SEARCH · Search NASA
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
The design and operating characteristics of valve assemblies used in liquid propellant rocket engines are discussed. The subjects considered are as follows: (1) valve selection parameters, (2) major design aspects, (3) design integration of valve subassemblies, and (4) assembly of components and functional tests. Information is provided on engine, stage, and spacecraft checkout procedures.
A monograph on valves for use with liquid rocket propellant engines is presented. The configurations of the various types of valves are described and illustrated. Design criteria and recommended practices for the various valves are explained. Tables of data are included to show the chief features of valve components in use on operational vehicles.
The research is reported concerning the development of a three-dimensional nonlinear nozzle admittance relation to be used as a boundary condition in the nonlinear combustion instability theories for liquid propellant rocket engines. The derivation of the nozzle wave equation and the application of the Galerkin method are discussed along with the nozzle response.
Explore the source record for details and available documents.
Turbopump assembly for modern liquid propellant rocket engine is complete system in itself. It consists of many components, some of which are themselves subsystems. Monograph deals with turbopump as system, covering selection of proper system type for each application and integration of components into working system.
A simple, semi-empirical performance correlation/prediction technique applicable to gaseous and liquid propellant rocket engines is presented. Excellent correlations were attained for over 100 test firings by adjusting the computation of the gaseous mixing of an unreactive, coaxial jet using a correlation factor, F, which resulted in prediction of the experimental combustion efficiency for each firing. Static pressure, mean velocity and turbulence intensity in the developing region of non-reactive coaxial jets, typical of those of coaxial injector elements were determined. Detailed profiles were obtained at twelve axial locations (extending from the nozzle exit for a distance of five diameters) downstream from a single element of the Bell Aerospace H2/O2 19-element coaxial injector. These data are compared with analytical predictions made using both eddy viscosity and turbulence kinetic energy mixing models and available computer codes. Comparisons were disappointing, demonstrating the necessity of developing improved turbulence models and computational techniques before detailed predictions of practical coaxial free jet flows are attempted.
A study to investigate the hydraulic atomization characteristics of several novel injector designs for use in liquid propellant rocket engines is presented. The injectors were manufactured from a series of thin stainless steel platelets through which orifices were very accurately formed by a photoetching process. These individual platelets were stacked together and the orifices aligned so as to produce flow passages of prescribed geometry. After alignment, the platelets were bonded into a single, 'platelet injector', unit by a diffusion bonding process. Because of the complex nature of the flow associated with platelet injectors, it was necessary to use experimental techniques, exclusively, throughout the study. Large scale models of the injectors were constructed from aluminum plates and the appropriate fluids were modeled using a glycerol-water solution. Stop-action photographs of test configurations, using spark-shadowgraph or stroboscopic back-lighting, are shown.
Accurate predictions of the thrust loss due to boundary layer effects and of the wall heat flux are very important to the design and performance evaluation of rocket nozzles. A method used in analytical procedures for liquid fuel rocket engine performance prediction and evaluation is presented. A computer program is examined that is a fast and accurate procedure for solving the set of boundary layer equation (momentum, energy, and species) for laminar or turbulent, chemically reacting flows with a wide variety of boundary conditions. Results of comparison of the various turbulent models are presented. A summary of the modifications and additions to the program is examined.
A system using a hot-wire transducer as an analog of a liquid droplet of propellant was employed to investigate the ingredients of the acoustic instability of liquid-propellant rocket engines. It was assumed that the combustion process was vaporization-limited and that the combustion chamber was acoustically similar to a closed-closed right-circular cylinder. Before studying the hot-wire closed-loop system (the analog system), a microphone closed-loop system, which used the response of a microphone as the source of a linear feedback exciting signal, was investigated to establish the characteristics of self-sustenance of acoustic fields. Self-sustained acoustic fields were found to occur only at resonant frequencies of the chamber. In the hot-wire closed-loop system, the response of hot-wire anemometer was used as the source of the feedback exciting signal. The self-sustained acoustic fields which developed in the system were always found to be harmonically distorted and to have as their fundamental frquency a resonant frequency for which there also existed a second resonant frequency which was approximately twice the fundamental frequency.
