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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 1,657 records · Page 92
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

Engineering analysis and test results of the pre-stage planetary gear trains for wrist rotation and pitch assembly and azimuth and elevation assembly of the extendable stiff arm manipulator kit assembly

In order to improve the performance capability of the Extendable Stiff Arm Manipulator (ESAM) it was necessary to increase the overall gear ratio by a factor of approximately four. This is accomplished with minimum effect to existing hardware by the interposition of a planetary gear transmission between the respective drive motors and the harmonic drive transmissions. The engineering analysis in support of this design approach and the subsequent no-load test results are reported.

Morris, R. N.↗
NASA NTRS2001

Parametric Data from a Wind Tunnel Test on a Rocket-Based Combined-Cycle Engine Inlet

A 40-percent scale model of the inlet to a rocket-based combined-cycle (RBCC) engine was tested in the NASA Glenn Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT). The full-scale RBCC engine is scheduled for test in the Hypersonic Tunnel Facility (HTF) at NASA Glenn's Plum Brook Station at Mach 5 and 6. This engine will incorporate the configuration of this inlet model which achieved the best performance during the present experiment. The inlet test was conducted at Mach numbers of 4.0, 5.0, 5.5, and 6.0. The fixed-geometry inlet consists of an 8 deg.. forebody compression plate, boundary layer diverter, and two compressive struts located within 2 parallel sidewalls. These struts extend through the inlet, dividing the flowpath into three channels. Test parameters investigated included strut geometry, boundary layer ingestion, and Reynolds number (Re). Inlet axial pressure distributions and cross-sectional Pitot-pressure surveys at the base of the struts were measured at varying back-pressures. Inlet performance and starting data are presented. The inlet chosen for the RBCC engine self-started at all Mach numbers from 4 to 6. Pitot-pressure contours showed large flow nonuniformity on the body-side of the inlet. The inlet provided adequate pressure recovery and flow quality for the RBCC cycle even with the flow separation.

Fernandez, Rene↗
NASA NTRS1981

Evaluation of the micro-carburetor

A prototype sonic, variable-venturi automotive carburetor was evaluated for its effects on vehicle performance, fuel economy, and exhaust emissions. A 350 CID Chevrolet Impala vehicle was tested on a chassis dynamometer over the 1975 Federal Test Procedure, urban driving cycle. The Micro-carburetor was tested and compared with stock and modified-stock engine configurations. Subsequently, the test vehicle's performance characteristics were examined with the stock carburetor and again with the Micro-carburetor in a series of on-road driveability tests. The test engine was then removed from the vehicle and installed on an engine dynamometer. Engine tests were conducted to compare the fuel economy, thermal efficiency, and cylinder-to-cylinder mixture distribution of the Micro-carburetor to that of the stock configuration. Test results show increases in thermal efficiency and improvements in fuel economy at all test conditions. Improve fuel/air mixture preparation is implied from the information presented. Further improvements in fuel economy and exhaust emissions are possible through a detailed recalibration of the Micro-carburetor.

Weiss, M. F.↗
NASA NTRS2001

Thin Film Sensors for Surface Measurements

Advanced thin film sensors that can provide accurate surface temperature, strain, and heat flux measurements have been developed at NASA Glenn Research Center. These sensors provide minimally intrusive characterization of advanced propulsion materials and components in hostile, high-temperature environments as well as validation of propulsion system design codes. The sensors are designed for applications on different material systems and engine components for testing in engine simulation facilities. Thin film thermocouples and strain gauges for the measurement of surface temperature and strain have been demonstrated on metals, ceramics and advanced ceramic-based composites of various component configurations. Test environments have included both air-breathing and space propulsion-based engine and burner rig environments at surface temperatures up to 1100 C and under high gas flow and pressure conditions. The technologies developed for these sensors as well as for a thin film heat flux gauge have been integrated into a single multifunctional gauge for the simultaneous real-time measurement of surface temperature, strain, and heat flux. This is the first step toward the development of smart sensors with integrated signal conditioning and high temperature electronics that would have the capability to provide feedback to the operating system in real-time. A description of the fabrication process for the thin film sensors and multifunctional gauge will be provided. In addition, the material systems on which the sensors have been demonstrated, the test facilities and the results of the tests to-date will be described. Finally, the results will be provided of the current effort to demonstrate the capabilities of the multifunctional gauge.

