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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 1,405 records · Page 78
All sourcesNASA NTRSDOE OSTI
NASA NTRS1953

Effect of Axially Staged Fuel Introduction on Performance of One-quarter Sector of Annular Turbojet Combustor

The design principle of injecting liquid fuel at more than one axial station in an annual turbojet combustor was investigated. Fuel was injected into the combustor as much as 5 inches downstream of the primary fuel injectors. Many fuel-injection configurations were examined and the performance results are presented for 11 configurations that best demonstrate the trends in performance obtained. The performance investigations were made at a constant combustor-inlet pressure of 15 inches of mercury absolute and at air flows up to 70 percent higher than values typical of current design practice. At these higher air flows, staging the fuel introduction improved the combustion efficiency considerably over that obtained in the combustor when no fuel staging was employed. At air flows currently encountered in turbojet engines, fuel staging was of minor value. Radial temperature distribution seemed relatively unaffected by the location of fuel-injection stations.

COMBUSTION - EFFECT OF ENGINE OPERATING CONDITIONS↗
NASA NTRS1999

High Reliability Engine Control Demonstrated for Aircraft Engines

For a dual redundant-control system, which is typical for short-haul aircraft, if a failure is detected in a control sensor, the engine control is transferred to a safety mode and an advisory is issued for immediate maintenance action to replace the failed sensor. The safety mode typically results in severely degraded engine performance. The goal of the High Reliability Engine Control (HREC) program was to demonstrate that the neural-network-based sensor validation technology can safely operate an engine by using the nominal closed-loop control during and after sensor failures. With this technology, engine performance could be maintained, and the sensor could be replaced as a conveniently scheduled maintenance action.

Guo, Ten-Huei↗
DOE OSTI2021

Global Power System Transformation Consortium

The Global Power System Transformation (G-PST) Consortium, an expert- and practitioner-driven initiative, engages key power system operators, research and educational institutes, governments, businesses, and stakeholders in all regions of the world to accelerate clean energy transitions at the ambitious scope and scale that is required. The G-PST Consortium provides coordinated, holistic support and knowledge infusion for system operators pursuing clean energy transitions, including performing cutting-edge research; providing implementation support for world-class engineering and operational solutions; supporting workforce development; building and disseminating open-access data and tools; and accelerating localized technology adoption, standards development, and testing programs.

consortium↗
NASA NTRS2010

Verification and Implementation of Operations Safety Controls for Flight Missions

There are several engineering disciplines, such as reliability, supportability, quality assurance, human factors, risk management, safety, etc. Safety is an extremely important engineering specialty within NASA, and the consequence involving a loss of crew is considered a catastrophic event. Safety is not difficult to achieve when properly integrated at the beginning of each space systems project/start of mission planning. The key is to ensure proper handling of safety verification throughout each flight/mission phase. Today, Safety and Mission Assurance (S&MA) operations engineers continue to conduct these flight product reviews across all open flight products. As such, these reviews help ensure that each mission is accomplished with safety requirements along with controls heavily embedded in applicable flight products. Most importantly, the S&MA operations engineers are required to look for important design and operations controls so that safety is strictly adhered to as well as reflected in the final flight product.

Smalls, James R.↗
NASA NTRS1983

Proceedings of the Flat-Plate Solar Array Project Research Forum on the Design of Flat-Plate Photovoltaic Arrays for Central Stations

The Flat Plate Solar Array Project, focuses on advancing technologies relevant to the design and construction of megawatt level central station systems. Photovoltaic modules and arrays for flat plate central station or other large scale electric power production facilities require the establishment of a technical base that resolves design issues and results in practical and cost effective configurations. Design, qualification and maintenance issues related to central station arrays derived from the engineering and operating experiences of early applications and parallel laboratory reserch activities are investigated. Technical issues are examined from the viewpoint of the utility engineer, architect/engineer and laboratory researcher. Topics on optimum source circuit designs, module insulation design for high system voltages, array safety, structural interface design, measurements, and array operation and maintenance are discussed.

