The ground vortex flow field associated with a jet in a cross flow impinging on a ground plane for uniform and annular turbulent axisymmetric jets. M.S. Thesis

NASA Technical Reports Server (NTRS)

Cavage, William M.; Kuhlman, John M.

1993-01-01

An experimental study was conducted of the impingement of a single circular jet on a ground plane in a cross flow. This geometry is a simplified model of the interaction of propulsive jet exhaust from a V/STOL aircraft with the ground in forward flight. Jets were oriented normal to the cross flow and ground plane. Jet size, cross flow-to-jet velocity ratio, ground plane-to-jet board spacing, and jet exit turbulence level and mean velocity profile shape were all varied to determine their effects on the size of the ground vortex interaction region which forms on the ground plane, using smoke injection into the jet. Three component laser Doppler velocimeter measurements were made with a commercial three color system for the case of a uniform jet with exit spacing equal to 5.5 diameters and cross flow-to-jet velocity ratio equal to 0.11. The flow visualization data compared well for equivalent runs of the same nondimensional jet exit spacing and the same velocity ratio for different diameter nozzles, except at very low velocity ratios and for the larger nozzle, where tunnel blockage became significant. Variation of observed ground vortex size with cross flow-to-jet velocity ratio was consistent with previous studies. Observed effects of jet size and ground plane-to-jet board spacing were relatively small. Jet exit turbulence level effects were also small. However, an annular jet with a low velocity central core was found to have a significantly smaller ground vortex than an equivalent uniform jet at the same values of cross flow-to-jet velocity ratio and jet exit-to-ground plane spacing. This may suggest a means of altering ground vortex behavior somewhat, and points out the importance of proper simulation of jet exit velocity conditions. LV data indicated unsteady turbulence levels in the ground vortex in excess of 70 percent.

Computational and Experimental Characterization of the Mach 6 Facility Nozzle Flow for the Enhanced Injection and Mixing Project at NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Drozda, Tomasz G.; Cabell, Karen F.; Passe, Bradley J.; Baurle, Robert A.

2017-01-01

Computational fluid dynamics analyses and experimental data are presented for the Mach 6 facility nozzle used in the Arc-Heated Scramjet Test Facility for the Enhanced Injection and Mixing Project (EIMP). This project, conducted at the NASA Langley Research Center, aims to investigate supersonic combustion ramjet (scramjet) fuel injection and mixing physics relevant to flight Mach numbers greater than 8. The EIMP experiments use a two-dimensional Mach 6 facility nozzle to provide the high-speed air simulating the combustor entrance flow of a scramjet engine. Of interest are the physical extent and the thermodynamic properties of the core flow at the nozzle exit plane. The detailed characterization of this flow is obtained from three-dimensional, viscous, Reynolds-averaged simulations. Thermodynamic nonequilibrium effects are also investigated. The simulations are compared with the available experimental data, which includes wall static pressures as well as in-stream static pressure, pitot pressure and total temperature obtained via in-stream probes positioned just downstream of the nozzle exit plane.

Emission spectroscopy and laser-induced fluorescence measurements on the plume from a 1-kW arcjet operated on simulated ammonia

NASA Technical Reports Server (NTRS)

Ruyten, Wilhelmus M.; Burtner, D.; Keefer, Dennis

1993-01-01

Spectroscopic and laser-induced fluorescence measurements were performed on the exhaust plume from a 1 kW NASA Lewis arcjet, operated on simulated ammonia. In particular, emissions were analyzed from the Balmer lines of atomic hydrogen and from one of the rotational bands of the NH radical. The laser-induced fluorescence measurements were performed on the Balmer-alpha line of atomic hydrogen. We find that exit plane temperatures are in the range 1500 to 3500 K and that the electron density upstream of the exit plane is on the order of 1.5 x 10(exp 14)/cu cm as determined by the Stark width of the Balmer-alpha line. Both emission spectroscopy and laser-induced fluorescence were used to measure the plume velocities of atomic hydrogen. Using either technique, velocities on the order of 4 km/sec were found at the exit plane and significant acceleration of the flow was observed in the first 2 mm beyond the exit plane. This result indicates that the design of the arcjet nozzle may not be optimum.

The Kinetic Nonequilibrium Processes in the Internal Flow and in the Plume of Subsonic and Supersonic Aircrafts

NASA Technical Reports Server (NTRS)

Starik, Alexander M.

1997-01-01

(1) Our results show that under combustion of thermal destruction products of n-C8H18, and other hydrocarbon fuels with air at the equivalent ratio -0.5 and less the chemical equilibrium is not realized at the exit plane of combustion chamber and in the gas turbine and nozzle for most of small components such as NO2, NO3, HNO, HNO2, HNO3, N(x)H(y), HO2, OH. The chemical equilibrium is not realized in the internal flow of ramjet hydrogen combustion engine too. So at the nozzle exit plane both of gas-turbine hydrocarbon combustion engine and of ramjet hydrogen combustion engine the relatively large values of concentration of such small components as NO3, HNO2, N2O, HNO3, HNO, NH, N2H, HO2, H2O2 may be realized. The exact definition of these component concentration as well as concentration of NO(x), OH, SO2, O, H, H2, H2O at the nozzle exit plane is very important for plume chemistry. (2) The results which were obtained for subsonic and hypersonic aircrafts indicate on the considerable change of the composition of the gas mixture along the plume. This change can be caused not only by the mixture of combustion products with the atmosphere air but by proceeding of whole complex of nonequilibrium photochemical reactions. The photodissociation processes begin to influence on the formation of the free atoms and radicals at flight altitude H greater than or equal to 18 km. Neglect of these processes can result in essential (up to 10(exp 4) times) mistakes of values gamma(sub OH), gamma(sub O), gamma(sub H), gamma(sub HSO3) and some products of CFC's disintegration. It was found that penetration of Cl-containing species from the atmosphere into the exhaust flow and its interaction with nitrogen oxides leads to essential increasing of the concentration of Cl, Cl2, ClO2, ClNO3, CH3Cl and sometimes HCl and the decreasing of ClO concentration by comparison with background values. The results of our analysis show that the plume aircraft with both hydrocarbon and hydrogen combustion engine may be source of various pollutant components such as HNO, HNO4,ClO2, CH3NO2, CH3NO3, CH2O, Cl, H2O2, but not only NO, NO2, HNO2, HNO3, N2O5, SO2, SO3, H2SO4 as it was supposed before.

Gaseous hydrogen/oxygen injector performance characterization

NASA Technical Reports Server (NTRS)

Degroot, W. A.; Tsuei, H. H.

1994-01-01

Results are presented of spontaneous Raman scattering measurements in the combustion chamber of a 110 N thrust class gaseous hydrogen/oxygen rocket. Temperature, oxygen number density, and water number density profiles at the injector exit plane are presented. These measurements are used as input profiles to a full Navier-Stokes computational fluid dynamics (CFD) code. Predictions of this code while using the measured profiles are compared with predictions while using assumed uniform injector profiles. Axial and radial velocity profiles derived from both sets of predictions are compared with Rayleigh scattering measurements in the exit plane of a 33:1 area ratio nozzle. Temperature and number density Raman scattering measurements at the exit plane of a test rocket with a 1:1.36 area ratio nozzle are also compared with results from both sets of predictions.

Investigation of impingement region and wall jets formed by the interaction of high aspect ratio lift jets and a ground plane

NASA Technical Reports Server (NTRS)

Kotansky, D. R.; Glaze, L. W.

1978-01-01

Flow characteristics of impinging jets emanating from rectangular exit area converging nozzles of exit area aspect ratio four, six, and eight were investigated. Azimuthal distributions of wall jet radial momentum flux in the ground plane were strongly directional and sensitive to rectangular nozzle exit area aspect ratio, jet impingement angle, and height above ground, H/D. Effects of jet exit velocity profile nonuniformities were also investigated. Data from the single nozzle rectangular jet impringement investigations were incorporated into an existing VTOL aircraft ground flow field computer program. It is suggested that this program together with the Douglas Neumann program modified for V/STOL applications may be used for the analysis and prediction of flow fields and resulting forces and moments on multijet V/STOL aircraft hovering in ground effect.

Mixing of Supersonic Streams

NASA Technical Reports Server (NTRS)

Hawk, C. W.; Landrum, D. B.; Muller, S.; Turner, M.; Parkinson, D.

1998-01-01

The Strutjet approach to Rocket Based Combined Cycle (RBCC) propulsion depends upon fuel-rich flows from the rocket nozzles and turbine exhaust products mixing with the ingested air for successful operation in the ramjet and scramjet modes. It is desirable to delay this mixing process in the air-augmented mode of operation present during low speed flight. A model of the Strutjet device has been built and is undergoing test to investigate the mixing of the streams as a function of distance from the Strutjet exit plane during simulated low speed flight conditions. Cold flow testing of a 1/6 scale Strutjet model is underway and nearing completion. Planar Laser Induced Fluorescence (PLIF) diagnostic methods are being employed to observe the mixing of the turbine exhaust gas with the gases from both the primary rockets and the ingested air simulating low speed, air augmented operation of the RBCC. The ratio of the pressure in the turbine exhaust duct to that in the rocket nozzle wall at the point of their intersection is the independent variable in these experiments. Tests were accomplished at values of 1.0, 1.5 and 2.0 for this parameter. Qualitative results illustrate the development of the mixing zone from the exit plane of the model to a distance of about 10 rocket nozzle exit diameters downstream. These data show the mixing to be confined in the vertical plane for all cases, The lateral expansion is more pronounced at a pressure ratio of 1.0 and suggests that mixing with the ingested flow would be likely beginning at a distance of 7 nozzle exit diameters downstream of the nozzle exit plane.

Measurement and analysis of a small nozzle plume in vacuum

NASA Technical Reports Server (NTRS)

Penko, P. F.; Boyd, I. D.; Meissner, D. L.; Dewitt, K. J.

1993-01-01

Pitot pressures and flow angles are measured in the plume of a nozzle flowing nitrogen and exhausting to a vacuum. Total pressures are measured with Pitot tubes sized for specific regions of the plume and flow angles measured with a conical probe. The measurement area for total pressure extends 480 mm (16 exit diameters) downstream of the nozzle exit plane and radially to 60 mm (1.9 exit diameters) off the plume axis. The measurement area for flow angle extends to 160 mm (5 exit diameters) downstream and radially to 60 mm. The measurements are compared to results from a numerical simulation of the flow that is based on kinetic theory and uses the direct-simulation Monte Carlo (DSMC) method. Comparisons of computed results from the DSMC method with measurements of flow angle display good agreement in the far-field of the plume and improve with increasing distance from the exit plane. Pitot pressures computed from the DSMC method are in reasonably good agreement with experimental results over the entire measurement area.

Experimental study of a wake behind a barrier

NASA Astrophysics Data System (ADS)

Tomáš, Dufek; Katarína, Ratkovská

2017-09-01

This article describes in detail an experiment which was carried out on a wind tunnel in the Laboratory of the Department of Power Machines, Faculty of Mechanical Engineering, at the University of West Bohemia (UWB), using Particle Image Velocimetry and Stereo Particle Image Velocimetry. PIV is a non-invasive method that allows you to simultaneously measure the flow velocity across the entire field under investigation. In the experiment, the field was located behind the exit of the wind tunnel. The experiment dealt with the measurement of the wake behind a barrier. Measurement with Stereo PIV was carried out in several vertical parallel planes perpendicular to the axis of the tunnel. Conventional PIV method was then used for a horizontal plane passing through the axis of the tunnel at half the height of the barrier. The velocities in the measured plane are expressed by a vector map. In areas not affected by the wake, the speed in the w direction is about 16 m / s. The wake is formed behind the barrier. A shear layer is formed at the boundary between the flowing air and the braked air. A backflow occurs in the area just behind the barrier. The highest speed in the area is achieved in places just behind the exit of the tunnel, where the current is not affected by the barrier. In the direction from the axis and the obstacle, the speed gradually rises from the negative values of the return flow through the zero speed. In addition to the velocity fields, the output from the experimental measurement was also the distribution of the sum of variances, standard deviation and correlation coefficient in the measured planes.

High resolution EUV monochromator/spectrometer

DOEpatents

Koike, Masako

1996-01-01

This invention is related to a monochromator which employs a spherical mirror, a traveling plane mirror with simultaneous rotation, and a varied spacing plane grating. The divergent beam from the entrance slit is converged by the spherical mirror located at the various positions in the monochromator depending of the inventive system. To provide the meaningful diffraction efficiencies and to reduce unwanted higher order lights, the deviation angle subtending the incidence and diffraction beams for the plane grating is varied with the position of the traveling plane mirror with simultaneous rotation located in the front or back of the plane grating with wavelength scanning. The outgoing beam from the monochromator goes through the fixed exit slit and has same beam direction regardless of the scanning wavelength. The combination of properly designed motions of the plane mirror and novel varied-spacing parameters of the inventive plane grating corrects the aberrations and focuses the monochromatic spectral image on the exit slit, enabling measurements at high spectral resolution.

High resolution EUV monochromator/spectrometer

DOEpatents

Koike, Masako

1996-06-18

This invention is related to a monochromator which employs a spherical mirror, a traveling plane mirror with simultaneous rotation, and a varied spacing plane grating. The divergent beam from the entrance slit is converged by the spherical mirror located at the various positions in the monochromator depending of the inventive system. To provide the meaningful diffraction efficiencies and to reduce unwanted higher order lights, the deviation angle subtending the incidence and diffraction beams for the plane grating is varied with the position of the traveling plane mirror with simultaneous rotation located in the front or back of the plane grating with wavelength scanning. The outgoing beam from the monochromator goes through the fixed exit slit and has same beam direction regardless of the scanning wavelength. The combination of properly designed motions of the plane mirror and novel varied-spacing parameters of the inventive plane grating corrects the aberrations and focuses the monochromatic spectral image on the exit slit, enabling measurements at high spectral resolution. 10 figs.

Saturn systems holddown acoustic efficiency and normalized acoustic power spectrum.

NASA Technical Reports Server (NTRS)

Gilbert, D. W.

1972-01-01

Saturn systems field acoustic data are used to derive mid- and far-field prediction parameters for rocket engine noise. The data were obtained during Saturn vehicle launches at the Kennedy Space Center. The data base is a sorted set of acoustic data measured during the period 1961 through 1971 for Saturn system launches SA-1 through AS-509. The model assumes hemispherical radiation from a simple source located at the intersection of the longitudinal axis of each booster and the engine exit plane. The model parameters are evaluated only during vehicle holddown. The acoustic normalized power spectrum and efficiency for each system are isolated as a composite from the data using linear numerical methods. The specific definitions of each allows separation. The resulting power spectra are nondimensionalized as a function of rocket engine parameters. The nondimensional Saturn system acoustic spectrum and efficiencies are compared as a function of Strouhal number with power spectra from other systems.

Active Combustion Control for Aircraft Gas Turbine Engines

NASA Technical Reports Server (NTRS)

DeLaat, John C.; Breisacher, Kevin J.; Saus, Joseph R.; Paxson, Daniel E.

2000-01-01

Lean-burning combustors are susceptible to combustion instabilities. Additionally, due to non-uniformities in the fuel-air mixing and in the combustion process, there typically exist hot areas in the combustor exit plane. These hot areas limit the operating temperature at the turbine inlet and thus constrain performance and efficiency. Finally, it is necessary to optimize the fuel-air ratio and flame temperature throughout the combustor to minimize the production of pollutants. In recent years, there has been considerable activity addressing Active Combustion Control. NASA Glenn Research Center's Active Combustion Control Technology effort aims to demonstrate active control in a realistic environment relevant to aircraft engines. Analysis and experiments are tied to aircraft gas turbine combustors. Considerable progress has been shown in demonstrating technologies for Combustion Instability Control, Pattern Factor Control, and Emissions Minimizing Control. Future plans are to advance the maturity of active combustion control technology to eventual demonstration in an engine environment.

Experiments on The Wakes of Multiple Nozzle Cusps

DTIC Science & Technology

1977-11-01

is sensitive to the Mach and Reynolds number; a "recovery factor calibration" is then an important prerequisite to using this probe. This sensor ...middle nozzle at its exit plane. Here the sensor was fixed first at y ■ 0 (the vertical center of the nozzle exit plane at x ■ 0) and then...traversed from z ■ 0.6 inches to -0.6 inches, which includes a little less than half the span (figure 2). By thus traversing the sensor parallel to the two

Mixing of Supersonic Streams

NASA Technical Reports Server (NTRS)

Hawk, C. W.; Landrum, D. B.; Muller, S.; Turner, M.; Parkinson, D.

1998-01-01

The Strutjet approach to Rocket Based Combined Cycle (RBCC) propulsion depends upon fuel-rich flows from the rocket nozzles and turbine exhaust products mixing with the ingested air for successful operation in the ramjet and scramjet modes. It is desirable to delay this mixing process in the air-augmented mode of operation present during low speed flight. A model of the Strutjet device has been built and is undergoing test to investigate the mixing of the streams as a function of distance from the Strutjet exit plane during simulated low speed flight conditions. Cold flow testing of a 1/6 scale Strutjet model is underway and nearing completion. Planar Laser Induced Fluorescence (PLIF) diagnostic methods are being employed to observe the mixing of the turbine exhaust gas with the gases from both the primary rockets and the ingested air simulating low speed, air augmented operation of the RBCC. The ratio of the pressure in the turbine exhaust duct to that in the rocket nozzle wall at the point of their intersection is the independent variable in these experiments. Tests were accomplished at values of 1.0, 1.5 and 2.0 for this parameter. Qualitative results illustrate the development of the mixing zone from the exit plane of the model to a distance of about 19 equivalent rocket nozzle exit diameters downstream. These data show the mixing to be confined in the vertical plane for all cases, The lateral expansion is more pronounced at a pressure ratio of 1.0 and suggests that mixing with the ingested flow would be likely beginning at a distance of 7 nozzle exit diameters downstream of the nozzle exit plane.

Method for Making Measurements of the Post-Combustion Residence Time in a Gas Turbine Engine

NASA Technical Reports Server (NTRS)

Miles, Jeffrey H (Inventor)

2015-01-01

A system and method of measuring a residence time in a gas-turbine engine is provided, whereby the method includes placing pressure sensors at a combustor entrance and at a turbine exit of the gas-turbine engine and measuring a combustor pressure at the combustor entrance and a turbine exit pressure at the turbine exit. The method further includes computing cross-spectrum functions between a combustor pressure sensor signal from the measured combustor pressure and a turbine exit pressure sensor signal from the measured turbine exit pressure, applying a linear curve fit to the cross-spectrum functions, and computing a post-combustion residence time from the linear curve fit.

Why Do Women Leave Science and Engineering? NBER Working Paper No. 15853

ERIC Educational Resources Information Center

Hunt, Jennifer

2010-01-01

I use the 1993 and 2003 National Surveys of College Graduates to examine the higher exit rate of women compared to men from science and engineering relative to other fields. I find that the higher relative exit rate is driven by engineering rather than science, and show that 60% of the gap can be explained by the relatively greater exit rate from…

Pressure measurements in a low-density nozzle plume for code verification

NASA Technical Reports Server (NTRS)

Penko, Paul F.; Boyd, Iain D.; Meissner, Dana L.; Dewitt, Kenneth J.

1991-01-01

Measurements of Pitot pressure were made in the exit plane and plume of a low-density, nitrogen nozzle flow. Two numerical computer codes were used to analyze the flow, including one based on continuum theory using the explicit MacCormack method, and the other on kinetic theory using the method of direct-simulation Monte Carlo (DSMC). The continuum analysis was carried to the nozzle exit plane and the results were compared to the measurements. The DSMC analysis was extended into the plume of the nozzle flow and the results were compared with measurements at the exit plane and axial stations 12, 24 and 36 mm into the near-field plume. Two experimental apparatus were used that differed in design and gave slightly different profiles of pressure measurements. The DSMC method compared well with the measurements from each apparatus at all axial stations and provided a more accurate prediction of the flow than the continuum method, verifying the validity of DSMC for such calculations.

IR signature study of aircraft engine for variation in nozzle exit area

NASA Astrophysics Data System (ADS)

Baranwal, Nidhi; Mahulikar, Shripad P.

2016-01-01

In general, jet engines operate with choked nozzle during take-off, climb and cruise, whereas unchoking occurs while landing and taxiing (when engine is not running at full power). Appropriate thrust in an aircraft in all stages of the flight, i.e., take-off, climb, cruise, descent and landing is achieved through variation in the nozzle exit area. This paper describes the effect on thrust and IR radiance of a turbojet engine due to variation in the exit area of a just choked converging nozzle (Me = 1). The variations in the nozzle exit area result in either choking or unchoking of a just choked converging nozzle. Results for the change in nozzle exit area are analyzed in terms of thrust, mass flow rate and specific fuel consumption. The solid angle subtended (Ω) by the exhaust system is estimated analytically, for the variation in nozzle exit area (Ane), as it affects the visibility of the hot engine parts from the rear aspect. For constant design point thrust, IR radiance is studied from the boresight (ϕ = 0°, directly from the rear side) for various percentage changes in nozzle exit area (%ΔAne), in the 1.9-2.9 μm and 3-5 μm bands.

Effect of Geometric Parameters on the Performance of Second Throat Annular Steam Ejectors

DTIC Science & Technology

1991-07-01

Cell Pressure versus Rake Average Exit Pitot Pressure . . . . . . . . . . . 42 15. Baseline Wall Pressure Profiles...diffuser exit plane pitot pressure rake . 2.5.2 Alternate Configurations Six alternate ejector diffuser configurations were tested. A summary of...along the walls of the diffusers to help characterize the flow. The ejector diffuser exit pitot pressure was measured with a 6-probe pitot pressure rake

Near Field Pressure Fluctuations in the Exit Plane of a Choked Axisymmetric Nozzle

NASA Technical Reports Server (NTRS)

Ponton, Michael K.; Seiner, John M.; Brown, Martha C.

1997-01-01

Nearfield pressure data are presented for an unheated jet issuing from an underexpanded sonic nozzle for two exit lip thicknesses of 0.200 and 0.625 nozzle diameters. Fluctuating measurements were obtained on the nozzle exit surface as well as in the acoustic nearfield. Narrowband spectra are presented for numerous operating conditions expressed in terms of the fully expanded Mach number based on nozzle pressure ratio.

Application of the RNS3D Code to a Circular-Rectangular Transition Duct With and Without Inlet Swirl and Comparison with Experiments

NASA Technical Reports Server (NTRS)

Cavicchi, Richard H.

1999-01-01

Circular-rectangular transition ducts are used between engine exhausts and nozzles with rectangular cross sections that are designed for high performance aircraft. NASA Glenn Research Center has made experimental investigations of a series of circular-rectangular transition ducts to provide benchmark flow data for comparison with numerical calculations. These ducts are all designed with superellipse cross sections to facilitate grid generation. In response to this challenge, the three-dimensional RNS3D code has been applied to one of these transition ducts. This particular duct has a length-to-inlet diameter ratio of 1.5 and an exit-plane aspect ratio of 3.0. The inlet Mach number is 0.35. Two GRC experiments and the code were run for this duct without inlet swirl. One GRC experiment and the code were also run with inlet swirl. With no inlet swirl the code was successful in predicting pressures and secondary flow conditions, including a pair of counter-rotating vortices at both sidewalls of the exit plane. All these phenomena have been reported from the two GRC experiments. However, these vortices were suppressed in the one experiment when inlet swirl was used; whereas the RNS3D code still predicted them. The experiment was unable to provide data near the sidewalls, the very region where the vortices were predicted.

