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  • Other Sources  (20)
  • Spacecraft Propulsion and Power  (20)
  • 1955-1959  (20)
  • 1
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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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    NASA TECHNICAL REPORTS
  • 2
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    Static Investigation of Paddle Vane Oscillating in Jet of 1,300-Pound-Thrust Rocket Motor (1958)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: The results of a static investigation conducted to measure the normal forces on the entire jet-vane assembly and the hinge moments on the jet vane produced by a paddle vane oscillating in the jet of a 1,500-pound-thrust rocket motor are presented for vane-deflection angles from -5 deg to 25 deg. A maximum average normal force of 71 pounds with a corresponding value for maximum average hinge moment of 228 inch-pounds was obtained with the maximum area of jet vane immersed at a jet-vane angle of 25 deg decrease in thrust caused by immersion of the jet vane varied from a maximum loss of about 58 pounds, or approximately 5 percent at maximum jet-vane angle of 25 deg, to zero loss at jet-vane angles less than approximately 10 deg.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA-MEMO-10-19-58L
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    A Five-Stage Solid-Fuel Sounding-Rocket System (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: A five-stage solid-fuel sounding-rocket system which can boost a payload of 25 pounds to an altitude of 525 nautical miles and that of 100 pounds to 300 nautical miles is described. Data obtained from a typical flight test of the system are discussed.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA-MEMO-3-6-59L
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    Effect of Ozone Addition on Combustion Efficiency of Hydrogen: Liquid-Oxygen Propellant in Small Rockets (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: An experimental study shows that 2 percent by weight ozone in oxygen has little effect on overall reactivity for a range of oxidant-fuel weight ratios from 1 to 6. This conclusion is based on characteristic-velocity measurements in 200-pound-thrust chambers at a pressure of 300 pounds per square inch absolute with low-efficiency injectors. The presence of 9 percent ozone in oxygen also did not affect performance in an efficient chamber. Explosions were encountered when equipment or procedure permitted ozone to concentrate locally. These experiments indicate that even small amounts of ozone in oxygen can cause operational problems.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA-MEMO-5-26-59E , E-327
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Materials Problems in Chemical Liquid-Propellant Rocket Systems (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: With the advent of the space age, new adjustments in technical thinking and engineering experience are necessary. There is an increasing and extensive interest in the utilization of materials for components to be used at temperatures ranging from -423 to over 3500 deg F. This paper presents a description of the materials problems associated with the various components of chemical liquid rocket systems. These components include cooled and uncooled thrust chambers, injectors, turbine drive systems, propellant tanks, and cryogenic propellant containers. In addition to materials limitations associated with these components, suggested research approaches for improving materials properties are made. Materials such as high-temperature alloys, cermets, carbides, nonferrous alloys, plastics, refractory metals, and porous materials are considered.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA-TM-X-89
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 6
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    Analytical Investigation of the Significance of Turbine-Inlet Temperature in High-Energy Rocket Turbodrive Applications (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The effect of turbine-inlet temperature on rocket gross weight was investigated for three high-energy long-range rockets in order to explore the desirability of turbine cooling in rocket turbodrive applications. Temperatures above and below the maximum that is permissible in uncooled turbines were included. Turbine bleed rate and stage number were considered as independent variables. The gross weight of the hydrogen-reactor system was more sensitive to changes in turbine-inlet temperature than either the hydrogen-oxygen or the hydrogen-fluorine systems. Gross weight of the hydrogen-reactor system could be reduced by 2.6 percent by the use of cooling and a turbine-inlet temperature of 3000 R. The reductions in the first stages of the hydrogen-oxygen and hydrogen-fluorine systems were 0.7 and 0.2 percent, respectively. The effect of turbine-inlet temperature on rocket gross weight was small because the resulting turbine weight and bleed rate variations were small. Since these small gains must be balanced against considerations of greater cost, weight, and complexity as well as lessened reliability with a system utilizing a cooled turbine, none of the systems investigated showed gains warranting the use of turbine cooling.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA-MEMO-1-6-59E
