SEARCH · Search NASA

Results for “Engineer”

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.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 1,333 records · Page 74

All sources NASA NTRS DOE OSTI

NASA NTRS1995

PVD TBC experience on GE aircraft engines

The higher performance levels of modern gas turbine engines present significant challenges in the reliability of materials in the turbine. The increased engine temperatures required to achieve the higher performance levels reduce the strength of the materials used in the turbine sections of the engine. Various forms of thermal barrier coatings (TBC's) have been used for many years to increase the reliability of gas turbine engine components. Recent experience with the physical vapor deposition (PVD) process using ceramic material has demonstrated success in extending the service life of turbine blades and nozzles. Engine test results of turbine components with a 125 micron (0.005 in) PVD TBC have demonstrated component operating temperatures of 56-83 C (100-150 F) lower than non-PVD TBC components. Engine testing has also revealed the TBC is susceptible to high angle particle impact damage. Sand particles and other engine debris impact the TBC surface at the leading edge of airfoils and fracture the PVD columns. As the impacting continues, the TBC erodes away in local areas. Analysis of the eroded areas has shown a slight increase in temperature over a fully coated area, however a significant temperature reduction was realized over an airfoil without TBC.

Antonio Maricocchi↗

NASA NTRS1996

Options for flight testing rocket-based combined-cycle (RBCC) engines

While NASA's current next-generation launch vehicle research has largely focused on advanced all-rocket single-stage-to-orbit vehicles (i.e. the X-33 and it's RLV operational follow-on), some attention is being given to advanced propulsion concepts suitable for 'next-generation-and-a-half' vehicles. Rocket-based combined-cycle (RBCC) engines combining rocket and airbreathing elements are one candidate concept. Preliminary RBCC engine development was undertaken by the United States in the 1960's. However, additional ground and flight research is required to bring the engine to technological maturity. This paper presents two options for flight testing early versions of the RBCC ejector scramjet engine. The first option mounts a single RBCC engine module to the X-34 air-launched technology testbed for test flights up to about Mach 6.4. The second option links RBCC engine testing to the simultaneous development of a small-payload (220 lb.) two-stage-to-orbit operational vehicle in the Bantam payload class. This launcher/testbed concept has been dubbed the W vehicle. The W vehicle can also serve as an early ejector ramjet RBCC launcher (albeit at a lower payload). To complement current RBCC ground testing efforts, both flight test engines will use earth-storable propellants for their RBCC rocket primaries and hydrocarbon fuel for their airbreathing modes. Performance and vehicle sizing results are presented for both options.

Olds, John↗

NASA NTRS1996

Optimization of Air-Breathing Engine Concept

The design optimization of air-breathing propulsion engine concepts has been accomplished by soft-coupling the NASA Engine Performance Program (NEPP) analyzer with the NASA Lewis multidisciplinary optimization tool COMETBOARDS. Engine problems, with their associated design variables and constraints, were cast as nonlinear optimization problems with thrust as the merit function. Because of the large number of mission points in the flight envelope, the diversity of constraint types, and the overall distortion of the design space; the most reliable optimization algorithm available in COMETBOARDS, when used by itself, could not produce satisfactory, feasible, optimum solutions. However, COMETBOARDS' unique features-which include a cascade strategy, variable and constraint formulations, and scaling devised especially for difficult multidisciplinary applications-successfully optimized the performance of subsonic and supersonic engine concepts. Even when started from different design points, the combined COMETBOARDS and NEPP results converged to the same global optimum solution. This reliable and robust design tool eliminates manual intervention in the design of air-breathing propulsion engines and eases the cycle analysis procedures. It is also much easier to use than other codes, which is an added benefit. This paper describes COMETBOARDS and its cascade strategy and illustrates the capabilities of the combined design tool through the optimization of a high-bypass- turbofan wave-rotor-topped subsonic engine and a mixed-flow-turbofan supersonic engine.

