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  • Other Sources  (22)
  • Spacecraft Propulsion and Power  (14)
  • STRUCTURAL MECHANICS  (8)
  • 1955-1959  (22)
  • 1959  (22)
  • 1
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    PRESS CONFERENCE- EXPLORER VII PROJECT BACKGROUND AND EXPERIMENTS (1959)
    In:  CASI
    Details
    Publication Date: 2006-10-26
    Keywords: STRUCTURAL MECHANICS
    Format: text
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  • 2
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    THERMAL STRESSES IN CIRCULAR DISCS. (1959)
    In:  CASI
    Details
    Publication Date: 2006-03-16
    Description: Thermal stress in circular discs
    Keywords: STRUCTURAL MECHANICS
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  • 3
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    Experimental and calculated supersonic flutter characteristics of models of the x-15 horizontal and vertical tails (1959)
    In:  CASI
    Details
    Publication Date: 2019-05-23
    Description: Experimental and calculated supersonic flutter characteristics of X-15 horizontal and vertical tails
    Keywords: STRUCTURAL MECHANICS
    Type: NASA-TM-X-176
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  • 4
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    Some measurements of the dynamic and static stability of two blunt-nosed, low-fineness- ratio bodies of revolution in free flight at mequal4 (1959)
    In:  CASI
    Details
    Publication Date: 2019-05-23
    Description: Dynamic and static stability of two blunt nosed low fineness ratio bodies of revolution in free flight - ballistics
    Keywords: STRUCTURAL MECHANICS
    Type: NASA-TM-X-20
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  • 5
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    The static and dynamic-rotary stability derivatives of a model of the x-15 research airplane at mach numbers from 1.55 to 3.50 (1959)
    In:  CASI
    Details
    Publication Date: 2019-05-23
    Description: Static and dynamic rotary stability derivatives for X-15 aircraft at supersonic speeds
    Keywords: STRUCTURAL MECHANICS
    Type: NASA-MEMO-12-23-58A
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  • 6
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    CONFERENCE REPORT, MEETINGS OF FEB. 6-7, 1959, NASA HEADQUARTERS WASHINGTON, D.C. VEGA MANAGEMENT DOCUMENT NO. 1 (1959)
    In:  CASI
    Details
    Publication Date: 2019-05-10
    Keywords: STRUCTURAL MECHANICS
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  • 7
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    Flutter Investigation at a Mach Number of 7.2 of Models of the Horizontal-tail and Vertical-tail Surfaces of the X-15 Airplane (1959)
    In:  CASI
    Details
    Publication Date: 2019-05-30
    Description: Flutter studies at Mach 7.2 of horizontal and vertical tail surface models of X-15 aircraft
    Keywords: STRUCTURAL MECHANICS
    Type: NASA-MEMO-4-14-59L
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  • 8
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    The static stability characteristics of several preliminary models of the x-15 research airplane at mach numbers of 2.98 and 4.01 (1959)
    In:  CASI
    Details
    Publication Date: 2019-05-23
    Description: Static stability characteristics of preliminary models of X-15 research aircraft at supersonic speeds
    Keywords: STRUCTURAL MECHANICS
    Type: NASA-TM-X-166
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  • 9
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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
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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
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  • 11
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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
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  • 12
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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
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  • 13
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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
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  • 14
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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
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  • 15
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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
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  • 16
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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
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  • 17
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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
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  • 18
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    Design of Axisymmetric Exhaust Nozzles by Method of Characteristics Incorporating a Variable Isentropic Exponent (1959)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: Limitations on fully developed laminar flows due to compressibility and property variations are examined. The cases, for liquids and for gases, wherein such motions are "exact" are determined and solutions are given. For more general conditions, not permitting an exact fully developed flow, limitations are set. Two cases arise depending on the size of the temperature variation across the channel. Both the forced and free flow are solved for the case of large temperature variation. Finally, there are described briefly some circumstances under which streamwise variations of velocity occur. The case where the velocity varies inversely with the square root of the distance is solved.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA-TR-R-33
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  • 19
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    Analysis of of Injection-Velocity Effects on Rocket Motor Dynamics and Stability (1959)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: A concept of combustion time lag that includes dependency on injection velocity is introduced. The concept is used in the formulation of chamber transfer functions and in an analysis of low-frequency combustion instability. Theoretical frequency responses and stability boundaries are compared with those obtained when the injection-velocity effect on the time lag to be an important consideration, in the theory of chamber dynamics and combustion instability
    Keywords: Spacecraft Propulsion and Power
    Type: NASA-TR-R-43
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  • 20
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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
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  • 21
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    Comparison of Rocket Performance using Exhaust Diffuser and Conventional Techniques for Altitude Simulation (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A rocket engine with an exhaust-nozzle area ratio of 25 was operated at a constant chamber pressure of 600 pounds per square inch absolute over a range of oxidant-fuel ratios at an altitude pressure corresponding to approximately 47,000 feet. At this condition, the nozzle flow is slightly underexpanded as it leaves the nozzle. The altitude simulation was obtained first through the use of an exhaust diffuser coupled with the rocket engine and secondly, in an altitude test chamber where separate exhauster equipment provided the altitude pressure. A comparison of performance data from these two tests has established that a diffuser used with a rocket engine operating at near-design nozzle pressure ratio can be a valid means of obtaining altitude performance data for rocket engines.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA-TM-X-100
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  • 22
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    Analytical Investigation of the Effect of Turbopump Design on Gross-Weight Characteristics of a Hydrogen-Propelled Nuclear Rocket (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The effect of turbopump design on rocket gross weight was investigated for a high-pressure bleed-type hydrogen-reactor long-range rocket with a fixed mission. Axial-flow, mixed-flow, and centrifugal pumps driven by single and twin turbines were considered. With an efficiency of 0.7 assumed for all pumps, the lowest rocket gross weights were obtained with an axial-flow or a mixed-flow pump driven by a single turbine of at least eight stages. All turbopump combinations could be used, however, with gross weight varying less than 8 percent for a given payload. Turbopump efficiencies have a significant effect on the ratio of gross weight to payload with the magnitude of the effect determined by the ratio of rocket structural weight to total propellant weight. One point in pump efficiency is worth 0.2 percent in gross weight for a given payload with a structural weight parameter of 0.1 and 0.6 percent with a structural weight parameter of 0.2. Turbine and pump weights are much less significant in terms of gross-to-pay weight ratio than the efficiencies of these components. One point in pump efficiency is equivalent to approximately 13 percent in pump weight, while 1 point in turbine efficiency is equivalent to about 7 percent in turbine weight.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA-MEMO-5-12-59E , E-215
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