Numerical Simulations of Noise Generated by High Aspect Ratio Supersonic Rectangular Jets - Validation

NASA Astrophysics Data System (ADS)

Viswanath, Kamal; Johnson, Ryan; Kailasanath, Kailas; Malla, Bhupatindra; Gutmark, Ephraim

2017-11-01

The noise from high performance jet engines of both civilian and military aircraft is an area of active concern. Asymmetric exhaust nozzle configurations, in particular rectangular, potentially offer a passive way of modulating the farfield noise and are likely to become more important in the future. High aspect ratio nozzles offer the further benefit of easier airframe integration. In this study we validate the far field noise for ideally and over expanded supersonic jets issuing from a high aspect ratio rectangular nozzle geometry. Validation of the acoustic data is performed against experimentally recorded sound pressure level (SPL) spectra for a host of observer locations around the asymmetric nozzle. Data is presented for a slightly heated jet case for both nozzle pressure ratios. The contrast in the noise profile from low aspect ratio rectangular and circular nozzle jets are highlighted, especially the variation in the azimuthal direction that shows ``quiet'' and ``loud'' planes in the farfield in the peak noise direction. This variation is analyzed in the context of the effect of mixing at the sharp corners, the sense of the vortex pairs setup in the exit plane, and the evolution of the high aspect ratio exit cross-section as it propagates downstream including possible axis-switching. Supported by Office of Naval Research (ONR) through the Computational Physics Task Area under the NRL 6.1 Base Program.

Overview of High-Fidelity Modeling Activities in the Numerical Propulsion System Simulations (NPSS) Project

NASA Technical Reports Server (NTRS)

Veres, Joseph P.

2002-01-01

A high-fidelity simulation of a commercial turbofan engine has been created as part of the Numerical Propulsion System Simulation Project. The high-fidelity computer simulation utilizes computer models that were developed at NASA Glenn Research Center in cooperation with turbofan engine manufacturers. The average-passage (APNASA) Navier-Stokes based viscous flow computer code is used to simulate the 3D flow in the compressors and turbines of the advanced commercial turbofan engine. The 3D National Combustion Code (NCC) is used to simulate the flow and chemistry in the advanced aircraft combustor. The APNASA turbomachinery code and the NCC combustor code exchange boundary conditions at the interface planes at the combustor inlet and exit. This computer simulation technique can evaluate engine performance at steady operating conditions. The 3D flow models provide detailed knowledge of the airflow within the fan and compressor, the high and low pressure turbines, and the flow and chemistry within the combustor. The models simulate the performance of the engine at operating conditions that include sea level takeoff and the altitude cruise condition.

Extractive sampling and optical remote sensing of F100 aircraft engine emissions.

PubMed

Cowen, Kenneth; Goodwin, Bradley; Joseph, Darrell; Tefend, Matthew; Satola, Jan; Kagann, Robert; Hashmonay, Ram; Spicer, Chester; Holdren, Michael; Mayfield, Howard

2009-05-01

The Strategic Environmental Research and Development Program (SERDP) has initiated several programs to develop and evaluate techniques to characterize emissions from military aircraft to meet increasingly stringent regulatory requirements. This paper describes the results of a recent field study using extractive and optical remote sensing (ORS) techniques to measure emissions from six F-15 fighter aircraft. Testing was performed between November 14 and 16, 2006 on the trim-pad facility at Tyndall Air Force Base in Panama City, FL. Measurements were made on eight different F100 engines, and the engines were tested on-wing of in-use aircraft. A total of 39 test runs were performed at engine power levels that ranged from idle to military power. The approach adopted for these tests involved extractive sampling with collocated ORS measurements at a distance of approximately 20-25 nozzle diameters downstream of the engine exit plane. The emission indices calculated for carbon dioxide, carbon monoxide, nitric oxide, and several volatile organic compounds showed very good agreement when comparing the extractive and ORS sampling methods.

Experimental investigation of five parallel plane jets with variation of Reynolds number and outlet conditions

NASA Astrophysics Data System (ADS)

Daubner, Tomas; Kizhofer, Jens; Dinulescu, Mircea

2018-06-01

This article describes an experimental investigation in the near field of five parallel plane jets. The study applies 2D Particle Image Velocimetry (PIV) for ventilated and unventilated jets, where ventilated means exiting into a duct with expansion ratio 3.5 and unventilated means exiting to the free atmosphere. Results are presented for Reynolds numbers 1408, 5857 and 10510. The Reynolds number is calculated for the middle channel and is based on the height of the nozzle (channel) equivalent diameter 2h. All characteristic regions of the methodology to describe multiple interacting jets are observed by the PIV measurements - converging, merging and combined. Each of the five parallel channels has an aspect ratio of 25 defined as nozzle width (w) to height (h). The channels have a length of 185 times the channel height guaranteeing a fully developed velocity profile at the exit from the channel. Spacing between the single plane jets is 3 times the channel height. The near field of multiple mixing jets is depended on outlet nozzle geometry. Blunt geometry of the nozzle was chosen (sudden contraction).

Species and temperature measurement in H2/O2 rocket flow fields by means of Raman scattering diagnostics

NASA Technical Reports Server (NTRS)

De Groot, Wim A.; Weiss, Jonathan M.

1992-01-01

Validation of CFD codes developed for prediction and evaluation of rocket performance is hampered by a lack of experimental data. Nonintrusive laser based diagnostics are needed to provide spatially and temporally resolved gas dynamic and fluid dynamic measurements. This paper reports the first nonintrusive temperature and species measurements in the plume of a 110 N gaseous hydrogen/oxygen thruster at and below ambient pressures, obtained with spontaneous Raman spectroscopy. Measurements at 10 mm downstream of the exit plane are compared with predictions from a numerical solution of the axisymmetric Navier-Stokes and species transport equations with chemical kinetics, which fully model the combustor-nozzle-plume flowfield. The experimentally determined oxygen number density at the centerline at 10 mm downstream of the exit plane is four times that predicted by the model. The experimental number density data fall between those numerically predicted for the exit and 10 mm downstream planes in both magnitude and radial gradient. The predicted temperature levels are within 10 to 15 percent of measured values.

Engineering Development Tests Airdrop Controlled Exit System (ACES)

DTIC Science & Technology

1980-09-01

AIRDROP CONTROLLED EXIT SYSTEM ( ACES ) RECOVERY PARACHUTES TELEMETERING DATA 20. D5TFAC c• Cat •u•u am revers e• ift n•ceesafy ad Ide•lityf by block...rTECHNICAL REPORT , NATICK /TR-82 /017 f C’n Engineering Development Tests Airdropý Controlled Exit System ( ACES ) COPY CLV40ble to DTIC doe’ io C...and,50.,,,10) s. TYPE OF REPORT A PERIOn COVEnEo Test Report ENCINEERTNG DEVELOPMENT TESTS Oct 79 - Apr 80 AIRDROP CONTROLLED EXIT SYSTEM ( ACES ) 6

Jet engine nozzle exit configurations and associated systems and methods

NASA Technical Reports Server (NTRS)

Mengle, Vinod G. (Inventor)

2011-01-01

Nozzle exit configurations and associated systems and methods are disclosed. An aircraft system in accordance with one embodiment includes a jet engine exhaust nozzle having an internal flow surface and an exit aperture, with the exit aperture having a perimeter that includes multiple projections extending in an aft direction. Aft portions of individual neighboring projections are spaced apart from each other by a gap, and a geometric feature of the multiple can change in a monotonic manner along at least a portion of the perimeter.

Jet Engine Nozzle Exit Configurations and Associated Systems and Methods

NASA Technical Reports Server (NTRS)

Mengle, Vinod G. (Inventor)

2013-01-01

Nozzle exit configurations and associated systems and methods are disclosed. An aircraft system in accordance with one embodiment includes a jet engine exhaust nozzle having an internal flow surface and an exit aperture, with the exit aperture having a perimeter that includes multiple projections extending in an aft direction. Aft portions of individual neighboring projections are spaced apart from each other by a gap, and a geometric feature of the multiple can change in a monotonic manner along at least a portion of the perimeter.

Tests of a D vented thrust deflecting nozzle behind a simulated turbofan engine

NASA Technical Reports Server (NTRS)

Watson, T. L.

1982-01-01

A D vented thrust deflecting nozzle applicable to subsonic V/STOL aircraft was tested behind a simulated turbofan engine in the verticle thrust stand. Nozzle thrust, fan operating characteristics, nozzle entrance conditions, and static pressures were measured. Nozzle performance was measured for variations in exit area and thrust deflection angle. Six core nozzle configurations, the effect of core exit axial location, mismatched core and fan stream nozzle pressure ratios, and yaw vane presence were evaluated. Core nozzle configuration affected performance at normal and engine out operating conditions. Highest vectored nozzle performance resulted for a given exit area when core and fan stream pressure were equal. Its is concluded that high nozzle performance can be maintained at both normal and engine out conditions through control of the nozzle entrance Mach number with a variable exit area.

Method for Making Measurements of the Post-Combustion Residence Time in a Gas Turbine Engine

NASA Technical Reports Server (NTRS)

Miles, Jeffrey H. (Inventor)

2017-01-01

A method of measuring a residence time in a gas-turbine engine is disclosed that includes measuring a combustor pressure signal at a combustor entrance and a turbine exit pressure signal at a turbine exit. The method further includes computing a cross-spectrum function between the combustor pressure signal and the turbine exit pressure signal, calculating a slope of the cross-spectrum function, shifting the turbine exit pressure signal an amount corresponding to a time delay between the measurement of the combustor pressure signal and the turbine exit pressure signal, and recalculating the slope of the cross-spectrum function until the slope reaches zero.

ETR WASTE GAS EXITED THE ETR COMPLEX FROM THE NORTH ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

ETR WASTE GAS EXITED THE ETR COMPLEX FROM THE NORTH SIDE THROUGH A TUNNEL AND THEN TO A FILTER PIT. TUNNEL EXIT IS UNDER CONSTRUCTION WHILE CONTROL BUILDING IS BEING FORMED BEYOND. CAMERA FACING WEST. INL NEGATIVE NO. 56-1238. Jack L. Anderson, Photographer, 4/17/1956 - Idaho National Engineering Laboratory, Test Reactor Area, Materials & Engineering Test Reactors, Scoville, Butte County, ID

Fabrication and Testing of Low Cost 2D Carbon-Carbon Nozzle Extensions at NASA/MSFC

NASA Technical Reports Server (NTRS)

Greene, Sandra Elam; Shigley, John K.; George, Russ; Roberts, Robert

2015-01-01

Subscale liquid engine tests were conducted at NASA/MSFC using a 1.2 Klbf engine with liquid oxygen (LOX) and gaseous hydrogen. Testing was performed for main-stage durations ranging from 10 to 160 seconds at a chamber pressure of 550 psia and a mixture ratio of 5.7. Operating the engine in this manner demonstrated a new and affordable test capability for evaluating subscale nozzles by exposing them to long duration tests. A series of 2D C-C nozzle extensions were manufactured, oxidation protection applied and then tested on a liquid engine test facility at NASA/MSFC. The C-C nozzle extensions had oxidation protection applied using three very distinct methods with a wide range of costs and process times: SiC via Polymer Impregnation & Pyrolysis (PIP), Air Plasma Spray (APS) and Melt Infiltration. The tested extensions were about 6" long with an exit plane ID of about 6.6". The test results, material properties and performance of the 2D C-C extensions and attachment features will be discussed.

Optimal trajectories for the aeroassisted flight experiment. Part 3: Formulation, results, and analysis

NASA Technical Reports Server (NTRS)

Miele, A.; Wang, T.; Lee, W. Y.; Zhao, Z. G.

1989-01-01

The determination of optimal trajectories for the aero-assisted flight experiment (AFE) is investigated. The intent of this experiment is to simulate a GEO-to-LEO transfer, where GEO denotes a geosynchronous Earth orbit and LEO denotes a low Earth orbit. The trajectories of an AFE spacecraft are analyzed in a 3D-space, employing the full system of 6 ODEs describing the atmospheric pass. The atmospheric entry conditions are given, and the atmospheric exit conditions are adjusted in such a way that the following conditions are satisfied: (1) the atmospheric velocity depletion is such that, after exiting, the AFE spacecraft first ascends to a specified apogee and then descends to a specified perigee; and (2) the exit orbital plane is identical with the entry orbital plane. The final maneuver, not analyzed here, includes the rendezvous with and the capture by the space shuttle.

Experimental investigation of jet-induced loads on a flat plate in hover out-of-ground effect

NASA Technical Reports Server (NTRS)

Kuhlman, J. M.; Warcup, R. W.

1979-01-01

Effects of varying jet decay rate on jet-induced loads on a flat plate located in the plane of the jet exit perpendicular to the jet axis were investigated using a small-scale laboratory facility. Jet decay rate has been varied through use of two cylindrical centerbodies having either a flat or hemispherical tip, which were submerged various distances below the flat plate jet exit plane. Increased jet decay rate, caused by the presence of a center-body or plug in the jet nozzle, led to an increased jet-induced lift loss on the flat plate. Jet-induced lift losses reached 1 percent of the jet thrust for the quickest jet decay rates for plate areas equal to 100 times the effective jet exit area. The observed lift loss versus jet decay rate trend agreed well with results of previous investigations.

Laser Transmission Measurements of Soot Extinction Coefficients in the Exhaust Plume of the X-34 60K-lb Thrust Fastrac Rocket Engine

NASA Technical Reports Server (NTRS)

Dobson, C. C.; Eskridge, R. H.; Lee, M. H.

2000-01-01

A four-channel laser transmissometer has been used to probe the soot content of the exhaust plume of the X-34 60k-lb thrust Fastrac rocket engine at NASA's Marshall Space Flight Center. The transmission measurements were made at an axial location approximately equal 1.65 nozzle diameters from the exit plane and are interpreted in terms of homogeneous radial zones to yield extinction coefficients from 0.5-8.4 per meter. The corresponding soot mass density, spatially averaged over the plume cross section, is, for Rayleigh particles, approximately equal 0.7 microgram/cc, and alternative particle distributions are briefly considered. Absolute plume radiance at the laser wavelength (515 nm) is estimated from the data at approximately equal 2,200 K equivalent blackbody temperature, and temporal correlations in emission from several spatial locations are noted.

Laser Transmission Measurements of Soot Extinction Coefficients in the Exhaust Plume of the X-34 60k-lb Thrust Fastrac Rocket Engine

NASA Technical Reports Server (NTRS)

Dobson, C. C.; Eskridge, R. H.; Lee, M. H.

2000-01-01

A four-channel laser transmissometer has been used to probe the soot content of the exhaust plume of the X-34 60k-lb thrust Fastrac rocket engine at NASA's Marshall Space Flight Center. The transmission measurements were made at an axial location about equal 1.65 nozzle diameters from the exit plane and are interpreted in terms of homogeneous radial zones to yield extinction coefficients from 0.5-8.4 per meter. The corresponding soot mass density, spatially averaged over the plume cross section, is, for Rayleigh particles, approximately equal to 0.7 micrograms/cubic cm and alternative particle distributions are briefly considered. Absolute plume radiance at the laser wavelength (515 nm) is estimated from the data at approximately equal to 2.200 K equivalent blackbody temperature, and temporal correlations in emission from several spatial locations are noted.

Mixing of Supersonic Jets in a RBCC Strutjet Propulsion System

NASA Technical Reports Server (NTRS)

Muller, S.; Hawk, Clark W.; Bakker, P. G.; Parkinson, D.; Turner, M.

1998-01-01

The Strutjet approach to Rocket Based Combined Cycle (RBCC) propulsion depends upon fuel-rich flows from the rocket nozzles and turbine exhaust products mixing with the ingested air for successful operation in the ramjet and scramjet modes. It is desirable to delay this mixing process in the air-augmented mode of operation present during take-off and low speed flight. A scale model of the Strutjet device was built and tested to investigate the mixing of the streams as a function of distance from the Strut exit plane in simulated sea level take-off conditions. The Planar Laser Induced Fluorescence (PLIF) diagnostic method has been employed to observe the mixing of the turbine exhaust gas with the gases from both the primary rockets and the ingested air. The ratio of the pressure in the turbine exhaust to that in the rocket nozzle wall at the point where the two jets meet, is the independent variable in these experiments. Tests were accomplished at values of 1.0 (the original design point), 1.5 and 2.0 for this parameter at 8 locations downstream of the rocket nozzle exit. The results illustrate the development of the mixing zone from the exit plane of the strut to a distance of about 18 equivalent rocket nozzle exit diameters downstream (18"). These images show the turbine exhaust to be confined until a short distance downstream. The expansion into the ingested air is more pronounced at a pressure ratio of 1.0 and 1.5 and shows that mixing with this air would likely begin at a distance of 2" downstream of the nozzle exit plane. Of the pressure ratios tested in this research, 2.0 is the best value for delaying the mixing at the operating conditions considered.

Aerodynamic Measurements of a Variable-Speed Power-Turbine Blade Section in a Transonic Turbine Cascade

NASA Technical Reports Server (NTRS)

Flegel, Ashlie B.

2014-01-01

The purpose of this thesis is to document the impact of incidence angle and Reynolds number variations on the three-dimensional flow field and midspan loss and turning of a two-dimensional section of a variable-speed power-turbine (VSPT) rotor blade. Aerodynamic measurements were obtained in a transonic linear cascade at NASA Glenn Research Center in Cleveland, Ohio. Steady-state data were obtained for 10 incidence angles ranging from +15.8deg to -51.0deg. At each angle, data were acquired at five flow conditions with the exit Reynolds number (based on axial chord) varying over an order-of-magnitude from 2.12×105 to 2.12×106. Data were obtained at the design exit Mach number of 0.72 and at a reduced exit Mach number of 0.35 as required to achieve the lowest Reynolds number. Midspan tota lpressure and exit flow angle data were acquired using a five-hole pitch/yaw probe surveyed on a plane located 7.0 percent axial-chord downstream of the blade trailing edge plane. The survey spanned three blade passages. Additionally, three-dimensional half-span flow fields were examined with additional probe survey data acquired at 26 span locations for two key incidence angles of +5.8deg and -36.7deg. Survey data near the endwall were acquired with a three-hole boundary-layer probe. The data were integrated to determine average exit total-pressure and flow angle as functions of incidence and flow conditions. The data set also includes blade static pressures measured on four spanwise planes and endwall static pressures.

Design and Analyses of High Aspect Ratio Nozzles for Distributed Propulsion Acoustic Measurements

NASA Technical Reports Server (NTRS)

Dippold, Vance F., III

2016-01-01

A series of three convergent, round-to-rectangular high aspect ratio (HAR) nozzles were designed for acoustic testing at the NASA Glenn Research Center Nozzle Acoustic Test Rig (NATR). The HAR nozzles had exit area aspect ratios of 8:1, 12:1, and 16:1. The nozzles were designed to mimic a distributed propulsion system array with a slot nozzle. The nozzle designs were screened using Reynolds-Averaged Navier-Stokes (RANS) simulations. In addition to meeting the geometric constraints required for testing in the NATR, the HAR nozzles were designed to be free of flow features that would produce unwanted noise (e.g., flow separations) and to have uniform flow at the nozzle exit. Multiple methods were used to generate HAR nozzle designs. The final HAR nozzle designs were generated in segments using a computer code that parameterized each segment. RANS screening simulations showed that intermediate nozzle designs suffered flow separation, a normal shockwave at the nozzle exit (caused by an aerodynamic throat produced by boundary layer growth), and non-uniform flow at the nozzle exit. The RANS simulations showed that the final HAR nozzle designs were free of flow separations, but were not entirely successful at producing a fully uniform flow at the nozzle exit. The final designs suffered a pair of counter-rotating vortices along the outboard walls of the nozzle. The 16:1 aspect ratio HAR nozzle had the least uniform flow at the exit plane; the 8:1 aspect ratio HAR nozzles had a fairly uniform flow at the nozzle exit plane.

Sound Radiation from a Supersonic Jet Passing Through a Partially Open Exhaust Duct

NASA Technical Reports Server (NTRS)

Kandula, Max

2011-01-01

The radiation of sound from a perfectly expanded Mach 2.5 cold supersonic jet of 25.4 mm exit diameter flowing through a partially open rigid-walled duct with an upstream i-deflector has been studied experimentally. In the experiments, the nozzle is mounted vertically, with the nozzle exit plane at a height of 73 jet diameters above ground level. Relative to the nozzle exit plane (NEP), the location of the duct inlet is varied at 10, 5, and -1 jet diameters. Far-field sound pressure levels were obtained at 54 jet diameters above ground with the aid of acoustic sensors equally spaced around a circular arc of radius equal to 80 jet diameters from the jet axis. Data on the jet acoustic field for the partially open duct were obtained and compared with those with a free jet and with a closed duct. The results suggest that for the partially open duct the overall sound pressure level (OASPL) decreases as the distance between the NEP and the duct inlet plane decreases, while the opposite trend is observed for the closed duct. It is also concluded that the observed peak frequency in the partially open duct increases above the free jet value as the angle from the duct axis is increased, and as the duct inlet plane becomes closer to the NEP.

Recent Results From Internal and Very-Near-Field Plasma Diagnostics of a High Specific Impulse Hall Thruster

NASA Technical Reports Server (NTRS)

Hofer, Richard R.; Gallimore, Alec D.; Jacobson, David (Technical Monitor)

2003-01-01

Floating potential and ion current density measurements were taken on the laboratory model NASA-173Mv2 in order to improve understanding of the physical processes affecting Hall thruster performance at high specific impulse. Floating potential was measured on discharge chamber centerline over axial positions spanning 10 mm from the anode to 100 mm downstream of the exit plane. Ion current density was mapped radially up to 300 mm from thruster centerline over axial positions in the very-near-field (10 to 250 mm from the exit plane). All data were collected using a planar probe in conjunction with a high-speed translation stage to minimize probe-induced thruster perturbations. Measurements of floating potential at a xenon flow rate of 10 mg/s have shown that the acceleration layer moved upstream 3 1 mm when the voltage increased from 300 to 600 V. The length of the acceleration layer was 14 2 mm and was approximately constant with voltage and magnetic field. Ion current density measurements indicated the annular ion beam crossed the thruster centerline 163 mm downstream of the exit plane. Radial integration of the ion current density at the cathode plane provided an estimate of the ion current fraction. At 500 V and 5 mg/s, the ion current fraction was calculated as 0.77.

S-Duct Engine Inlet Flow Control Using SDBD Plasma Streamwise Vortex Generators

NASA Astrophysics Data System (ADS)

Kelley, Christopher; He, Chuan; Corke, Thomas

2009-11-01

The results of a numerical simulation and experiment characterizing the performance of plasma streamwise vortex generators in controlling separation and secondary flow within a serpentine, diffusing duct are presented. A no flow control case is first run to check agreement of location of separation, development of secondary flow, and total pressure recovery between the experiment and numerical results. Upon validation, passive vane-type vortex generators and plasma streamwise vortex generators are implemented to increase total pressure recovery and reduce flow distortion at the aerodynamic interface plane: the exit of the S-duct. Total pressure recovery is found experimentally with a pitot probe rake assembly at the aerodynamic interface plane. Stagnation pressure distortion descriptors are also presented to show the performance increase with plasma streamwise vortex generators in comparison to the baseline no flow control case. These performance parameters show that streamwise plasma vortex generators are an effective alternative to vane-type vortex generators in total pressure recovery and total pressure distortion reduction in S-duct inlets.

Computational Study of the CC3 Impeller and Vaneless Diffuser Experiment

NASA Technical Reports Server (NTRS)

Kulkarni, Sameer; Beach, Timothy A.; Skoch, Gary J.

2013-01-01

Centrifugal compressors are compatible with the low exit corrected flows found in the high pressure compressor of turboshaft engines and may play an increasing role in turbofan engines as engine overall pressure ratios increase. Centrifugal compressor stages are difficult to model accurately with RANS CFD solvers. A computational study of the CC3 centrifugal impeller in its vaneless diffuser configuration was undertaken as part of an effort to understand potential causes of RANS CFD mis-prediction in these types of geometries. Three steady, periodic cases of the impeller and diffuser were modeled using the TURBO Parallel Version 4 code: 1) a k-epsilon turbulence model computation on a 6.8 million point grid using wall functions, 2) a k-epsilon turbulence model computation on a 14 million point grid integrating to the wall, and 3) a k-omega turbulence model computation on the 14 million point grid integrating to the wall. It was found that all three cases compared favorably to data from inlet to impeller trailing edge, but the k-epsilon and k-omega computations had disparate results beyond the trailing edge and into the vaneless diffuser. A large region of reversed flow was observed in the k-epsilon computations which extended from 70% to 100% span at the exit rating plane, whereas the k-omega computation had reversed flow from 95% to 100% span. Compared to experimental data at near-peak-efficiency, the reversed flow region in the k-epsilon case resulted in an under-prediction in adiabatic efficiency of 8.3 points, whereas the k-omega case was 1.2 points lower in efficiency.