Patnaik, Surya N.↗

NASA NTRS1997

RL10A-3-3A Rocket Engine Modeling Project

Two RL10A-3-3A rocket engines comprise the main propulsion system for the Centaur upper stage vehicle. Centaur is used with bod Titan and Atlas launch vehicles, carrying military and civilian payloads from high altitudes into orbit and beyond. The RL10 has delivered highly reliable service for the past 30 years. Recently, however, there have been two in-flight failures which have refocused attention on the RL10. This heightened interest has sparked a desire for an independent RL10 modeling capability within NASA and th Air Force. Pratt & Whitney, which presently has the most detailed model of the RL10, also sees merit in having an independent model which could be used as a cross-check with their own simulations. The Space Propulsion Technology Division (SPTD) at the NASA Lewis Research Center has developed a computer model of the RL10A-3-3A. A project team was formed, consisting of experts in the areas of turbomachinery, combustion, and heat transfer. The overall goal of the project was to provide a model of the entire RL10 rocket engine for government use. In the course of the project, the major engine components have been modeled using a combination of simple correlations and detailed component analysis tools (computer codes). The results of these component analyses were verified with data provided by Pratt & Whitney. Select modeling results and test data curves were then integrated to form the RL10 engine system model The purpose of this report is to introduce the reader to the RL10 rocket engine and to describe the engine system model. The RL10 engine and its application to U.S. launch vehicles are described first, followed by a summary of the SPTD project organization, goals, and accomplishments. Simulated output from the system model are shown in comparison with test and flight data for start transient, steady state, and shut-down transient operations. Detailed descriptions of all component analyses, including those not selected for integration with the system model, are included as appendices.

Binder, Michael↗

NASA NTRS1997

Metallized Gelled Propellants: Oxygen/RP-1/Aluminum Rocket Engine Calorimeter Heat Transfer Measurements and Analysis

A set of analyses was conducted to determine the heat transfer characteristics of metallized gelled liquid propellants in a rocket engine. The analyses used the data from experiments conducted with a small 30- to 40-lbf thrust engine composed of a modular injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-wt %, 5-wt%, and 55-wt% loadings of aluminum with silicon dioxide gellant, and gaseous oxygen as the oxidizer. Heat transfer was computed based on measurements using calorimeter rocket chamber and nozzle hardware with a total of 31 cooling channels. A gelled fuel coating formed in the 0-, 5- and 55-wt% engines, and the coating was composed of unburned gelled fuel and partially combusted RP-1. The coating caused a large decrease in calorimeter engine heat flux in the last half of the chamber for the 0- and 5-wt% RP-1/Al. This heat flux reduction effect was analyzed by comparing engine runs and the changes in the heat flux during a run as well as from run to run. Heat transfer and time-dependent heat flux analyses and interpretations are provided. The 5- and 55-wt% RP-1/Al fueled engines had the highest chamber heat fluxes, with the 5-wt% fuel having the highest throat flux. This result is counter to the predicted result, where the 55 wt% fuel has the highest combustion and throat temperature, and therefore implies that it would deliver the highest throat heat flux. The 5-wt% RP-1/Al produced the most influence on the engine heat transfer and the heat flux reduction was caused by the formation of a gelled propellant layer in the chamber and nozzle.

Palaszewski, Bryan↗

NASA NTRS1997

Dynamic Simulation of a Wave Rotor Topped Turboshaft Engine

The dynamic behavior of a wave rotor topped turboshaft engine is examined using a numerical simulation. The simulation utilizes an explicit, one-dimensional, multi-passage, CFD based wave rotor code in combination with an implicit, one-dimensional, component level dynamic engine simulation code. Transient responses to rapid fuel flow rate changes and compressor inlet pressure changes are simulated and compared with those of a similarly sized, untopped, turboshaft engine. Results indicate that the wave rotor topped engine responds in a stable, and rapid manner. Furthermore, during certain transient operations, the wave rotor actually tends to enhance engine stability. In particular, there is no tendency toward surge in the compressor of the wave rotor topped engine during rapid acceleration. In fact, the compressor actually moves slightly away from the surge line during this transient. This behavior is precisely the opposite to that of an untopped engine. The simulation is described. Issues associated with integrating CFD and component level codes are discussed. Results from several transient simulations are presented and discussed.

