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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 685 records · Page 38
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Pulse performance analysis for small hypergolic propellant rocket engines

Small rocket engine tests were conducted for the purpose of obtaining pulse performance data to aid in preliminary design and evaluation of attitude control systems. Both monopropellant and hypergolic bipropellant engines of thrust levels from 5 to 445 N (1 to 100 lb) were tested. The performance data for the hypergolic propellant rockets are compared with theoretical performance calculated from idealized chamber filling and evacuation characteristics. Electromechanical delays in valve response and heat transfer characteristics were found to cause substantial deviation between theoretical performance and test performance. The theoretical analysis is modified to obtain a semiempirical model for hypergolic propellant rockets.

Smith, G. W.↗
NASA NTRS1994

System Sensitivity Analysis Applied to the Conceptual Design of a Dual-Fuel Rocket SSTO

This paper reports the results of initial efforts to apply the System Sensitivity Analysis (SSA) optimization method to the conceptual design of a single-stage-to-orbit (SSTO) launch vehicle. SSA is an efficient, calculus-based MDO technique for generating sensitivity derivatives in a highly multidisciplinary design environment. The method has been successfully applied to conceptual aircraft design and has been proven to have advantages over traditional direct optimization methods. The method is applied to the optimization of an advanced, piloted SSTO design similar to vehicles currently being analyzed by NASA as possible replacements for the Space Shuttle. Powered by a derivative of the Russian RD-701 rocket engine, the vehicle employs a combination of hydrocarbon, hydrogen, and oxygen propellants. Three primary disciplines are included in the design - propulsion, performance, and weights & sizing. A complete, converged vehicle analysis depends on the use of three standalone conceptual analysis computer codes. Efforts to minimize vehicle dry (empty) weight are reported in this paper. The problem consists of six system-level design variables and one system-level constraint. Using SSA in a 'manual' fashion to generate gradient information, six system-level iterations were performed from each of two different starting points. The results showed a good pattern of convergence for both starting points. A discussion of the advantages and disadvantages of the method, possible areas of improvement, and future work is included.

Olds, John R.↗
NASA NTRS2014

A Modular Aerospike Engine Design Using Additive Manufacturing

A modular aerospike engine concept has been developed with the objective of demonstrating the viability of the aerospike design using additive manufacturing techniques. The aerospike system is a self-compensating design that allows for optimal performance over the entire flight regime and allows for the lowest possible mass vehicle designs. At low altitudes, improvements in Isp can be traded against chamber pressure, staging, and payload. In upper stage applications, expansion ratio and engine envelope can be traded against nozzle efficiency. These features provide flexibility to the System Designer optimizing a complete vehicle stage. The aerospike concept is a good example of a component that has demonstrated improved performance capability, but traditionally has manufacturing requirements that are too expensive and complex to use in a production vehicle. In recent years, additive manufacturing has emerged as a potential method for improving the speed and cost of building geometrically complex components in rocket engines. It offers a reduction in tooling overhead and significant improvements in the integration of the designer and manufacturing method. In addition, the modularity of the engine design provides the ability to perform full scale testing on the combustion devices outside of the full engine configuration. The proposed design uses a hydrocarbon based gas-generator cycle, with plans to take advantage of existing powerhead hardware while focusing DDT&E resources on manufacturing and sub-system testing of the combustion devices. The major risks for the modular aerospike concept lie in the performance of the propellant feed system, the structural integrity of the additive manufactured components, and the aerodynamic efficiency of the exhaust flow.

Peugeot, John↗
NASA NTRS1973

Space vehicle engine and heat shield environment review. Volume 1: Engineering analysis

Methods for predicting the base heating characteristics of a multiple rocket engine installation are discussed. The environmental data is applied to the design of adequate protection system for the engine components. The methods for predicting the base region thermal environment are categorized as: (1) scale model testing, (2) extrapolation of previous and related flight test results, and (3) semiempirical analytical techniques.