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 7
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    Residual Fuel Expulsion from a Simulated 50,000 Pound Thrust Liquid-Propellant Rocket Engine Having a Continuous Rocket-Type Igniter (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Tests have been conducted to determine the starting characteristics of a 50,000-pound-thrust rocket engine with the conditions of a quantity of fuel lying dormant in the simulated main thrust chamber. Ignition was provided by a smaller rocket firing rearwardly along the center line. Both alcohol-water and anhydrous ammonia were used as the residual fuel. The igniter successfully expelled the maximum amount of residual fuel (3 1/2 gal) in 2.9 seconds when the igniter.was equipped with a sonic discharge nozzle operating at propellant flow rates of 3 pounds per second. Lesser amounts of residual fuel required correspondingly lower expulsion times. When the igniter was equipped with a supersonic exhaust nozzle operating at a flow of 4 pounds per second, a slightly less effective expulsion rate was encountered.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA-MEMO-2-1-59H , H-101
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 8
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    Experimental and Analytical Study of Rolling-Velocity Amplification during the Thrusting Process for Two 10-Inch-Diameter Spherical Rocket Motors in Free Flight (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Two 10-inch-diameter spherical rocket motors have been flight tested at the NASA Wallops Station. These tests were conducted to measure "spin-up" or amplification of the spinning velocity of the motor during the thrusting process due to internal swirling of the exhaust gases. Model 1, a heavy-wall motor, experienced an increase in spin rate during thrusting of about 10 percent, whereas model 2, a flight-type motor with a lightweight motor case, experienced an increase of about 19 percent. The propellant weight and geometry were the same for both motors. A simple relationship for "spin-up" which satisfies these measured results is reported herein. Both models were spin stabilized throughout their flights. A theoretical method of predicting spin-up was derived and used to extend the measured 10-inch-motor results to spherical rocket motors of other sizes having a similar propellant geometry. This method is presented and its predictions are shown to compare favorably with the measured flight results.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA-TM-X-75
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 9
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    A 20,000-Kilowatt Nuclear Turboelectric Power Supply for Manned Space Vehicles (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A conceptual design of a nuclear turboelectric powerplant, producing 20,000 kilowatts of power suitable for manned space vehicles is presented. The study indicates that the radiator necessary for rejecting cycle waste heat is the dominant weight, and emphasis is placed on the selection of cycle operating conditions in order to reduce this weight. A thermodynamic cycle using sodium vapor as the working fluid and operating at a turbine-inlet temperature of 2500 R was selected. The total powerplant weight was calculated to be approximately 6 pounds per kilowatt. The radiator contributes approximately 2.1 pounds per kilowatt to the total weight and the reactor and reactor shield contribute approximately 0.24 and 1.2 pounds per kilowatt, respectively. The generator, turbine, and piping add significantly to the total weight (between 0.5 and 0.6 lb/kw), but the heat exchanger, pumps, and so on are less important. Several important research areas associated with the development of a reliable nuclear turboelectric powerplant of the type analyzed are discussed.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA-MEMO-2-20-59E , E-156
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 10
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    Experimental Altitude Performance of JP-4 Fuel and Liquid-Oxygen Rocket Engine with an Area Ratio of 48 (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The performance for four altitudes (sea-level, 51,000, 65,000, and 70,000 ft) of a rocket engine having a nozzle area ratio of 48.39 and using JP-4 fuel and liquid oxygen as a propellant was evaluated experimentally by use of a 1000-pound-thrust engine operating at a chamber pressure of 600 pounds per square inch absolute. The altitude environment was obtained by a rocket-ejector system which utilized the rocket exhaust gases as the pumping fluid of the ejector. Also, an engine having a nozzle area ratio of 5.49 designed for sea level was tested at sea-level conditions. The following table lists values from faired experimental curves at an oxidant-fuel ratio of 2.3 for various approximate altitudes.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA-MEMO-5-14-59E
    Format: application/pdf
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