Greendyke, R. B.↗

NASA NTRS1998

1-D Numerical Analysis of RBCC Engine Performance

An RBCC engine combines air breathing and rocket engines into a single engine to increase the specific impulse over an entire flight trajectory. Considerable research pertaining to RBCC propulsion was performed during the 1960's and these engines were revisited recently as a candidate propulsion system for either a single-stage-to-orbit (SSTO) or two-stage-to-orbit (TSTO) launch vehicle. There are a variety of RBCC configurations that had been evaluated and new designs are currently under development. However, the basic configuration of all RBCC systems is built around the ejector scramjet engine originally developed for the hypersonic airplane. In this configuration, a rocket engine plays as an ejector in the air-augmented initial acceleration mode, as a fuel injector in scramjet mode and the rocket in all rocket mode for orbital insertion. Computational fluid dynamics (CFD) is a useful tool for the analysis of complex transport processes in various components in RBCC propulsion systems. The objective of the present research was to develop a transient 1-D numerical model that could be used to predict flow behavior throughout a generic RBCC engine following a flight path.

Han, Samuel S.↗

NASA NTRS1999

X-33/RLV Program Aerospike Engines

Substantial progress was made during the past year in support of the X-33/RLV program. X-33 activity was directed towards completing the remaining design work and building hardware to support test activities. RLV work focused on the nozzle ramp and powerpack technology tasks and on supporting vehicle configuration studies. On X-33, the design activity was completed to the detail level and the remainder of the drawings were released. Component fabrication and engine assembly activity was initiated, and the first two powerpacks and the GSE and STE needed to support powerpack testing were completed. Components fabrication is on track to support the first engine assembly schedule. Testing activity included powerpack testing and component development tests consisting of thrust cell single cell testing, CWI system spider testing, and EMA valve flow and vibration testing. Work performed for RLV was divided between engine system and technology development tasks. Engine system activity focused on developing the engine system configuration and supporting vehicle configuration studies. Also, engine requirements were developed, and engine performance analyses were conducted. In addition, processes were developed for implementing reliability, mass properties, and cost controls during design. Technology development efforts were divided between powerpack and nozzle ramp technology tasks. Powerpack technology activities were directed towards the development of a prototype powerpack and a ceramic turbine technology demonstrator (CTTD) test article which will allow testing of ceramic turbines and a close-coupled gas generator design. Nozzle technology efforts were focused on the selection of a composite nozzle supplier and on the fabrication and test of composite nozzle coupons.

Source record↗

NASA NTRS1999

Rocketdyne RBCC Engine Concept Development

Boeing Rocketdyne is pursuing the development of Rocket Based Combined Cycle (RBCC), propulsion systems as demonstrated by significant contract work in the hypersonic arena (ART, NASP, SCT, system studies) and over 12 years of steady company discretionary investment. The Rocketdyne concept is a fixed geometry integrated rocket, ramjet, scramjet which is hydrogen fueled and uses hydrogen regenerative cooling. The baseline engine structural configuration uses an integral structure that eliminates panel seals. seal purge gas, and closeout side attachments. Rocketdyne's experimental RBCC engine (Engine A5) was constructed under contract with the NASA Marshall Space Flight Center. Engine A5 models the complete flight engine flowpath consisting of an inlet, isolator, airbreathing combustor and nozzle. High performance rocket thrusters are integrated into the engine to enable both air-augmented rocket (AAR) and pure rocket operation. Engine A5 was tested in CASL's new FAST facility as an air-augmented rocket, a ramjet and a pure rocket. Measured performance demonstrated vision vehicle performance levels for Mach 3 AAR operation and ramjet operation from Mach 3 to 4. Rocket mode performance was above predictions. For the first time. testing also demonstrated transition from AAR operation to ramjet operation. This baseline configuration has also been shown, in previous testing, to perform well in the scramjet mode.

G. Ratckin↗

NASA NTRS1966

Thermal Control of Liquid H2/O2 Rocket Engines and Vehicle Impact

Cryogenic propulsion systems are subjected to various engine start environments during ground testing, first burn, and restart after orbital coast. These operational modes necessitate careful engine-stage integration, particularly in thermal control of propellant feed systems and engine subsystems. The wide range of operational conditions imposed on propellant feed systems include varying heat loads and acceleration forces for static test first burn and orbital restart. One method used to provide Net Positive Suctign,Head (NPSH) and maintain turbopump I is thermal control propellant circulation through the feed system prior to engine start. The attendant design and operational criteria are inter-related with stage and vehicle systems, including for example, auxiliary propulsion and propellant ullaging methods such as propellant tank propulsive venting. Thermally sensitive engine components require environmental control, protection, or special sequencing in order to assure repeatable engine start performance. An example is thrust chamber pre-chill to avoid pump stall during rapid start transient. Trends in engine development are toward expanding system flexibility to accommodate the wide variety of environments without special pre-start conditioning.