Mcanelly, W. B.↗
NASA NTRS1973

Space Shuttle Main Engine program status

Performance, weight, and reusability of the Space Shuttle Main Engine, the next generation rocket engine, are important factors in achieving the Space Shuttle mission. The design features of the engine contributing to the achievement of these engine demands are discussed. Many new fabrication methods, material applications, and innovations incorporated into the engine are presented. The engine program is now in the second year of the 47 months development period. The current status of activities in engineering, manufacturing, and test are reviewed. The Design Verification Specification approach to engine development is compared to the development programs for Saturn/Apollo engines.

Kirby, F. M.↗
NASA NTRS1988

Space transportation main engines for two-stage shuttles

Current studies are examining two-stage fully reusable vehicles, heavy-lift vehicles, and rocket engines for these vehicles. A previous study showed that several advanced engine designs could reduce the dry mass of a single-stage vehicle relative to designs with the baseline hydrocarbon and hydrogen engines. This paper reports the results of a similar study for two-stage vehicles. The results indicate that the advanced engines can reduce the dry mass of two-stage vehicles, and that the same engine can be used for both stages.

Martin, James A.↗
NASA NTRS1988

Rocket engine control and monitoring expert system

This paper focuses on the application of expert systems technology to the automatic detection, verification and correction of anomalous rocket engine operations through interfacing with an intelligent adaptive control system. The design of a reliable and intelligent propulsion control and monitoring system is outlined which includes the architecture of an Integrated Expert System (IES) serving as the core component. The IES functions include automatic knowledge acquisition, integrated knowledge base, and fault diagnosis and prediction methodology. The results of fault analysis and diagnostic techniques are presented for an example fault in the SSME main combustion chamber injectors.

Ali, Moonis↗
NASA NTRS1991

Report of the Nuclear Propulsion Mission Analysis, Figures of Merit Subpanel: Quantifiable figures of merit for nuclear thermal propulsion

The results of an inquiry by the Nuclear Propulsion Mission Analysis, Figures of Merit subpanel are given. The subpanel was tasked to consider the question of what are the appropriate and quantifiable parameters to be used in the definition of an overall figure of merit (FoM) for Mars transportation system (MTS) nuclear thermal rocket engines (NTR). Such a characterization is needed to resolve the NTR engine design trades by a logical and orderly means, and to provide a meaningful method for comparison of the various NTR engine concepts. The subpanel was specifically tasked to identify the quantifiable engine parameters which would be the most significant engine factors affecting an overall FoM for a MTS and was not tasked with determining 'acceptable' or 'recommended' values for the identified parameters. In addition, the subpanel was asked not to define an overall FoM for a MTS. Thus, the selection of a specific approach, applicable weighting factors, to any interrelationships, for establishing an overall numerical FoM were considered beyond the scope of the subpanel inquiry.

Haynes, Davy A.↗
NASA NTRS1998

Booster Main Engine Selection Criteria for the Liquid Fly-Back Booster

The Liquid Fly-Back Booster (LFBB) Program seeks to enhance the Space Shuttle system safety, performance and economy of operations through the use of an advanced, liquid propellant Booster Main Engine (BME). There are several viable BME candidates that could be suitable for this application. The objective of this study was to identify the key Criteria to be applied in selecting among these BME candidates. This study involved an assessment of influences on the overall LFBB utility due to variations in the candidate rocket-engines characteristics. This includes BME impacts on vehicle system weight, performance, design approaches, abort modes, margins of safety, engine-out operations, and maintenance and support concepts. Systems engineering analyses and trade studies were performed to identify the LFBB system level sensitivities to a wide variety of BME related parameters. This presentation summarizes these trade studies and the resulting findings of the LFBB design teams regarding the BME characteristics that most significantly affect the LFBB system. The resulting BME choice should offer the best combination of reliability, performance, reusability, robustness, cost, and risk for the LFBB program.