Martin, Lisa C.↗
NASA NTRS1995

The Role of Analysis and Testing in the Service Life Assessment of Ion Engines

Experience from tests and flights and engineering analysis represent the two sources of information on which to base conclusions on the reliability or failure risk of aerospace flight systems. It is rarely feasible to establish high reliability at high confidence by testing aerospace flight systems or components. The limitations of testing in evaluating failure risk are discussed and an alternate statistical approach which relies on both test experience and analysis to quantitatively assess reliability is outlined. The implementation of this methodology in the service life assessment of ion thrusters is discussed and examples of failure modes being addressed in the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) program are given.

Ion↗
NASA NTRS2013

Combustion Stability of the Gas Generator Assembly from J-2X Engine E10001 and Powerpack Tests

Testing of a powerpack configuration (turbomachinery and gas generator assembly) and the first complete engine system of the liquid oxygen/liquid hydrogen propellant J-2X rocket engine have been completed at the NASA Stennis Space Center. The combustion stability characteristics of the gas generator assemblies on these two systems are of interest for reporting since considerable effort was expended to eliminate combustion instability during early development of the gas generator assembly with workhorse hardware. Comparing the final workhorse gas generator assembly development test data to the powerpack and engine system test data provides an opportunity to investigate how the nearly identical configurations of gas generator assemblies operate with two very different propellant supply systems one the autonomous pressure-fed test configuration on the workhorse development test stand, the other the pump-fed configurations on the powerpack and engine systems. The development of the gas generator assembly and the elimination of the combustion instability on the pressure-fed workhorse test stand have been reported extensively in the two previous Liquid Propulsion Subcommittee meetings 1-7. The powerpack and engine system testing have been conducted from mid-2011 through 2012. All tests of the powerpack and engine system gas generator systems to date have been stable. However, measureable dynamic behavior, similar to that observed on the pressure-fed test stand and reported in Ref. [6] and attributed to an injection-coupled response, has appeared in both powerpack and engine system tests. As discussed in Ref. [6], these injection-coupled responses are influenced by the interaction of the combustion chamber with a branch pipe in the hot gas duct that supplies gaseous helium to pre-spin the turbine during the start transient. This paper presents the powerpack and engine system gas generator test data, compares these data to the development test data, and provides additional combustion stability analyses of the configurations.

Hulka, J. R.↗
NASA NTRS2011

Project Morpheus Main Engine Development and Preliminary Flight Testing

A LOX/Methane rocket engine was developed for a prototype terrestrial lander and then used to fly the lander at Johnson Space Center. The development path of this engine is outlined, including unique items such as variable acoustic damping and variable film cooling.

Morehead, Robert L.↗
NASA NTRS1999

Testing of the Engineering Model Electrical Power Control Unit for the Fluids and Combustion Facility