Computational Study of the CC3 Impeller and Vaneless Diffuser Experiment

NASA Technical Reports Server (NTRS)

Kulkarni, Sameer; Beach, Timothy A.; Skoch, Gary J.

2013-01-01

Centrifugal compressors are compatible with the low exit corrected flows found in the high pressure compressor of turboshaft engines and may play an increasing role in turbofan engines as engine overall pressure ratios increase. Centrifugal compressor stages are difficult to model accurately with RANS CFD solvers. A computational study of the CC3 centrifugal impeller in its vaneless diffuser configuration was undertaken as part of an effort to understand potential causes of RANS CFD mis-prediction in these types of geometries. Three steady, periodic cases of the impeller and diffuser were modeled using the TURBO Parallel Version 4 code: (1) a k-e turbulence model computation on a 6.8 million point grid using wall functions, (2) a k-e turbulence model computation on a 14 million point grid integrating to the wall, and (3) a k-? turbulence model computation on the 14 million point grid integrating to the wall. It was found that all three cases compared favorably to data from inlet to impeller trailing edge, but the k-e and k-? computations had disparate results beyond the trailing edge and into the vaneless diffuser. A large region of reversed flow was observed in the k-e computations which extended from 70 to 100 percent span at the exit rating plane, whereas the k-? computation had reversed flow from 95 to 100 percent span. Compared to experimental data at near-peak-efficiency, the reversed flow region in the k-e case resulted in an underprediction in adiabatic efficiency of 8.3 points, whereas the k-? case was 1.2 points lower in efficiency.

14 CFR 27.1557 - Miscellaneous markings and placards.

Code of Federal Regulations, 2011 CFR

2011-01-01

...; (iii) For turbine engine powered rotorcraft, the permissible fuel designations; and (iv) For pressure...) Emergency exit placards. Each placard and operating control for each emergency exit must be red. A placard must be near each emergency exit control and must clearly indicate the location of that exit and its...

14 CFR 27.1557 - Miscellaneous markings and placards.

Code of Federal Regulations, 2013 CFR

2013-01-01

...; (iii) For turbine engine powered rotorcraft, the permissible fuel designations; and (iv) For pressure...) Emergency exit placards. Each placard and operating control for each emergency exit must be red. A placard must be near each emergency exit control and must clearly indicate the location of that exit and its...

14 CFR 27.1557 - Miscellaneous markings and placards.

Code of Federal Regulations, 2012 CFR

2012-01-01

...; (iii) For turbine engine powered rotorcraft, the permissible fuel designations; and (iv) For pressure...) Emergency exit placards. Each placard and operating control for each emergency exit must be red. A placard must be near each emergency exit control and must clearly indicate the location of that exit and its...

14 CFR 27.1557 - Miscellaneous markings and placards.

Code of Federal Regulations, 2014 CFR

2014-01-01

...; (iii) For turbine engine powered rotorcraft, the permissible fuel designations; and (iv) For pressure...) Emergency exit placards. Each placard and operating control for each emergency exit must be red. A placard must be near each emergency exit control and must clearly indicate the location of that exit and its...

Study of Plume Impingement Effects in the Lunar Lander Environment

NASA Technical Reports Server (NTRS)

Marichalar, Jeremiah; Prisbell, A.; Lumpkin, F.; LeBeau, G.

2010-01-01

Plume impingement effects from the descent and ascent engine firings of the Lunar Lander were analyzed in support of the Lunar Architecture Team under the Constellation Program. The descent stage analysis was performed to obtain shear and pressure forces on the lunar surface as well as velocity and density profiles in the flow field in an effort to understand lunar soil erosion and ejected soil impact damage which was analyzed as part of a separate study. A CFD/DSMC decoupled methodology was used with the Bird continuum breakdown parameter to distinguish the continuum flow from the rarefied flow. The ascent stage analysis was performed to ascertain the forces and moments acting on the Lunar Lander Ascent Module due to the firing of the main engine on take-off. The Reacting and Multiphase Program (RAMP) method of characteristics (MOC) code was used to model the continuum region of the nozzle plume, and the Direct Simulation Monte Carlo (DSMC) Analysis Code (DAC) was used to model the impingement results in the rarefied region. The ascent module (AM) was analyzed for various pitch and yaw rotations and for various heights in relation to the descent module (DM). For the ascent stage analysis, the plume inflow boundary was located near the nozzle exit plane in a region where the flow number density was large enough to make the DSMC solution computationally expensive. Therefore, a scaling coefficient was used to make the DSMC solution more computationally manageable. An analysis of the effectiveness of this scaling technique was performed by investigating various scaling parameters for a single height and rotation of the AM. Because the inflow boundary was near the nozzle exit plane, another analysis was performed investigating three different inflow contours to determine the effects of the flow expansion around the nozzle lip on the final plume impingement results.

Evaluation of Federal Aviation Administration Engine Exhaust Sampling Rake.

DTIC Science & Technology

1977-06-01

Engine exit total pressure - Pt7 " High turbine discharge temperature - Tt65,6 . Engine exit total temperature - Tt7 - Burner pressure - Ps4 0 Total...illustrated by Figure 22, a plot of Ps4 /Pt7 (expansion thru the turbine) vs. EPR (Pt7/Pt2). This plot indicates that the ECCP rake has no measurable effect on

1998 Calibration of the Mach 4.7 and Mach 6 Arc-Heated Scramjet Test Facility Nozzles

NASA Technical Reports Server (NTRS)

Witte, David W.; Irby, Richard G.; Auslender, Aaron H.; Rock, Kenneth E.

2004-01-01

A calibration of the Arc-Heated Scramjet Test Facility (AHSTF) Mach 4.7 and Mach 6 nozzles was performed in 1998. For each nozzle, three different typical facility operating test points were selected for calibration. Each survey consisted of measurements, at 340 separate locations across the 11 inch square nozzle exit plane, of pitot pressure, static pressure, and total temperature. Measurement density was higher (4/inch) in the boundary layer near the nozzle wall than in the core nozzle flow (1/inch). The results generated for each of these calibration surveys were contour plots at the nozzle exit plane of the measured and calculated flow properties which completely defined the thermodynamic state of the nozzle exit flow. An area integration of the mass flux at the nozzle exit for each survey was compared to the AHSTF mass flow meter results to provide an indication of the overall quality of the calibration performed. The percent difference between the integrated nozzle exit mass flow and the flow meter ranged from 0.0 to 1.3 percent for the six surveys. Finally, a comparison of this 1998 calibration was made with the 1986 calibration. Differences of less than 10 percent were found within the nozzle core flow while in the boundary layer differences on the order of 20 percent were quite common.

Exit Exam as Academic Performance Indicator

ERIC Educational Resources Information Center

Al Ahmad, Mahmoud; Al Marzouqi, Ali H.; Hussien, Mousa

2014-01-01

This paper focuses on the impact of exit exams on different elements of the educational process, namely: curriculum development, students and instructors. A 50-question multiple-choice Exit Exam was prepared by Electrical Engineering (EE) faculty members covering a poll of questions from EE core courses. A copy of the Exit Exam applied during each…

Pressure transfer function of a JT15D nozzle due to acoustic and convected entropy fluctuations

NASA Astrophysics Data System (ADS)

Miles, J. H.

An acoustic transmission matrix analysis of sound propagation in a variable area duct with and without flow is extended to include convected entropy fluctuations. The boundary conditions used in the analysis are a transfer function relating entropy and pressure at the nozzle inlet and the nozzle exit impedance. The nozzle pressure transfer function calculated is compared with JT15D turbofan engine nozzle data. The one dimensional theory for sound propagation in a variable area nozzle with flow but without convected entropy is good at the low engine speeds where the nozzle exit Mach number is low (M=0.2) and the duct exit impedance model is good. The effect of convected entropy appears to be so negligible that it is obscured by the inaccuracy of the nozzle exit impedance model, the lack of information on the magnitude of the convected entropy and its phase relationship with the pressure, and the scatter in the data. An improved duct exit impedance model is required at the higher engine speeds where the nozzle exit Mach number is high (M=0.56) and at low frequencies (below 120 Hz).

Catalytic reactor for low-Btu fuels

DOEpatents

Smith, Lance; Etemad, Shahrokh; Karim, Hasan; Pfefferle, William C.

2009-04-21

An improved catalytic reactor includes a housing having a plate positioned therein defining a first zone and a second zone, and a plurality of conduits fabricated from a heat conducting material and adapted for conducting a fluid therethrough. The conduits are positioned within the housing such that the conduit exterior surfaces and the housing interior surface within the second zone define a first flow path while the conduit interior surfaces define a second flow path through the second zone and not in fluid communication with the first flow path. The conduit exits define a second flow path exit, the conduit exits and the first flow path exit being proximately located and interspersed. The conduits define at least one expanded section that contacts adjacent conduits thereby spacing the conduits within the second zone and forming first flow path exit flow orifices having an aggregate exit area greater than a defined percent of the housing exit plane area. Lastly, at least a portion of the first flow path defines a catalytically active surface.

Species and temperature measurement in H2/O2 rocket flow fields by means of Raman scattering diagnostics

NASA Technical Reports Server (NTRS)

Degroot, Wim A.; Weiss, Jonathan M.

1992-01-01

Validation of Computational Fluid Dynamics (CFD) codes developed for prediction and evaluation of rocket performance is hampered by a lack of experimental data. Non-intrusive laser based diagnostics are needed to provide spatially and temporally resolved gas dynamic and fluid dynamic measurements. This paper reports the first non-intrusive temperature and species measurements in the plume of a 110 N gaseous hydrogen/oxygen thruster at and below ambient pressures, obtained with spontaneous Raman spectroscopy. Measurements at 10 mm downstream of the exit plane are compared with predictions from a numerical solution of the axisymmetric Navier-Stokes and species transport equations with chemical kinetics, which fully model the combustor-nozzle-plume flowfield. The experimentally determined oxygen number density at the centerline at 10 mm downstream of the exit plane is four times that predicted by the model. The experimental number density data fall between those numerically predicted for the exit and 10 mm downstream planes in both magnitude and radial gradient. The predicted temperature levels are within 10 to 15 percent of measured values. Some of the discrepancies between experimental data and predictions result from not modeling the three dimensional core flow injection mixing process, facility back pressure effects, and possible diffuser-thruster interactions.

Design, Activation, and Operation of the J2-X Subscale Simulator (JSS)

NASA Technical Reports Server (NTRS)

Saunders, Grady P.; Raines, Nickey G.; Varner, Darrel G.

2009-01-01

The purpose of this paper is to give a detailed description of the design, activation, and operation of the J2-X Subscale Simulator (JSS) installed in Cell 1 of the E3 test facility at Stennis Space Center, MS (SSC). The primary purpose of the JSS is to simulate the installation of the J2-X engine in the A3 Subscale Rocket Altitude Test Facility at SSC. The JSS is designed to give aerodynamically and thermodynamically similar plume properties as the J2-X engine currently under development for use as the upper stage engine on the ARES I and ARES V spacecraft. The JSS is a scale pressure fed, LOX/GH fueled rocket that is geometrically similar to the J2-X from the throat to the nozzle exit plane (NEP) and is operated at the same oxidizer to fuel ratios and chamber pressures. This paper describes the heritage hardware used as the basis of the JSS design, the newly designed rocket hardware, igniter systems used, and the activation and operation of the JSS.

The investigation of time dependent flame structure by ionization probes

NASA Technical Reports Server (NTRS)

Ventura, J. M. P.; Suzuki, T.; Yule, A. J.; Ralph, S.; Chigier, N. A.

1980-01-01

Ionization probes were used to measure mean ionization current and frequency spectra, auto-correlations and cross-correlations in jet flames with variation in the initial Reynolds numbers and equivalence ratios. Special attention was paid to the transitional region between the burner exit plane and the plane of onset of turbulence.

Pollutant emissions from and within a model gas turbine combustor at elevated pressures and temperatures

NASA Technical Reports Server (NTRS)

Drennan, S. A.; Peterson, C. O.; Khatib, F. M.; Sowa, W. A.; Samuelsen, G. S.

1993-01-01

Conventional and advanced gas turbine engines are coming under increased scrutiny regarding pollutant emissions. This, in turn, has created a need to obtain in-situ experimental data at practical conditions, as well as exhaust data, and to obtain the data in combustors that reflect modern designs. The in-situ data are needed to (1) assess the effects of design modifications on pollutant formation, and (2) develop a detailed data base on combustor performance for the development and verification of computer modeling. This paper reports on a novel high pressure, high temperature facility designed to acquire such data under controlled conditions and with access (optical and extractive) for in-situ measurements. To evaluate the utility of the facility, a model gas turbine combustor was selected which features practical hardware design, two rows of jets (primary and dilution) with four jets in each row, and advanced wall cooling techniques with laser drilled effusive holes. The dome is equipped with a flat-vaned swirler with vane angles of 60 degrees. Data are obtained at combustor pressures ranging from 2 to 10 atmospheres of pressure, levels of air preheat to 427 C, combustor reference velocities from 10.0 to 20.0 m/s, and an overall equivalence ratio of 0.3. Exit plane and in-situ measurements are presented for HC, O2, CO2, CO, and NO(x). The exit plane emissions of NO(x) correspond to levels reported from practical combustors and the in-situ data demonstrate the utility and potential for detailed flow field measurements.

Jet engine nozzle exit configurations, including projections oriented relative to pylons, and associated systems and methods

NASA Technical Reports Server (NTRS)

Mengle, Vinod G. (Inventor); Thomas, Russell H. (Inventor)

2012-01-01

Nozzle exit configurations and associated systems and methods are disclosed. An aircraft system in accordance with one embodiment includes a jet engine exhaust nozzle having an internal flow surface and an exit aperture, with the exit aperture having a perimeter that includes multiple projections extending in an aft direction. Aft portions of individual neighboring projections are spaced apart from each other by a gap, and a geometric feature of the multiple can change in a monotonic manner along at least a portion of the perimeter. Projections near a support pylon and/or associated heat shield can have particular configurations, including greater flow immersion than other projections.

Cooling Air Inlet and Exit Geometries on Aircraft Engine Installations

NASA Technical Reports Server (NTRS)

Katz, Joseph; Corsiglia, Victor R.; Barlow, Philip R.

1982-01-01

A semispan wing and nacelle of a typical general aviation twin-engine aircraft was tested to evaluate the cooling capability and drag or several nacelle shapes; the nacelle shapes included cooling air inlet and exit variations. The tests were conducted in the Ames Research Center 40 x 80-ft Wind Tunnel. It was found that the cooling air inlet geometry of opposed piston engine installations has a major effect on inlet pressure recovery, but only a minor effect on drag. Exit location showed large effect on drag, especially for those locations on the sides of the nacelle where the suction characteristics were based on interaction with the wing surface pressures.

Hall thruster microturbulence under conditions of modified electron wall emission

NASA Astrophysics Data System (ADS)

Tsikata, S.; Héron, A.; Honoré, C.

2017-05-01

In recent numerical, theoretical, and experimental papers, the short-scale electron cyclotron drift instability (ECDI) has been studied as a possible contributor to the anomalous electron current observed in Hall thrusters. In this work, features of the instability, in the presence of a zero-electron emission material at the thruster exit plane, are analyzed using coherent Thomson scattering. Limiting the electron emission at the exit plane alters the localization of the accelerating electric field and the expected drift velocity profile, which in turn modifies the amplitude and localization of the ECDI. The resulting changes to the standard thruster operation are expected to favor an increased contribution by the ECDI to electron current. Such an operation is associated with a degradation of thruster performance and stability.

Evaluation of water cooled supersonic temperature and pressure probes for application to 2000 F flows

NASA Technical Reports Server (NTRS)

Lagen, Nicholas T.; Seiner, John M.

1990-01-01

The development of water cooled supersonic probes used to study high temperature jet plumes is addressed. These probes are: total pressure, static pressure, and total temperature. The motivation for these experiments is the determination of high temperature supersonic jet mean flow properties. A 3.54 inch exit diameter water cooled nozzle was used in the tests. It is designed for exit Mach 2 at 2000 F exit total temperature. Tests were conducted using water cooled probes capable of operating in Mach 2 flow, up to 2000 F total temperature. Of the two designs tested, an annular cooling method was chosen as superior. Data at the jet exit planes, and along the jet centerline, were obtained for total temperatures of 900 F, 1500 F, and 2000 F, for each of the probes. The data obtained from the total and static pressure probes are consistent with prior low temperature results. However, the data obtained from the total temperature probe was affected by the water coolant. The total temperature probe was tested up to 2000 F with, and without, the cooling system turned on to better understand the heat transfer process at the thermocouple bead. The rate of heat transfer across the thermocouple bead was greater when the coolant was turned on than when the coolant was turned off. This accounted for the lower temperature measurement by the cooled probe. The velocity and Mach number at the exit plane and centerline locations were determined from the Rayleigh-Pitot tube formula.

Sound radiation from a flanged inclined duct.

PubMed

McAlpine, Alan; Daymond-King, Alex P; Kempton, Andrew J

2012-12-01

A simple method to calculate sound radiation from a flanged inclined duct is presented. An inclined annular duct is terminated by a rigid vertical plane. The duct termination is representative of a scarfed exit. The concept of a scarfed duct has been examined in turbofan aero-engines as a means to, potentially, shield a portion of the radiated sound from being transmitted directly to the ground. The sound field inside the annular duct is expressed in terms of spinning modes. Exterior to the duct, the radiated sound field owing to each mode can be expressed in terms of its directivity pattern, which is found by evaluating an appropriate form of Rayleigh's integral. The asymmetry is shown to affect the amplitude of the principal lobe of the directivity pattern, and to alter the proportion of the sound power radiated up or down. The methodology detailed in this article provides a simple engineering approach to investigate the sound radiation for a three-dimensional problem.

Analysis of the Magneto-Hydrodynamic (MHD) Energy Bypass Engine for High-Speed Air-Breathing Propulsion

NASA Technical Reports Server (NTRS)

Riggins, David W.

2002-01-01

The performance of the MHD energy bypass air-breathing engine for high-speed propulsion is analyzed in this investigation. This engine is a specific type of the general class of inverse cycle engines. In this paper, the general relationship between engine performance (specific impulse and specific thrust) and the overall total pressure ratio through an engine (from inlet plane to exit plane) is first developed and illustrated. Engines with large total pressure decreases, regardless of cause or source, are seen to have exponentially decreasing performance. The ideal inverse cycle engine (of which the MHD engine is a sub-set) is then demonstrated to have a significant total pressure decrease across the engine; this total pressure decrease is cycle-driven, degrades rapidly with energy bypass ratio, and is independent of any irreversibility. The ideal MHD engine (inverse cycle engine with no irreversibility other than that inherent in the MHD work interaction processes) is next examined and is seen to have an additional large total pressure decrease due to MHD-generated irreversibility in the decelerator and the accelerator. This irreversibility mainly occurs in the deceleration process. Both inherent total pressure losses (inverse cycle and MHD irreversibility) result in a significant narrowing of the performance capability of the MHD bypass engine. The fundamental characteristics of MHD flow acceleration and flow deceleration from the standpoint of irreversibility and second-law constraints are next examined in order to clarify issues regarding flow losses and parameter selection in the MM modules. Severe constraints are seen to exist in the decelerator in terms of allowable deceleration Mach numbers and volumetric (length) required for meaningful energy bypass (work interaction). Considerable difficulties are also encountered and discussed due to thermal/work choking phenomena associated with the deceleration process. Lastly, full engine simulations utilizing inlet shock systems, finite-rate chemistry, wall cooling with thermally balanced engine (fuel heat sink), fuel injection and mixing, friction, etc. are shown and discussed for both the MHD engine and the conventional scramjet. The MHD bypass engine has significantly lower performance in all categories across the Mach number range (8 to 12.2). The lower performance is attributed to the combined effects of 1) additional irreversibility and cooling requirements associated with the MHD components and 2) the total pressure decrease associated with the inverse cycle itself.

A first scramjet study

NASA Technical Reports Server (NTRS)

Emanuel, George

1989-01-01

A variety of related scramjet engine topics are examined. The flow is assumed to be 1-D, the gas is thermally and calorically perfect, and focus is on low hypersonic Mach numbers. The thrust and lift of an exposed half nozzle, which is used on the aerospace plane, is evaluated as well as a fully confined nozzle. A rough estimate of the drag of an aerospace plane is provided. Thermal effects and shock waves are next discussed. A parametric scramjet model is then presented based on the influence coefficient method, which evaluates the dominant scramjet processes. The independent parameters are the ratio of specific heats, a nondimensional heat addition parameter, and four Mach numbers. The total thrust generated by the combustor and nozzle is shown to be independent of the heat release distribution and the combustor exit Mach number, providing thermal choking is avoided. An operating condition for the combustor is found that maximizes the thrust. An alternative condition is explored when this optimum is no longer realistic. This condition provides a favorable pressure gradient and a reasonable area ratio for the combustor. Parametric results based on the model is provided.

Determining when a set of compute nodes participating in a barrier operation on a parallel computer are ready to exit the barrier operation

DOEpatents

Blocksome, Michael A [Rochester, MN

2011-12-20

Methods, apparatus, and products are disclosed for determining when a set of compute nodes participating in a barrier operation on a parallel computer are ready to exit the barrier operation that includes, for each compute node in the set: initializing a barrier counter with no counter underflow interrupt; configuring, upon entering the barrier operation, the barrier counter with a value in dependence upon a number of compute nodes in the set; broadcasting, by a DMA engine on the compute node to each of the other compute nodes upon entering the barrier operation, a barrier control packet; receiving, by the DMA engine from each of the other compute nodes, a barrier control packet; modifying, by the DMA engine, the value for the barrier counter in dependence upon each of the received barrier control packets; exiting the barrier operation if the value for the barrier counter matches the exit value.

Performance and technological feasibility of rocket powered HTHL-SSTO with take-off assist (aerospace plane/ekranoplane)

NASA Astrophysics Data System (ADS)

Tomita, Nobuyuki; Nebylov, Alexander V.; Sokolov, Victor V.; Ohkami, Yoshiaki

It might be said that it is common understanding that rocket-powered single stage to orbit (SSTO) aerospace planes will become feasible with near-term technology as described in [1] (Koelle, D. E. Survey and comparison of winged launch vehicle options, ISTS 94-g-11 V 1994) and [2] (Bekey, I. Why SSTO rocket launch vehicles are now feasible and practical, IAF-94-V.1.524 1994). Among two methods of launching aerospace planes into orbit, vertical take-off (VT) and horizontal take-off (HT), it seems that VT takes the lead from HT [1, 2]. The decision for the X-33 program by NASA, also, seems to favor VT. In retrospect, almost all of the launch vehicles in the past have been VT, mainly because VT solved the problem of exit from atmosphere to space. However, broadening the range of requirements for space transportation systems from military to commercial and unmanned to manned seems to favor the need for HT. In this paper, the authors are going to prove that aerospace plane/ekranoplane system, which is a reusable launch vehicle system based on the HT concept, with ekranoplane as a take-off and possibly, landing assist, could be competitive with the VT concept from both technological and economical view points. Ekranoplane is a wing-in-ground-effect craft (WIG), which moves at a speed of approximately 0.5 M, carrying heavy loads above the sea surface. Combination of high initial velocity and high performance tri-propellant engine for aerospace plane makes it possible to configure an aerospace plane which is competitive with VT. Other specific features of HT in comparison with VT are discussed.

