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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 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Focused technology: Nuclear propulsion

The topics presented are covered in viewgraph form and include: nuclear thermal propulsion (NTP), which challenges (1) high temperature fuel and materials, (2) hot hydrogen environment, (3) test facilities, (4) safety, (5) environmental impact compliance, and (6) concept development, and nuclear electric propulsion (NEP), which challenges (1) long operational lifetime, (2) high temperature reactors, turbines, and radiators, (3) high fuel burn-up reactor fuels, and designs, (4) efficient, high temperature power conditioning, (5) high efficiency, and long life thrusters, (6) safety, (7) environmental impact compliance, and (8) concept development.

Miller, Thomas J.↗
NASA NTRS1994

NASA Lewis advanced subsonic transport program overview

An Advanced Subsonic Transport (AST) research program has been initiated by NASA to develop a new generation of superior U.S. aircraft and engines, and to help U.S. aeronautics industry increase their market share and competitiveness. NASA Lewis has the lead responsibility for the AST propulsion element. The top priority is to achieve a 70% reduction of NOx at all operating conditions including subsonic cruise. Since the technology for the NOx reductions does not exist, an aggressive research program has been established for its evolution. To date the program has been initiated with study contracts with industry as well as the initiation of analytical and experimental research. To conduct the high pressure tests, a new national facility with a high temperature capability for continuous operation at 60:1 pressure ratios is being constructed at Lewis and will be operational in 1996. The general flow of the program is to first conduct fundamental experiments and analytical studies to define combustor concepts, demonstrate the 70% NOx reduction in combustors in 1999, conduct a test-bed core engine demonstration in 2001, and demonstrate in 2002 the required operability and altitude relight for an Entry Into Service of the technology by industry in 2005.

Lee, Chi-Ming↗
NASA NTRS1995

NASA Lewis Research Center's combustor test facilities and capabilities

NASA Lewis Research Center (LeRC) presently accommodates a total of six combustor test facilities with unique capabilities. The facilities are used to evaluate combustor and afterburner concepts for future engine applications, and also to test the survivability and performance of innovative high temperature materials, new instrumentation, and engine components in a realistic jet engine environment. The facilities provide a variety of test section interfaces and lengths to allow for flametube, sector and component testing. The facilities can accommodate a wide range of operating conditions due to differing capabilities in the following areas: inlet air pressure, temperature, and flow; fuel flow rate, pressure, and fuel storage capacity; maximum combustion zone temperature; cooling water flow rate and pressure; types of exhaust - atmospheric or altitude; air heater supply pressure; and types of air heaters - vitiated or nonvitiated. All of the facilities have provisions for standard gas (emissions) analysis, and a few of the facilities are equipped with specialized gas analysis equipment, smoke and particle size measurement devices, and a variety of laser systems. This report will present some of the unique features of each of the high temperature/high pressure combustor test facilities at NASA LeRC.

Bianco, Jean↗
NASA NTRS2019

Progress Towards Modeling the Mars Science Laboratory PICA-NuSil Heatshield

The data collected by the Mars Science Laboratory (MSL) Entry, Descent and Landing Instrumentation, MEDLI, have become an established reference to assess the performance of engineering models of the Phenolic Impregnated Carbon Ablator (PICA) and to validate hypersonic computational fluid dynamics (CFD) tools for entry systems. MEDLI measurements are also extensively used as validation reference for current developments of high-fidelity material response models for PICA. So large has been the scientific output and impact of MEDLI that a follow-up instrumentation suite MEDLI2 is underway for the upcoming Mars 2020 mission.A feature neglected thus far in the modeling of the MSL heatshield, is the presence of a silicone-based room temperature vulcanizing coating designated NuSil CV-1144-0. NuSil was used to coat the entire MSL heatshield, including the MEDLI plugs, to mitigate the spread of phenolic dust from PICA, and limit contamination during clean room operations. NuSil CV-1144-0 is a space grade siloxane copolymer, designed as an oxygen protection barrier for extreme low temperature environment.Assessments conducted during MSL development demonstrated that the presence of NuSil had no adverse effect on the performance of PICA. However, evidence from ground testing of PICA-NuSil (PICA-N) models in the HyMETS arc-jet test facility suggests that the silicone changes the high temperature response of PICA. It is therefore critical to assess the importance of modeling the coating in ongoing code validation efforts.

