1 atmosphere is 325.7 for 70.37 percent fluorine in the oxidant as compared with 284.9 and 305.1 for 100 percent oxygen and 100 percent fluorine, respectively.
    Keywords: Spacecraft Propulsion and Power
    Type: NACA-RM-E57K22
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  • 4
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    Theoretical Rocket Performance of Liquid Methane with Several Fluorine-Oxygen Mixtures Assuming Frozen Composition (1958)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Theoretical rocket performance for frozen composition during expansion was calculated for liquid methane with several fluorine-oxygen mixtures for a range of pressure ratios and oxidant-fuel ratios. The parameters included are specific impulse, combustion-chamber temperature, nozzle-exit temperature molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, and thermal conductivity. The maximum calculated value of specific impulse for a chamber pressure of 600 pounds per square inch absolute (40.827atm) and an exit pressure of 1 atmosphere is 315.3 for 79.67 percent fluorine in the oxidant.
    Keywords: Spacecraft Propulsion and Power
    Type: NACA-RM-E58B20
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  • 5
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    Near-Field and Far-Field Noise Surveys of Solid-Fuel Rocket Engines for a Range of Nozzle Exit Pressures (1959)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Measurements of near- and far-field noise pressures are presented for a 1,500-pound-thrust engine and for several 5,000-pound-thrust engines for which the nozzle exit pressure was changed systematically in order to study its effects on the noise level and spectra. Near-field surveys indicated that the highest noise pressure occurred at about 20 exit diameters downstream if the nozzle near the transition from super-sonic to subsonic flow. The acoustical power radiated from all engines averaged about 0.5 percent of the mechanical power of the exhaust stream, the least noise being radiated by the nozzle having an exit pressure less than atmospheric. The rocket engines of these tests radiate more power per cycle at the lower frequencies than arte reported for subsonic jets in other related studies.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA-TN-D-21
    Format: application/pdf
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  • 6
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    An Oil-Stream Photomicrographic Aeroscope for Obtaining Cloud Liquid-Water Content and Droplet Size Distributions in Flight (1956)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An airborne cloud aeroscope by which droplet size, size distribution, and liquid-water content of clouds can be determined has been developed and tested in flight and in wind tunnels with water sprays. In this aeroscope the cloud droplets are continuously captured in a stream of oil, which Is then photographed by a photomicrographic camera. The droplet size and size distribution can be determined directly from the photographs. With the droplet size distribution known, the liquid-water content of the cloud can be computed from the geometry of the aeroscope, the airspeed, and the oil-flow rate. The aeroscope has the following features: Data are obtained semi-automatically, and permanent data are taken in the form of photographs. A single picture usually contains a sufficient number of droplets to establish the droplet size distribution. Cloud droplets are continuously captured in the stream of oil, but pictures are taken at Intervals. The aeroscope can be operated in icing and non-icing conditions. Because of mixing of oil in the instrument, the droplet-distribution patterns and liquid-water content values from a single picture are exponentially weighted average values over a path length of about 3/4 mile at 150 miles per hour. The liquid-water contents, volume-median diameters, and distribution patterns obtained on test flights and in the Lewis icing tunnel are similar to previously published data.
    Keywords: INSTRUMENTATION AND PHOTOGRAPHY
    Type: NACA-TN-3592
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  • 7
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    A heated-wire liquid-water-content instrument and results of initial flight tests in icing conditions (1955)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Keywords: INSTRUMENTATION AND PHOTOGRAPHY
    Type: NACA-RM-A54I23
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  • 8
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    Methods for the control of interfering currents originating at the input of an electrostatic fluxmeter during its operation in a conducting medium (1959)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Keywords: INSTRUMENTATION AND PHOTOGRAPHY
    Type: NASA-TT-F-16 , Artificial Earth Satellite, V. 3; 10 p
    Format: text
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  • 9
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    Injection Principles from Combustion Studies in a 200-Pound-Thrust Rocket Engine Using Liquid Oxygen and Heptane (1955)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: The importance of atomizing and mixing liquid oxygen and heptane was studied in a 200-pound-thrust rocket engine. Ten injector elements were used with both steel and transparent chambers. Characteristic velocity was measured over a range of mixture ratios. Combustion gas-flow and luminosity patterns within the chamber were obtained by photographic methods. The results show that, for efficient combustion, the propellants should be both atomized and mixed. Heptane atomization controlled the combustion rate to a much larger extent than oxygen atomization. Induced mixing, however, was required to complete combustion in the smallest volume. For stable, high-efficiency combustion and smooth engine starts, mixing after atomization was most promising.
    Keywords: Spacecraft Propulsion and Power
    Type: NACA-RM-E55C22 , L-3572
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  • 10
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    Analysis of Flow-System Starting Dynamics of Turbopump-Fed Liquid-Propellant Rocket (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: Two rocket configurations with turbopump drive were investigated analytically. In one configuration the inlet pressure to the turbine was fixed at the design value. The second configuration employed a "bootstrap" technique for supplying energy to the turbine. An injector was the chief resistance between the pump and the rocket combustion chamber. From the analysis two parameters were developed from which the speed response time of the turbopump, the flow response time, and the maximum dynamic line loss could be evaluated. These parameters were functions of turbopump moment of inertia, design performance of the turbine, and flow-system geometry. The moment of inertia of the turbopump and the ratio of turbine torque at zero speed to design torque had the most influence on the starting dynamics of the flow system. These parameters were also applicable to the bootstrap configuration as long as the inlet pressure to the turbine exceeded half the design value.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA-MEMO-4-21-59E
    Format: application/pdf
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