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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 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Alternate vehicles for engine/vehicle optimization

The space shuttle main engine, full flow staged combustion, integrated modular engine concept, split expander, and expander bleed rocket engine cycles were studied for use on Single-Stage-to-Orbit fully reusable vehicles. The vehicle uses vertical take-off and horizontal landing and liquid hydrogen rockets for propulsion. The PROPSIZE computer code was modified for use on the available computers. Vehicle sizing was done after trajectory optimization for various engine designs. The various rocket engine cycles were compared and it was found that the full flow staged combustion cycle had the lowest vehicle dry mass. The split expander cycle could have advantages in safety and cooling which could make it the preferred choice in some cases. Comparison to previous work with two-stage heavy-lift vehicles showed that the same ranking of the cycles was found for both vehicle concepts.

Venkatasubramanyam, G.↗
NASA NTRS2000

Engine System Loads Development for the Fastrac 60K Flight Engine

Early implementation of structural dynamics finite element analyses for calculation of design loads is considered common design practice for high volume manufacturing industries such as automotive and aeronautical industries. However, with the rarity of rocket engine development programs starts, these tools are relatively new to the design of rocket engines. In the new Fastrac engine program, the focus has been to reduce the cost to weight ratio; current structural dynamics analysis practices were tailored in order to meet both production and structural design goals. Perturbation of rocket engine design parameters resulted in a number of Fastrac load cycles necessary to characterize the impact due to mass and stiffness changes. Evolution of loads and load extraction methodologies, parametric considerations and a discussion of load path sensitivities are discussed.

Frady, Greg↗
NASA NTRS1988

Space Software for Automotive Design

John Thousand of Wolverine Western Corp. put his aerospace group to work on an unfamiliar job, designing a brake drum using computer design techniques. Computer design involves creation of a mathematical model of a product and analyzing its effectiveness in simulated operation. Technique enables study of performance and structural behavior of a number of different designs before settling on a final configuration. Wolverine employees attacked a traditional brake drum problem, the sudden buildup of heat during fast and repeated braking. Part of brake drum not confined tends to change its shape under combination of heat, physical pressure and rotational forces, a condition known as bellmouthing. Since bellmouthing is a major factor in braking effectiveness, a solution of problem would be a major advance in automotive engineering. A former NASA employee, now a Wolverine employee, knew of a series of NASA computer programs ideally suited to confronting bellmouthing. Originally developed as aids to rocket engine nozzle design, it's capable of analyzing problems generated in a rocket engine or automotive brake drum by heat, expansion, pressure and rotational forces. Use of these computer programs led to new brake drum concept featuring a more durable axle, and heat transfer ribs, or fins, on hub of drum.

Source record↗
NASA NTRS2002

Engine System Loads Analysis Compared to Hot-Fire Data

Early implementation of structural dynamics finite element analyses for calculation of design loads is considered common design practice for high volume manufacturing industries such as automotive and aeronautical industries. However with the rarity of rocket engine development programs starts, these tools are relatively new to the design of rocket engines. In the NASA MC-1 engine program, the focus was to reduce the cost-to-weight ratio. The techniques for structural dynamics analysis practices, were tailored in this program to meet both production and structural design goals. Perturbation of rocket engine design parameters resulted in a number of MC-1 load cycles necessary to characterize the impact due to mass and stiffness changes. Evolution of loads and load extraction methodologies, parametric considerations and a discussion of load path sensitivities are important during the design and integration of a new engine system. During the final stages of development, it is important to verify the results of an engine system model to determine the validity of the results. During the final stages of the MC-1 program, hot-fire test results were obtained and compared to the structural design loads calculated by the engine system model. These comparisons are presented in this paper.