Bessire, Brody K.↗
NASA NTRS2024

Radiant Heater Test Facility (RHTF) NASA Johnson Space Center

The RHTF provides the capability to perform multizone, high-temperature radiant testing of large spacecraft TPS and associated structures in a controlled pressure environment to simulate reentry thermal profiles and thermal gradients.

Tien Nguyen↗
NASA NTRS2004

High Speed Operation and Testing of a Fault Tolerant Magnetic Bearing

Research activities undertaken to upgrade the fault-tolerant facility, continue testing high-speed fault-tolerant operation, and assist in the commission of the high temperature (1000 degrees F) thrust magnetic bearing as described. The fault-tolerant magnetic bearing test facility was upgraded to operate to 40,000 RPM. The necessary upgrades included new state-of-the art position sensors with high frequency modulation and new power edge filtering of amplifier outputs. A comparison study of the new sensors and the previous system was done as well as a noise assessment of the sensor-to-controller signals. Also a comparison study of power edge filtering for amplifier-to-actuator signals was done; this information is valuable for all position sensing and motor actuation applications. After these facility upgrades were completed, the rig is believed to have capabilities for 40,000 RPM operation, though this has yet to be demonstrated. Other upgrades included verification and upgrading of safety shielding, and upgrading control algorithms. The rig will now also be used to demonstrate motoring capabilities and control algorithms are in the process of being created. Recently an extreme temperature thrust magnetic bearing was designed from the ground up. The thrust bearing was designed to fit within the existing high temperature facility. The retrofit began near the end of the summer, 04, and continues currently. Contract staff authored a NASA-TM entitled "An Overview of Magnetic Bearing Technology for Gas Turbine Engines", containing a compilation of bearing data as it pertains to operation in the regime of the gas turbine engine and a presentation of how magnetic bearings can become a viable candidate for use in future engine technology.

DeWitt, Kenneth↗
NASA NTRS2022

Design of Hypervelocity Flow Generator (HFG) and Its Diagnostics

Ground facilities for hypersonic research are a key element for successful development of aerodynamically proven hypervelocity vehicles. Design concepts and diagnostics of a hypervelocity flow generator (HFG) were made as a test platform for hypersonic/hypervelocity spacecraft models at the NASA Langley Research Center. The HFG is a hypersonic flow field generator using optically heated gas which is blown into an 80 cu m vacuum chamber. The vacuum chamber is kept at a stable vacuum pressure with a combination of three large vacuum pumps, while the HFG is in the test mode. The HFG provides a relatively small test section with approximately a 20 cm window. This facility was designed to generate 2.45 km/sec of flow speed, and potentially generate a continuous flow with the nozzle and vacuum system. The window in test section provides a direct view of the shock wave around a model in order to measure temperature, pressure, and density profile within the shock layer. One of the key test goals under this project is to understand why the emission spectra from the standing shockwave plasma predicted by Lora-Loran codes are significantly different from the measured emission spectra from the Flight Investigation Reentry Environment (FIRE II) Flight. The correct estimate of the thermal loading on the leading edge of hypersonic vehicles greatly affects the aerodynamic design, the material selection for the vehicle, and the cooling requirement and can be obtained by the precise modeling of emission spectra from shockwave plasma. However, the estimation of thermal loading is not an easy task due to complex non-equilibrium radiative process within high temperature shock layers that still falls into a category of cold plasma. Direct flight experiments are the most desirable, but not a cost-effective approach. Analysis by computational fluid dynamics (CFD) offers many test flexibilities. However, the CFD codes must be fully tested and validated with experimental data before the codes are effectively used for practical design. Large discrepancies between experiments and codes appear in hypersonic/hypervelocity flow regimes at high altitudes of 60 km ~ 90 km. This HFG facility offers some important parameters for CFD code validation, such as collision cross-sections, relaxation times, reaction rate coefficients and transportation coefficients. The HFG test facility is based on the ejection flow of high temperature gas heated over to 3500 K through a nozzle. The tungsten gas chamber of the HFG is heated up to a desired temperature by a 60 kW optical power beam source. This system consists of an optical power source, a thermal chamber, an expansion nozzle, a test section, and an 80 cu m vacuum tank. 60 kW optical input power is obtained from the 150 kW Vortek arc lamp system (by Vortek Industries, Vancouver, Canada). This optical beam is focused to heat the gas chamber within which a flow media is heated. The maximum achievable temperature of the flow medium reached approximately 3500 K or even higher but is limited by the melting point of the chamber material used. The exhaust velocity through the nozzle was determined by the stagnation temperature and the molecular weight of the working medium at the test section. To provide design parameters, a NASA Chemical Equilibrium with Application (CEA) computer program is used for the simulation of aerothermal data. This CEA program can calculate chemical equilibrium and properties of complex mixtures using shock tube parameters. For nitrogen gas at 2666 K stagnation temperature, the maximum achievable velocity at the test section is approximately 2.45 km/sec which is within the range of the thermal velocity of 8000 K shock layer. Based on the calculation through the CEA program, the design parameters of a HFG were determined and implemented for the test section that includes an expansion nozzle. The installed test section of the experimental facility can sustain a condition of a re-entry vehicle from the Space at an altitude of 60-90 km. After installation of the HFG, the system was fully tested and its operational parameters were measured. An 80 cu m vacuum chamber of HFG was set at 1 torr level to keep a stable downstream condition. The pumping time to reach the minimum vacuum pressure (~ 1 torr) at the test chamber from atmospheric pressure was approximately an hour using Kinney (MBV-14000/MB -1600/KT-300) and Stokes (Model 1772 and 412) pumps. Such a setup condition allowed a continued stable operation of the HFG experiment with flowrates through 1-, 3-, or 5-mm diameters of nozzle throat. The flow characteristics of the HFG for various operating conditions were performed using a focused Schlieren method. At 0.8 torr chamber pressure, a barrel shock was observed at the test section. With a sphere obstacle of 12 mm diameter in the flow, a bow shock (~ 2-mm thickness) was observed by a focused Schlieren visualization method.