Winstead, Thomas W.↗

NASA NTRS1999

Wave-Rotor-Enhanced Gas Turbine Engine Demonstrator

The U.S. Army Research Laboratory, NASA Glenn Research Center, and Rolls-Royce Allison are working collaboratively to demonstrate the benefits and viability of a wave-rotor-topped gas turbine engine. The self-cooled wave rotor is predicted to increase the engine overall pressure ratio and peak temperature by 300% and 25 to 30%. respectively, providing substantial improvements in engine efficiency and specific power. Such performance improvements would significantly reduce engine emissions and the fuel logistics trails of armed forces. Progress towards a planned demonstration of a wave-rotor-topped Rolls-Royce Allison model 250 engine has included completion of the preliminary design and layout of the engine, the aerodynamic design of the wave rotor component and prediction of its aerodynamic performance characteristics in on- and off-design operation and during transients, and the aerodynamic design of transition ducts between the wave rotor and the high pressure turbine. The topping cycle increases the burner entry temperature and poses a design challenge to be met in the development of the demonstrator engine.

Welch, Gerard E.↗

NASA NTRS2000

Parametric Model of an Aerospike Rocket Engine

A suite of computer codes was assembled to simulate the performance of an aerospike engine and to generate the engine input for the Program to Optimize Simulated Trajectories. First an engine simulator module was developed that predicts the aerospike engine performance for a given mixture ratio, power level, thrust vectoring level, and altitude. This module was then used to rapidly generate the aerospike engine performance tables for axial thrust, normal thrust, pitching moment, and specific thrust. Parametric engine geometry was defined for use with the engine simulator module. The parametric model was also integrated into the iSIGHT multidisciplinary framework so that alternate designs could be determined. The computer codes were used to support in-house conceptual studies of reusable launch vehicle designs.

Korte, J. J.↗

NASA NTRS2000

Parametric Model of an Aerospike Rocket Engine

A suite of computer codes was assembled to simulate the performance of an aerospike engine and to generate the engine input for the Program to Optimize Simulated Trajectories. First an engine simulator module was developed that predicts the aerospike engine performance for a given mixture ratio, power level, thrust vectoring level, and altitude. This module was then used to rapidly generate the aerospike engine performance tables for axial thrust, normal thrust, pitching moment, and specific thrust. Parametric engine geometry was defined for use with the engine simulator module. The parametric model was also integrated into the iSIGHTI multidisciplinary framework so that alternate designs could be determined. The computer codes were used to support in-house conceptual studies of reusable launch vehicle designs.

Korte, J. J.↗

NASA NTRS2000

Numerical Zooming Between a NPSS Engine System Simulation and a One-Dimensional High Compressor Analysis Code

Within NASA's High Performance Computing and Communication (HPCC) program, NASA Glenn Research Center is developing an environment for the analysis/design of aircraft engines called the Numerical Propulsion System Simulation (NPSS). NPSS focuses on the integration of multiple disciplines such as aerodynamics, structures, and heat transfer along with the concept of numerical zooming between zero-dimensional to one-, two-, and three-dimensional component engine codes. In addition, the NPSS is refining the computing and communication technologies necessary to capture complex physical processes in a timely and cost-effective manner. The vision for NPSS is to create a "numerical test cell" enabling full engine simulations overnight on cost-effective computing platforms. Of the different technology areas that contribute to the development of the NPSS Environment, the subject of this paper is a discussion on numerical zooming between a NPSS engine simulation and higher fidelity representations of the engine components (fan, compressor, burner, turbines, etc.). What follows is a description of successfully zooming one-dimensional (row-by-row) high-pressure compressor analysis results back to a zero-dimensional NPSS engine simulation and a discussion of the results illustrated using an advanced data visualization tool. This type of high fidelity system-level analysis, made possible by the zooming capability of the NPSS, will greatly improve the capability of the engine system simulation and increase the level of virtual test conducted prior to committing the design to hardware.

