Experimental performance of a high-area-ratio rocket nozzle at high combustion chamber pressure
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Experimental performance of a high-area-ratio rocket nozzle at high combustion chamber pressure

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Published by National Aeronautics and Space Administration, Office of Management, Scientific and Technical Information Program, National Technical Information Service, distributor in [Washington, D.C.], [Springfield, Va .
Written in English

Subjects:

  • Rocket nozzles.,
  • Exhaust velocity.,
  • Specific impulse.,
  • Nozzle thrust coefficients.,
  • Nozzle geometry.,
  • Thrust chamber pressure.,
  • Nozzle efficiency.,
  • Divergent nozzles.

Book details:

Edition Notes

Other titlesExperimental performance of a high area ratio rocket nozzle at high combustion chamber pressure
StatementRobert S. Jankovsky and John M. Kazaroff, Albert J. Pavli.
SeriesNASA techical paper -- 3576, NASA technical paper -- 3576.
ContributionsKazaroff, John M., Pavli, Albert J., United States. National Aeronautics and Space Administration. Scientific and Technical Information Program
The Physical Object
FormatMicroform
Pagination1 v.
ID Numbers
Open LibraryOL17126935M

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Experimental data were obtained on an optimally contoured nozzle with an area ratio of and on a truncated version of this nozzle with an area ratio of. High-Performance Rocket Nozzle Concept Article (PDF Available) in Journal of Propulsion and Power 26(5) September with 2, Reads How we measure 'reads'. December version of the TDK program for the high- area-ratio rocket engine regime. The calibration was accom- plished by modeling the performance of a 1 rocket nozzle tested at NASA Lewis Research Center. A detailed description of the test conditions and TDK input parameters is given. the chamber pressure and chamber mixture ratio, and the wall temperature. The nozzle ambient pressure was used as a boundary condition when the normalized contravari-ent velocity was less than the speed of sound. At the noz-zle inlet, the flow was assumed to be in chemical equilibrium. Constant reference-plane characteristic relations wereFile Size: 1MB.

The engine's nozzle is in ( m) long with a diameter of inches ( m) at its throat and inches ( m) at its exit. The nozzle is a bell-shaped extension bolted to the main combustion chamber, referred to as a de Laval RS nozzle has an unusually large expansion ratio (about ) for the chamber pressure. At sea level, a nozzle of this ratio would normally Associated L/V: Space Shuttle, Space Launch System. only possible with high combustion chamber pressures () Advanced Rocket Engines. In A rather interesting aspect about the impact of the combustion chamber pressure on the : Oskar Haidn.   Reacting flow simulation of high area ratio rocket nozzles is done using an indigenously developed Point Implicit Unstructured Finite Volume Solver. A numerical solution procedure to solve turbulent-reacting nozzle flow field is developed, which is based on the implicit treatment of chemical source terms by preconditioning and then explicitly Author: S. Shyji, N. Asok Kumar, T. Jayachandran, M. Deepu.   Effects of Nozzle Throat and Combustion Chamber Design Variables on Divergent Portion of the Nozzle. Performance of high-area-ratio nozzle for a small rocket thruster. Spike nozzle contour for optimum thrust. Planetary and Space Science, Vol. 4. Book by:

was low for a nozzle expansion ratio of 12–35 using air as the working fluid [13]. In the present work, the feasibility of an arc-based design method for generation of high-area-ratio nozzle contours has been determined ijusing a numerical approach iacross jthe range of flow regimes expected to occur in a core stage rocket engine. Viability ofCited by: 5. Recent hot- ring tests performed in Russia with a modi ed RD engine, equipped with a secondary nozzle insert, revealed a signi cant performance gain of 12% during the sealevel operation at % chamber pressure, compared with the original RD performance Nominal chamber pressure of this engine is p c = bar, with an area ratio of. Full text of "Theoretical and Experimental Investigations of Ignition, Combustion and Expansion Processes of Hypergolic Liquid Fuel Combinations at Gas Temperatures up to K" See other formats. A solid-propellant rocket or solid rocket is a rocket with a rocket engine that uses solid propellants (fuel/oxidizer).The earliest rockets were solid-fuel rockets powered by gunpowder; they were used in warfare by the Chinese, Indians, Mongols and Persians, as early as the 13th century.. All rockets used some form of solid or powdered propellant up until the 20th century, when liquid.