Experimental Assessment of the Emissions Control Potential of a Rich/Quench/Lean Combustor for High Speed Civil Transport Aircraft Engines

NASA Technical Reports Server (NTRS)

Rosfjord, T. J.; Padget, F. C.; Tacina, Robert R. (Technical Monitor)

2001-01-01

In support of Pratt & Whitney efforts to define the Rich burn/Quick mix/Lean burn (RQL) combustor for the High Speed Civil Transport (HSCT) aircraft engine, UTRC conducted a flametube-scale study of the RQL concept. Extensive combustor testing was performed at the Supersonic Cruise (SSC) condition of a HSCT engine cycle, Data obtained from probe traverses near the exit of the mixing section confirmed that the mixing section was the critical component in controlling combustor emissions. Circular-hole configurations, which produced rapidly-, highly-penetrating jets, were most effective in limiting NOx. The spatial profiles of NOx and CO at the mixer exit were not directly interpretable using a simple flow model based on jet penetration, and a greater understanding of the flow and chemical processes in this section are required to optimize it. Neither the rich-combustor equivalence ratio nor its residence time was a direct contributor to the exit NOx. Based on this study, it was also concluded that (1) While NOx formation in both the mixing section and the lean combustor contribute to the overall emission, the NOx formation in the mixing section dominates. The gas composition exiting the rich combustor can be reasonably represented by the equilibrium composition corresponding to the rich combustor operating condition. Negligible NOx exits the rich combustor. (2) At the SSC condition, the oxidation processes occurring in the mixing section consume 99 percent of the CO exiting the rich combustor. Soot formed in the rich combustor is also highly oxidized, with combustor exit SAE Smoke Number <3. (3) Mixing section configurations which demonstrated enhanced emissions control at SSC also performed better at part-power conditions. Data from mixer exit traverses reflected the expected mixing behavior for off-design jet to crossflow momentum-flux ratios. (4) Low power operating conditions require that the RQL combustor operate as a lean-lean combustor to achieve low CO and high efficiency. (5) A RQL combustor can achieve the emissions goal of EINOX = 5 at the Supersonic Cruise operating condition for a HSCT engine.

Experimental Assessment of the Emissions Control Potential of a Rich/Quench/ Lean Combustor for High Speed Civil Transport Aircraft Engines

NASA Technical Reports Server (NTRS)

Tacina, Robert R. (Technical Monitor); Rosfjord, T. J.; Padget, F. C.

2001-01-01

In support of Pratt & Whitney efforts to define the Rich burn/Quick mix/Lean burn (RQL) combustor for the High Speed Civil Transport (HSCT) aircraft engine, UTRC conducted a flametube-scale study of the RQL concept. Extensive combustor testing was performed at the Supersonic Cruise (SSC) condition of an HSCT engine cycle. Data obtained from probe traverses near the exit of the mixing section confirmed that the mixing section was the critical component in controlling combustor emissions. Circular-hole configurations, which produced rapidly-, highly-penetrating jets, were most effective in limiting NO(x). The spatial profiles of NO(x) and CO at the mixer exit were not directly interpretable using a simple flow model based on jet penetration, and a greater understanding of the flow and chemical processes in this section are required to optimize it. Neither the rich-combustor equivalence ratio nor its residence time was a direct contributor to the exit NO(x). Based on this study, it was also concluded that: (1) While NO(x) formation in both the mixing section and the lean combustor contribute to the overall emission, the NOx formation in the mixing section dominates. The gas composition exiting the rich combustor can be reasonably represented by the equilibrium composition corresponding to the rich combustor operating condition. Negligible NO(x) exits the rich combustor. (2) At the SSC condition, the oxidation processes occurring in the mixing section consume 99 percent of the CO exiting the rich combustor. Soot formed in the rich combustor is also highly oxidized, with combustor exit SAE Smoke Number <3. (3) Mixing section configurations which demonstrated enhanced emissions control at SSC also performed better at part-power conditions. Data from mixer exit traverses reflected the expected mixing behavior for off-design jet to crossflow momentum-flux ratios. (4) Low power operating conditions require that the RQL combustor operate as a lean-lean combustor to achieve low CO and high efficiency. (5) An RQL combustor can achieve the emissions goal of EINO(x) = 5 at the Supersonic Cruise operating condition for an HSCT engine.

Investigation of the flow-field of two parallel round jets impinging normal to a flat surface

NASA Astrophysics Data System (ADS)

Myers, Leighton M.

The flow-field features of dual jet impingement were investigated through sub-scale model experiments. The experiments were designed to simulate the environment of a Short Takeoff, and Vertical Landing, STOVL, aircraft performing a hover over the ground, at different heights. Two different dual impinging jet models were designed, fabricated, and tested. The Generation 1 Model consisted of two stainless-steel nozzles, in a tandem configuration, each with an exit diameter of approximately 12.7 mm. The front convergent nozzle was operated at the sonic Mach number of 1.0, while the rear C-D nozzle was generally operated supersonically. The nozzles were embedded in a rectangular flat plate, referred to as the lift plate, which represents a generic lifting surface. The lift plate was instrumented with 36 surface pressure taps, which were used to examine the flow entrainment and recirculation patterns caused by varying the stand-off distance from the nozzle exits to a flat ground surface. The stand-off distance was adjusted with a sliding rail frame that the ground plane was mounted to. Typical dimensionless stand-off distances (ground plane separation) were H/DR = 2 to 24. A series of measurements were performed with the Generation 1 model, in the Penn State High Speed Jet Aeroacoustics Laboratory, to characterize the basic flow phenomena associated with dual jet impingement. The regions of interest in the flow-field included the vertical jet plume(s), near impingement/turning region, and wall jet outwash. Other aspects of interest included the loss of lift (suckdown) that occurs as the ground plane separation distance becomes small, and azimuthal variation of the acoustic noise radiation. Various experimental methods and techniques were used to characterize the flow-field, including flow-visualization, pressure rake surveys, surface mounted pressure taps, laser Doppler velocimetry, and acoustic microphone arrays. A second dual impinging jet scale model, Generation 2, was designed and fabricated with a 50% increase in nozzle exit diameter. The primary design improvement is the ability to quickly and easily exchange the nozzles of the model. This allowed experiments to be performed with rapid-prototyped nozzles that feature more realistic geometry to that of tactical military aircraft engines. One such nozzle, which was designed and demonstrated by previous researchers to reduce jet noise in a free-jet, was incorporated into the model. The nozzle, featuring deflected seals, was installed in the Generation 2 model and its effect on suckdown was evaluated.

Plume mass flow and optical damage distributions for an MMH/N2O4 RCS thruster. [exhaust plume contamination of spacecraft components

NASA Technical Reports Server (NTRS)

Spisz, E. W.; Bowman, R. L.; Jack, J. R.

1973-01-01

The data obtained from two recent experiments conducted in a continuing series of experiments at the Lewis Research Center into the contamination characteristics of a 5-pound thrust MMH/N2O4 engine are presented. The primary objectives of these experiments were to establish the angular distribution of condensible exhaust products within the plume and the corresponding optical damage angular distribution of transmitting optical elements attributable to this contaminant. The plume mass flow distribution was measured by five quartz crystal microbalances (QCM's) located at the engine axis evaluation. The fifth QCM was located above the engine and 15 deg behind the nozzle exit plane. The optical damage was determined by ex-situ transmittance measurements for the wavelength range from 0.2 to 0.6 microns on 2.54 cm diameter fused silica discs also located at engine centerline elevation. Both the mass deposition and optical damage angular distributions followed the expected trend of decreasing deposition and damage as the angle between sensor or sample and the nozzle axis increased. A simple plume gas flow equation predicted the deposition distribution reasonably well for angles of up to 55 degrees. The optical damage measurements also indicated significant effects at large angles.

Aerodynamic Measurements of a Variable-Speed Power-Turbine Blade Section in a Transonic Turbine Cascade at Low Inlet Turbulence

NASA Technical Reports Server (NTRS)

Flegel-McVetta, Ashlie B.; Giel, Paul W.; Welch, Gerard E.

2013-01-01

Aerodynamic measurements obtained in a transonic linear cascade were used to assess the impact of large incidence angle and Reynolds number variations on the 3-D flow field and midspan loss and turning of a 2-D section of a variable-speed power-turbine (VSPT) rotor blade. Steady-state data were obtained for ten incidence angles ranging from +15.8 deg to -51.0 deg. At each angle, data were acquired at five flow conditions with the exit Reynolds number (based on axial chord) varying over an order-of-magnitude from 2.12×10(exp 5) to 2.12×10(exp 6). Data were obtained at the design exit Mach number of 0.72 and at a reduced exit Mach number of 0.35 as required to achieve the lowest Reynolds number. Midspan total-pressure and exit flow angle data were acquired using a five-hole pitch/yaw probe surveyed on a plane located 7.0 percent axial chord downstream of the blade trailing edge plane. The survey spanned three blade passages. Additionally, three-dimensional half-span flow fields were examined with additional probe survey data acquired at 26 span locations for two key incidence angles of +5.8 deg and -36.7 deg. Survey data near the endwall were acquired with a three-hole boundary-layer probe. The data were integrated to determine average exit total-pressure and flow angle as functions of incidence and flow conditions. The data set also includes blade static pressures measured on four spanwise planes and endwall static pressures. Tests were conducted in the NASA Glenn Transonic Turbine Blade Cascade Facility. The measurements reflect strong secondary flows associated with the high aerodynamic loading levels at large positive incidence angles and an increase in loss levels with decreasing Reynolds number. The secondary flows decrease with negative incidence as the blade becomes unloaded. Transitional flow is admitted in this low inlet turbulence dataset, making it a challenging CFD test case. The dataset will be used to advance understanding of the aerodynamic challenges associated with maintaining efficient power turbine operation over a wide shaft-speed range. deg

Performance of a Model Rich Burn-quick Mix-lean Burn Combustor at Elevated Temperature and Pressure

NASA Technical Reports Server (NTRS)

Peterson, Christopher O.; Sowa, William A.; Samuelsen, G. S.

2002-01-01

As interest in pollutant emission from stationary and aero-engine gas turbines increases, combustor engineers must consider various configurations. One configuration of increasing interest is the staged, rich burn - quick mix - lean burn (RQL) combustor. This report summarizes an investigation conducted in a recently developed high pressure gas turbine combustor facility. The model RQL combustor was plenum fed and modular in design. The fuel used for this study is Jet-A which was injected from a simplex atomizer. Emission (CO2, CO, O2, UHC, NOx) measurements were obtained using a stationary exit plane water-cooled probe and a traversing water-cooled probe which sampled from the rich zone exit and the lean zone entrance. The RQL combustor was operated at inlet temperatures ranging from 367 to 700 K, pressures ranging from 200 to 1000 kPa, and combustor reference velocities ranging from 10 to 20 m/s. Variations were also made in the rich zone and lean zone equivalence ratios. Several significant trends were observed. NOx production increased with reaction temperature, lean zone equivalence ratio and residence time and decreased with increased rich zone equivalence ratio. NOx production in the model RQL combustor increased to the 0.4 power with increased pressure. This correlation, compared to those obtained for non-staged combustors (0.5 to 0.7), suggests a reduced dependence on NOx on pressure for staged combustors. Emissions profiles suggest that rich zone mixing is not uniform and that the rich zone contributes on the order of 16 percent to the total NOx produced.

CHARACTERIZATION OF ROTATING-WING AIRCRAFT EMISSIONS

DOE Office of Scientific and Technical Information (OSTI.GOV)

Cheng, Mengdawn; Corporan, E.; Mahurin, Shannon Mark

2007-01-01

Rotating-wing aircraft or helicopters are heavily used by the US military to transport cargo, troops and personnel, and perform combat missions. Similar helicopter engines (those from the Chinook helicopter, for example) are being used by civilian companies to lift and transport heavy loads. Emissions data for this type of engines are limited but are important for development and design of air quality control strategy for military installations and bases in the country that are surrounded by cities and metropolitan areas. Major gaseous, selected air toxics, and particulate emissions data from helicopters were measured for T700-GE-700 and T700-GE-701C running JP-8 andmore » Fischer-Tropsch fuels in separate engine exhaust tests. Each engine-fuel combination test was run at three engine power levels from idle to maximum in sequence in each test in June 2007 at Hunter Army Airfield (HAAF) in Savannah, GA. The emissions from these engines were smaller than those (T33 and T56) tested earlier in terms of gas concentrations and particulate mass/number concentration. The mode diameter of a particle size distribution obtained from a test run throughout the whole campaign was smaller than 100 nm by a research-grade fast scanning mobility particle sizer, which was confirmed by a commercial scanning mobility particle sizer taking sample from a collocated position right at the engine exhaust exit plane. Use of FT fuel led to reduced particulate and gaseous emissions as compared to the use of JP-8 fuel on the same engine. Production of nanoparticles (with mobility diameter smaller than 20 nm) by the engine running on JP-8 fuel was clearly observed using a nano-DMA equipped scanning mobility particle sizer a few meters downstream from the engine exhaust plane. The production was proportional to the engine power setting, and likely to be caused by the sulfur content in the JP-8 fuel. Sulfate/sulfur data measured at the engine exhaust and the same downstream location supports such a hypothesis. Such a production was not observed when FT-fuel was used that further strengthens the hypothesis, since the sulfur content of the FT-fuel was zero. This work was supported by the Department of Defense Strategic Environmental Research and Defense Program (SERDP) under project number WP 1401.« less

Theoretical study on the effect of the design of small (milli-Newton) thruster jets on molecular contamination for the space station

NASA Technical Reports Server (NTRS)

Riley, B. R.

1986-01-01

The self-induced molecular contamination around the space station could have adverse effects on space station components (for example solar panels) as well as scientific experiments that might be done on or near the space station. Aerospace engineers need to design a space station (SS) propulsion system that keeps the SS in a stable orbit and at the same time does not allow the propellant gases to interfere with the experiments of the user. One scenario that might accomplish the above requirements is to use an electrothermal propulsion system, resistojet, that will thrust continuously in the hundreds of milli-Newton range which will provide a constant altitude for the SS with a low g environment. As a first attempt to understand the contamination from such a propulsion system, a point source model was developed. The numerical results of the point source model are given. Number column densities for CO2 are presented as a function of direction of observation (line of sight), temperature of the exit gas, and mean exit velocity. All the results are for a constant exhaust rate of 5,000 kg/year. In addition, a mathematical model to study the effect of nozzle design on the induced molecular environment around the space station produced by simple gas propellants is described. The mathematical model would allow one to follow the expansion of the gas from the throat of a nozzle to the nozzle exit plane and then into the space external to the nozzle.

Plasma Potential and Langmuir Probe Measurements in the Near-field Plume of the NASA-457Mv2 Hall Thruster

NASA Technical Reports Server (NTRS)

Shastry, Rohit; Huang, Wensheng; Herman, Daniel A.; Soulas, George C.; Kamhawi, Hani

2012-01-01

In order to further the design of future high-power Hall thrusters and provide experimental validation for ongoing modeling efforts, plasma potential and Langmuir probe measurements were performed on the 50-kW NASA-457Mv2. An electrostatic probe array comprised of a near-field Faraday probe, single Langmuir probe, and emissive probe was used to interrogate the near-field plume from approximately 0.1 - 2.0 mean thruster diameters downstream of the thruster exit plane at the following operating conditions: 300 V, 400 V and 500 V at 30 kW and 500 V at 50 kW. Results have shown that the acceleration zone is limited to within 0.4 mean thruster diameters of the exit plane while the high-temperature region is limited to 0.25 mean thruster diameters from the exit plane at all four operating conditions. Maximum plasma potentials in the near-field at 300 and 400 V were approximately 50 V with respect to cathode potential, while maximum electron temperatures varied from 24 - 32 eV, depending on operating condition. Isothermal lines at all operating conditions were found to strongly resemble the magnetic field topology in the high-temperature regions. This distribution was found to create regions of high temperature and low density near the magnetic poles, indicating strong, thick sheath formation along these surfaces. The data taken from this study are considered valuable for future design as well as modeling validation.

Far Noise Field of Air Jets and Jet Engines

NASA Technical Reports Server (NTRS)

Callaghan, Edmund E; Coles, Willard D

1957-01-01

An experimental investigation was conducted to study and compare the acoustic radiation of air jets and jet engines. A number of different nozzle-exit shapes were studied with air jets to determine the effect of exit shape on noise generation. Circular, square, rectangular, and elliptical convergent nozzles and convergent-divergent and plug nozzles were investigated. The spectral distributions of the sound power for the engine and the air jet were in good agreement for the case where the engine data were not greatly affected by reflection or jet interference effects. Such power spectra for a subsonic or slightly choked engine or air jet show that the peaks of the spectra occur at a Strouhal number of 0.3.

AALC Fan Model Test Program

DTIC Science & Technology

1979-05-08

Prediction. Scaling. Fan Design. 29AOSY14ACY (Cettm.. a P#`04 rv1s. *It UO004N 01- dew 16001 bYstI 610 01c Nh A 12-inch-diameter centrifugal fan impeller...Performance, 3500 RPM84............ 53 Inlet Bellmouth Velocity Survey , Oper Point A, Config 3. 54 Inlet Bellmouth Velocity Survey, Oper Point B...Config 3. 8& 55 Inlet Bellmouth Velocity Survey, Oper Point C, Config 3. ss So Volute Exit Plane Press. Measurement Locations, Config 3 89 57 Volute Exit

A Compendium of Solar Dish/Stirling Technology

DTIC Science & Technology

1994-01-01

systems and Plataforma Solar in Almeria, Spain, with the goal being plans to produce fourteen 7.5-kWe systems for testing to test the system’s long-term...the sun is not a point source, its rays 21 Chapter 3 (a) (b) - N Mounting Ring and CollaraI/ / I/\\ I / Virtual Exit I / Target S• Entrance I 0 L...tptical \\ I Real Exit / Virtual Target \\ Aperture\\ / Cooling \\ / I Coils N - Focal - - - - " Plane 4. Figure 3-2. A secondary concentrator with side view (a

Prediction of sound radiation from different practical jet engine inlets

NASA Technical Reports Server (NTRS)

Zinn, B. T.; Meyer, W. L.

1982-01-01

The computer codes necessary for this study were developed and checked against exact solutions generated by the point source method using the NASA Lewis QCSEE inlet geometry. These computer codes were used to predict the acoustic properties of the following five inlet configurations: the NASA Langley Bellmouth, the NASA Lewis JT15D-1 Ground Test Nacelle, and three finite hyperbolic inlets of 50, 70 and 90 degrees. Thirty-five computer runs were done for the NASA Langley Bellmouth. For each of these computer runs, the reflection coefficient at the duct exit plane was calculated as was the far field radiation pattern. These results are presented in both graphical and tabular form with many of the results cross plotted so that trends in the results verses cut-off ratio (wave number) and tangential mode number may be easily identified.

Unsteady loads due to propulsive lift configurations. Part D: The development of an experimental facility for the investigation of scaling effects on propulsive lift configurations

NASA Technical Reports Server (NTRS)

Haviland, J. K.; Herling, W. W.

1978-01-01

The design and construction of an experimental facility for the investigation of scaling effects in propulsive lift configurations are described. The facility was modeled after an existing full size NASA facility which consisted of a coaxial turbofan jet engine with a rectangular nozzle in a blown surface configuration. The flow field of the model facility was examined with and without a simulated wing surface in place at several locations downstream of the nozzle exit plane. Emphasis was placed on obtaining pressure measurements which were made with static probes and surface pressure ports connected via plastic tubing to condenser microphones for fluctuating measurements. Several pressure spectra were compared with those obtained from the NASA facility, and were used in a preliminary evaluation of scaling laws.

Core noise measurements on a YF-102 turbofan engine

NASA Technical Reports Server (NTRS)

Reshotko, M.; Karchmer, A. M.; Penko, P. F.; Mcardle, J. G.

1977-01-01

Core noise from a YF-102 high bypass ratio turbofan engine was investigated through the use of simultaneous measurements of internal fluctuating pressures and far field noise. Acoustic waveguide probes, located in the engine at the compressor exit, in the combustor, at the turbine exit, and in the core nozzle, were employed to measure internal fluctuating pressures. Spectra showed that the internal signals were free of tones, except at high frequency where machinery noise was present. Data obtained over a wide range of engine conditions suggest that below 60% of maximum fan speed the low frequency core noise contributes significantly to the far field noise.

Viscous flow computations for elliptical two-duct version of the SSME hot gas manifold

NASA Technical Reports Server (NTRS)

Roger, R. P.

1986-01-01

The objective of the effort was to numerically simulate viscous subsonic flow in a proposed elliptical two-duct version of the fuel side Hot Gas Manifold (HGM) for the Space Shuttle Main Engine (SSME). The numerical results were to complement both water flow and air flow experiments in the two-duct geometry performed at NASA-MSFC and Rocketdyne. The three-dimensional character of the HGM consists of two essentially different geometries. The first part of the construction is a concentric shell duct structure which channels the gases from a turbine exit into the second part comprised of two cylindrically shaped transfer ducts. The initial concentric shell portion can be further subdivided into a turnaround section and a bowl section. The turnaround duct (TAD) changes the direction of the mean flow by 180 degress from a smaller radius to a larger radius duct which discharges into the bowl. The cylindrical transfer ducts are attached to the bowl on one side thus providing a plane of symmetry midway between the two. Centerline flow distance from the TAD inlet to the transfer duct exit is approximately two feet. Details of the approach used to numerically simulate laminar or turbulent flow in the HGM geometry are presented. Computational results are presented and discussed.

Turbofan forced mixer lobe flow modeling. 1: Experimental and analytical assessment

NASA Technical Reports Server (NTRS)

Barber, T.; Paterson, R. W.; Skebe, S. A.

1988-01-01

A joint analytical and experimental investigation of three-dimensional flowfield development within the lobe region of turbofan forced mixer nozzles is described. The objective was to develop a method for predicting the lobe exit flowfield. In the analytical approach, a linearized inviscid aerodynamical theory was used for representing the axial and secondary flows within the three-dimensional convoluted mixer lobes and three-dimensional boundary layer analysis was applied thereafter to account for viscous effects. The experimental phase of the program employed three planar mixer lobe models having different waveform shapes and lobe heights for which detailed measurements were made of the three-dimensional velocity field and total pressure field at the lobe exit plane. Velocity data was obtained using Laser Doppler Velocimetry (LDV) and total pressure probing and hot wire anemometry were employed to define exit plane total pressure and boundary layer development. Comparison of data and analysis was performed to assess analytical model prediction accuracy. As a result of this study a planar mixed geometry analysis was developed. A principal conclusion is that the global mixer lobe flowfield is inviscid and can be predicted from an inviscid analysis and Kutta condition.

A Three-Dimensional CFD Investigation of Secondary Flow in an Accelerating, 90 deg Elbow

NASA Technical Reports Server (NTRS)

Cavicchi, Richard H.

2001-01-01

NASA Glenn Research Center has recently applied the WIND National Code flow solver to an accelerating elbow with a 90 deg. bend to reveal aspects of secondary flow. This elbow was designed by NACA in the early 1950's such that flow separation would be avoided. Experimental testing was also done at that time. The current three dimensional CFD investigation shows that separation has indeed been avoided. Using its three-dimensional capability, this investigation provides various viewpoints in several planes that display the inception, development, and final location of a passage vortex. Its shape first becomes discernible as a vortex near the exit of the bend. This rendition of the exit passage vortex compares well with that found in the experiments. The viewpoints show that the passage vortex settles on the suction surface at the exit about one-third of the distance between the plane wall and midspan. Furthermore, it projects into the mainstream to about one-third of the channel width. Of several turbulence models used in this investigation, the Spalart Alimaras, Baldwin Lomax, and SST (Shear Stress Transport) models were by far the most successful in matching the experiments.

Transition duct system with arcuate ceramic liner for delivering hot-temperature gases in a combustion turbine engine

DOE Office of Scientific and Technical Information (OSTI.GOV)

Wiebe, David J.