Frady, Gregory P.↗
NASA NTRS1993

Design and test of a small two stage counter-rotating turbine for rocket engine application

The aerodynamic design and rig test evaluation of a small counter-rotating turbine system is described. The technology represented by this turbine is being developed for application in an advanced upper stage rocket engine turbopump. This engine will employ an oxygen/hydrogen expander cycle and achieve high performance through efficient combustion, high combustion pressure, and high area ratio exhaust nozzle expansion. Engine performance goals require that the turbopump drive turbines achieve high efficiency at low gas flow rates. The low flow rates result in very small airfoil diameter, height and chord. The high efficiency and small size requirements present a challenging turbine design problem. The unconventional approach employed to meet this challenge is described, along with the detailed design process and resulting airfoil configurations. The method and results of full scale aerodynamic performance evaluation testing of both one and two stage configurations, as well as operation without the secondary stage stator are presented. The overall results of this effort illustrate that advanced aerodynamic design tools and hardware fabrication techniques have provided improved capability to produce small high performance turbines for advanced rocket engines.

Huber, F. W.↗
NASA NTRS1973

Liquid rocket engine centrifugal flow turbopumps

Design criteria and recommended practices are discussed for the following configurations selected from the design sequence of a liquid rocket engine centrifugal flow turbopump: (1) pump performance including speed, efficiency, and flow range; (2) impeller; (3) housing; and (4) thrust balance system. Hydrodynamic, structural, and mechanical problems are addressed for the achievement of required pump performance within the constraints imposed by the engine/turbopump system. Materials and fabrication specifications are also discussed.

Source record↗
NASA NTRS1991

Space engine safety system

A rocket engine safety system was designed to initiate control procedures to minimize damage to the engine or vehicle or test stand in the event of an engine failure. The features and the implementation issues associated with rocket engine safety systems are discussed, as well as the specific concerns of safety systems applied to a space-based engine and long duration space missions. Examples of safety system features and architectures are given, based on recent safety monitoring investigations conducted for the Space Shuttle Main Engine and for future liquid rocket engines. Also, the general design and implementation process for rocket engine safety systems is presented.

Maul, William A.↗
NASA NTRS1975

A fast sampling device for the mass spectrometric analysis of liquid rocket engine exhaust

The design of a device to obtain compositional data on rocket exhaust by direct sampling of reactive flow exhausts into a mass spectrometer is presented. Sampling at three stages differing in pressure and orifice angle and diameter is possible. Results of calibration with pure gases and gas mixtures are erratic and of unknown accuracy for H2, limiting the usefulness of the apparatus for determining oxidizer/fuel ratios from combustion product analysis. Deposition effects are discussed, and data obtained from rocket exhaust spectra are analyzed to give O/F ratios and mixture ratio distribution. The O/F ratio determined spectrometrically is insufficiently accurate for quantitative comparison with cold flow data. However, a criterion for operating conditions with improved mixing of fuel and oxidizer which is consistent with cold flow results may be obtained by inspection of contour plots. A chemical inefficiency in the combustion process when oxidizer is in excess is observed from reactive flow measurements. Present results were obtained with N2O4/N2H4 propellants.

Ryason, P. R.↗
NASA NTRS1991

Space engine safety system

A rocket engine safety system is designed to initiate control procedures which will minimize damage to the engine and vehicle or test stand in the event of an engine failure. This report describes the features and the implementation issues associated with rocket engine safety systems. Specific concerns of safety systems applied to a space-based engine and long duration space missions are discussed. Examples of safety system features and architectures are given from recent safety monitoring investigations conducted for the Space Shuttle Main Engine and for future liquid rocket engines. Also, a general design and implementation process for rocket engine safety systems is presented.

Maul, William A.↗
NASA NTRS1973

An injector design model for predicting rocket engine performance and heat transfer

A model is formulated for estimating the performance and chamber heat transfer in rocket injectors/chambers operating with gaseous H2-O2 propellants. The model quantifies the combustion performance and chamber heat flux for variables such as chamber length, element type, element area ratio, impingement angle, thrust/element, mixture ratio, moment ratio, element spacing, and physical size. Design equations are given and curves are plotted for evaluation of combustion performance in injectors comprised of F-O-F triplet, premix, coaxial and swirl coaxial element types. Curve plots and equations are also included for estimation of the chamber wall heat fluxes generated by these element types.