The John H. Glenn Research Center at Lewis Field (GRC) in Cleveland, OH and the Sundstrand Corporation in Rockford, IL have designed and developed an Engineering Model (EM) Electrical Power Control Unit (EPCU) for the Fluids Combustion Facility, (FCF) experiments to be flown on the International Space Station (ISS). The EPCU will be used as the power interface to the ISS power distribution system for the FCF's space experiments'test and telemetry hardware. Furthermore. it is proposed to be the common power interface for all experiments. The EPCU is a three kilowatt 12OVdc-to-28Vdc converter utilizing three independent Power Converter Units (PCUs), each rated at 1kWe (36Adc @ 28Vdc) which are paralleled and synchronized. Each converter may be fed from one of two ISS power channels. The 28Vdc loads are connected to the EPCU output via 48 solid-state and current-limiting switches, rated at 4Adc each. These switches may be paralleled to supply any given load up to the 108Adc normal operational limit of the paralleled converters. The EPCU was designed in this manner to maximize allocated-power utilization. to shed loads autonomously, to provide fault tolerance. and to provide a flexible power converter and control module to meet various ISS load demands. Tests of the EPCU in the Power Systems Facility testbed at GRC reveal that the overall converted-power efficiency, is approximately 89% with a nominal-input voltage of 12OVdc and a total load in the range of 4O% to 110% rated 28Vdc load. (The PCUs alone have an efficiency of approximately 94.5%). Furthermore, the EM unit passed all flight-qualification level (and beyond) vibration tests, passed ISS EMI (conducted, radiated. and susceptibility) requirements. successfully operated for extended periods in a thermal/vacuum chamber, was integrated with a proto-flight experiment and passed all stability and functional requirements.

Kimnach, Greg L.↗
NASA NTRS1980

AiResearch QCGAT engine performance and emissions tests

Results of aerodynamic performance and emission tests, conducted on a specially designed QCGAT engine in the 17,793-N (4,000 lb) thrust class, are presented. Performance of the AiResearch QCGAT engine was excellent throughout all testing. No serious mechanical malfunctions were encountered, and no significant test time was lost due to engine-related problems. Emissions were drastically reduced over similar engines, and the engine exhibited good smoke performance.

Norgren, W. M.↗
NASA NTRS1983

Fluctuating pressure measurements on the fan blades of a turbofan engine during ground and flight tests

As part of a program to investigate fan noise generated by a turbofan engine, miniature pressure transducers were used to measure the fluctuating pressures on the surface of a fan blade on the JT15D engine. Tests were conducted on the ground, with tunnel and in flight. Using narrowband spectra analysis and signal enhancement techniques, comparisons for these various test conditions are presented. Results on the blade pressures are also presented from a flight test in which 41 rods were placed in front of the fan face to excite a known spinning mode in the duct. Results showed that with care, ground tests can be conducted in which the pressure loading on the blades are similar to those in flight. Comparison of pressure transducers on the blade surface also showed that a single transducer would provide much of the information available from this type of measurement.

Schoenster, J. A.↗
NASA NTRS2017

Near Earth Asteroid Solar Sail Engineering Development Unit Test Program

The Near Earth Asteroid (NEA) Scout project is a 30x20x10cm (6U) cubesat reconnaissance mission to investigate a near Earth asteroid utilizing an 86m2 solar sail as the primary propulsion system. This will be the largest solar sail NASA will launch to date. NEA Scout is a secondary payload currently manifested on the maiden voyage of the Space Launch System in 2018. In development of the solar sail subsystem, design challenges were identified and investigated for packaging within a 6U form factor and deployment in cis-lunar space. Analysis furthered understanding of thermal, stress, and dynamics of the stowed system and matured an integrated sail membrane model for deployed flight dynamics. This paper will address design, fabrication, and lessons learned from the NEA Scout solar sail subsystem engineering development unit. From optical properties of the sail material to folding and spooling the single 86m2 sail, the team has developed a robust deployment system for the solar sail. This paper will also address expected and received test results from ascent vent, random vibration, and deployment tests.

Lockett, Tiffany Russell↗
NASA NTRS2016

Full-Scale Turbofan Engine Noise-Source Separation Using a Four-Signal Method

Contributions from the combustor to the overall propulsion noise of civilian transport aircraft are starting to become important due to turbofan design trends and expected advances in mitigation of other noise sources. During on-ground, static-engine acoustic tests, combustor noise is generally sub-dominant to other engine noise sources because of the absence of in-flight effects. Consequently, noise-source separation techniques are needed to extract combustor-noise information from the total noise signature in order to further progress. A novel four-signal source-separation method is applied to data from a static, full-scale engine test and compared to previous methods. The new method is, in a sense, a combination of two- and three-signal techniques and represents an attempt to alleviate some of the weaknesses of each of those approaches. This work is supported by the NASA Advanced Air Vehicles Program, Advanced Air Transport Technology Project, Aircraft Noise Reduction Subproject and the NASA Glenn Faculty Fellowship Program.