Source record↗
NASA NTRS2010

Operation Program for the Spatially Phase-Shifted Digital Speckle Pattern Interferometer - SPS-DSPI

SPS-DSPI software has been revised so that Goddard optical engineers can operate the instrument, instead of data programmers. The user interface has been improved to view the data collected by the SPS-DSPI, with a real-time mode and a play-back mode. The SPS-DSPI has been developed by NASA/GSFC to measure the temperature distortions of the primary-mirror backplane structure for the James Webb Space Telescope. It requires a team of computer specialists to run successfully, because, at the time of this reporting, it just finished the prototype stage. This software improvement will transition the instrument to become available for use by many programs that measure distortion

Blake, Peter N.↗
NASA NTRS2003

Towards an Automated Full-Turbofan Engine Numerical Simulation

The objective of this study was to demonstrate the high-fidelity numerical simulation of a modern high-bypass turbofan engine. The simulation utilizes the Numerical Propulsion System Simulation (NPSS) thermodynamic cycle modeling system coupled to a high-fidelity full-engine model represented by a set of coupled three-dimensional computational fluid dynamic (CFD) component models. Boundary conditions from the balanced, steady-state cycle model are used to define component boundary conditions in the full-engine model. Operating characteristics of the three-dimensional component models are integrated into the cycle model via partial performance maps generated automatically from the CFD flow solutions using one-dimensional meanline turbomachinery programs. This paper reports on the progress made towards the full-engine simulation of the GE90-94B engine, highlighting the generation of the high-pressure compressor partial performance map. The ongoing work will provide a system to evaluate the steady and unsteady aerodynamic and mechanical interactions between engine components at design and off-design operating conditions.

Reed, John A.↗
NASA NTRS2003

Initial closed operation of the CELSS Test Facility Engineering Development Unit

As part of the NASA Advanced Life Support Flight Program, a Controlled Ecological Life Support System (CELSS) Test Facility Engineering Development Unit has been constructed and is undergoing initial operational testing at NASA Ames Research Center. The Engineering Development Unit (EDU) is a tightly closed, stringently controlled, ground-based testbed which provides a broad range of environmental conditions under which a variety of CELSS higher plant crops can be grown. Although the EDU was developed primarily to provide near-term engineering data and a realistic determination of the subsystem and system requirements necessary for the fabrication of a comparable flight unit, the EDU has also provided a means to evaluate plant crop productivity and physiology under controlled conditions. This paper describes the initial closed operational testing of the EDU, with emphasis on the hardware performance capabilities. Measured performance data during a 28-day closed operation period are compared with the specified functional requirements, and an example of inferring crop growth parameters from the test data is presented. Plans for future science and technology testing are also discussed. Published by Elsevier Science Ltd on behalf of COSPAR.

Lettuce/growth & development/metabolism↗
NASA NTRS2005

Single-String Integration Test Measurements of the NEXT Ion Engine Plume

Measurements were made of a 40 cm ion-thruster plume as part of the single-string-integration-test (SSIT) activity of Phase I of the NASA's Evolutionary Xenon Thruster (NEXT) project. The NEXT ion engine incorporates design improvements that extend NSTAR power levels and efficiencies. During SSIT, an engineering model (EM2) 40 cm engine was operated using an advanced xenon propellant system in combination with either a GRC power console or advanced power processing unit. Integral goals of the single-string phase were to characterize engine performance over the full input power range and to detail thruster operation within the specification of the NEXT throttle table. Plume diagnostics measurements of relative Xe(+) and Xe(++) currents were made using near-field and far-field ExB probes. Planar geometry faraday probes were used to obtain beam current density profiles. This paper reports on the characterization of the EM2 plume over a range of SSIT operating conditions, first with the advanced propellant management system teamed with the GRC power console and then with the power-processing unit.

Snyder, Aaron↗
NASA NTRS2010

CECE: Expanding the Envelope of Deep Throttling Technology in Liquid Oxygen/Liquid Hydrogen Rocket Engines for NASA Exploration Missions

As one of the first technology development programs awarded by NASA under the Vision for Space Exploration, the Pratt & Whitney Rocketdyne (PWR) Deep Throttling, Common Extensible Cryogenic Engine (CECE) program was selected by NASA in November 2004 to begin technology development and demonstration toward a deep throttling, cryogenic engine supporting ongoing trade studies for NASA s Lunar Lander descent stage. The CECE program leverages the maturity and previous investment of a flight-proven hydrogen/oxygen expander cycle engine, the PWR RL10, to develop and demonstrate an unprecedented combination of reliability, safety, durability, throttlability, and restart capabilities in high-energy, cryogenic, in-space propulsion. The testbed selected for the deep throttling demonstration phases of this program was a minimally modified RL10 engine, allowing for maximum current production engine commonality and extensibility with minimum program cost. Four series of demonstrator engine tests have been successfully completed between April 2006 and April 2010, accumulating 7,436 seconds of hot fire time over 47 separate tests. While the first two test series explored low power combustion (chug) and system instabilities, the third test series investigated and was ultimately successful in demonstrating several mitigating technologies for these instabilities and achieved a stable throttling ratio of 13:1. The fourth test series significantly expanded the engine s operability envelope by successfully demonstrating a closed-loop control system and extensive transient modeling to enable lower power engine starting, faster throttle ramp rates, and mission-specific ignition testing. The final hot fire test demonstrated a chug-free, minimum power level of 5.9%, corresponding to an overall 17.6:1 throttling ratio achieved. In total, these tests have provided an early technology demonstration of an enabling cryogenic propulsion concept with invaluable system-level technology data acquisition toward design and development risk mitigation for future lander descent main engines.