Ryan, Richard M.↗
NASA NTRS1998

Booster Main Engine Selection Criteria for the Liquid Fly-Back Booster

The Liquid Fly-Back Booster (LFBB) Program seeks to enhance the Space Shuttle system safety performance and economy of operations through the use of an advanced, liquid propellant Booster Main Engine (BME). There are several viable BME candidates that could be suitable for this application. The objective of this study was to identify the key criteria to be applied in selecting among these BME candidates. This study involved an assessment of influences on the overall LFBB utility due to variations in the candidate rocket engines' characteristics. This includes BME impacts on vehicle system weight, perfortnance,design approaches, abort modes, margins of safety, engine-out operations, and maintenance and support concepts. Systems engineering analyses and trade studies were performed to identify the LFBB system level sensitivities to a wide variety of BME related parameters. This presentation summarizes these trade studies and the resulting findings of the LFBB design teams regarding the BME characteristics that most significantly affect the LFBB system. The resulting BME choice should offer the best combination of reliability, performance, reusability, robustness, cost, and risk for the LFBB program.

Ryan, Richard M.↗
NASA NTRS2005

REIMR: A Process for Utilizing Propulsion-Oriented 'Lessons-Learned' to Mitigate Development Risk

This paper is a summary overview of a study conducted a t the NASA Marshall Space Flight Center (MSFC) during the initial phases of the Space Launch Initiative (SLI) program to evaluate a large number of technical problems associated with the design, development, test, evaluation and operation of several major liquid propellant rocket engine systems (i.e., SSME, Fastrac, J-2, F-1). The results of this study was the identification of the "Fundamental Root Causes" that enabled the technical problems to manifest, and practices that can be implemented to prevent them from recurring in future engine development efforts. This paper will discus the Fundamental Root Causes, cite some examples of how the technical problems arose from them, and provide a discussion of how they can be mitigated or avoided.

Ballard, Richard O.↗
NASA NTRS2006

REIMR - A Process for Utilizing Liquid Rocket Propulsion-Oriented 'Lessons Learned' to Mitigate Development Risk in Nuclear Thermal Propulsion

This paper is a summary overview of a study conducted at the NASA Marshall Space Flight Center (NASA MSFC) during the initial phases of the Space Launch Initiative (SLI) program to evaluate a large number of technical problems associated with the design, development, test, evaluation and operation of several major liquid propellant rocket engine systems (i.e., SSME, Fastrac, J-2, F-1). One of the primary results of this study was the identification of the Fundamental Root Causes that enabled the technical problems to manifest, and practices that can be implemented to prevent them from recurring in future propulsion system development efforts, such as that which is currently envisioned in the field of nuclear thermal propulsion (NTF). This paper will discuss the Fundamental Root Causes, cite some examples of how the technical problems arose from them, and provide a discussion of how they can be mitigated or avoided in the development of an NTP system

Ballard, RIchard O.↗
NASA NTRSDate not supplied

NASA Progress on the Development and Qualification of a 12-kW Hall-Effect, Solar Electric Propulsion Thruster

I. Motivation and Background Beginning in 2014, the National Aeronautics and Space Administration (NASA) Space Technology Mission Directorate (STMD) began a project to increase the state of the art for the Hall-Effect Solar Electric Propulsion (SEP) technology. The resulting Advanced Electric Propulsion System (AEPS) project has developed a 12 kW Hall Current Thruster in support of the NASA mission to establish a permanent human presence in lunar orbit and to land the next American astronauts on the South Pole of the Moon. The project is led by the NASA Glenn Research Center, supported by the Jet Propulsion Laboratory and development, qualification & flight hardware all provided by L3 Harris Aerojet Rocketdyne (AR).The AEPS project has completed the development testing of a high power, solar electric propulsion Hall Current thruster that will be used on the NASA Power & Propulsion Element (PPE) of the Gateway space station. NASA initially built three Technology Development Units to understand key characteristics of the hall-effect rocket with magnetic shielding. The design led to development testing on two Engineering Test Unit Thrusters and multiple critical components. The project has begun production of the three flight thrusters and entered qualification testing at the component and thruster levels. II. Approach NASA and AR teams completed all development phases of the project, including full development and integration testing of the Engineering Model hardware, Critical Design Review, and ground test equipment validation, as well as fabrication and acceptance testing of the initial qualification thruster. Qualification and verification of the environmental and life requirements of the AEPS design was initiated in the Fall of 2023 and will be accomplished on two thruster units and using a series of component-level tests during 2024 and 2025. Environmental testing will incorporate functional reference firings, shock, vibration, and Thermal Vacuum (TVAC) testing. Life verification will assess the thruster wear and performance over the lifetime of the Gateway spacecraft. Critical component qualification tests include cathode heater, magnet coils, magnet heaters, temperature sensors, and a cathode assembly that will undergo life cycle testing on multiple units. Flight thrusters will complete assembly and acceptance testing and be delivered to the PPE program in early 2025. III. Preliminary and Anticipated Results The program has completed the acceptance testing, including dynamic testing and hot fire characterization, of the first qualification thruster. In the Fall of 2023, the program entered the environmental qualification phase for thruster testing. This paper will present an overview of the AEPS thruster project, thruster capabilities and flight design, preliminary results from the thruster acceptance and qualification testing, component life cycle testing and flight hardware status.