Hypervelocity flow↗
NASA NTRS2005

Cooled Ceramic Matrix Composite Propulsion Structures Demonstrated

NASA's Next Generation Launch Technology (NGLT) Program has successfully demonstrated cooled ceramic matrix composite (CMC) technology in a scramjet engine test. This demonstration represented the world s largest cooled nonmetallic matrix composite panel fabricated for a scramjet engine and the first cooled nonmetallic composite to be tested in a scramjet facility. Lightweight, high-temperature, actively cooled structures have been identified as a key technology for enabling reliable and low-cost space access. Tradeoff studies have shown this to be the case for a variety of launch platforms, including rockets and hypersonic cruise vehicles. Actively cooled carbon and CMC structures may meet high-performance goals at significantly lower weight, while improving safety by operating with a higher margin between the design temperature and material upper-use temperature. Studies have shown that using actively cooled CMCs can reduce the weight of the cooled flow-path component from 4.5 to 1.6 lb/sq ft and the weight of the propulsion system s cooled surface area by more than 50 percent. This weight savings enables advanced concepts, increased payload, and increased range. The ability of the cooled CMC flow-path components to operate over 1000 F hotter than the state-of-the-art metallic concept adds system design flexibility to space-access vehicle concepts. Other potential system-level benefits include smaller fuel pumps, lower part count, lower cost, and increased operating margin.

Jaskowiak, Martha H.↗
NASA NTRS1998

Rayleigh Scattering Diagnostic for Measurement of Temperature and Velocity in Harsh Environments

A molecular Rayleigh scattering system for temperature and velocity measurements in unseeded flows is described. The system is capable of making measurements in the harsh environments commonly found in aerospace test facilities, which may have high acoustic sound levels, varying temperatures, and high vibration levels. Light from an argon-ion laser is transmitted via an optical fiber to a remote location where two flow experiments were located. One was a subsonic free air jet; the second was a low-speed heated airjet. Rayleigh scattered light from the probe volume was transmitted through another optical fiber from the remote location to a controlled environment where a Fabry-Perot interferometer and cooled CCD camera were used to analyze the Rayleigh scattered light. Good agreement between the measured velocity and the velocity calculated from isentropic flow relations was demonstrated (less than 5 m/sec). The temperature measurements, however, exhibited systematic errors on the order of 10-15%.