A transition duct system (10) for delivering hot-temperature gases from a plurality of combustors in a combustion turbine engine is provided. The system includes an exit piece (16) for each combustor. The exit piece may include an arcuate connecting segment (36). An arcuate ceramic liner (60) may be inwardly disposed onto a metal outer shell (38) along the arcuate connecting segment of the exit piece. Structural arrangements are provided to securely attach the ceramic liner in the presence of substantial flow path pressurization. Cost-effective serviceability of the transition duct systems is realizable since the liner can be readily removed andmore » replaced as needed.« less

Flow and clog in a silo with oscillating exit

NASA Astrophysics Data System (ADS)

To, Kiwing; Tai, Hsiang-Ting

2017-09-01

When grains flow out of a silo, flow rate W increases with exit size D . If D is too small, an arch may form and the flow may be blocked at the exit. To recover from clogging, the arch has to be destroyed. Here we construct a two-dimensional silo with movable exit and study the effects of exit oscillation (with amplitude A and frequency f ) on flow rate, clogging, and unclogging of grains through the exit. We find that, if exit oscillates, W remains finite even when D (measured in unit of grain diameter) is only slightly larger than one. Surprisingly, while W increases with oscillation strength Γ ≡4 π2A f2 as expected at small D , W decreases with Γ when D ≥5 due to induced random motion of the grains at the exit. When D is small and oscillation speed v ≡2 π A f is slow, temporary clogging events cause the grains to flow intermittently. In this regime, W depends only on v —a feature consistent to a simple arch breaking mechanism, and the phase boundary of intermittent flow in the D -v plane is consistent to either a power law: D ∝v-7 or an exponential form: D ∝e-D /0.55 . Furthermore, the flow time statistic is Poissonian whereas the recovery time statistic follows a power-law distribution.

Chemical composition and photochemical reactivity of exhaust from aircraft turbine engines

NASA Astrophysics Data System (ADS)

Spicer, C. W.; Holdren, M. W.; Riggin, R. M.; Lyon, T. F.

1994-10-01

Assessment of the environmental impact of aircraft emissions is required by planners and policy makers. Seveal areas of concern are: 1. exposure of airport workers and urban residents to toxic chemicals emitted when the engines operate at low power (idle and taxi) on the ground; 2. contributions to urban photochemical air pollution of aircraft volatile organic and nitrogen oxides emissions from operations around airports; and 3. emissions of nitrogen oxides and particles during high-altitude operation. The environmental impact of chemicals emitted from jet aircraft turbine engines has not been firmly established due to lack of data regarding emission rates and identities of the compounds emitted. This paper describes an experimental study of two different aircraft turbine engines designed to determine detailed organic emissions, as well as emissions of inorganic gases. Emissions were measured at several engine power settings. Measurements were made of detailed organic composition from C1 through C17, CO, CO2, NO, NOx, and polycyclic aromatic hydrocarbons. Measurements were made using a multi-port sampling pro be positioned directly behind the engine in the exhaust exit plane. The emission measurements have been used to determine the organic distribution by carbon number and the distribution by compound class at each engine power level. The sum of the organic species was compared with an independent measurement of total organic carbon to assess the carbon mass balance. A portion of the exhaust was captured and irradiated in outdoor smog chambers to assess the photochemical reactivity of the emissions with respect to ozone formation. The reactivity of emissions from the two engines was apportioned by chemical compound class.

Axial vane-type swirler performance characteristics. M.S. Thesis

NASA Technical Reports Server (NTRS)

Sander, G. F.

1983-01-01

The performance of an axial vane-type swirler was investigated to aid in computer modeling of gas turbine combustor flowfields and in evaluation of turbulence models for swirling confined jet flow. The swirler studied is annular with a hub-to-swirler diameter ratio of 0.25 and ten adjustable vanes of pitch-to-chord ratio 0.68. Measurements of time-mean axial, radial, and tangential velocities were made at the swirler exit plane using a five-hole pitot probe technique with computer data reduction. Nondimensionalized velocities from both radial and azimuthal traverses are tabulated and plotted for a range of swirl vane angles phi from 0 to 70 degrees. A study was done of idealized exit-plane velocity profiles relating the swirl numbers S and S' to the ratio of maximum swirl and axial velocities for each idealized case, and comparing the idealized swirl numbers with ones calculated from measured profiles.

Radial magnetic compression in the expelled jet of a plasma deflagration accelerator

DOE Office of Scientific and Technical Information (OSTI.GOV)

Loebner, Keith T. K., E-mail: kloebner@stanford.edu; Underwood, Thomas C.; Mouratidis, Theodore

2016-02-29

A spectroscopic study of a pulsed plasma deflagration accelerator is carried out that confirms the existence of a strong compression in the emerging jet at the exit plane of the device. An imaging spectrometer is used to collect broadened Hα emission from a transaxial slice of the emerging jet at high spatial resolution, and the radial plasma density profile is computed from Voigt fits of the Abel inverted emissivity profiles. The plasma temperature, determined via Doppler broadening of impurity line emission, is compared against the temperature predictions of a radial magnetohydrodynamic equilibrium model applied to the measured density profiles. Empiricalmore » scaling laws developed for the plasma density, combined with the measured and predicted temperatures, indicate that a radially equilibrated Z-pinch is formed within the expelled plasma jet at the exit plane during the deflagration process.« less

Navier-Stokes analysis and experimental data comparison of compressible flow within ducts

NASA Technical Reports Server (NTRS)

Harloff, G. J.; Reichert, B. A.; Sirbaugh, J. R.; Wellborn, S. R.

1992-01-01

Many aircraft employ ducts with centerline curvature or changing cross-sectional shape to join the engine with inlet and exhaust components. S-ducts convey air to the engine compressor from the intake and often decelerate the flow to achieve an acceptable Mach number at the engine compressor by increasing the cross-sectional area downstream. Circular-to-rectangular transition ducts are used on aircraft with rectangular exhaust nozzles to connect the engine and nozzle. To achieve maximum engine performance, the ducts should minimize flow total pressure loss and total pressure distortion at the duct exit. Changes in the curvature of the duct centerline or the duct cross-sectional shape give rise to streamline curvature which causes cross stream pressure gradients. Secondary flows can be caused by deflection of the transverse vorticity component of the boundary layer. This vortex tilting results in counter-rotating vortices. Additionally, the adverse streamwise pressure gradient caused by increasing cross-sectional area can lead to flow separation. Vortex pairs have been observed in the exit planes of both duct types. These vortices are due to secondary flows induced by pressure gradients resulting from streamline curvature. Regions of low total pressure are produced when the vortices convect boundary layer fluid into the main flow. The purpose of the present study is to predict the measured flow field in a diffusing S-duct and a circular-to-rectangular transition duct with a full Navier-Stokes computer program, PARC3D, and to compare the numerical predictions with new detailed experimental measurements. The work was undertaken to extend previous studies and to provide additional CFD validation data needed to help model flows with strong secondary flow and boundary layer separation. The S-duct computation extends the study of Smith et al, and Harloff et al, which concluded that the computation might be improved by using a finer grid and more advanced turbulence models. The present study compares results for both the Baldwin-Lomas and k-epsilon turbulence models and is conducted with a refined grid. For the transition duct, two inlet conditions were considered, the first with straight flow and the second with swirling flow. The first case permits examination of the effects of the geometric transition on the flow field, while the second case includes the rotational flow effect characteristic of a gas turbine engine.

High freestream turbulence studies on a scaled-up stator vane

NASA Astrophysics Data System (ADS)

Radomsky, Roger William, Jr.

2000-10-01

Today's gas turbine engines are operating at combustor exit temperatures far exceeding the maximum temperatures of the component alloys downstream of the combustor. These higher temperatures are necessary to increase the efficiency of the engine, and, as such, durability of the downstream components becomes an issue. The highly turbulent flowfield that exists at the exit of the combustor complicates issues further by increasing heat transfer from the hot gas to the component surface. To account for the high heat transfer rates, and provide a better prediction of the applied heat loads, detailed heat transfer and flowfield information is needed at turbulence levels representative those exiting a combustor. Flowfield measurements at high freestream turbulence levels indicated that turbulence, which was isotropic at the inlet, became highly anisotropic in the test section as a result of surface curvature and strain. Turbulent kinetic energy levels were shown to increase in the passage by as much as 131% and 31% for the 10% and 19.5% turbulence levels. Although the turbulent kinetic energy was high, the turbulence level based upon local velocity decreased quickly to levels of 3% and 6% near the suction surface for the 10% and 19.5% turbulence levels. For the pressure surface, local turbulence levels were as high as 10% and 16% for the 10% and 19.5% turbulence levels. High local turbulence levels and heat transfer augmentation were observed near the stagnation location, by as much as 50%, and along the pressure surface, by as much as 80%, where airfoil geometries have shown degradation after prolonged usage. Endwall flowfield measurements on a plane at the stagnation location showed that a horseshoe vortex developed in the juncture region of the vane at high freestream. turbulence similar to that at low freestream turbulence. Measurements near the center of the vortex indicated that the vortex was highly unsteady. In regions where strong secondary flows (horseshoe and passage vortex) were present, these vortices dominated the heat transfer and the augmentations due to high freestream turbulence were small.

Flow in out-of-plane double S-bonds

NASA Technical Reports Server (NTRS)

Schmidt, M. C.; Whitelaw, J. H.; Yianneskis, M.

1986-01-01

Developing flows in two out-of-plane double S-bend configurations have been measured by laser-Doppler anemometry. The first duct had a rectangular cross-section 40mmx40mm at the inlet and consisted of a uniform area 22.5 deg. - 22.5 deg. S-duct upstream with a 22.5 deg.- 22.5 deg. S- diffuser downstream. The second duct had a circular cross-section and consisted of a 45 deg. - 45 deg. uniform area S-duct upstream with a 22.5 deg. -22.5 deg. S-diffuser downstream. In both configurations the ratio of the mean radius of curvature to the inlet hydraulic diameter was 7.0, the exit-to-inlet area ratio of the diffusers was 1.5 and the ducts were connected so that the centerline of the S-duct lay in a plane normal to that of the S-diffuser. Streamwise and cross-stream velocity components were measured in laminar flow for the rectangular duct and in turbulent flow for both configurations; measurements of the turbulence levels, cross-correlations and wall static pressures were also made in the turbulent flow cases. Secondary flows of the first kind are present in the first S-duct and they are complemented or counteracted by the secondary flows generated by the area expansion and by the curvature of the S-diffusers downstream. Cross-stream velocities with magnitudes up to 0.19 and 0.11 of the bulk velocity were measured in the laminar and turbulent flows respectively in the rectangular duct and six cross-flow vortices were evident at the exit of the duct in both flow cases. The turbulent flow in the circular duct was qualitatively similar to that in the rectangular configuration, but the cross-stream velocities measured at the exit plane were smaller in the circular geometry. The results are presented in sufficient detail and accuracy for the assessment of numerical calculation methods and are listed in tabular form for this purpose.

Undulated Nozzle for Enhanced Exit Area Mixing

NASA Technical Reports Server (NTRS)

Seiner, John M. (Inventor); Gilinsky, Mikhail M. (Inventor)

2000-01-01

A nozzle having an undulating surface for enhancing the mixing of a primary flow with a secondary flow or ambient air, without requiring an ejector. The nozzle includes a nozzle structure and design for introducing counter-rotating vorticity into the primary flow either through (i) internal surface corrugations where an axisymmetric line through each corrugation is coincident with an axisymmetric line through the center of the flow passageway or (ii) through one or more sets of alternating convexities and cavities in the internal surface of the nozzle where an axisymmetric line through each convexity and cavity is coincident with an axisymmetric line through the center of the flow passageway, and where the convexities contract from the entrance end towards the exit end. Exit area mixing is also enhanced by one or more chevrons attached to the exit edge of the nozzle. The nozzle is ideally suited for application as a jet engine nozzle. When used as a jet engine nozzle, noise suppression with simultaneous thrust augmentation/minimal thrust loss is achieved.

Tuned intake air inlet for a rotary engine

DOE Office of Scientific and Technical Information (OSTI.GOV)

Corbett, W.D.; Sheaffer, B.L.

This patent describes, in a rotary internal combustion engine, an improved assembly for providing a balanced flow of combustion air to the fuel supply inlet. It comprises: a plenum chamber attached to the engine block, the plenum chamber including an air inlet adapted to receive air from the cooling air exit passage and an air outlet for the discharge of air; and an outlet conduit connecting the air outlet and the fuel supply inlet. The conduit disposed to partially surround the plenum chamber to provide a conduit length substantially greater than the distance from the cooling air exit passage totmore » he fuel supply inlet.« less

Transition duct system with straight ceramic liner for delivering hot-temperature gases in a combustion turbine engine

DOE Office of Scientific and Technical Information (OSTI.GOV)

Wiebe, David J.

A transition duct system (10) for delivering hot-temperature gases from a plurality of combustors in a combustion turbine engine is provided. The system includes an exit piece (16) for each combustor. The exit piece may include a straight path segment (26) for receiving a gas flow from a respective combustor. A straight ceramic liner (40) may be inwardly disposed onto a metal outer shell (38) along the straight path segment of the exit piece. Structural arrangements are provided to securely attach the ceramic liner in the presence of substantial flow path pressurization. Cost-effective serviceability of the transition duct systems ismore » realizable since the liner can be readily removed and replaced as needed.« less

Mode Transitions in Hall Effect Thrusters

DTIC Science & Technology

2013-07-01

Al2O3), silicon carbide ( SiC ) and graphite (C). The significant differences being ion bombardment sputter yield and secondary electron emission...channel cross-section is radially symmetric about ( mirrored above and below) discharge channel centerline from the anode to the exit plane, whereas

Characterization of Hall effect thruster propellant distributors with flame visualization

NASA Astrophysics Data System (ADS)

Langendorf, S.; Walker, M. L. R.

2013-01-01

A novel method for the characterization and qualification of Hall effect thruster propellant distributors is presented. A quantitative measurement of the azimuthal number density uniformity, a metric which impacts propellant utilization, is obtained from photographs of a premixed flame anchored on the exit plane of the propellant distributor. The technique is demonstrated for three propellant distributors using a propane-air mixture at reservoir pressure of 40 psi (gauge) (377 kPa) exhausting to atmosphere, with volumetric flow rates ranging from 15-145 cfh (7.2-68 l/min) with equivalence ratios from 1.2 to 2.1. The visualization is compared with in-vacuum pressure measurements 1 mm downstream of the distributor exit plane (chamber pressure held below 2.7 × 10-5 Torr-Xe at all flow rates). Both methods indicate a non-uniformity in line with the propellant inlet, supporting the validity of the technique of flow visualization with flame luminosity for propellant distributor characterization. The technique is applied to a propellant distributor with a manufacturing defect in a known location and is able to identify the defect and characterize its impact. The technique is also applied to a distributor with numerous small orifices at the exit plane and is able to resolve the resulting non-uniformity. Luminosity data are collected with a spatial resolution of 48.2-76.1 μm (pixel width). The azimuthal uniformity is characterized in the form of standard deviation of azimuthal luminosities, normalized by the mean azimuthal luminosity. The distributors investigated achieve standard deviations of 0.346 ± 0.0212, 0.108 ± 0.0178, and 0.708 ± 0.0230 mean-normalized luminosity units respectively, where a value of 0 corresponds to perfect uniformity and a value of 1 represents a standard deviation equivalent to the mean.

14 CFR 29.1557 - Miscellaneous markings and placards.

Code of Federal Regulations, 2011 CFR

2011-01-01

...; (iii) For turbine-engine-powered rotorcraft, the permissible fuel designations, except that if... placard and operating control for each emergency exit must differ in color from the surrounding fuselage surface as prescribed in § 29.811(h)(2). A placard must be near each emergency exit control and must...

14 CFR 29.1557 - Miscellaneous markings and placards.

Code of Federal Regulations, 2014 CFR

2014-01-01

...; (iii) For turbine-engine-powered rotorcraft, the permissible fuel designations, except that if... placard and operating control for each emergency exit must differ in color from the surrounding fuselage surface as prescribed in § 29.811(h)(2). A placard must be near each emergency exit control and must...

14 CFR 29.1557 - Miscellaneous markings and placards.

Code of Federal Regulations, 2013 CFR

2013-01-01

...; (iii) For turbine-engine-powered rotorcraft, the permissible fuel designations, except that if... placard and operating control for each emergency exit must differ in color from the surrounding fuselage surface as prescribed in § 29.811(h)(2). A placard must be near each emergency exit control and must...

14 CFR 29.1557 - Miscellaneous markings and placards.

Code of Federal Regulations, 2010 CFR

2010-01-01

...; (iii) For turbine-engine-powered rotorcraft, the permissible fuel designations, except that if... placard and operating control for each emergency exit must differ in color from the surrounding fuselage surface as prescribed in § 29.811(h)(2). A placard must be near each emergency exit control and must...

14 CFR 29.1557 - Miscellaneous markings and placards.

Code of Federal Regulations, 2012 CFR

2012-01-01

...; (iii) For turbine-engine-powered rotorcraft, the permissible fuel designations, except that if... placard and operating control for each emergency exit must differ in color from the surrounding fuselage surface as prescribed in § 29.811(h)(2). A placard must be near each emergency exit control and must...

DOE Office of Scientific and Technical Information (OSTI.GOV)

M Severson; M Bissen; M Fisher

SRC has recently commissioned a new Varied Line-Spacing Plane Grating Monochromator (VLS-PGM) utilizing as its source a 1 m long APPLE II insertion device in short-straight-section 9 of the Aladdin storage ring. The insertion device reliably delivers horizontal, vertical, and right and left circularly polarized light to the beamline. Measurements from an in situ polarimeter can be used for undulator corrections to compensate for depolarizing effects of the beamline. The beamline has only three optical elements and covers the energy range from 11.1 to 270 eV using two varied line-spacing gratings. A plane mirror rotates to illuminate the gratings atmore » the correct angle to cancel the defocus term at all photon energies. An exit slit and elliptical-toroid refocusing mirror complete the beamline. Using a 50 {mu}m exit slit, the beamline provides moderate to high resolution, with measured flux in the mid 10{sup 12} (photons/s/200 mA) range, and a spot size of 400 {mu}m horizontal by 30 {mu}m vertical.« less

Full-Scale Tests of a New Type NACA Nose-Slot Cowling

NASA Technical Reports Server (NTRS)

Theodorsen, Theodore; Brevoort, M J; Stickle, George W; Gough, M N

1937-01-01

An extended experimental study has been made in regard to the various refinements in the design of engine cowlings as related to the propeller-nacelle unit as a whole, under conditions corresponding to take-off, climb, and normal flight. The tests were all conducted at full scale in the 20-foot wind tunnel. This report presents the results of a novel type of engine cowling, characterized by the fact that the exit opening discharging the cooling air is not, as usual, located behind the engine but at the foremost extremity or nose of the cowling. The efficiency is found to be high, owing to the fact that higher velocities may be used in the exit opening.

Electronic Engineering Technology Program Exit Examination as an ABET and Self-Assessment Tool

ERIC Educational Resources Information Center

Thomas, Gary; Darayan, Shahryar

2018-01-01

Every engineering, computing, and engineering technology program accredited by the Accreditation Board for Engineering and Technology (ABET) has formulated many and varied self-assessment methods. Methods used to assess a program for ABET accreditation and continuous improvement are for keeping programs current with academic and industrial…

Transition duct system with metal liners for delivering hot-temperature gases in a combustion turbine engine

DOE Office of Scientific and Technical Information (OSTI.GOV)

Wiebe, David J.

A transition duct system (10) for delivering hot-temperature gases from a plurality of combustors in a combustion turbine engine is provided. The system includes an exit piece (16) for each combustor. The exit piece may include a straight path segment (26) and an arcuate connecting segment (36). A respective straight metal liner (92) and an arcuate metal liner (94) may be each inwardly disposed onto a metal outer shell (38) along the straight path segment and the arcuate connecting segment (36) of the exit piece. Structural arrangements are provided to securely attach the respective liners in the presence of substantialmore » flow path pressurization. Cost-effective serviceability of the transition duct systems is realizable since the liners can be readily removed and replaced as needed.« less

An experimental investigation of the aerodynamics and cooling of a horizontally-opposed air-cooled aircraft engine installation

NASA Technical Reports Server (NTRS)

Miley, S. J.; Cross, E. J., Jr.; Owens, J. K.; Lawrence, D. L.

1981-01-01

A flight-test based research program was performed to investigate the aerodynamics and cooling of a horizontally-opposed engine installation. Specific areas investigated were the internal aerodynamics and cooling mechanics of the installation, inlet aerodynamics, and exit aerodynamics. The applicable theory and current state of the art are discussed for each area. Flight-test and ground-test techniques for the development of the cooling installation and the solution of cooling problems are presented. The results show that much of the internal aerodynamics and cooling technology developed for radial engines are applicable to horizontally opposed engines. Correlation is established between engine manufacturer's cooling design data and flight measurements of the particular installation. Also, a flight-test method for the development of cooling requirements in terms of easily measurable parameters is presented. The impact of inlet and exit design on cooling and cooling drag is shown to be of major significance.

Static investigation of a two-dimensional convergent-divergent exhaust nozzle with multiaxis thrust-vectoring capability

NASA Technical Reports Server (NTRS)

Taylor, John G.

1990-01-01

An investigation was conducted in the Static Test Facility of the NASA Langley 16-Foot Transonic Tunnel to determine the internal performance of two-dimensional convergent-divergent nozzles designed to have simultaneous pitch and yaw thrust vectoring capability. This concept utilized divergent flap rotation of thrust vectoring in the pitch plane and deflection of flat yaw flaps hinged at the end of the sidewalls for yaw thrust vectoring. The hinge location of the yaw flaps was varied at four positions from the nozzle exit plane to the throat plane. The yaw flaps were designed to contain the flow laterally independent of power setting. In order to eliminate any physical interference between the yaw flap deflected into the exhaust stream and the divergent flaps, the downstream corners of both upper and lower divergent flaps were cut off to allow for up to 30 deg of yaw flap deflection. The impact of varying the nozzle pitch vector angle, throat area, yaw flap hinge location, yaw flap length, and yaw flap deflection angle on nozzle internal performance characteristics, was studied. High-pressure air was used to simulate jet exhaust at nozzle pressure ratios up to 7.0. Static results indicate that configurations with the yaw flap hinge located upstream of the exit plane provide relatively high levels of thrust vectoring efficiency without causing large losses in resultant thrust ratio. Therefore, these configurations represent a viable concept for providing simultaneous pitch and yaw thrust vectoring.

A dual two dimensional electronic portal imaging device transit dosimetry model based on an empirical quadratic formalism

PubMed Central

Metwaly, M; Glegg, M; Baggarley, S P; Elliott, A

2015-01-01

Objective: This study describes a two dimensional electronic portal imaging device (EPID) transit dosimetry model that can predict either: (1) in-phantom exit dose, or (2) EPID transit dose, for treatment verification. Methods: The model was based on a quadratic equation that relates the reduction in intensity to the equivalent path length (EPL) of the attenuator. In this study, two sets of quadratic equation coefficients were derived from calibration dose planes measured with EPID and ionization chamber in water under reference conditions. With two sets of coefficients, EPL can be calculated from either EPID or treatment planning system (TPS) dose planes. Consequently, either the in-phantom exit dose or the EPID transit dose can be predicted from the EPL. The model was tested with two open, five wedge and seven sliding window prostate and head and neck intensity-modulated radiation therapy (IMRT) fields on phantoms. Results were analysed using absolute gamma analysis (3%/3 mm). Results: The open fields gamma pass rates were >96.8% for all comparisons. For wedge and IMRT fields, comparisons between predicted and TPS-computed in-phantom exit dose resulted in mean gamma pass rate of 97.4% (range, 92.3–100%). As for the comparisons between predicted and measured EPID transit dose, the mean gamma pass rate was 97.5% (range, 92.6–100%). Conclusion: An EPID transit dosimetry model that can predict in-phantom exit dose and EPID transit dose was described and proven to be valid. Advances in knowledge: The described model is practical, generic and flexible to encourage widespread implementation of EPID dosimetry for the improvement of patients' safety in radiotherapy. PMID:25969867

Computer Modeling of a Rotating Detonation Engine in a Rocket Configuration

DTIC Science & Technology

2015-03-01

than the ambient pressure P0, the nozzle was fully supersonic . If the calculated pressure P9 after the normal shock was less than the ambient...18 Gas Properties...66 vii Nomenclature Variable Definition 3∗ Entrance to RDE 4 RDE exit 8 Nozzle 9 Nozzle exit A Area a Speed of

A method for calculating strut and splitter plate noise in exit ducts: Theory and verification

NASA Technical Reports Server (NTRS)

Fink, M. R.