Clayton Kachele↗
NASA NTRSDate not supplied

High Performance Rotating Detonation Rocket Engine for Mars Interplanetary Transport

NASA has successfully evaluated the performances and survivability of two additively manufactured (AM) continuously rotating detonation cycle rocket engine (RDRE) thrust chambers. The AM hardware was subjected to long duration continuous detonative environments in excess of 100 seconds. Performance trends were directly compared to theory and similar scale constant pressure engines with identical propellants. The performance benefits of RDREs include higher combustion efficiency at a more compact design trade space, which allow advantages for future NASA missions. A major finding of this work is that the heat fluxes experienced are substantially higher than traditional liquid rocket engines at the same operating conditions. In addition, completion of combustion occurs rapidly in an annular geometry at ¼ that of the L* and L’ of an equivalent state-of-the-art thrust chamber. Finally, all hardware survived the extreme environments and achieved the major goal of the test project by demonstrating hot fire conditions with detonation modes up to 133 seconds in duration. Follow-on efforts funded by STMD Early Career Initiative (ECI) award are now focused on achieving higher performances than the previous designs and closing the remaining technology risks.

Thomas Teasley↗
NASA NTRS1972

Experimental investigation of combustor effects on rocket thrust chamber performance

A design and experimental program to develop special instrumentation systems, design engine hardware, and conduct tests using LOX/GH2 propellants in which the propellant flow stratification was controlled is described. The mixture ratio was varied from 4.6 to 6 overall. The mixture ratios in the core and outer zone were varied from 3.5 to 6 and 5 to 8, respectively. The range in boundary layer coolant was from 0 to 10 percent of the fuel. The nominal chamber pressure and thrust were 225 psia and 7000 pounds, respectively. Pressure and heat flux profiles as well as gas sampling of the exhaust products were obtained. Specific impulse efficiencies of approximately 94 percent and characteristic velocity efficiencies of approximately 97 percent were obtained during the experiments.

Source record↗
NASA NTRS2000

Global Design Optimization for Aerodynamics and Rocket Propulsion Components

Modern computational and experimental tools for aerodynamics and propulsion applications have matured to a stage where they can provide substantial insight into engineering processes involving fluid flows, and can be fruitfully utilized to help improve the design of practical devices. In particular, rapid and continuous development in aerospace engineering demands that new design concepts be regularly proposed to meet goals for increased performance, robustness and safety while concurrently decreasing cost. To date, the majority of the effort in design optimization of fluid dynamics has relied on gradient-based search algorithms. Global optimization methods can utilize the information collected from various sources and by different tools. These methods offer multi-criterion optimization, handle the existence of multiple design points and trade-offs via insight into the entire design space, can easily perform tasks in parallel, and are often effective in filtering the noise intrinsic to numerical and experimental data. However, a successful application of the global optimization method needs to address issues related to data requirements with an increase in the number of design variables, and methods for predicting the model performance. In this article, we review recent progress made in establishing suitable global optimization techniques employing neural network and polynomial-based response surface methodologies. Issues addressed include techniques for construction of the response surface, design of experiment techniques for supplying information in an economical manner, optimization procedures and multi-level techniques, and assessment of relative performance between polynomials and neural networks. Examples drawn from wing aerodynamics, turbulent diffuser flows, gas-gas injectors, and supersonic turbines are employed to help demonstrate the issues involved in an engineering design context. Both the usefulness of the existing knowledge to aid current design practices and the need for future research are identified.

Shyy, Wei↗