Application of the MIRA 150A variable thrust rocket engine to manned lunar exploration flying systems.
MIRA 150A variable thrust rocket engine applied to manned lunar exploration flying systems
ENGINE TESTING↗
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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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Low thrust rocket test facility
A low thrust chemical rocket test facility has recently become operational at the NASA-Lewis. The new facility is used to conduct both long duration and performance tests at altitude over a thruster's operating envelope using hydrogen and oxygen gas for propellants. The facility provides experimental support for a broad range of objectives, including fundamental modeling of fluids and combustion phenomena, the evaluation of thruster components, and life testing of full rocket designs. The major mechanical and electrical systems are described along with aspects of the various optical diagnostics available in the test cell. The electrical and mechanical systems are designed for low down time between tests and low staffing requirements for test operations. Initial results are also presented which illustrate the various capabilities of the cell.
Arrington, Lynn A.↗
Low thrust rocket test facility
A low thrust chemical rocket test facility has recently become operational at the NASA-Lewis. The new facility is used to conduct both long duration and performance tests at altitude over a thruster's operating envelope using hydrogen and oxygen gas for propellants. The facility provides experimental support for a broad range of objectives, including fundamental modeling of fluids and combustion phenomena, the evaluation of thruster components, and life testing of full rocket designs. The major mechanical and electrical systems are described along with aspects of the various optical diagnostics available in the test cell. The electrical and mechanical systems are designed for low down time between tests and low staffing requirements for test operations. Initial results are also presented which illustrate the various capabilities of the cell.
Arrington, Lynn A.↗
Orbital transfer rocket engine technology 7.5K-LB thrust rocket engine preliminary design
A preliminary design of an advanced LOX/LH2 expander cycle rocket engine producing 7,500 lbf thrust for Orbital Transfer vehicle missions was completed. Engine system, component and turbomachinery analysis at both on design and off design conditions were completed. The preliminary design analysis results showed engine requirements and performance goals were met. Computer models are described and model outputs are presented. Engine system assembly layouts, component layouts and valve and control system analysis are presented. Major design technologies were identified and remaining issues and concerns were listed.
Harmon, T. J.↗
Transient and steady performance of an ion rocket thrust augmentor Summary report
Transient and steady performance of ion rocket thrust augmentor
IONIC PROPULSION↗
Rocket thrust chamber thermal barrier coatings
A research program was conducted to generate data and develop analytical techniques to predict the performance and reliability of ceramic thermal barrier coatings in high heat flux environments. A finite element model was used to analyze the thermomechanical behavior of coating systems in rocket thrust chambers. Candidate coating systems (using a copper substrate, NiCrAlY bond coat and ZrO2.8Y2O3 ceramic overcoat) were selected for detailed study based on photomicrographic evaluations of experimental test specimens. The effects of plasma spray application parameters on the material properties of these coatings were measured and the effects on coating performance evaluated using the finite element model. Coating design curves which define acceptable operating envelopes for seleted coating systems were constructed based on temperature and strain limitations. Spray gun power levels was found to have the most significant effect on coating structure. Three coating systems were selected for study using different power levels. Thermal conductivity, strain tolerance, density, and residual stress were measured for these coatings. Analyses indicated that extremely thin coatings ( 0.02 mm) are required to accommodate the high heat flux of a rocket thrust chamber and ensure structural integrity.
Batakis, A. P.↗
A STUDY OF INJECTION PROCESSES FOR 15-PERCENT FLUORINE-85- PERCENT OXYGEN AND HEPTANE IN A 200-POUND-THRUST ROCKET ENGINE
Injection techniques for fluorine, oxygen, and heptane in 200-pound thrust rocket engine
FUEL INJECTION↗
Perturbation solutions for low-thrust rocket trajectories.
Perturbation solutions of differential equations of motion for low-thrust rocket trajectories
PERTURBATION THEORY↗
Applications of a mathematical programming technique to finite thrust rocket trajectory optimization
Iterative solution of least squares programming problem for finite thrust rocket trajectory optimization
Armstrong, E. S.↗
Three dimensional thermal analysis of rocket thrust chambers
A numerical model for the three dimensional thermal analysis of rocket thrust chambers and nozzles has been developed. The input to the model consists of the composition of the fuel/oxidant mixture and flow rates, chamber pressure, coolant entrance temperature and pressure, dimensions of the engine, materials and the number of nodes in different parts of the engine. The model allows for temperature variation in three dimensions: axial, radial and circumferential directions and by implementing an iterative scheme, it provides nodal temperature distribution, rates of heat transfer, hot gas and coolant thermal and transport properties.
Naraghi, M. H. N.↗
EXPERIMENTAL INVESTIGATION OF HEAT TRANSFER RATES IN ROCKET THRUST CHAMBERS
Analysis of heat transfer in rocket engine thrust chambers that use convergent-divergent nozzles
ROCKET ENGINE↗
Some factors influencing heat transfer to liquid-propellant rocket-thrust-chamber coolant channels
Wall temperature and heat transfer patterns in liquid-propellant rocket thrust chamber coolant channels
LIQUID PROPELLANT ROCKET ENGINE↗
Some numerical comparisons of three-body trajectories with patched two body trajectories for low thrust rockets
Comparisons of three body trajectories with patched two body trajectories for low thrust rockets
Strack, W. C.↗
Three dimensional thermal analysis of rocket thrust chambers
A numerical model for the three-dimensional thermal analysis of rocket thrust chambers and nozzles has been developed. The input to the model consists of the composition of the fuel/oxidant mixture and flow rates, chamber pressure, coolant entrance temperature and pressure, dimensions of the engine, materials and the number of nodes in different parts of the engine. The model allows for temperature variation in three dimensions: axial, radial and circumferential directions and by implementing an iterative scheme, it provides nodal temperature distribution, rates of heat transfer, hot gas and coolant thermal and transport properties.
Naraghi, M. H. N.↗
Evaluation of Injector Principles in a 2400-pound-thrust Rocket Engine Using Liquid Oxygen and Liquid Ammonia
Injector evaluation in 2400-pound-thrust rocket engine using liquid oxygen and liquid ammonia
ROCKET ENGINE↗
A two-dimensional finite difference program for thermal analysis of rocket thrust chambers
A two-dimensional finite difference computer model for thermal analysis of rocket thrust chambers has been developed. The model uses an iterative scheme for calculating the temperature distribution within the chamber wall and implements a successive overrelaxation formula for a quick convergence. The inputs of the models are the dimensions of the thrust chamber wall, types of materials used, heat transfer coefficients and temperatures of the hot gas and the coolant. The resulting output of the program consists of the nodal temperature distribution, heat transfer to the coolant and heat transfer from the hot gas.
Naraghi, Mohammad H.↗
Heat pipe technology for advanced rocket thrust chambers Interim report, 3 Jan. 1969 - 2 May 1970
Heat pipe technology for advanced rocket thrust chambers
Rousar, D. C.↗
Some effects of cyclic induced deformation in rocket thrust chambers
A test program to investigate the deformation process observed in the hot gas wall of rocket thrust chambers was conducted using three different liner materials. Five thrust chambers were cycled to failure using hydrogen and oxygen as propellants at a chamber pressure of 4.14 MN/m square (600 psia). The deformation was observed nondestructively at midlife points and destructively after failure occurred. The cyclic life results are presented with an accompanying discussion about the types of failure encountered. Data indicating the deformation of the thrust chamber liner as cycles are accumulated are presented for each of the test thrust chambers.
Hannum, N. P.↗