1978-01-01

Portions of a four-year analytical and experimental investigation relative to noise radiation from engine internal components in turbulent flow are summarized. Spectra measured for such airfoils over a range of chord, thickness ratio, flow velocity, and turbulence level were compared with predictions made by an available rigorous thin-airfoil analytical method. This analysis included the effects of flow compressibility and source noncompactness. Generally good agreement was obtained. This noise calculation method for isolated airfoils in turbulent flow was combined with a method for calculating transmission of sound through a subsonic exit duct and with an empirical far-field directivity shape. These three elements were checked separately and were individually shown to give close agreement with data. This combination provides a method for predicting engine internally generated aft-radiated noise from radial struts and stators, and annular splitter rings. Calculated sound power spectra, directivity, and acoustic pressure spectra were compared with the best available data. These data were for noise caused by a fan exit duct annular splitter ring, larger-chord stator blades, and turbine exit struts.

Effect of exit beam phase aberrations on coherent x-ray reconstructions of Au nanocrystals

NASA Astrophysics Data System (ADS)

Hruszkewycz, Stephan; Harder, Ross; Fuoss, Paul

2010-03-01

Current studies in coherent x-ray diffractive imaging (CXDI) are focusing on in-situ imaging under a variety of environmental conditions. Such studies often involve environmental sample chambers through which the x-ray beam must pass before and after interacting with the sample: i.e. cryostats or high pressure cells. Such sample chambers usually contain polycrystalline x-ray windows with structural imperfections that can in turn interact with the diffracted beam. A phase object in the near field that interacts with the beam exiting the sample can introduce distortions at the detector plane that may affect coherent reconstructions. We investigate the effects of a thin beryllium membrane on the coherent exit beam of a gold nanoparticle. We compare three dimensional reconstructions from experimental diffraction patterns measured with and without a 380 micron thick Be dome and find that the reconstructions are reproducible within experimental errors. Simulated near-field distortions of the exit beam consistent with micron sized voids in Be establish a ``worst case scenario'' where distorted diffraction patterns inhibit accurate inversions.

Device for installing rocket engines

NASA Technical Reports Server (NTRS)

George, T. R., Jr. (Inventor)

1976-01-01

A device for installing rocket engines is reported that is supported at a cant relative to vertical by an axially extensible, tiltable pedestal. A lifting platform supports the rocket engine at its thrust chamber exit, including a mount having a concentric base characterized by a concave bearing surface, a plurality of uniformly spaced legs extended radially from the base, and an annular receiver coaxially aligned with the base and affixed to the distal ends of said legs for receiving the thrust chamber exit. The lifting platform rests on a seat concentrically related to the pedestal and affixed to an extended end portion thereof having a convex bearing surface mated in sliding engagement with the concave bearing surface of the annular base for accommodating a rocking motion of the platform.

Theoretical analysis of an augmentor wing for a VTOL fighter

NASA Technical Reports Server (NTRS)

Dillenius, M. F. E.; Mendenhall, M. R.

1979-01-01

A method based on potential flow theory was developed for predicting forces and moments acting on augmentor wings for prescribed ejector jet characteristics. A three dimensional nonplanar vortex lattice is laid out on the chordal planes of the augmentor wing components. Jet induced effects are included in the boundary condition from which the horseshoe vortex strengths are obtained. The jet within the diffusor is made to expand from the primary nozzles to the diffusor exit and is represented by a distribution of vorticity on the jet boundary to provide proper entrainment. The jet downstream of the diffusor exit is modeled by a vorticity distribution and blockage panels and its centerline location and spreading rate are taken from experimental data. The vortex lattice and jet models are used in an iterative manner until the predicted diffusor exit velocity matches the specified one. Some comparisons with available data show good agreement at lower power settings.

An ultra-high vacuum chamber for scattering experiments featuring in-vacuum continuous in-plane variation of the angle between entrance and exit vacuum ports

DOE Office of Scientific and Technical Information (OSTI.GOV)

Englund, Carl-Johan; Agåker, Marcus, E-mail: marcus.agaker@physics.uu.se; Fredriksson, Pierre

2015-09-15

A concept that enables in-vacuum continuous variation of the angle between two ports in one plane has been developed and implemented. The vacuum chamber allows for measuring scattering cross sections as a function of scattering angle and is intended for resonant inelastic X-ray scattering experiments. The angle between the ports can be varied in the range of 30°-150°, while the pressure change is less than 2 × 10{sup −10} mbars.

Nonlinear Modeling and Control of a Propellant Mixer

NASA Technical Reports Server (NTRS)

Barbieri, Enrique; Richter, Hanz; Figueroa, Fernando

2003-01-01

A mixing chamber used in rocket engine combustion testing at NASA Stennis Space Center is modeled by a second order nonlinear MIMO system. The mixer is used to condition the thermodynamic properties of cryogenic liquid propellant by controlled injection of the same substance in the gaseous phase. The three inputs of the mixer are the positions of the valves regulating the liquid and gas flows at the inlets, and the position of the exit valve regulating the flow of conditioned propellant. The outputs to be tracked and/or regulated are mixer internal pressure, exit mass flow, and exit temperature. The outputs must conform to test specifications dictated by the type of rocket engine or component being tested downstream of the mixer. Feedback linearization is used to achieve tracking and regulation of the outputs. It is shown that the system is minimum-phase provided certain conditions on the parameters are satisfied. The conditions are shown to have physical interpretation.

Preliminary study of temperature measurement techniques for Stirling engine reciprocating seals

NASA Technical Reports Server (NTRS)

Wilcock, D. F.; Hoogenboom, L.; Meinders, M.; Winer, W. O.

1981-01-01

Methods of determining the contact surface temperature in reciprocating seals are investigated. Direct infrared measurement of surface temperatures of a rod exiting a loaded cap seal or simulated seal are compared with surface thermocouple measurements. Significant cooling of the surface requires several milliseconds so that exit temperatures may be considered representative of internal contact temperatures.

Novel Sampling Techniques for Measurement of Turbine Engine Total Particulate Matter Emissions

EPA Science Inventory

This is the first progress report of a study to evaluate two different condensation devices for the measurement of the total (volatile + non-volatile) particulate matter (PM) emissions from aircraft turbine engines by direct sampling at the engine exit. The characteristics of th...

Heat Transfer in a Superelliptic Transition Duct

NASA Technical Reports Server (NTRS)

Poinsatte, Philip; Thurman, Douglas; Hippensteele, Steven

2008-01-01

Local heat transfer measurements were experimentally mapped using a transient liquid-crystal heat transfer technique on the surface of a circular-to-rectangular transition duct. The transition duct had a length-to-diameter ratio of 1.5 and an exit-plane aspect ratio of 3. The crosssectional geometry was defined by the equation of a superellipse. The cross-sectional area was the same at the inlet and exit but varied up to 15 percent higher through the transition. The duct was preheated to a uniform temperature (nominally 64 C) before allowing room temperature air to be suddenly drawn through it. As the surface cooled, the resulting isothermal contours on the duct surface were revealed using a surface coating of thermochromic liquid crystals that display distinctive colors at particular temperatures. A video record was made of the surface temperature and time data for all points on the duct surfaces during each test. Using this surface temperature-time data together with the temperature of the air flowing through the model and the initial temperature of the model wall, the heat transfer coefficient was calculated by employing the classic one-dimensional, semi-infinite wall heat transfer conduction model. Test results are reported for inlet diameter-based Reynolds numbers ranging from 0.4x106 to 2.4x106 and two grid-generated freestream turbulence intensities of about 1 percent, which is typical of wind tunnels, and up to 16 percent, which may be more typical of real engine conditions.

Investigation of the Rocket Induced Flow Field in a Rectangular Duct

NASA Technical Reports Server (NTRS)

Landrum, D. Brian; Thames, Mignon; Parkinson, Doug; Gautney, Serena; Hawk, Clark

1999-01-01

Several tests were performed on a one-sixth scale Rocket Based Combined Cycle (RBCC) engine model at the University of Alabama in Huntsville. The UAH RBCC facility consists of a rectangular duct with a vertical strut mounted in the center. The scaled strut consists of two supersonic rocket nozzles with an embedded vertical turbine between the rocket nozzles. The tests included mass flow, flow visualization and horizontal pressure traverses. The mass flow test indicated a c:hoked condition when the rocket chamber pressure is between 200 psi and 300 psi. The flow visualization tests narrowed the rocket chamber pressure range from, 250 psi to 300 psi. Also, from this t.est, an assumption of a minimum

Application of composite materials to turbofan engine fan exit guide vanes

NASA Technical Reports Server (NTRS)

Smith, G. T.

1980-01-01

A program was conducted by NASA with the JT9D engine manufacturer to develop a lightweight, cost effective, composite material fan exit guide vane design having satisfactory structural durability for commerical engine use. Based on the results of a previous company supported program, eight graphite/epoxy and graphite-glass/epoxy guide vane designs were evaluated and four were selected for fabrication and testing. Two commercial fabricators each fabricated 13 vanes. Fatigue tests were used to qualify the selected design configurations under nominally dry, 38 C (100 F) and fully wet and 60 C (140 F) environmental conditions. Cost estimates for a production rate of 1000 vanes per month ranged from 1.7 to 2.6 times the cost of an all aluminum vane. This cost is 50 to 80 percent less than the initial program target cost ratio which was 3 times the cost of an aluminum vane. Application to the JT9D commercial engine is projected to provide a weight savings of 236 N (53 lb) per engine.

Stationary Plasma Thruster Plume Emissions

NASA Technical Reports Server (NTRS)

Manzella, David H.

1994-01-01

The emission spectrum from a xenon plasma produced by a Stationary Plasma Thruster provided by the Ballistic Missile Defense Organization (BMDO) was measured. Approximately 270 individual Xe I, Xe II, and XE III transitions were identified. A total of 250 mW of radiated optical emission was estimated from measurements taken at the thruster exit plane. There was no evidence of erosion products in the emission signature. Ingestion and ionization of background gas at elevated background pressure was detected. The distribution of excited states could be described by temperatures ranging from fractions of 1 eV to 4 eV with a high degree of uncertainty due to the nonequilibrium nature of this plasma. The plasma was over 95 percent ionized at the thruster exit plane. Between 10 and 20 percent of the ions were doubly charged. Two modes of operation were identified. The intensity of plasma emission increased by a factor of two during operation in an oscillatory mode. The transfer between the two modes of operation was likely related to unidentified phenomena occurring on a time scale of minutes.

Development of a Tetrameric Streptavidin Mutein with Reversible Biotin Binding Capability: Engineering a Mobile Loop as an Exit Door for Biotin

PubMed Central

O'Sullivan, Valerie J.; Barrette-Ng, Isabelle; Hommema, Eric; Hermanson, Greg T.; Schofield, Mark; Wu, Sau-Ching; Honetschlaeger, Claudia; Ng, Kenneth K.-S.; Wong, Sui-Lam

2012-01-01

A novel form of tetrameric streptavidin has been engineered to have reversible biotin binding capability. In wild-type streptavidin, loop3–4 functions as a lid for the entry and exit of biotin. When biotin is bound, interactions between biotin and key residues in loop3–4 keep this lid in the closed state. In the engineered mutein, a second biotin exit door is created by changing the amino acid sequence of loop7–8. This door is mobile even in the presence of the bound biotin and can facilitate the release of biotin from the mutein. Since loop7–8 is involved in subunit interactions, alteration of this loop in the engineered mutein results in an 11° rotation between the two dimers in reference to wild-type streptavidin. The tetrameric state of the engineered mutein is stabilized by a H127C mutation, which leads to the formation of inter-subunit disulfide bonds. The biotin binding kinetic parameters (koff of 4.28×10−4 s−1 and Kd of 1.9×10−8 M) make this engineered mutein a superb affinity agent for the purification of biotinylated biomolecules. Affinity matrices can be regenerated using gentle procedures, and regenerated matrices can be reused at least ten times without any observable reduction in binding capacity. With the combination of both the engineered mutein and wild-type streptavidin, biotinylated biomolecules can easily be affinity purified to high purity and immobilized to desirable platforms without any leakage concerns. Other potential biotechnological applications, such as development of an automated high-throughput protein purification system, are feasible. PMID:22536357

Performance of a high-work, low-aspect-ratio turbine stator tested with a realistic inlet radial temperature gradient

NASA Technical Reports Server (NTRS)

Stabe, Roy G.; Schwab, John R.

1991-01-01

A 0.767-scale model of a turbine stator designed for the core of a high-bypass-ratio aircraft engine was tested with uniform inlet conditions and with an inlet radial temperature profile simulating engine conditions. The principal measurements were radial and circumferential surveys of stator-exit total temperature, total pressure, and flow angle. The stator-exit flow field was also computed by using a three-dimensional Navier-Stokes solver. Other than temperature, there were no apparent differences in performance due to the inlet conditions. The computed results compared quite well with the experimental results.

Design and two dimensional cascade test of a jet-flap turbine stator blade with ratio of axial chord to spacing of 0.5

NASA Technical Reports Server (NTRS)

Stabe, R. G.

1971-01-01

A jet-flap blade was designed for a velocity diagram typical of the first-stage stator of a jet engine turbine and was tested in a simple two-dimensional cascade of six blades. The principal measurements were blade surface static pressure and cross-channel surveys of exit total pressure, static pressure, and flow angle. The results of the experimental investigation include blade loading, exit angle, flow, and loss data for a range of exit critical velocity ratios and three jet flow conditions.

Develop preliminary engineering design and study the benefits of providing an access to the Indiana toll road at State Road 327 near Orland, IN.

DOT National Transportation Integrated Search

2016-10-05

A new toll road exit is proposed at the intersection of I-80/90 and SR 327 in Orland, : Indiana. The new exit is needed to facilitate travel for a proposed manufacturing plant to be : established on a 500-acre site east of SR 327 and south of I-80/90...

Design and Test of a Liquid Oxygen / Liquid Methane Thruster with Cold Helium Pressurization Heat Exchanger

NASA Technical Reports Server (NTRS)

Melcher, John C.; Morehead, Robert L.; Atwell, Matthew J.; Hurlbert, Eric A.

2015-01-01

A liquid oxygen / liquid methane 2,000 lbf thruster was designed and tested in conjuction with a nozzle heat exchanger for cold helium pressurization. Cold helium pressurization systems offer significant spacecraft vehicle dry mass savings since the pressurant tank size can be reduced as the pressurant density is increased. A heat exchanger can be incorporated into the main engine design to provide expansion of the pressurant supply to the propellant tanks. In order to study the systems integration of a cold-helium pressurization system, a 2,000 lbf thruster with a nozzle heat exchanger was designed for integration into the Project Morpheus vehicle at NASA Johnson Space Center. The testing goals were to demonstrate helium loading and initial conditioning to low temperatures, high-pressure/low temperature storage, expansion through the main engine heat exchanger, and propellant tank injection/pressurization. The helium pressurant tank was an existing 19 inch diameter composite-overwrap tank, and the targert conditions were 4500 psi and -250 F, providing a 2:1 density advantage compared to room tempatrue storage. The thruster design uses like-on-like doublets in the injector pattern largely based on Project Morpheus main engine hertiage data, and the combustion chamber was designed for an ablative chamber. The heat exchanger was installed at the ablative nozzle exit plane. Stand-alone engine testing was conducted at NASA Stennis Space Center, including copper heat-sink chambers and highly-instrumented spoolpieces in order to study engine performance, stability, and wall heat flux. A one-dimensional thermal model of the integrated system was completed. System integration into the Project Morpheus vehicle is complete, and systems demonstrations will follow.

Analysis of nulling phase functions suitable to image plane coronagraphy

NASA Astrophysics Data System (ADS)

Hénault, François; Carlotti, Alexis; Vérinaud, Christophe

2016-07-01

Coronagraphy is a very efficient technique for identifying and characterizing extra-solar planets orbiting in the habitable zone of their parent star, especially in a space environment. An important family of coronagraphs is actually based on phase plates located at an intermediate image plane of the optical system, and spreading the starlight outside the "Lyot" exit pupil plane of the instrument. In this commutation we present a set of candidate phase functions generating a central null at the Lyot plane, and study how it propagates to the image plane of the coronagraph. These functions include linear azimuthal phase ramps (the well-known optical vortex), azimuthally cosine-modulated phase profiles, and circular phase gratings. Nnumerical simulations of the expected null depth, inner working angle, sensitivity to pointing errors, effect of central obscuration located at the pupil or image planes, and effective throughput including image mask and Lyot stop transmissions are presented and discussed. The preliminary conclusion is that azimuthal cosine functions appear as an interesting alternative to the classical optical vortex of integer topological charge.

Oscillating flow loss test results in Stirling engine heat exchangers

NASA Technical Reports Server (NTRS)

Koester, G.; Howell, S.; Wood, G.; Miller, E.; Gedeon, D.

1990-01-01

The results are presented for a test program designed to generate a database of oscillating flow loss information that is applicable to Stirling engine heat exchangers. The tests were performed on heater/cooler tubes of various lengths and entrance/exit configurations, on stacked and sintered screen regenerators of various wire diameters and on Brunswick and Metex random fiber regenerators. The test results were performed over a range of oscillating flow parameters consistent with Stirling engine heat exchanger experience. The tests were performed on the Sunpower oscillating flow loss rig which is based on a variable stroke and variable frequency linear drive motor. In general, the results are presented by comparing the measured oscillating flow losses to the calculated flow losses. The calculated losses are based on the cycle integration of steady flow friction factors and entrance/exit loss coefficients.

SU-F-T-258: Efficacy of Exit Fluence-Based Dose Calculation for Prostate Radiation Therapy

DOE Office of Scientific and Technical Information (OSTI.GOV)

Siebers, J; Gardner, J; Neal, B

Purpose: To investigate the efficacy of exit-fluence-based dose computation for prostate radiotherapy by determining if it estimates true dose more accurately than the original planning dose. Methods: Virtual exit-fluencebased dose computation was performed for 19 patients, each with 9–12 repeat CT images. For each patient, a 78 Gy treatment plan was created utilizing 5 mm CTV-to-PTV and OAR-to-PRV margins. A Monte Carlo framework was used to compute dose and exit-fluence images for the planning image and for each repeat CT image based on boney-anatomyaligned and prostate-centroid-aligned CTs. Identical source particles were used for the MC dose-computations on the planning andmore » repeat CTs to maximize correlation. The exit-fluence-based dose and image were computed by multiplying source particle weights by FC(x,y)=FP(x,y)/FT(x,y), where (x,y) are the source particle coordinates projected to the exit-fluence plane and we denote the dose/fluence from the plan by (DP,FP), from the repeat-CT as (DT,FT), and the exit-fluence computation by (DFC,FFC). DFC mimics exit-fluence backprojection through the planning image as FT=FFC. Dose estimates were intercompared to judge the efficacy of exit-fluence-based dose computation. Results: Boney- and prostate-centroid aligned results are combined as there is no statistical difference between them, yielding 420 dose comparisons per dose-volume metric. DFC is more accurate than DP for 46%, 33%, and 44% of cases in estimating CTV D98, D50, and D2 respectively. DFC improved rectum D50 and D2 estimates 54% and 49% respectively and bladder D50 and D2 47 and 49% respectively. While averaged over all patients and images DFC and DP were within 3.1% of DT, they differed from DT by as much as 22% for GTV D98, 71% for the Bladder D50, 17% for Bladder D2, 19% for Rectum D2. Conclusion: Exit-fluence based dose computations infrequently improve CTV or OAR dose estimates and should be used with caution. Research supported in part by Varian Medical Systems.« less

SU-F-T-229: A Novel Method for EPID-Based In-Vivo Exit Dose Verification for Intensity Modulated Radiotherapy

DOE Office of Scientific and Technical Information (OSTI.GOV)

Wu, Z; Wang, J; Peng, J

Purpose: Electronic portal imaging device (EPID) can be used to acquire a two-dimensional exit dose distribution during treatment delivery, thus allowing the in-vivo verification of the dose delivery through a comparison of measured portal images to predicted portal dose images (PDI). The aim of this study was to present a novel method to easily and accurately predict PDI, and to establish an EPID-based in-vivo dose verification method during IMRT treatments. Methods: We developed a model to determine the predicted portal dose at the same plane of the EPID detector location. The Varian EPID (aS1000) positions at 150cm source-to-detector-distance (SDD), andmore » can be used to acquire in-vivo exit dose using Portal Dosimetry (PD) function. Our model was generated to make an equivalent water thickness represent the buildup plate of EPID. The exit dose at extend SDD plane with patient CT data in the beam can be calculated as the predicted PDI in the treatment planning system (TPS). After that, the PDI was converted to the fluence at SDD of 150cm using the inverse square law coded in MATLAB. Five head-and-neck and prostate IMRT patient plans contain 32 fields were investigated to evaluate the feasibility of this new method. The measured EPID image was compared with PDI using the gamma analysis. Results: The average results for cumulative dose comparison were 81.9% and 91.6% for 3%, 3mm and 4%, 4mm gamma criteria, respectively. Results indicate that the patient transit dosimetry predicted algorithm compares well with EPID measured PD doses for test situations. Conclusion: Our new method can be used as an easy and feasible tool for online EPID-based in-vivo dose delivery verification for IMRT treatments. It can be implemented for fast detecting those obvious treatment delivery errors for individual field and patient quality assurance.« less

Exact Solution of the Two-Dimensional Problem on an Impact Ideal-Liquid Jet

NASA Astrophysics Data System (ADS)

Belik, V. D.

2018-05-01

The two-dimensional problem on the collision of a potential ideal-liquid jet, outflowing from a reservoir through a nozzle, with an infinite plane obstacle was considered for the case where the distance between the nozzle exit section and the obstacle is finite. An exact solution of this problem has been found using methods of the complex-variable function theory. Simple analytical expressions for the complex velocity of the liquid, its flow rate, and the force of action of the jet on the obstacle have been obtained. The velocity distributions of the liquid at the nozzle exit section, in the region of spreading of the jet, and at the obstacle have been constructed for different distances between the nozzle exit section and the obstacle. Analytical expressions for the thickness of the boundary layer and the Nusselt number at the point of stagnation of the jet have been obtained. A number of distributions of the local friction coefficient and the Nusselt number of the indicated jet are presented.

Simultaneous high-speed gas property measurements at the exhaust gas recirculation cooler exit and at the turbocharger inlet of a multicylinder diesel engine using diode-laser-absorption spectroscopy.

PubMed

Jatana, Gurneesh S; Magee, Mark; Fain, David; Naik, Sameer V; Shaver, Gregory M; Lucht, Robert P

2015-02-10

A diode-laser-absorption-spectroscopy-based sensor system was used to perform high-speed (100 Hz to 5 kHz) measurements of gas properties (temperature, pressure, and H(2)O vapor concentration) at the turbocharger inlet and at the exhaust gas recirculation (EGR) cooler exit of a diesel engine. An earlier version of this system was previously used for high-speed measurements of gas temperature and H(2)O vapor concentration in the intake manifold of the diesel engine. A 1387.2 N m tunable distributed feedback diode laser was used to scan across multiple H(2)O absorption transitions, and the direct absorption signal was recorded using a high-speed data acquisition system. Compact optical connectors were designed to conduct simultaneous measurements in the intake manifold, the EGR cooler exit, and the turbocharger inlet of the engine. For measurements at the turbocharger inlet, these custom optical connectors survived gas temperatures as high as 800 K using a simple and passive arrangement in which the temperature-sensitive components were protected from high temperatures using ceramic insulators. This arrangement reduced system cost and complexity by eliminating the need for any active water or oil cooling. Diode-laser measurements performed during steady-state engine operation were within 5% of the thermocouple and pressure sensor measurements, and within 10% of the H(2)O concentration values derived from the CO(2) gas analyzer measurements. Measurements were also performed in the engine during transient events. In one such transient event, where a step change in fueling was introduced, the diode-laser sensor was able to capture the 30 ms change in the gas properties; the thermocouple, on the other hand, required 7.4 s to accurately reflect the change in gas conditions, while the gas analyzer required nearly 600 ms. To the best of our knowledge, this is the first implementation of such a simple and passive arrangement of high-temperature optical connectors as well as the first documented application of diode-laser absorption for high-speed gas dynamics measurements in the turbocharger inlet and EGR cooler exit of a diesel engine.

