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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 451 records · Page 25
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Application of polyimide actuator rod seals

Development of polyimide two-stage hydraulic actuator rod seals for application in high-performance aircraft was accomplished. The significant portion of the effort was concentrated on optimization of the chevron and K-section second-stage seal geometries to satisfy the requirements for operation at 450 K (350 F) with dynamic pressure loads varying between 200 psig steady-state and 1500 psig impulse cycling. Particular significance was placed on reducing seal gland dimension by efficiently utilizing the fatigue allowables of polyimide materials. Other objectives included investigation of pressure balancing techniques for first-stage polyimide rod seals for 4000 psig 450 K(350 F) environment and fabrication of a modular retainer for the two-stage combination. Seals were fabricated in 0.0254 m (1.0in.) and 0.0635 m (2.5in.) sizes and tested for structural integrity, frictional resistance, and endurance life. Test results showed that carefully designed second stages using polyimides could be made to satisfy the dynamic return pressure requirements of applications in high-performance aircraft. High wear under full system pressure indicated that further research is necessary to obtain an acceptable first-stage design. The modular retainer was successfully tested and showed potential for new actuator applications.

Watermann, A. W.↗
DOE OSTI2022

Modeling of streamflow in a 30 km long reach spanning 5 years using OpenFOAM 5.x

Abstract. Developing accurate and efficient modeling techniques for streamflow at the tens-of-kilometers spatial scale and multi-year temporal scale is critical for evaluating and predicting the impact of climate- and human-induced discharge variations on river hydrodynamics. However, achieving such a goal is challenging because of limited surveys of streambed hydraulic roughness, uncertain boundary condition specifications, and high computational costs. We demonstrate that accurate and efficient three-dimensional (3-D) hydrodynamic modeling of natural rivers at 30 km and 5-year scales is feasible using the following three techniques within OpenFOAM, an open-source computational fluid dynamics platform: (1) generating a distributed hydraulic roughness field for the streambed by integrating water-stage observation data, a rough wall theory, and a local roughness optimization and adjustment strategy; (2) prescribing the boundary condition for the inflow and outflow by integrating precomputed results of a one-dimensional (1-D) hydraulic model with the 3-D model; and (3) reducing computational time using multiple parallel runs constrained by 1-D inflow and outflow boundary conditions. Streamflow modeling for a 30 km long reach in the Columbia River (CR) over 58 months can be achieved in less than 6 d using 1.1 million CPU hours. The mean error between the modeled and the observed water stages for our simulated CR reach ranges from −16 to 9 cm (equivalent to approximately ±7 % relative to the average water depth) at seven locations during most of the years between 2011 and 2019. We can reproduce the velocity distribution measured by the acoustic Doppler current profiler (ADCP). The correlation coefficients of the depth-averaged velocity between the model and ADCP measurements are in the range between 0.71 and 0.83 at 75 % of the survey cross sections. With the validated model, we further show that the relative importance of dynamic pressure versus hydrostatic pressure varies with discharge variations and topography heterogeneity. Given the model's high accuracy and computational efficiency, the model framework provides a generic approach to evaluate and predict the impacts of climate- and human-induced discharge variations on river hydrodynamics at tens-of-kilometers and decadal scales.

58 GEOSCIENCES↗
DOE OSTI2021

Dynamic compression of TiO 2 to 221 GPa

The high-pressure dynamic response of titanium dioxide (TiO 2 ) is not only of interest because of its numerous industrial applications but also because of its structural similarities to silica (SiO 2 ). We performed plate impact experiments in a two-stage light gas gun, at peak stresses from 64 to 221 GPa to determine the TiO 2 response along the Hugoniot. The lower stress experiment at 64 GPa shows an elastic behavior followed by an elastic–plastic transition, whereas the high stress experiments above 64 GPa show a single wave structure. Previous shock studies have shown the presence of high-pressure phases (HPP) I (26 GPa) and HPP II (100 GPa); however, our data suggest that the HPP I phase is stable up to 150 GPa. Using a combination of data from our current study and our previous Z-data, we determine that TiO 2 likely melts on the Hugoniot at 157 GPa. Furthermore, our data confirm that TiO 2 is not highly incompressible as shown by a previous study.