Calhoon, D. F.↗
NASA NTRS1974

Liquid rocket engine turbines

Criteria for the design and development of turbines for rocket engines to meet specific performance, and installation requirements are summarized. The total design problem, and design elements are identified, and the current technology pertaining to these elements is described. Recommended practices for achieving a successful design are included.

Source record↗
NASA NTRS1988

3-D thermal analysis using finite difference technique with finite element model for improved design of components of rocket engine turbomachines for Space Shuttle Main Engine SSME

Three-dimensional finite element models were generated and transferred into three-dimensional finite difference models to perform transient thermal analyses for the SSME high pressure fuel turbopump's first stage nozzles and rotor blades. STANCOOL was chosen to calculate the heat transfer characteristics (HTCs) around the airfoils, and endwall effects were included at the intersections of the airfoils and platforms for the steady-state boundary conditions. Free and forced convection due to rotation effects were also considered in hollow cores. Transient HTCs were calculated by taking ratios of the steady-state values based on the flow rates and fluid properties calculated at each time slice. Results are presented for both transient plots and three-dimensional color contour isotherm plots; they were also converted into universal files to be used for FEM stress analyses.

Sohn, Kiho D.↗
NASA NTRS2001

Multiobjective Optimization of Rocket Engine Pumps Using Evolutionary Algorithm

A design optimization method for turbopumps of cryogenic rocket engines has been developed. Multiobjective Evolutionary Algorithm (MOEA) is used for multiobjective pump design optimizations. Performances of design candidates are evaluated by using the meanline pump flow modeling method based on the Euler turbine equation coupled with empirical correlations for rotor efficiency. To demonstrate the feasibility of the present approach, a single stage centrifugal pump design and multistage pump design optimizations are presented. In both cases, the present method obtains very reasonable Pareto-optimal solutions that include some designs outperforming the original design in total head while reducing input power by one percent. Detailed observation of the design results also reveals some important design criteria for turbopumps in cryogenic rocket engines. These results demonstrate the feasibility of the EA-based design optimization method in this field.

Oyama, Akira↗
NASA NTRS1986

Cryogenic gear technology for an orbital transfer vehicle engine and tester design

Technology available for gears used in advanced Orbital Transfer Vehicle rocket engines and the design of a cryogenic adapted tester used for evaluating advanced gears are presented. The only high-speed, unlubricated gears currently in cryogenic service are used in the RL10 rocket engine turbomachinery. Advanced rocket engine gear systems experience operational load conditions and rotational speed that are beyond current experience levels. The work under this task consisted of a technology assessment and requirements definition followed by design of a self-contained portable cryogenic adapted gear test rig system.

Calandra, M.↗
NASA NTRS1966

Commerical application of rocket engine turbopump technology

This paper describes how commercial pump technology was modified for rocket engine application by trading long life for low weight. The component life of typical turbopump components is presented, and the modes of failure will be discussed. A comparison of the basic commercial and rocket engine pump requirements such as flow, pressure, etc. is made. It is concluded that rocket engine pump technology can be applied directly only if similar basic requirements exist in emergency or stand-by equipment. The specific design requirements for rocket engine turbopumps are presented and the need for high speeds is explained by the weight-speed relationship and the optimum NPSH of the tank and pump system. The trend of required rocket engine pump discharge pressure and future operational requirements is predicted. The need for inducers and boost pumps in rocket engines is explained, and the results of the flat-plate-inducer cavitation therein are briefly presented. Design solutions and material selections for high-pressure pump housings, ultra-high tip speed impellers, and pump packaging are demonstrated. As a typical example of utilizing rocket engine technology for commercial application, a helicopter stand-by engine i shown. This standby engine is designed from existing rocket engine components, using JATO rockets as the prime energy source.

Turbopump↗