Seasholtz, Richard G.↗
NASA NTRS2011

Advanced High-Temperature Flexible TPS for Inflatable Aerodynamic Decelerators

Typical entry vehicle aeroshells are limited in size by the launch vehicle shroud. Inflatable aerodynamic decelerators allow larger aeroshell diameters for entry vehicles because they are not constrained to the launch vehicle shroud diameter. During launch, the hypersonic inflatable aerodynamic decelerator (HIAD) is packed in a stowed configuration. Prior to atmospheric entry, the HIAD is deployed to produce a drag device many times larger than the launch shroud diameter. The large surface area of the inflatable aeroshell provides deceleration of high-mass entry vehicles at relatively low ballistic coefficients. Even for these low ballistic coefficients there is still appreciable heating, requiring the HIAD to employ a thermal protection system (TPS). This TPS must be capable of surviving the heat pulse, and the rigors of fabrication handling, high density packing, deployment, and aerodynamic loading. This paper provides a comprehensive overview of flexible TPS tests and results, conducted over the last three years. This paper also includes an overview of each test facility, the general approach for testing flexible TPS, the thermal analysis methodology and results, and a comparison with 8-foot High Temperature Tunnel, Laser-Hardened Materials Evaluation Laboratory, and Panel Test Facility test data. Results are presented for a baseline TPS layup that can withstand a 20 W/cm2 heat flux, silicon carbide (SiC) based TPS layup, and polyimide insulator TPS layup. Recent work has focused on developing material layups expected to survive heat flux loads up to 50 W/cm2 (which is adequate for many potential applications), future work will consider concepts capable of withstanding more than 100 W/cm2 incident radiant heat flux. This paper provides an overview of the experimental setup, material layup configurations, facility conditions, and planned future flexible TPS activities.

DelCorso, Joseph A.↗
NASA NTRS2002

Fail-Safe Operation of a High-Temperature Magnetic Bearing Investigated for Gas Turbine Engine Applications

The Structural Mechanics and Dynamics Branch at the NASA Glenn Research Center has developed a three-axis high-temperature magnetic bearing suspension rig to enhance the safety of the bearing system up to 1000 F. This test rig can accommodate thrust and radial bearings up to a 22.84 cm (9 in.) diameter with a maximum axial loading of 22.25 kN (5000 lb) and a maximum radial loading up to 4.45 kN (1000 lb). The test facility was set up to test magnetic bearings under high-temperature (1100 F) and high-speed (20,000 rpm) conditions. The magnetic bearing is located at the center of gravity of the rotor between two high-temperature grease-packed mechanical ball bearings. The drive-end duplex angular contact ball bearing, which is in full contact, acts as a moment release and provides axial stability. The outboard end ball bearing has a 0.015-in. radial clearance between the rotor to act as a backup bearing and to compensate for axial thermal expansion. There is a 0.020-in. radial air gap between the stator pole and the rotor. The stator was wrapped with three 1-kW band heaters to create a localized hot section; the mechanical ball bearings were outside this section. Eight threaded rods supported the stator. These incorporated a plunger and Bellville washers to compensate for radial thermal expansion and provide rotor-to-stator alignment. The stator was instrumented with thermocouples and a current sensor for each coil. Eight air-cooled position sensors were mounted outside the hot section to monitor the rotor. Another sensor monitored this rotation of the outboard backup bearing. Ground fault circuit interrupts were incorporated into all power amplifier loops for personnel safety. All instrumentation was monitored and recorded on a LabView-based data acquisition system. Currently, this 12-pole heteropolar magnetic bearing has 13 thermal cycles and over 26 hr of operation at 1000 F.

Choi, Benjamin B.↗
NASA NTRS1979

Capability for aerothermal-structural tests of large-to-full-scale components of future space transportation systems

The design of viable, low-mass thermostructural concepts for future space transportation systems requires accurate definition of the localized heat flux, pressures, and flow-surface interaction parameters for complex flow regions of detailed areas of large vehicles, such as wing-elevon coves, thermal protection system tile gaps, and corrugated metallic surfaces. This paper discusses investigations recently conducted in and planned for two high-energy wind tunnels at the Langley Research Center - the 8-foot High-temperatures Tunnel and the Thermal Protection System Test Facility. The data obtained on large-to-full-scale vehicle components, comparisons of the experimental data with theory, and the large, reusable, generalized test apparatus available at or planned for these facilities are discussed.