The application of advanced liquid-bipropellant rocket engine analysis techniques has been utilized for prediction of the potential delivered performance and the design of thruster wall cooling schemes for laser-heated rocket thrusters. Delivered specific impulse values greater than 1000 lbf-sec/lbm are potentially achievable based on calculations for thrusters designed for 10-kW and 5000-kW laser beam power levels. A thruster wall-cooling technique utilizing a combination of regenerative cooling and a carbon-seeded hydrogen boundary layer is presented. The flowing carbon-seeded hydrogen boundary layer provides radiation absorption of the heat radiated from the high-temperature plasma. Also described is a forced convection thruster wall cooling design for an experimental test thruster.
Fluid-flow components in a liquid propellant rocket engine and the rocket vehicle which it propels are interconnected by lines, bellows, and flexible hoses. Elements involved in the successful design of these components are identified and current technologies pertaining to these elements are reviewed, assessed, and summarized to provide a technology base for a checklist of rules to be followed by project managers in guiding a design or assessing its adequacy. Recommended procedures for satisfying each of the design criteria are included.
A program for predicting the linear stability of liquid propellant rocket engines is presented. The underlying model assumptions and analytical steps necessary for understanding the program and its input and output are also given. The rocket engine is modeled as a right circular cylinder with an injector with a concentrated combustion zone, a nozzle, finite mean flow, and an acoustic admittance, or the sensitive time lag theory. The resulting partial differential equations are combined into two governing integral equations by the use of the Green's function method. These equations are solved using a successive approximation technique for the small amplitude (linear) case. The computational method used as well as the various user options available are discussed. Finally, a flow diagram, sample input and output for a typical application and a complete program listing for program MODULE are presented.
Practices are outlined for the design, installation, checkout, calibration, and operation of a thrust measurement system to be used during tests of a liquid monopropellant rocket engine. Appendixes include: (1) thrust measurement system elemental uncertainties; (2) short- and long-term thrust measurement system uncertainty; and (3) shunt calibration of force transducers.
Practices are outlined for the design, installation, checkout, calibration, and operation of a system to be used for measuring propellant flow in a liquid monopropellant rocket engine. Design guidelines rather than detailed specifications are provided for the critical components of each portion of the system. System elemental uncertainties are presented in an appendix.
Practices are outlined for the design, installation, checkout, calibration, and operation of a pressure measuring system to be used during tests of a liquid monopropellant rocket engine. Appendixes include: (1) pressure measurement system elemental uncertainties; (2) short- and long-term pressure measurement system uncertainty; (3) shunt calibration of pressure transducers; (4) special considerations for vacuum measurement; and (5) methods of determining the dynamic characteristics of pressure transducers. Design guidelines are provided for the critical components of each portion of the system to provide a pressure measurement system which meets the performance criteria specified.
The design, installation, checkout, calibration, and operation of a temperature measuring system to be used during tests of a liquid monopropellant rocket engine are discussed. Appendixes include: (1) temperature measurement system elemental uncertainties, and (2) tables and equations for use with thermocouples and resistance thermometers. Design guidelines are given for the critical components of each portion of the system to provide an optimum temperature measurement system which meets the performance criteria specified.
The design, installation, checkout, and operation of an exhaust gas composition measurement system for collecting and analyzing the exhaust gas from a liquid monopropellant rocket engine are described. Design guidelines are given for the critical components of each portion of the system to provide an exhaust gas composition measurement which meets the performance criteria specified.
The amplification of a Reynolds number dependent process by wave distortion and the possibility of applying the results to other similar Reynolds number dependent processes were investigated. The process investigated was that associated with the operation of a constant-temperature hot-wire anemometer. The application of vaporization limited combustion, the type of combustion typically associated with liquid propellant rocket engines, was studied. A series of experiments were carried out to determine the effect of wave distortion on a Reynolds number dependent process and to establish the analogy between the anemometer process and the combustion process. Parametric trends, behavior common to different chamber geometries, and stability boundaries were identified. The results indicate a high degree of similarity between the two processes and the possibility of using the anemometer system to investigate combustion instability. The nonlinear aspects of a Reynolds number dependent process appear to be the dominant mechanisms controlling instability.