Giuliano, Victor J.↗
NASA NTRS2011

Advanced Development of a Compact 5-15 lbf Lox/Methane Thruster for an Integrated Reaction Control and Main Engine Propulsion System

This paper describes the advanced development and testing of a compact 5 to 15 lbf LOX/LCH4 thruster for a pressure-fed integrated main engine and RCS propulsion system to be used on a spacecraft "vertical" test bed (VTB). The ability of the RCS thruster and the main engine to operate off the same propellant supply in zero-g reduces mass and improves mission flexibility. This compact RCS engine incorporates several features to dramatically reduce mass and parts count, to ease manufacturing, and to maintain acceptable performance given that specific impulse (Isp) is not the driver. For example, radial injection holes placed on the chamber body for easier drilling, and high temperature Haynes 230 were selected for the chamber over other more expensive options. The valve inlets are rotatable before welding allowing different orientations for vehicle integration. In addition, the engine design effort selected a coil-on-plug ignition system which integrates a relay and coil with the plug electrode, and moves some exciter electronics to avionics driver board. The engine injector design has small dribble volumes to target minimum pulse widths of 20 msec. and an efficient minimum impulse bit of less than 0.05 lbf-sec. The propellants, oxygen and methane, were chosen because together they are a non-toxic, Mars-forward, high density, space storable, and high performance propellant combination that is capable of pressure-fed and pump-fed configurations and integration with life support and power subsystems. This paper will present the results of the advanced development testing to date of the RCS thruster and the integration with a vehicle propulsion system.

Hurlbert, Eric A.↗
NASA NTRS1957

Free-jet Tests of a 1.1-inch-diameter Supersonic Ram-jet Engine

Results are presented of free-jet tests of a 1.1-inch-diameter hydrogen-burning ram-jet engine over a Mach number range from 1.42 to 2.28 and a Reynolds number range from 6.01 times 10 to the 6th power to 15.78 times 10 to the 6th power. Tests demonstrated the reliability and wide operating range of the engine and showed its suitability for installation on wind-tunnel models of airplane and missile configurations. A comparison of engine operation with combustion-chamber lengths of 3.33 and 1.51 engine diameters was made at a Mach number of 2.06. A maximum test thrust coefficient of 0.905 was obtained at fuel-air ratio of 0.034 and a Mach number of 2.06 for the engine with the 3.33-diameter combustion chamber.

ENGINES, RAM-JET↗
NASA NTRS2012

Space Environments and Spacecraft Effects Organization Concept

The National Aeronautics and Space Administration (NASA) is embarking on a course to expand human presence beyond Low Earth Orbit (LEO) while also expanding its mission to explore the solar system. Destinations such as Near Earth Asteroids (NEA), Mars and its moons, and the outer planets are but a few of the mission targets. Each new destination presents an opportunity to increase our knowledge of the solar system and the unique environments for each mission target. NASA has multiple technical and science discipline areas specializing in specific space environments disciplines that will help serve to enable these missions. To complement these existing discipline areas, a concept is presented focusing on the development of a space environments and spacecraft effects (SENSE) organization. This SENSE organization includes disciplines such as space climate, space weather, natural and induced space environments, effects on spacecraft materials and systems and the transition of research information into application. This space environment and spacecraft effects organization will be composed of Technical Working Groups (TWG). These technical working groups will survey customers and users, generate products, and provide knowledge supporting four functional areas: design environments, engineering effects, operational support, and programmatic support. The four functional areas align with phases in the program mission lifecycle and are briefly described below. Design environments are used primarily in the mission concept and design phases of a program. Engineering effects focuses on the material, component, sub-system and system-level selection and the testing to verify design and operational performance. Operational support provides products based on real time or near real time space weather to mission operators to aid in real time and near-term decision-making. The programmatic support function maintains an interface with the numerous programs within NASA, other federal government agencies, and the commercial sector to ensure that communications are well established and the needs of the programs are being met. The programmatic support function also includes working in coordination with the program in anomaly resolution and generation of lessons learned documentation. The goal of this space environment and spacecraft effects organization is to develop decision-making tools and engineering products to support all mission phases from mission concept through operations by focusing on transitioning research to application. Products generated by this space environments and effects application are suitable for use in anomaly investigations. This paper will describe the scope of the TWGs and their relationship to the functional areas, and discuss an organizational structure for this space environments and spacecraft effects organization.