36 MATERIALS SCIENCE↗
NASA NTRS1983

An Automated Stall-Speed Warning System

The NASA Dryden Flight Research Facility embarked upon a project with the United States Army Aviation Engineering Flight Activity (USAAEFA) to develop and test a stall-speed warning system. NASA designed and built an automated stall-speed warning system which presents both airspeed and stall speed to the pilot. The airspeed and stall speed are computed in real time by monitoring the basic aerodynamic parameters (dynamic pressure, horizontal and vertical accelerations, and pressure altitude) and other parameters (elevator and flap positions, engine torques, and fuel flow). In addition, an aural warning at predetermined stall margins is presented to the pilot through a voice synthesizer. Once the system was designed and installed in the aircraft, a flight-test program of less than 20 hr was anticipated to determine the stall-speed software coefficients. These coefficients would then be inserted in the system's software and then test flown over a period of about 10 hr for the purposes of evaluation.

Dynamic Pressure↗
NASA NTRS1984

An Automated Stall-Speed Warning System

The NASA Dryden Flight Research Facility embarked upon a project with the United States Army Aviation Engineering Flight Activity (USAAEFA) to develop and test a stall-speed warning system. NASA designed and built an automated stall-speed warning system which presents both airspeed and stall speed to the pilot. The airspeed and stall speed are computed in real time by monitoring the basic aerodynamic parameters {dynamic pressure, horizontal and vertical accelerations, and pressure altitude) and other parameters (elevator and flap positions, engine torques, and fuel flow). In addition, an aural warning at predetermined stall margins is presented to the pilot through a voice synthesizer. Once the system was designed and installed in the aircraft, a flight-test program of less than 20 hr was anticipated to determine the stall-speed software coefficients. These coefficients would then be inserted in the system's software and then test flown over a period of about 10 hr for the purposes of evaluation.

Pilot↗
NASA NTRS1990

Analytical and experimental investigations of the oblique detonation wave engine concept

Wave combustors, which include the oblique detonation wave engine (ODWE), are attractive propulsion concepts for hypersonic flight. These engines utilize oblique shock or detonation waves to rapidly mix, ignite, and combust the air-fuel mixture in thin zones in the combustion chamber. Benefits of these combustion systems include shorter and lighter engines which require less cooling and can provide thrust at higher Mach numbers than conventional scramjets. The wave combustor's ability to operate at lower combustor inlet pressures may allow the vehicle to operate at lower dynamic pressures which could lessen the heating loads on the airframe. The research program at NASA-Ames includes analytical studies of the ODWE combustor using Computational Fluid Dynamics (CFD) codes which fully couple finite rate chemistry with fluid dynamics. In addition, experimental proof-of-concept studies are being performed in an arc heated hypersonic wind tunnel. Several fuel injection design were studied analytically and experimentally. In-stream strut fuel injectors were chosen to provide good mixing with minimal stagnation pressure losses. Measurements of flow field properties behind the oblique wave are compared to analytical predictions.

Menees, Gene P.↗
DOE OSTI2024

Large eddy simulation of a pressurized reactive fluid–particle system: single-particle dynamics analysis

Pressurized oxy-combustion (POC) is a promising candidate to reduce carbon emissions in power generation. Designing an effective burner plays a vital role in developing new combustion technologies. Because of the high pressure, the volume fraction of particles in a pressurized oxy-combustor could be higher than that in the conventional combustor, therefore, the particle dynamics in the pressurized vessel need to be evaluated accordingly. The present computational work is to predict the influence of the particle size and particle injection location on the particle trajectory. The Stokes number based on turbulence fluctuation, St, is adopted to evaluate the impact of the gas phase on the particulate phase. A large eddy simulation (LES) with the environment pressure of 15 bar and the thermal input of the fuel particle of 100 kWth has been performed by means of the commercial CFD package. It is shown that that in this pressurized environment St << 1 for the 25 µm particles, St ~ 1 for the 100 µm particles, while for the particles exceeding 200 µm St >> 1. The trajectories of the sampling particles from the LES results support this conclusion: the particles less than 100 µm tend to be uniformly distributed in the space while the particles exceeding 200 µm tend to concentrate in the combustor.