Follen, Gregory↗

NASA NTRS1999

Oxygen/Alcohol Dual Thrust RCS Engines

A non-toxic dual thrust RCS engine offers significant operational, safety, and performance advantages to the space shuttle and the next generation RLVs. In this concept, a single engine produces two thrust levels of 25 and 870 lbf. The low thrust level is provided by the spark torch igniter, which, with the addition of 2 extra valves, can also be made to function as a vernier. A dual thrust RCS engine allows 38 verniers to be packaged more efficiently on a vehicle. These 38 vemiers improve translation and reduce cross coupling, thereby providing more pure roll, pitch, and yaw maneuvers of the vehicle. Compared to the 6 vemiers currently on the shuttle, the 38 dual thrust engines would be 25 to 40% more efficient for the same maneuvers and attitude control. The vernier thrust level also reduces plume impingement and contamination concerns. Redundancy is also improved, thereby improving mission success reliability. Oxygen and ethanol are benign propellants which do not create explosive reaction products or contamination, as compared to hypergolic propellants. These characteristics make dual-thrust engines simpler to implement on a non-toxic reaction control system. Tests at WSTF in August 1999 demonstrated a dual-thrust concept that is successful with oxygen and ethanol. Over a variety of inlet pressures and mixture ratios at 22:1 area ratio, the engine produced between 230 and 297 sec Isp, and thrust levels from 8 lbf. to 50 lbf. This paper describes the benefits of dual-thrust engines and the recent results from tests at WSTF.

Angstadt, Tara↗

NASA NTRS2001

Robust Strategy for Rocket Engine Health Monitoring

Monitoring the health of rocket engine systems is essentially a two-phase process. The acquisition phase involves sensing physical conditions at selected locations, converting physical inputs to electrical signals, conditioning the signals as appropriate to establish scale or filter interference, and recording results in a form that is easy to interpret. The inference phase involves analysis of results from the acquisition phase, comparison of analysis results to established health measures, and assessment of health indications. A variety of analytical tools may be employed in the inference phase of health monitoring. These tools can be separated into three broad categories: statistical, rule based, and model based. Statistical methods can provide excellent comparative measures of engine operating health. They require well-characterized data from an ensemble of "typical" engines, or "golden" data from a specific test assumed to define the operating norm in order to establish reliable comparative measures. Statistical methods are generally suitable for real-time health monitoring because they do not deal with the physical complexities of engine operation. The utility of statistical methods in rocket engine health monitoring is hindered by practical limits on the quantity and quality of available data. This is due to the difficulty and high cost of data acquisition, the limited number of available test engines, and the problem of simulating flight conditions in ground test facilities. In addition, statistical methods incur a penalty for disregarding flow complexity and are therefore limited in their ability to define performance shift causality. Rule based methods infer the health state of the engine system based on comparison of individual measurements or combinations of measurements with defined health norms or rules. This does not mean that rule based methods are necessarily simple. Although binary yes-no health assessment can sometimes be established by relatively simple rules, the causality assignment needed for refined health monitoring often requires an exceptionally complex rule base involving complicated logical maps. Structuring the rule system to be clear and unambiguous can be difficult, and the expert input required to maintain a large logic network and associated rule base can be prohibitive.

Santi, L. Michael↗

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 NTRS1946

Performance of the Modified V-1710-93 Engine-Stage Supercharger with a Constant-Area Vaneless Diffuser

As part of an investigation to increase the power output of the V-1710-93 engine at altitude, the engine-stage supercharger was combined with a constant-area vaneless diffuser designed to improve the performance of the engine-stage supercharger at the rated engine operating point. The performance of the modified supercharger was investigated in a variable-component supercharger test rig and compared with that of the standard supercharger with an 8-vaned diffuser. A separate evaluation of the component efficiencies and a study of the flow characteristics of the modified supercharger was made possible by internal diffuser instrumentation. At the volume flow required by the engine for rated operating conditions, the modified supercharger increased the over-all adiabatic efficiency 0.05 and the over-all pressure coefficient 0.035. Furthermore, the capacity of the engine-stage supercharger was increased by replacing the standard 8-vaned diffuser with the vaneless diffuser. The peak over-all adiabatic efficiency for the modified supercharger, however, was 0.05 to 0.07 lower than that of the standard unit over the range of tip speeds investigated. The improved performance of the modified supercharger at rated engine operating conditions resulted from a shift of the point of peak adiabatic efficiency and pressure coefficient of the standard supercharger to a higher flow. The energy loss through the vaneless diffuser was found to be small. Because of the restricted diffuser diameter, however, diffusion was inadequate, which resulted in a relatively small static-pressure rise through the diffuser, high diffuser-exit velocities, and excessive collector-case losses.

Douglas, John E.↗