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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 1,351 records · Page 75
All sourcesNASA NTRSDOE OSTI
NASA NTRS1953

Theoretical performance of liquid ammonia and liquid fluorine as a rocket propellant

Theoretical values of performance parameters for liquid ammonia and liquid fluorine as a rocket propellant were calculated on the assumption of equilibrium composition during the expansion process for a wide range of fuel-oxidant and expansion ratios. The parameters included were specific impulse, combustion chamber temperature, nozzle-exit temperature, equilibrium composition, mean molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, coefficient of viscosity, and coefficient of thermal conductivity. The maximum value of specific impulse was 311.5 pound-seconds per pound for a chamber pressure of 300 pounds per square inch absolute (20.41 atm) and an exit pressure of 1 atmosphere.

COMBUSTION - ROCKET ENGINES↗
NASA NTRS1961

Rocket-UV Spectrometer

This is a report of the second three months (March-May 1961) of a program to design and fabricate an ultraviolet spectrometer and fine guidance control for use in an Aerobee-Hi rocket. The work is being done under Subcontract No. 1 under Contract NASr-3 between Princeton University and the National Aeronautical and Space Administration. The design goals for this instrument are described in the report for the first three month period (Perkin-Elmer Engineering Report No. 5906). The accomplishments of the second three month period include the solution of certain design problems, the establishment of an effective liaison system with personnel at Goddard Space Flight Center, and the completion of certain design and fabrication tasks.

ULTRAVIOLET SPECTROMETER↗
NASA NTRS1961

Effects of Ionizing Radiation on Solid Rocket Motor Components

Problems which the solid propellant rocket engineer will encounter in designing for long-term storage in a radiation environment are discussed. A summary of present knowledge of the radiation environment is given. Mechanisms of radiation degradation and its effects on tensile properties of propellant binders are discussed qualitatively. Data from a program of irradiation of several propellants is presented. Properties of two of the propellants were changed significantly by doses of the order of 4 x 10(exp 6) rads.

RADIATION↗
NASA NTRS1962

Performance Evaluation of Fixed- and Variable-Area Rocket Exhaust Diffusers Using Single and Clustered Nozzles with and Without Gimbaling

An investigation of exhaust diffusers used for altitude simulation in testing rocket engines was conducted with model diffusers and gaseous nitrogen as the working fluid. This investigation was conducted to eval­uate the effects on performance of gimbaling clustered nozzles in a single fixed-area exhaust diffuser. A two-nozzle cluster was gimbaled in all attitudes in four different exhaust-diffuser configurations: (1) a straight circular tube, (2) a straight figure-eight tube, (3) a circular tube with a second throat, and (4) a figure-eight tube with a second throat. Gimbaling clustered nozzles had little effect on the operating pressure ratio for either straight-tube or second-throat diffusers, but caused a large increase in the starting-pressure-ratio requirements in some straight-tube exhaust diffusers. The performance of two or four clustered nozzles with no gimbaling was compared with single-nozzle per­formance on the basis of the ratio of diffuser area to nozzle-throat area. The performance was found to be dependent on the diffuser- to nozzle-throat-area ratio and independent of both the number of primary nozzles and the nozzle-area ratio. A separate investigation was also conducted to evaluate the per­formance improvement obtainable with a variable-area exhaust diffuser. This type of diffuser achieved a 26-percent reduction in starting pressure ratio, a 40-percent reduction in operating pressure ratio, and a 50-percent reduction in the overall diffuser length over the values attainable with a fixed-area second-throat diffuser.

ROCKET↗
NASA NTRS1962

PAYLOAD SEPARATION AND THRUST TERMINATION IN A SOLID-PROPELLANT ROCKET MOTOR

A command method for thrust reduction and/or thrust termination of a solid-propellant rocket motor-called project Shavetail-was demonstrated successfully in flight by Jet Propulsion Laboratory. Thrust reduction is achieved by the separation of a portion of the nozzle to increase the throat area. The chamber pressure is lowered and motor continues to burn at a vernier level of thrust. At any time before motor burnout, upon command from guidance, thrust cancellation and motor separation from the payload are achieved by opening an orifice in the head end of the motor.

VELOCITY↗