Li, Lei↗
NASA NTRS1974

Simulation of real-gas effects for Mars entry

The changes in aerodynamic characteristics due to real-gas effects associated with high speed flight (characterized by large shock density ratios) are primarily the result of changes in surface pressures acting on the forebody. The surface pressures are affected by a change in shock density ratio (real-gas effects) in two ways. First, the level of pressure at the stagnation point relative to freestream dynamic pressure is changed, and second, the distribution of surface pressure relative to stagnation-point pressure is changed. The density-ratio effect on the stagnation point pressure level can be estimated by considering the flow of a perfect gas about a blunt body.

Hunt, J. L.↗
NASA NTRS1976

A simple total energy sensor

This research concerns a simple technique for total energy compensation having the primary requirement of good compensation with a low cost, easy to make sensor. A comprehensive library search and a small wind tunnel were used to explore fundamentals and to develop a number of probes which were flight tested with satisfactory results. Simple probe configurations were tested using the characteristics of laminar flow separation around a small cylinder to produce a sensor pressure having the desired relationship between static and dynamic pressures. Wind tunnel and flight tests confirmed the concepts for a wide range of speeds, altitudes, and flow directions encountered in soaring. Data and findings are presented, along with a discussion of factors important to total energy compensation.

Nicks, O. W.↗
NASA NTRS1985

Experimental study of forward-located jet-interaction nozzles

The effects of forward-mounted sonic jet-interaction nozzles with one, three equal, or one large and two small orifices on the transitional or turbulent boundary-layer flow over a biconic (9-deg-25-arcmin/6-deg) model at angle of attack -10 to the 15 deg and Reynolds number 8.10 x 10 to the 6th/ft are investigated experimentally using pressure taps, force measurements, oil-flow visualization, and schlieren photography in the 20-in. Mach-6 wind tunel at NASA Langley. The injection flow rates are 0.055 and 0.026 lbm/s, with stagnation pressure 500 psia, stagnation temperature 1000 R, and dynamic pressure 8.7 psia. The results are presented in diagrams, graphs, and photographs and characterized. Moment amplification factors of 1.0 or better (maximum 1.8 at angle of attack 15 deg and flow rate 0.055 lbm/s) are observed for all nozzle types, but the configuration with three unequal orifices is found to give improvements of 10 percent over the single-orifice nozzle at a given flow rate.

Cavalleri, R. J.↗
NASA NTRS1943

Determination of general relations for the behavior of turbulent boundary layers

An analysis has been made of a considerable amount of data for turbulent boundary layers along wings and bodies of various shapes in order to determine the fundamental variables that control the development of turbulent boundary layers. It was found that the type of velocity distribution in the boundary layer could be expressed in terms of a single parameter. This parameter was chosen as the ratio of the displacement thickness to the momentum thickness of the boundary layer. The variables that control the development of the turbulent boundary layer apparently are: (1) the ratio of the nondimensional pressure gradient, expressed in terms of the local dynamic pressure outside the boundary layer and boundary-layer thickness, to the local skin-friction coefficient and (2) the shape of the boundary layer. An empirical equation has been developed in terms of these variables that, when used with the momentum equation and the skin-friction relation, makes it possible to trace the development of the turbulent boundary layer to the separation point.

Von Doenhoff, Albert E↗
NASA NTRS1931

The use of slots for increasing the lift of airplane wings

This report discusses in detail the results obtained with devices for preventing the separation of the boundary layer. In order to give an idea of the order of magnitude of the positive and negative pressures involved, we made a diagram of them, as measured at 15 degrees angle of attack on the Gottigen profile 387. The pressure scale is graduated in terms of the dynamic pressure q, which enables the evaluation of the positive and negative pressures at all velocities.

Haus, FR↗
NASA NTRS2001

Alleviation of Facility/Engine Interactions in an Open-Jet Scramjet Test Facility