Couch, L. M.↗
NASA NTRS2001

High Altitude Small Engine Test Techniques at the NASA Glenn Propulsion Systems Lab

A High Altitude Test was performed in the Propulsion Systems Lab (PSL) at the NASA Glenn Research Center using a Pratt and Whitney Canada PW545 jet engine. This engine was tested to develop a highaltitude database on small, high-bypass ratio, engine performance and operability. Industry is interested in the use of high-bypass engines for Uninhabited Aerial Vehicles (UAV's) to perform high altitude surveillance. The tests were a combined effort between Pratt & Whitney Canada (PWC) and NASA Glenn Research Center. A large portion of this test activity was to collect performance data with a highly instrumented low-pressure turbine. Low-pressure turbine aerodynamic performance at low Reynolds numbers was collected and compared to analytical models developed by NASA and PWC. This report describes the test techniques implemented to obtain high accuracy turbine performance data in an altitude test facility, including high accuracy airflow at high altitudes, very low mass flow, and low air temperatures. Major accomplishments from this test activity were to collect accurate and repeatable turbine performance data at high altitudes to within 1 percent. Data were collected at 19,800m, 16,750m, and 13,700m providing documentation of diminishing LPT performance with reductions in Reynolds number in an actual engine flight environment. The test provided a unique database for the development of engine analysis codes to be used for future LPT performance improvements.

Castner, Raymond↗
NASA NTRS2007

Overview of Engineering Design and Analysis at the NASA John C. Stennis Space Center

A wide range of rocket propulsion test work occurs at the NASA John C. Stennis Space Center (SSC) including full-scale engine test activities at test facilities A-1, A-2, B-1 and B-2 as well as combustion device research and development activities at the E-Complex (E-1, E-2, E-3 and E-4) test facilities. The propulsion test engineer at NASA SSC faces many challenges associated with designing and operating a test facility due to the extreme operating conditions (e.g., cryogenic temperatures, high pressures) of the various system components and the uniqueness of many of the components and systems. The purpose of this paper is to briefly describe the NASA SSC Engineering Science Directorate s design and analysis processes, experience, and modeling techniques that are used to design and support the operation of unique rocket propulsion test facilities.

Ryan, Harry↗
NASA NTRS1995

NASA Lewis Research Center's Preheated Combustor and Materials Test Facility

The Preheated Combustor and Materials Test Facility (PCMTF) in the Engine Research Building (ERB) at the NASA Lewis Research Center is one of two unique combustor facilities that provide a nonvitiated air supply to two test stands, where the air can be used for research combustor testing and high-temperature materials testing. Stand A is used as a research combustor stand, whereas stand B is used for cyclic and survivability tests of aerospace materials at high temperatures. Both stands can accommodate in-house and private industry research programs. The PCMTF is capable of providing up to 30 lb/s (pps) of nonvitiated, 450 psig combustion air at temperatures ranging from 850 to 1150 g F. A 5000 gal tank located outdoors adjacent to the test facility can provide jet fuel at a pressure of 900 psig and a flow rate of 11 gal/min (gpm). Gaseous hydrogen from a 70,000 cu ft (CF) tuber is also available as a fuel. Approximately 500 gpm of cooling water cools the research hardware and exhaust gases. Such cooling is necessary because the air stream reaches temperatures as high as 3000 deg F. The PCMTF provides industry and Government with a facility for studying the combustion process and for obtaining valuable test information on advanced materials. This report describes the facility's support systems and unique capabilities.

Nemets, Steve A.↗
NASA NTRS2004

NASA Glenn's Advanced Subsonic Combustion Rig Supported the Ultra-Efficient Engine Technology Project's Emissions Reduction Test

The Advanced Subsonic Combustor Rig (ASCR) is NASA Glenn Research Center's unique high-pressure, high-temperature combustor facility supporting the emissions reduction element of the Ultra-Efficient Engine Technology (UEET) Project. The facility can simulate combustor inlet test conditions up to a pressure of 900 psig and a temperature of 1200 F (non-vitiated). ASCR completed three sector tests in fiscal year 2003 for General Electric, Pratt & Whitney, and Rolls-Royce North America. This will provide NASA and U.S. engine manufacturers the information necessary to develop future low-emission combustors and will help them to better understand durability and operability at these high pressures and temperatures.

Beltran, Luis R.↗
NASA NTRS1982

Life prediction and constitutive behavior: Overview

The evolution of programs to investigate high temperature consititutive behavior and develop cyclic life prediction methods is reviewed. Contracts granted for developing and verifying workable engineering methods for the calculation, in advance of service, of the local stress-strain response at the critical life governing location in typical hot section components as well as the resultant cyclic crack initiation and crack growth lifetimes are listed. The Langley fatigue facility is being upgraded to include: (1) a servocontrolled testing machine for high temperature crack growth; (2) three servocontrolled tension/torsion machines for biaxial studies; (3) a HOST/satellite computer for data acquisition, processing, storage, and retrieval; and (4) HCV/LCF machines for cumulative damage studies.

Halford, G. R.↗