Design of the soft x-ray tomography beamline at Taiwan photon source

DOE Office of Scientific and Technical Information (OSTI.GOV)

Su, Yi-Jr, E-mail: su.yj@nsrrc.org.tw; Fu, Huang-Wen; Chung, Shih-Chun

2016-07-27

The optical design of the varied-line-spacing plane-grating monochromator for transmission full-field imaging of frozen-hydrated biological samples at NSRRC is presented. This monochromator consists of a plane mirror and three interchangeable gratings with groove densities 600, 1200 and 2400 l/mm to cover the energy range 260 – 2600 eV. The groove parameters of the varied-line-spacing plane gratings are designed to minimize the effect of coma and spherical aberration to maintain the exit slit in focus for any value of incident angle. All parameters of optical components at the beamline are verified with a ray-tracing method. In the beamline design, the calculatedmore » results from the ray-tracing codes and the expected performances are discussed.« less

Short frontal plane fractures involving the dorsoproximal articular surface of the proximal phalanx: Description of the injury and a technique for repair.

PubMed

Wright, I M; Minshall, G J

2018-01-01

Chip fractures of the dorsoproximal articular margin of the proximal phalanx are common injuries in racehorses. Large fractures can extend distal to the joint capsule insertion and have been described as dorsal frontal fractures. To report the location and morphology of short frontal plane fractures involving the dorsoproximal articular surface of the proximal phalanx and describe a technique for repair under arthroscopic and radiographic guidance. Single centre retrospective case study. Case records of horses with frontal plane fractures restricted to the dorsoproximal epiphysis and metaphysis of the proximal phalanx referred to Newmarket Equine Hospital were retrieved, images reviewed and lesion morphology described. A technique for repair and the results obtained are reported. A total of 22 fractures in 21 horses commencing at the proximal articular surface exited the dorsal cortex of the proximal phalanx distal to the metacarpophalangeal/metatarsophalangeal joint capsule in 17 hind- and five forelimbs. All were in Thoroughbred racehorses. In 16 cases these were acute racing or training injuries; 20 fractures were medial, one lateral and one was midline. All were repaired with a single lag screw using arthroscopic and radiographically determined landmarks. A total of 16 horses raced after surgery with performance data similar to their preinjury levels. The study demonstrates substantial morphological similarities between individual lesions supporting a common pathophysiology, but does not identify precise causation. There are no cases managed differently that might permit assessment of the comparative efficacy of the treatment described. Short frontal plane fractures involving the dorsoproximal margin of the proximal phalanx that exit the bone distal to the metacarpophalangeal/metatarsophalangeal joint capsule have substantial morphological similarities, are amenable to minimally invasive repair and carry a good prognosis for return to training and racing. © 2017 EVJ Ltd.

Supersonic Injection of Aerated Liquid Jet

NASA Astrophysics Data System (ADS)

Choudhari, Abhijit; Sallam, Khaled

2016-11-01

A computational study of the exit flow of an aerated two-dimensional jet from an under-expanded supersonic nozzle is presented. The liquid sheet is operating within the annular flow regime and the study is motivated by the application of supersonic nozzles in air-breathing propulsion systems, e.g. scramjet engines, ramjet engines and afterburners. The simulation was conducted using VOF model and SST k- ω turbulence model. The test conditions included: jet exit of 1 mm and mass flow rate of 1.8 kg/s. The results show that air reaches transonic condition at the injector exit due to the Fanno flow effects in the injector passage. The aerated liquid jet is alternately expanded by Prandtl-Meyer expansion fan and compressed by oblique shock waves due to the difference between the back (chamber) pressure and the flow pressure. The process then repeats itself and shock (Mach) diamonds are formed at downstream of injector exit similar to those typical of exhaust plumes of propulsion system. The present results, however, indicate that the flow field of supersonic aerated liquid jet is different from supersonic gas jets due to the effects of water evaporation from the liquid sheet. The contours of the Mach number, static pressure of both cases are compared to the theory of gas dynamics.

Energy Optimization Audit at Humphreys Engineer Center

DTIC Science & Technology

2008-09-01

EPDM (ethylene propylene diene M- class [ rubber ]). Doors There are three pairs of doors to the interior terrace (Figure 6) and exit with a high...System EISA Energy Independence and Security Act EPAct Energy Policy Act EPDM EPDM (ethylene propylene diene M-class [ rubber ]) ERDC Engineer

Investigation of the feasibility of optical diagnostic measurements at the exit of the SSME

NASA Technical Reports Server (NTRS)

Shirley, John A.; Boedeker, Laurence R.

1993-01-01

Under Contract NAS8-36861 sponsored by NASA Marshall Space Flight Center, the United Technologies Research Center is conducting an investigation of the feasibility of remote optical diagnostics to measure temperature, species concentration and velocity at the exit of the Space Shuttle Main Engine (SSME). This is a two phase study consisting of a conceptual design phase followed by a laboratory experimental investigation. The first task of the conceptual design studies is to screen and evaluate the techniques which can be used for the measurements. The second task is to select the most promising technique or techniques, if as expected, more than one type of measurement must be used to measure all the flow variables of interest. The third task is to examine in detail analytically the capabilities and limitations of the selected technique(s). The results of this study are described in the section of this report entitled Conceptual Design Investigations. The conceptual design studies identified spontaneous Raman scattering and photodissociative flow-tagging for measurements respectively of gas temperature and major species concentration and for velocity. These techniques and others that were considered are described in the section describing the conceptual design. The objective of the second phase of investigations was to investigate experimentally the techniques identified in the first phase. The first task of the experimental feasibility study is to design and assemble laboratory scale experimental apparatus to evaluate the best approaches for SSME exit optical diagnostics for temperature, species concentrations and velocity, as selected in the Phase I conceptual design study. The second task is to evaluate performance, investigate limitations, and establish actual diagnostic capabilities, accuracies and precision for the selected optical systems. The third task is to evaluate design requirements and system trade-offs of conceptual instruments. Spontaneous Raman scattering excited by a KrF excimer laser pulse was investigated for SSME exit plane temperature and major species concentration measurements. The relative concentrations of molecular hydrogen and water vapor would be determined by measuring the integrated Q-branch scattering signals through narrow bandpass filters in front of photomultipliers. The temperature would be determined by comparing the signal from a single hydrogen rotational Raman line to the total hydrogen Q-branch signal. The rotational Raman line would be isolated by a monochromator and detected with a PMT.

The effect of nozzle-exit-channel shape on resultant fiber diameter in melt-electrospinning

NASA Astrophysics Data System (ADS)

Esmaeilirad, Ahmad; Ko, Junghyuk; Rukosuyev, Maxym V.; Lee, Jason K.; Lee, Patrick C.; Jun, Martin B. G.

2017-01-01

In recent decades, electrospinning using a molten poly (ε-caprolactone) resin has gained attention for creating fibrous tissue scaffolds. The topography and diameter control of such electrospun microfibers is an important issue for their different applications in tissue engineering. Charge density, initial nozzle-exit-channel cross-sectional area, nozzle to collector distance, viscosity, and processing temperature are the most important input parameters that affect the final electrospun fiber diameters. In this paper we will show that the effect of nozzle-exit-channel shape is as important as the other effective parameters in a resultant fiber diameter. However, to the best of our knowledge, the effect of nozzle-exit-channel shapes on a resultant fiber diameter have not been studied before. Comparing rectangular and circular nozzles with almost the same exit-channel cross-sectional areas in a similar processing condition showed that using a rectangular nozzle resulted in decreasing final fiber diameter up to 50%. Furthermore, the effect of processing temperature on the final fiber topography was investigated.

Aerodynamic Performance of a Compact, High Work-Factor Centrifugal Compressor at the Stage and Subcomponent Level

NASA Technical Reports Server (NTRS)

Braunscheidel, Edward P.; Welch, Gerard E.; Skoch, Gary J.; Medic, Gorazd; Sharma, Om P.

2015-01-01

The measured aerodynamic performance of a compact, high work-factor, single-stage centrifugal compressor, comprising an impeller, diffuser, 90deg-bend, and exit guide vane is reported. Performance levels are based on steady-state total-pressure and total-temperature rake and angularity-probe data acquired at key machine rating planes during recent testing at NASA Glenn Research Center. Aerodynamic performance at the stage level is reported for operation between 70 to 105 percent of design corrected speed, with subcomponent (impeller, diffuser, and exit-guide-vane) flow field measurements presented and discussed at the 100 percent design-speed condition. Individual component losses from measurements are compared with pre-test CFD predictions on a limited basis.

Aerodynamic Performance of a Compact, High Work-Factor Centrifugal Compressor at the Stage and Subcomponent Level

NASA Technical Reports Server (NTRS)

Braunscheidel, Edward P.; Welch, Gerard E.; Skoch, Gary J.; Medic, Gorazd; Sharma, Om P.

2014-01-01

The measured aerodynamic performance of a compact, high work-factor, single-stage centrifugal compressor, comprising an impeller, diffuser, 90º-bend, and exit guide vane is reported. Performance levels are based on steady-state total-pressure and total-temperature rake and angularity-probe data acquired at key machine rating planes during recent testing at NASA Glenn Research Center. Aerodynamic performance at the stage level is reported for operation between 70 to 105% of design corrected speed, with subcomponent (impeller, diffuser, and exit-guide-vane) flow field measurements presented and discussed at the 100% design-speed condition. Individual component losses from measurements are compared with pre-test CFD predictions on a limited basis.

POD- Mapping and analysis of hydroturbine exit flow dynamics

NASA Astrophysics Data System (ADS)

Kjeldsen, Morten; Finstad, Pal Henrik

2012-11-01

Pairwise radial dynamic measurements of the swirling draft tube flow have been made at the 25 MW Svorka power plant in Surnadal operating at 48% load at 6 radial and 7 angular positions. The data is analyzed with traditional methods as well as with POD. The measurements were made in the turbine draft tube/exit flow in an axial measurement plane about 1200mm downstream the turbine runner. The draft tube diameter in the measurement plane is about 1300mm. The flow rate during measurements was close to 5.8m3/s. Two probes were used; both of length Le=700 mm and made of stainless steel with an outer diameter of Do=20 mm and inner diameter Di=4mm. At the end of each probe a full bridge cylindrical KULITE xcl152, 0-3.5, was mounted. 90 seconds samples at 10 kS/s were taken. The POD analysis largely follows that of Tutkun et al. (see e.g. AIAA J., 45,5,2008). The analysis shows that 26% of the pressure pulsation energy can be addressed to azimuthal mode 1. The work has been supported by Energy Norway.

Accounting for the fringe magnetic field from the bending magnet in a Monte Carlo accelerator treatment head simulation.

PubMed

O'Shea, Tuathan P; Foley, Mark J; Faddegon, Bruce A

2011-06-01

Monte Carlo (MC) simulation can be used for accurate electron beam treatment planning and modeling. Measurement of large electron fields, with the applicator removed and secondary collimator wide open, has been shown to provide accurate simulation parameters, including asymmetry in the measured dose, for the full range of clinical field sizes and patient positions. Recently, disassembly of the treatment head of a linear accelerator has been used to refine the simulation of the electron beam, setting tightly measured constraints on source and geometry parameters used in simulation. The simulation did not explicitly include the known deflection of the electron beam by a fringe magnetic field from the bending magnet, which extended into the treatment head. Instead, the secondary scattering foil and monitor chamber were unrealistically laterally offset to account for the beam deflection. This work is focused on accounting for this fringe magnetic field in treatment head simulation. The magnetic field below the exit window of a Siemens Oncor linear accelerator was measured with a Tesla-meter from 0 to 12 cm from the exit window and 1-3 cm off-axis. Treatment head simulation was performed with the EGSnrc/BEAMnrc code, modified to incorporate the effect of the magnetic field on charged particle transport. Simulations were used to analyze the sensitivity of dose profiles to various sources of asymmetry in the treatment head. This included the lateral spot offset and beam angle at the exit window, the fringe magnetic field and independent lateral offsets of the secondary scattering foil and electron monitor chamber. Simulation parameters were selected within the limits imposed by measurement uncertainties. Calculated dose distributions were then compared with those measured in water. The magnetic field was a maximum at the exit window, increasing from 0.006 T at 6 MeV to 0.020 T at 21 MeV and dropping to approximately 5% of the maximum at the secondary scattering foil. It was up to three times higher in the bending plane, away from the electron gun, and symmetric within measurement uncertainty in the transverse plane. Simulations showed the magnetic field resulted in an offset of the electron beam of 0.80 cm (mean) at the machine isocenter for the exit window only configuration. The fringe field resulted in a 3.5%-7.6% symmetry and 0.25-0.35 cm offset of the clinical beam R(max) profiles. With the magnetic field included in simulations, a single (realistic) position of the secondary scattering foil and monitor chamber was selected. Measured and simulated dose profiles showed agreement to an average of 2.5%/0.16 cm (maximum: 3%/0.2 cm), which is a better match than previously achieved without incorporating the magnetic field in the simulation. The undulations from the 3 stepped layers of the secondary scattering foil, evident in the measured profiles of the higher energy beams, are now aligned with those in the simulated beam. The simulated fringe magnetic field had negligible effect on the central axis depth dose curves and cross-plane dose profiles. The fringe magnetic field is a significant contributor to the electron beam in-plane asymmetry. With the magnetic field included explicitly in the simulation, realistic monitor chamber and secondary scattering foil positions have been achieved, and the calculated fluence and dose distributions are more accurate.

Hypersonic research engine/aerothermodynamic integration model, experimental results. Volume 2: Mach 6 performance

NASA Technical Reports Server (NTRS)

Andrews, E. H., Jr.; Mackley, E. A.

1976-01-01

Computer program performance results of a Mach 6 hypersonic research engine during supersonic and subsonic combustion modes were presented. The combustion mode transition was successfully performed, exit surveys made, and effects of altitude, angle of attack, and inlet spike position were determined during these tests.

The 3D modeling of high numerical aperture imaging in thin films

NASA Technical Reports Server (NTRS)

Flagello, D. G.; Milster, Tom

1992-01-01

A modelling technique is described which is used to explore three dimensional (3D) image irradiance distributions formed by high numerical aperture (NA is greater than 0.5) lenses in homogeneous, linear films. This work uses a 3D modelling approach that is based on a plane-wave decomposition in the exit pupil. Each plane wave component is weighted by factors due to polarization, aberration, and input amplitude and phase terms. This is combined with a modified thin-film matrix technique to derive the total field amplitude at each point in a film by a coherent vector sum over all plane waves. Then the total irradiance is calculated. The model is used to show how asymmetries present in the polarized image change with the influence of a thin film through varying degrees of focus.

A Hydrogen Peroxide Hot-Jet Simulator for Wind-Tunnel Tests of Turbojet-Exit Models

NASA Technical Reports Server (NTRS)

Runckel, Jack F.; Swihart, John M.

1959-01-01

A turbojet-engine-exhaust simulator which utilizes a hydrogen peroxide gas generator has been developed for powered-model testing in wind tunnels with air exchange. Catalytic decomposition of concentrated hydrogen peroxide provides a convenient and easily controlled method of providing a hot jet with characteristics that correspond closely to the jet of a gas turbine engine. The problems associated with simulation of jet exhausts in a transonic wind tunnel which led to the selection of a liquid monopropellant are discussed. The operation of the jet simulator consisting of a thrust balance, gas generator, exit nozzle, and auxiliary control system is described. Static-test data obtained with convergent nozzles are presented and shown to be in good agreement with ideal calculated values.

Advanced 35 W Free-Piston Stirling Engine for Space Power Applications

NASA Astrophysics Data System (ADS)

Wood, J. Gary; Lane, Neill

2003-01-01

This paper presents the projected performance and overall design characteristics of a high efficiency, low mass 35 W free-piston Stirling engine design. Overall (engine plus linear alternator) thermodynamic performance greater than 50% of Carnot, with a specific power close to 100 W/kg appears to be a reasonable goal at this small power level. Supporting test data and analysis results from exiting engines are presented. Design implications of high specific power in relatively low power engines is presented and discussed.

Static internal performance of ventral and rear nozzle concepts for short-takeoff and vertical-landing aircraft

NASA Technical Reports Server (NTRS)

Re, Richard J.; Carson, George T., Jr.

1991-01-01

The internal performance of two exhaust system concepts applicable to single-engine short-take-off and vertical-landing tactical fighter configurations was investigated. These concepts involved blocking (or partially blocking) tailpipe flow to the rear (cruise) nozzle and diverting it through an opening to a ventral nozzle exit for vertical thrust. A set of variable angle vanes at the ventral nozzle exit were used to vary ventral nozzle thrust angle between 45 and 110 deg relative to the positive axial force direction. In the vertical flight mode the rear nozzle (or tailpipe flow to it) was completely blocked. In the transition flight mode flow in the tailpipe was split between the rear and ventral nozzles and the flow was vectored at both exits for aircraft control purposes through this flight regime. In the cruise flight mode the ventral nozzle was sealed and all flow exited through the rear nozzle.

Fine particle and organic vapor emissions from staged tests of an in-use aircraft engine

NASA Astrophysics Data System (ADS)

Presto, Albert A.; Nguyen, Ngoc T.; Ranjan, Manish; Reeder, Aaron J.; Lipsky, Eric M.; Hennigan, Christopher J.; Miracolo, Marissa A.; Riemer, Daniel D.; Robinson, Allen L.

2011-07-01

Staged tests were conducted to measure the particle and vapor emissions from a CFM56-2B1 gas-turbine engine mounted on a KC-135T Stratotanker airframe at different engine loads. Exhaust was sampled using a rake inlet installed 1-m downstream of the engine exit plane of a parked and chocked aircraft and a dilution sampler and portable smog chamber were used to investigate the particulate matter (PM) emissions. Total fine PM mass emissions were highest at low (4%) and high (85%) load and lower at intermediate loads (7% and 30%). PM mass emissions at 4% load are dominated by organics, while at 85% load elemental carbon is dominant. Quantifying the primary organic aerosol (POA) emissions is complicated by substantial filter sampling artifacts. Partitioning experiments reveal that the majority of the POA is semivolatile; for example, the POA emission factor changed by a factor of two when the background organic aerosol concentration was increased from 0.7 to 4 μg m -3. Therefore, one cannot define a single non-volatile PM emission factor for aircraft exhaust. The gas- and particle-phase organic emissions were comprehensively characterized by analyzing canister, sorbent and filter samples with gas-chromatography/mass-spectrometry. Vapor-phase organic emissions are highest at 4% load and decrease with increasing load. Low-volatility organics (less volatile than a C 12n-alkane) contributed 10-20% of the total organic emissions. The low-volatility organic emissions contain signatures of unburned fuel and aircraft lubricating oil but are dominated by an unresolved complex mixture (UCM) of presumably branched and cyclic alkanes. Emissions at all loads contain more low-volatility organic vapors than POA; thus secondary organic aerosol formation in the aging plume will likely exceed POA emissions.

An Initial Study of the Fundamentals of Ice Crystal Icing Physics in the NASA Propulsion Systems Laboratory

NASA Technical Reports Server (NTRS)

Struk, Peter; Bartkus, Tadas; Tsao, Jen-Ching; Bencic, Timothy; King, Michael; Ratvasky, Thomas; Van Zante, Judith

2017-01-01

This presentation shows results from an initial study of the fundamental physics of ice-crystal ice accretion using the NASA Propulsion Systems Lab (PSL). Ice accretion due to the ingestion of ice-crystals is being attributed to numerous jet-engine power-loss events. The NASA PSL is an altitude jet-engine test facility which has recently added a capability to inject ice particles into the flow. NASA is evaluating whether this facility, in addition to full-engine and motor-driven-rig tests, can be used for more fundamental ice-accretion studies that simulate the different mixed-phase icing conditions along the core flow passage of a turbo-fan engine compressor. The data from such fundamental accretion tests will be used to help develop and validate models of the accretion process. The present study utilized a NACA0012 airfoil. The mixed-phase conditions were generated by partially freezing the liquid-water droplets ejected from the spray bars. This presentation shows data regarding (1) the freeze out characteristics of the cloud, (2) changes in aerothermal conditions due to the presence of the cloud, and (3) the ice accretion characteristics observed on the airfoil model. The primary variable in this test was the PSL plenum humidity which was systematically varied for two duct-exit-plane velocities (85 and 135 ms) as well as two particle size clouds (15 and 50 m MVDi). The observed clouds ranged from fully glaciated to fully liquid, where the liquid clouds were at least partially supercooled. The air total temperature decreased at the test section when the cloud was activated due to evaporation. The ice accretions observed ranged from sharp arrow-like accretions, characteristic of ice-crystal erosion, to cases with double-horn shapes, characteristic of supercooled water accretions.

An Initial Study of the Fundamentals of Ice Crystal Icing Physics in the NASA Propulsion Systems Laboratory

NASA Technical Reports Server (NTRS)

Struk, Peter M.; Ratvasky, Thomas P.; Bencic, Timothy J.; Van Zante, Judith F.; King, Michael C.; Tsao, Jen-Ching; Bartkus, Tadas P.

2017-01-01

This paper presents results from an initial study of the fundamental physics of ice-crystal ice accretion using the NASA Propulsion Systems Lab (PSL). Ice accretion due to the ingestion of ice-crystals is being attributed to numerous jet-engine power-loss events. The NASA PSL is an altitude jet-engine test facility which has recently added a capability to inject ice particles into the flow. NASA is evaluating whether this facility, in addition to full-engine and motor-driven-rig tests, can be used for more fundamental ice-accretion studies that simulate the different mixed-phase icing conditions along the core flow passage of a turbo-fan engine compressor. The data from such fundamental accretion tests will be used to help develop and validate models of the accretion process. The present study utilized a NACA0012 airfoil. The mixed-phase conditions were generated by partially freezing the liquid-water droplets ejected from the spray bars. This paper presents data regarding (1) the freeze out characteristics of the cloud, (2) changes in aerothermal conditions due to the presence of the cloud, and (3) the ice accretion characteristics observed on the airfoil model. The primary variable in this test was the PSL plenum humidity which was systematically varied for two duct-exit-plane velocities (85 and 135 ms) as well as two particle size clouds (15 and 50 m MVDi). The observed clouds ranged from fully glaciated to fully liquid, where the liquid clouds were at least partially supercooled. The air total temperature decreased at the test section when the cloud was activated due to evaporation. The ice accretions observed ranged from sharp arrow-like accretions, characteristic of ice-crystal erosion, to cases with double-horn shapes, characteristic of supercooled water accretions.

Electromagnetic Effects in the Near Field Plume Exhaust of a Micro-Pulsed Plasma Thruster

DTIC Science & Technology

2002-06-12

plasma focus is developed at a few millimeters from the thruster exit plane at the axis. This plasma focus exists during the entire pulse, but the plasma density in the focus decreases from about 2x10(exp 22)/cu m at the beginning of the pulse down to 0.3x10(exp 22)/cu m at 5 microsec.