Combustor Noise↗
NASA NTRS1978

Definition study of a Variable Cycle Experimental Engine (VCEE) and associated test program and test plan

The Definition Study of a Variable Cycle Experimental Engine (VCEE) and Associated Test Program and Test Plan, was initiated to identify the most cost effective program for a follow-on to the AST Test Bed Program. The VCEE Study defined various subscale VCE's based on different available core engine components, and a full scale VCEE utilizing current technology. The cycles were selected, preliminary design accomplished and program plans and engineering costs developed for several program options. In addition to the VCEE program plans and options, a limited effort was applied to identifying programs that could logically be accomplished on the AST Test Bed Program VCE to extend the usefulness of this test hardware. Component programs were provided that could be accomplished prior to the start of a VCEE program.

Allan, R. D.↗
NASA NTRS1974

The 727 airplane target thrust reverser static performance model test for refanned JT8D engines

The results of a scale model static performance test of target thrust reverser configurations for the Pratt and Whitney Aircraft JT8D-100 series engine are presented. The objective of the test was to select a series of suitable candidate reverser configurations for the subsequent airplane model wind tunnel ingestion and flight controls tests. Test results indicate that adequate reverse thrust performance with compatible engine airflow match is achievable for the selected configurations. Tapering of the lips results in loss of performance and only minimal flow directivity. Door pressure surveys were conducted on a selected number of lip and fence configurations to obtain data to support the design of the thrust reverser system.

Chow, C. T. P.↗
DOE OSTI2017

Gas Atomization Equipment Statement of Work and Specification for Engineering design, Fabrication, Testing, and Installation

The Gas Atomization Equipment will be used to fabricate metallic powder suitable for Powder Bed Fusion additive Manufacturing material to support Lawrence Livermore National Laboratory (LLNL) research and development. The project will modernize our capabilities to develop spherical reactive, refractory, and radioactive powders in the 10-75 μm diameter size range at LLNL.

42 ENGINEERING↗
NASA NTRS1987

RL10A-3-3B high mixture ratio qualification program

The results of the high mixture ratio qualification testing of the RL10 engine for the Shuttle/Centaur program are presented. The objective of the engine qualification test was to demonstrate the suitability of the RL10A-3-3B engine for space vehicle flight by subjecting it to the testing specified in RL10A-3-3B Model Specification Number 2295 dated February 1986. The applicable section of the specification is presented. Due to payload volume advantages which can be achieved by increasing the operating mixture ratio of the RL10, a decision was made to qualify the engine to run at a higher mixture ratio. A program was created to qualify the RL10 engine for operation at 15,000 pounds thrust and a nominal 6.0 to 1 mixture ratio. This model of the engine was designated the RL10A-3-3B. The qualification program included three test series as follows: (1) hardware durability and limits test in which the engine completed 23 firings and 4605.7 seconds with 1588.7 seconds at less than 6.6 mixture ratio; (2) preliminary qualification test in which the engine completed 26 firings and 5750 seconds; and (3) qualification test in which the engine completed 26 hot firings and 5693.4 seconds with 905.9 seconds at 6.7 mixture ratio. Several changes in engine hardware were required for operation of the RL10A-3-3B engine in the Space Shuttle which include a duel pressure switch ignition, an oxidizer flow control, and helium plumbing changes.

Vogel, T.↗
NASA NTRS1977

Outer planet probe engineering model thermal vacuum test

A thermal vacuum test was performed on the engineering model of the outer planets atmospheric entry probe. Steady state runs at three simulated radioisotope heating unit loads and one transient run simulating the pre-entry power profile were made to determine the thermal characteristics of the engineering model. An analytic simulation of the model was correlated to the test data. Several steady state and one transient run were made with the model attached to the spacecraft adapter to determine the thermal interface between the model and the adapter.

Grote, M. G.↗