Results of a series of shakedown tests to eliminate facility/engine interactions in an open-jet scramjet test facility are presented. The tests were conducted with the NASA DFX (Dual-Fuel eXperimental scramjet) engine in the NASA Langley Combustion Heated Scramjet Test Facility (CHSTF) in support of the Hyper-X program, The majority of the tests were conducted at a total enthalpy and pressure corresponding to Mach 5 flight at a dynamic pressure of 734 psf. The DFX is the largest engine ever tested in the CHSTF. Blockage, in terms of the projected engine area relative to the nozzle exit area, is 81% with the engine forebody leading edge aligned with the upper edge of the facility nozzle such that it ingests the nozzle boundary layer. The blockage increases to 95% with the engine forebody leading edge positioned 2 in. down in the core flow. Previous engines successfully tested in the CHSTF have had blockages of no more than 51%. Oil flow studies along with facility and engine pressure measurements were used to define flow behavior. These results guided modifications to existing aeroappliances and the design of new aeroappliances. These changes allowed fueled tests to be conducted without facility interaction effects in the data with the engine forebody leading edge positioned to ingest the facility nozzle boundary layer. Interaction effects were also reduced for tests with the engine forebody leading edge positioned 2 in. into the core flow, however some interaction effects were still evident in the engine data. A new shroud and diffuser have been designed with the goal of allowing fueled tests to be conducted with the engine forebody leading edge positioned in the core without facility interaction effects in the data. Evaluation tests of the new shroud and diffuser will be conducted once ongoing fueled engine tests have been completed.

Albertson, Cindy W.↗
NASA NTRS1991

Analytical and experimental investigations of the oblique detonation wave engine concept

Wave combustors, which include the Oblique Detonation Wave Engine (ODWE), are attractive propulsion concepts for hypersonic flight. These engines utilize oblique shock or detonation waves to rapidly mix, ignite, and combust the air-fuel mixture in thin zones in the combustion chamber. Benefits of these combustion systems include shorter and lighter engines which will require less cooling and can provide thrust at higher Mach numbers than conventional scramjets. The wave combustor's ability to operate at lower combustor inlet pressures may allow the vehicle to operate at lower dynamic pressures which could lessen the heating loads on the airframe. The research program at NASA-Ames includes analytical studies of the ODWE combustor using CFD codes which fully couple finite rate chemistry with fluid dynamics. In addition, experimental proof-of-concept studies are being carried out in an arc heated hypersonic wind tunnel. Several fuel injection designs were studied analytically and experimentally. In-stream strut fuel injectors were chosen to provide good mixing with minimal stagnation pressure losses. Measurements of flow field properties behind the oblique wave are compared to analytical predictions.

Menees, Gene P.↗
NASA NTRS2015

Concept Demonstration of Dopant Selective Reactive Etching (DSRIE) in Silicon Carbide

Accurate quantification of combustor pressure dynamics for the primary purpose of experimental validation of computational fluid dynamics (CFD) codes requires the use of robust, reliable and sensitive pressure sensors that can resolve sub--pound-per-square-inch pressure levels in high temperature environments (i.e., combustor). The state of the art microfabricated piezoresistive silicon carbide (SiC) pressure sensors that we have developed are capable of operating reliably at 600 degrees Centigrade. This technology was used in support of the ARMD ISRP-ERA (NASA's Aeronautics Research Mission Directorate, Integrated System Research Project - Environmentally Responsible Aviation) program to quantify combustor thermoacoustic instabilities. The results showed that while the SiC pressure sensors survived the high temperature and measured instabilities, the diaphragm (force collector) was not thin enough to be sensitive in resolving sub-pound-per-square-inch pressures; 30 meters is the thinnest diaphragm achievable with conventional reactive ion etching (RIE) processes. Therefore, this precludes its use for sub-pound-per-square-inch pressure measurement with high fidelity. In order to effectively resolve sub-pound-per-square-inch pressures, a thinner more sensitive diaphragm (10 meters) is needed. To achieve this would require a new and innovative fabrication process technique.

static↗
NASA NTRS1966

Acceleration-compensated pressure transducer has fast response

Flush-diaphragm transducer accurately measures small dynamic pressures when it is subjected to high accelerations and severe temperature environments. The transducer uses piezoelectric crystals for measuring the pressure and balancing out acceleration forces.

Source record↗
NASA NTRS1980

Transonic unsteady airloads on an energy efficient transport wing with oscillating control surfaces

An aspect ratio 10.8 supercritical wing with oscillating control surfaces is described. The wing is instrumental with 252 static orifices and 164 in situ dynamic pressure transducers for studying the effects of control surface deflection on steady and unsteady pressures at transonic speeds. Results from initial wind tunnel tests conducted in the Langley Transonic Dynamics Tunnel are discussed. Unsteady pressure results are presented for two trailing edge control surfaces oscillating separately at the design Mach number of 0.78. Some experimental results are compared with analytical results obtained by using linear lifting surface theory.

Sandford, M. C.↗