Edwards, David L.↗
NASA NTRS1973

Advanced space engine preliminary design

A preliminary design was completed for an O2/H2, 89 kN (20,000 lb) thrust staged combustion rocket engine that has a single-bell nozzle with an overall expansion ratio of 400:1. The engine has a best estimate vacuum specific impulse of 4623.8 N-s/kg (471.5 sec) at full thrust and mixture ratio = 6.0. The engine employs gear-driven, low pressure pumps to provide low NPSH capability while individual turbine-driven, high-speed main pumps provide the system pressures required for high-chamber pressure operation. The engine design dry weight for the fixed-nozzle configuration is 206.9 kg (456.3 lb). Engine overall length is 234 cm (92.1 in.). The extendible nozzle version has a stowed length of 141.5 cm (55.7 in.). Critical technology items in the development of the engine were defined. Development program plans and their costs for development, production, operation, and flight support of the ASE were established for minimum cost and minimum time programs.

Cuffe, J. P. B.↗
NASA NTRS1948

Simulated Altitude Performance of Combustor of Westinghouse 19XB-1 Jet-Propulsion Engine

A 19XB-1 combustor was operated under conditions simulating zero-ram operation of the 19XB-1 turbojet engine at various altitudes and engine speeds. The combustion efficiencies and the altitude operational limits were determined; data were also obtained on the character of the combustion, the pressure drop through the combustor, and the combustor-outlet temperature and velocity profiles. At altitudes about 10,000 feet below the operational limits, the flames were yellow and steady and the temperature rise through the combustor increased with fuel-air ratio throughout the range of fuel-air ratios investigated. At altitudes near the operational limits, the flames were blue and flickering and the combustor was sluggish in its response to changes in fuel flow. At these high altitudes, the temperature rise through the combustor increased very slowly as the fuel flow was increased and attained a maximum at a fuel-air ratio much leaner than the over-all stoichiometric; further increases in fuel flow resulted in decreased values of combustor temperature rise and increased resonance until a rich-limit blow-out occurred. The approximate operational ceiling of the engine as determined by the combustor, using AN-F-28, Amendment-3, fuel, was 30,400 feet at a simulated engine speed of 7500 rpm and increased as the engine speed was increased. At an engine speed of 16,000 rpm, the operational ceiling was approximately 48,000 feet. Throughout the range of simulated altitudes and engine speeds investigated, the combustion efficiency increased with increasing engine speed and with decreasing altitude. The combustion efficiency varied from over 99 percent at operating conditions simulating high engine speed and low altitude operation to less than 50 percent at conditions simulating operation at altitudes near the operational limits. The isothermal total pressure drop through the combustor was 1.82 times as great as the inlet dynamic pressure. As expected from theoretical considerations, a straight-line correlation was obtained when the ratio of the combustor total pressure drop to the combustor-inlet dynamic pressure was plotted as a function of the ratio of the combustor-inlet air density to the combustor-outlet gas density. The combustor-outlet temperature profiles were, in general, more uniform for runs in which the temperature rise was low and the combustion efficiency was high. Inspection of the combustor basket after 36 hours of operation showed very little deterioration and no appreciable carbon deposits.

Childs, J. Howard↗
NASA NTRS1973

Report to the administrator by the NASA Aerospace Safety Advisory Panel on the Skylab program. Volume 1: Summary report

Contractor and NASA technical management for the development and manufacture of the Skylab modules is reviewed with emphasis on the following management controls: configuration and interface management; vendor control; and quality control of workmanship. A review of the modified two-stage Saturn V launch vehicle which focused on modifications to accommodate the Skylab payload; resolution of prior flight anomalies; and changes in personnel and management systems is presented along with an evaluation of the possible age-life and storage problems for the Saturn 1-B launch vehicle. The NASA program management's visibility and control of contractor operations, systems engineering and integration, the review process for the evaluation of design and flight hardware, and the planning process for mission operations are investigated. It is concluded that the technical management system for development and fabrication of the modules, spacecraft, and launch vehicles, the process of design and hardware acceptance reviews, and the risk assessment activities are satisfactory. It is indicated that checkout activity, integrated testing, and preparations for and execution of mission operation require management attention.

Source record↗