Instrumentation Working Group Summary

NASA Technical Reports Server (NTRS)

Zaller, Michelle; Miake-Lye, Richard

1999-01-01

The Instrumentation Working Group compiled a summary of measurement techniques applicable to gas turbine engine aerosol precursors and particulates. An assessment was made of the limits, accuracy, applicability, and technology readiness of the various techniques. Despite advances made in emissions characterization of aircraft engines, uncertainties still exist in the mechanisms by which aerosols and particulates are produced in the near-field engine exhaust. To adequately assess current understanding of the formation of sulfuric acid aerosols in the exhaust plumes of gas turbine engines, measurements are required to determine the degree and importance of sulfur oxidation in the turbine and at the engine exit. Ideally, concentrations of all sulfur species would be acquired, with emphasis on SO2 and SO3. Numerous options exist for extractive and non-extractive measurement of SO2 at the engine exit, most of which are well developed. SO2 measurements should be performed first to place an upper bound on the percentage of SO2 oxidation. If extractive and non-extractive techniques indicate that a large amount of the fuel sulfur is not detected as SO2, then efforts are needed to improve techniques for SO3 measurements. Additional work will be required to account for the fuel sulfur in the engine exhaust. Chemical Ionization Mass Spectrometry (CI-MS) measurements need to be pursued, although a careful assessment needs to be made of the sampling line impact on the extracted sample composition. Efforts should also be placed on implementing non-intrusive techniques and extending their capabilities by maximizing exhaust coverage for line-of-sight measurements, as well as development of 2-D techniques, where feasible. Recommendations were made to continue engine exit and combustor measurements of particulates. Particulate measurements should include particle size distribution, mass fraction, hydration properties, and volatile fraction. However, methods to ensure that unaltered samples are obtained need to be developed. Particulate speciation was also assigned a high priority for quantifying the fractions of carbon soot, PAH, refractory materials, metals, sulfates, and nitrates. High priority was also placed on performing a comparison of particle sizing instruments. Concern was expressed by the workshop attendees who routinely make particulate measurements about the variation in number density measured during in-flight tests by different instruments. In some cases, measurements performed by different groups of researchers during the same flight tests showed an order of magnitude variation. Second priority was assigned to measuring concentrations of odd hydrogen and oxidizing species. Since OH, HO2, H2O2, and O are extremely reactive, non-extractive measurements are recommended. A combination of absorption and fluorescence is anticipated to be effective for OH measurements in the combustor and at the engine exit. Extractive measurements of HO2 have been made in the stratosphere, where the ambient level of OH is relatively low. Use of techniques that convert HO2 to OH for combustor and engine exit measurements needs to be evaluated, since the ratio of HO2/OH may be 1% or less at both the combustor and engine exit. CI-MS might be a viable option for H2O2, subject to sampling line conversion issues. However, H2O2 is a low priority oxidizing species in the combustor and at the engine exit. Two candidates for atomic oxygen measurements are Resonance Enhanced Multi-Photon Ionization (REMPI) and Laser-Induced Fluorescence (LIF). Particulate measurement by simultaneous extractive and non-extractive techniques was given equal priority to the oxidizer measurements. Concern was expressed over the ability of typical ground test sampling lines to deliver an unaltered sample to a remotely located instrument. It was suggested that the sampling probe and line losses be checked out by attempting measurements using an optical or non-extractive technique immediately upstream of the sampling probe. This is a possible application for Laser Induced Incandescence (LII) as a check on the volume fraction of soot. Optical measurements of size distribution are not well developed for ultrafine particles less than about 20 nm in diameter, so a non-extractive technique for particulate size distribution cannot be recommended without further development. Carbon dioxide measurements need to be made to complement other extractive measurement techniques. CO2 measurements enable conversion of other species concentrations to emission indices. Carbon monoxide, which acts as a sink for oxidizing species, should be measured using non-extractive techniques. CO can be rapidly converted to CO2 in extractive probes, and a comparison between extractive and non-extractive measurements should be performed. Development of non-extractive techniques would help to assess the degree of CO conversion, and might be needed to improve the concentration measurement accuracy. Measurements of NO(x) will continue to be critical due to the role of NO and NO2 in atmospheric chemistry, and their influence on atmospheric ozone. Time-resolved measurements of temperature, velocity, and species concentrations were included on the list of desired measurement. Thermocouples are typically adequate for engine exit measurements. PIV and LDV are well established for obtaining velocity profiles. The techniques are listed in the accompanying table; are divided into extractive and non-extractive techniques. Efforts were made to include a measurement uncertainty for each technique. An assessment of the technology readiness was included.

Minimum fuel coplanar aeroassisted orbital transfer using collocation and nonlinear programming

NASA Technical Reports Server (NTRS)

Shi, Yun Yuan; Young, D. H.

1991-01-01

The fuel optimal control problem arising in coplanar orbital transfer employing aeroassisted technology is addressed. The mission involves the transfer from high energy orbit (HEO) to low energy orbit (LEO) without plane change. The basic approach here is to employ a combination of propulsive maneuvers in space and aerodynamic maneuvers in the atmosphere. The basic sequence of events for the coplanar aeroassisted HEO to LEO orbit transfer consists of three phases. In the first phase, the transfer begins with a deorbit impulse at HEO which injects the vehicle into a elliptic transfer orbit with perigee inside the atmosphere. In the second phase, the vehicle is optimally controlled by lift and drag modulation to satisfy heating constraints and to exit the atmosphere with the desired flight path angle and velocity so that the apogee of the exit orbit is the altitude of the desired LEO. Finally, the second impulse is required to circularize the orbit at LEO. The performance index is maximum final mass. Simulation results show that the coplanar aerocapture is quite different from the case where orbital plane changes are made inside the atmosphere. In the latter case, the vehicle has to penetrate deeper into the atmosphere to perform the desired orbital plane change. For the coplanar case, the vehicle needs only to penetrate the atmosphere deep enough to reduce the exit velocity so the vehicle can be captured at the desired LEO. The peak heating rates are lower and the entry corridor is wider. From the thermal protection point of view, the coplanar transfer may be desirable. Parametric studies also show the maximum peak heating rates and the entry corridor width are functions of maximum lift coefficient. The problem is solved using a direct optimization technique which uses piecewise polynomial representation for the states and controls and collocation to represent the differential equations. This converts the optimal control problem into a nonlinear programming problem which is solved numerically by using a modified version of NPSOL. Solutions were obtained for the described problem for cases with and without heating constraints. The method appears to be more robust than other optimization methods. In addition, the method can handle complex dynamical constraints.

Design and Evaluation of Dual-Expander Aerospike Nozzle Upper Stage Engine

DTIC Science & Technology

2014-09-18

Nozzle , taken from Martin [2] . . . . . 19 2.3 Typical Liquid Rocket Engine Cycles from Huzel and Huang[3], credit J. Hall[4] 21 2.4 Liquid Rocket Engine...giving the maximum thrust. For steady, supersonic flow (no separation from the nozzle ) the exit pressure is constant for a given engine plus nozzle ...performance independent of a rocket’s nozzle . Assuming one-dimensional, steady, and isentropic flow of a perfect gas gives the definition for characteristic

How to Enter, Fly In, and Exit the A-Train Constellation

NASA Technical Reports Server (NTRS)

Vincent, Mark A.

2015-01-01

The collaborative science obtained from the satellites in the A-Train is an unparalleled success. The constellation framework that has evolved is well-formulated and documented by its international members. Communication between teams is enhanced by a web-based Constellation Coordination System. Safety and correlated observations are ensured by defining independent control boxes with buffers in between. Each mission stays within its control box by regular drag makeup maneuvers. Annual inclination adjustments are coordinated by all missions to maintain their absolute and relative Mean Local Time of Ascending Node (MLTAN). Since the satellites are in different orbit planes their separation involves a three-dimensional triad made up of the along track separations, reference groundtracks and MLTAN's. For further safety, a Constellation Envelope has been defined to determine safe entry and exit orbits.

Characterization of particulate matter and gaseous emissions of a C-130H aircraft.

PubMed

Corporan, Edwin; Quick, Adam; DeWitt, Matthew J

2008-04-01

The gaseous and nonvolatile particulate matter (PM) emissions of two T56-A-15 turboprop engines of a C-130H aircraft stationed at the 123rd Airlift Wing in the Kentucky Air National Guard were characterized. The emissions campaign supports the Strategic Environmental Research and Development Program (SERDP) project WP-1401 to determine emissions factors from military aircraft. The purpose of the project is to develop a comprehensive emissions measurement program using both conventional and advanced techniques to determine emissions factors of pollutants, and to investigate the spatial and temporal evolutions of the exhaust plumes from fixed and rotating wing military aircraft. Standard practices for the measurement of gaseous emissions from aircraft have been well established; however, there is no certified methodology for the measurement of aircraft PM emissions. In this study, several conventional instruments were used to physically characterize and quantify the PM emissions from the two turboprop engines. Emissions samples were extracted from the engine exit plane and transported to the analytical instrumentation via heated lines. Multiple sampling probes were used to assess the spatial variation and obtain a representative average of the engine emissions. Particle concentrations, size distributions, and mass emissions were measured using commercially available aerosol instruments. Engine smoke numbers were determined using established Society of Automotive Engineers (SAE) practices, and gaseous species were quantified via a Fourier-transform infrared-based gas analyzer. The engines were tested at five power settings, from idle to take-off power, to cover a wide range of operating conditions. Average corrected particle numbers (PNs) of (6.4-14.3) x 10(7) particles per cm3 and PN emission indices (EI) from 3.5 x 10(15) to 10.0 x 10(15) particles per kg-fuel were observed. The highest PN EI were observed for the idle power conditions. The mean particle diameter varied between 50 nm at idle to 70 nm at maximum engine power. PM mass EI ranged from 1.6 to 3.5 g/kg-fuel for the conditions tested, which are in agreement with previous T56 engine measurements using other techniques. Additional PM data, smoke numbers, and gaseous emissions will be presented and discussed.

Validation of Supersonic Film Cooling Modeling for Liquid Rocket Engine Applications

NASA Technical Reports Server (NTRS)

Morris, Christopher I.; Ruf, Joseph H.

2010-01-01

Topics include: upper stage engine key requirements and design drivers; Calspan "stage 1" results, He slot injection into hypersonic flow (air); test articles for shock generator diagram, slot injector details, and instrumentation positions; test conditions; modeling approach; 2-d grid used for film cooling simulations of test article; heat flux profiles from 2-d flat plate simulations (run #4); heat flux profiles from 2-d backward facing step simulations (run #43); isometric sketch of single coolant nozzle, and x-z grid of half-nozzle domain; comparison of 2-d and 3-d simulations of coolant nozzles (run #45); flowfield properties along coolant nozzle centerline (run #45); comparison of 3-d CFD nozzle flow calculations with experimental data; nozzle exit plane reduced to linear profile for use in 2-d film-cooling simulations (run #45); synthetic Schlieren image of coolant injection region (run #45); axial velocity profiles from 2-d film-cooling simulation (run #45); coolant mass fraction profiles from 2-d film-cooling simulation (run #45); heat flux profiles from 2-d film cooling simulations (run #45); heat flux profiles from 2-d film cooling simulations (runs #47, #45, and #47); 3-d grid used for film cooling simulations of test article; heat flux contours from 3-d film-cooling simulation (run #45); and heat flux profiles from 3-d and 2-d film cooling simulations (runs #44, #46, and #47).

Simulated Beam Extraction Performance Characterization of a 50-cm Ion Thruster Discharge

NASA Technical Reports Server (NTRS)

Foster, John E.; Hubble, Aimee; Nowak-Gucker, Sarah; Davis, Chris; Peterson, Peter; Viges, Eric; Chen, Dave

2013-01-01

A 50 cm ion thruster is being developed to operate at >65 percent total efficiency at 11 kW, 2700 s Isp and over 25 kW, 4500 s Isp at a total efficiency of >75 percent. The engine is being developed to address the need for a multimode system that can provide a range of thrust-to- power to service national and commercial near-earth onboard propulsion needs such as station-keeping and orbit transfer. Operating characteristics of the 50 cm ion thruster were measured under simulated beam extraction. The discharge current distribution at the various magnet rings was measured over a range of operating conditions. The relationship between the anode current distribution and the resulting plasma uniformity and ion flux measured at the thruster exit plane is discussed. The thermal envelope will also be investigated through the monitoring of magnet temperatures over the range of discharge powers investigated. Discharge losses as a function of propellant utilization was also characterized at multiple simulated beam currents. Bulk plasma conditions such as electron temperature and electron density near engine centerline was measured over a range of operating conditions using an internal Langmuir probe. Sensitivity of discharge performance to chamber length is also discussed. This data acquired from this discharge study will be used in the refinement of a throttle table in anticipation for eventual beam extraction testing.

HOT CELL BUILDING, TRA632, INTERIOR. CELL 3, "HEAVY" CELL. CAMERA ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

HOT CELL BUILDING, TRA-632, INTERIOR. CELL 3, "HEAVY" CELL. CAMERA FACES WEST TOWARD BUILDING EXIT. OBSERVATION WINDOW AT LEFT EDGE OF VIEW. INL NEGATIVE NO. HD46-28-4. Mike Crane, Photographer, 2/2005 - Idaho National Engineering Laboratory, Test Reactor Area, Materials & Engineering Test Reactors, Scoville, Butte County, ID

Optical Measurements at the Combustor Exit of the HIFiRE 2 Ground Test Engine

NASA Technical Reports Server (NTRS)

Brown, Michael S.; Herring, Gregory C.; Cabell, Karen; Hass, Neal; Barhorst, Todd F.; Gruber, Mark

2012-01-01

The development of optical techniques capable of measuring in-stream flow properties of air breathing hypersonic engines is a goal of the Aerospace Propulsion Division at AFRL. Of particular interest are techniques such as tunable diode laser absorption spectroscopy that can be implemented in both ground and flight test efforts. We recently executed a measurement campaign at the exit of the combustor of the HIFiRE 2 ground test engine during Phase II operation of the engine. Data was collected in anticipation of similar data sets to be collected during the flight experiment. The ground test optical data provides a means to evaluate signal processing algorithms particularly those associated with limited line of sight tomography. Equally important, this in-stream data was collected to compliment data acquired with surface-mounted instrumentation and the accompanying flowpath modeling efforts-both CFD and lower order modeling. Here we discuss the specifics of hardware and data collection along with a coarse-grained look at the acquired data and our approach to processing and analyzing it.

Aerodynamic Measurements of an Incidence Tolerant Blade in a Transonic Turbine Cascade

NASA Technical Reports Server (NTRS)

McVetta, Ashlie B.; Giel, Paul W.

2012-01-01

An overview of the recent facility modifications to NASA s Transonic Turbine Blade Cascade Facility and aerodynamic measurements on the VSPT incidence-tolerant blade are presented. This work supports the development of variable-speed power turbine (VSPT) speed-change technology for the NASA Large Civil Tilt Rotor (LCTR) vehicle. In order to maintain acceptable main rotor propulsive efficiency, the VSPT operates over a nearly 50% speed range from takeoff to altitude cruise. This results in 50 or more variations in VSPT blade incidence angles. The Transonic Turbine Blade Cascade Facility has the ability to operate over a wide range of Reynolds numbers and Mach numbers, but had to be modified in order to accommodate the negative incidence angle variation required by the LCTR VSPT operation. Details of the modifications are described. An incidence-tolerant blade was developed under an RTPAS study contract and tested in the cascade to look at the effects of large incidence angle and Reynolds number variations. Recent test results are presented which include midspan exit total pressure and flow angle measurements obtained at three inlet angles representing the cruise, take-off, and maximum incidence flight mission points. For each inlet angle, data were obtained at five flow conditions with exit Reynolds numbers varying from 2.12 106 to 2.12 105 and two isentropic exit Mach numbers of 0.72 and 0.35. Three-dimensional flowfield measurements were also acquired at the cruise and take-off points. The flowfield measurements were acquired using a five-hole and three-hole pneumatic probe located in a survey plane 8.6% axial chord downstream of the blade trailing edge plane and covering three blade passages. Blade and endwall static pressure distributions were also acquired for each flow condition.

Identification of Noise Sources During Rocket Engine Test Firings and a Rocket Launch Using a Microphone Phased-Array

NASA Technical Reports Server (NTRS)

Panda, Jayanta; Mosher, Robert N.; Porter, Barry J.

2013-01-01

A 70 microphone, 10-foot by 10-foot, microphone phased array was built for use in the harsh environment of rocket launches. The array was setup at NASA Wallops launch pad 0A during a static test firing of Orbital Sciences' Antares engines, and again during the first launch of the Antares vehicle. It was placed 400 feet away from the pad, and was hoisted on a scissor lift 40 feet above ground. The data sets provided unprecedented insight into rocket noise sources. The duct exit was found to be the primary source during the static test firing; the large amount of water injected beneath the nozzle exit and inside the plume duct quenched all other sources. The maps of the noise sources during launch were found to be time-dependent. As the engines came to full power and became louder, the primary source switched from the duct inlet to the duct exit. Further elevation of the vehicle caused spilling of the hot plume, resulting in a distributed noise map covering most of the pad. As the entire plume emerged from the duct, and the ondeck water system came to full power, the plume itself became the loudest noise source. These maps of the noise sources provide vital insight for optimization of sound suppression systems for future Antares launches.

Submarine Construction (Unterseebootsbau)

DTIC Science & Technology

1972-08-01

PIPE FOR THE SNORKEL EXHAUST MAST 11 AIR EXIT (GENERALLY TO MAIN AIR INDUCTION LINE) 12 EXHAUST GAS INLET FROM EXHAUST GAS LINE SIDE VIEW (MAST...Electric Engine 76 Diesel Engines 79 Air Intake and Gas Exhaust Systems for the Diesel Engines 79 Diesel Fuel and Pressurized Water System 82...Lines of a Submarine ■. 31 Figure 6 - Lines of a Submersible 31 Figure 7 - Twin- Screw Stern Configurations 34 Figure 8 - Single- Screw Stern

A study of the round jet/plane wall flow field

NASA Technical Reports Server (NTRS)

Foss, J. F.; Kleis, S. J.

1971-01-01

Impingement angles, between the axisymmetric jet axis and the plane wall, from zero to 15 degrees have been examined for nozzle heights of 0.75, 1.0, 1.5 and 2.0 diameters and for: (1) a fully developed pipe flow, and (2) a relatively uniform exit velocity condition. Velocity measurements have been used to define isotach contours and to determine mass, momentum and energy flux values for the near field (within five diameters) of the jet. Surface pressure measurements have been used to define surface pressure forces and jet centerline trajectories. The geometric and flow conditions examined and the interpretation of the results have been motivated by the externally blown flap STOL aircraft application.

Uprating the Frontal Thrust of a Spherical Gas-Dynamical Resonator-Pulse Amplifier

NASA Astrophysics Data System (ADS)

Bogdanov, V. I.; Khantalin, D. S.

2017-01-01

Calculations were carried out with application of current numerical methods and with the use of scientific-technical developments of a gas-dynamical resonator-thrust amplifier. The possibility of creating an exit device with a resonator for a small-size gas-turbine engine that in flight provides for thrust uprating by no less than 6% via gas mass attachment in the pulsating process is shown. In this way the size-mass characteristics of the exit device are preserved in the process.

Cold-air performance of compressor-drive turbine of Department of Energy upgraded automobile gas turbine engine. 1: Volute-manifold and stator performance

NASA Technical Reports Server (NTRS)

Roelke, R. J.; Haas, J. E.

1981-01-01

The aerodynamic performance of the inlet manifold and stator assembly of the compressor drive turbine was experimentally determined with cold air as the working fluid. The investigation included measurements of mass flow and stator-exit fluid torque as well as radial surveys of total pressure and flow angle at the stator inlet and annulus surveys of total pressure and flow angle at the stator exit. The stator-exit aftermixed flow conditions and overall stator efficiency were obtained and compared with their design values and the experimental results from three other stators. In addition, an analysis was made to determine the constituent aerodynamic losses that made up the stator kinetic energy loss.

Expedition 9 Landing

NASA Image and Video Library

2004-10-24

Expedition 9 Flight Engineer Michael Fincke exits the Russian search and rescue helicopter in Kustanay, Kazakhstan after the 2 hour flight from the landing site, Sunday, October 24, 2004. Photo Credit: (NASA/Bill Ingalls)

Electron energy distribution function in a low-power Hall thruster discharge and near-field plume

NASA Astrophysics Data System (ADS)

Tichý, M.; Pétin, A.; Kudrna, P.; Horký, M.; Mazouffre, S.

2018-06-01

Electron temperature and plasma density, as well as the electron energy distribution function (EEDF), have been obtained inside and outside the dielectric channel of a 200 W permanent magnet Hall thruster. Measurements were carried out by means of a cylindrical Langmuir probe mounted onto a compact fast moving translation stage. The 3D particle-in cell numerical simulations complement experiments. The model accounts for the crossed electric and magnetic field configuration in a weakly collisional regime where only electrons are magnetized. Since only the electron dynamics is of interest in this study, an artificial mass of ions corresponding to mi = 30 000me was used to ensure ions could be assumed at rest. The simulation domain is located at the thruster exit plane and does not include the cathode. The measured EEDF evidences a high-energy electron population that is superimposed onto the low energy bulk population outside the channel. Inside the channel, the EEDF is close to Maxwellian. Both the experimental and numerical EEDF depart from an equilibrium distribution at the channel exit plane, a region of high magnetic field. We therefore conclude that the fast electron group found in the experiment corresponds to the electrons emitted by the external cathode that reach the thruster discharge without experiencing collision events.

Plasma Potential and Langmuir Probe Measurements in the Near-field Plume of the NASA 300M Hall Thruster

NASA Technical Reports Server (NTRS)

Herman, Daniel A; Shastry, Rohit; Huang, Wensheng; Soulas, George C.; KamHawi, Hani

2012-01-01

In order to aid in the design of high-power Hall thrusters and provide experimental validation for existing modeling efforts, plasma potential and Langmuir probe measurements were performed in the near-field plume of the NASA 300M Hall thruster. A probe array consisting of a Faraday probe, Langmuir probe, and emissive probe was used to interrogate the plume from approximately 0.1 - 2.0 DT,m downstream of the thruster exit plane at four operating conditions: 300 V, 400 V, and 500 V at 20 kW as well as 300 V at 10 kW. Results show that the acceleration zone and high-temperature region were contained within 0.3 DT,m from the exit plane at all operating conditions. Isothermal lines were shown to strongly follow magnetic field lines in the nearfield, with maximum temperatures ranging from 19 - 27 eV. The electron temperature spatial distribution created large drops in measured floating potentials in front of the magnetic pole surfaces where the plasma density was small, which suggests strong sheaths at these surfaces. The data taken have provided valuable information for future design and modeling validation, and complements ongoing internal measurement efforts on the NASA 300 M.

How Student Written Communication Skills Benefit during Participation in an Industry-Sponsored Civil Engineering Capstone Course

ERIC Educational Resources Information Center

Fries, Ryan; Cross, Brad; Zhou, Jianpeng; Verbais, Chad

2017-01-01

Because many engineering programs use capstone design courses and value strong communication abilities, authors sought to identify how student written communication skills changed because of industry-sponsored capstone design projects. A student exit survey was collected at the end of the capstone design course during faculty-led projects and…

Dry ultralow NO{sub x} Green Thumb combustor for Allison`s 501-K series industrial engines

DOE Office of Scientific and Technical Information (OSTI.GOV)

Puri, R.; Stansel, D.M.; Smith, D.A.

1997-01-01

This paper describes the progress made in developing an external ultralow oxides of nitrogen (NO{sub x}) Green Thumb combustor for the Allison Engine Company`s 501-K series engines. A lean premixed approach is being pursued to meet the emissions goals of 9 ppm NO{sub x}, 50 ppm carbon monoxide (CO), and 10 ppm unburned hydrocarbon (UHC). Several lean premixed (LPM) module configurations were identified computationally for the best NO{sub x}-CO trade-off by varying the location of fuel injection and the swirl angle of the module. These configurations were fabricated and screened under atmospheric conditions by direct visualization through a quartz liner;more » measurement of the stoichiometry at lean blow out (LBO); measurement of the fuel-air mixing efficiency at the module exit; and emissions measurements at the combustor exit, as well as velocity measurements. The influence of linear residence time on emissions was also examined. An LPM module featuring a radial inflow swirler demonstrated efficient fuel-air mixing and subsequent low NO{sub x} and CO production in extensive atmospheric bench and simulated engine testing. Measurements show the fuel concentration distribution at the module exit impacts the tradeoff between NO{sub x} and CO emissions. The effect of varying the swirl angle of the module also has a similar effect with the gains in NO{sub x} emissions reduction being traded for increased CO emissions. A uniform fuel-air mixture ({+-}2.5% azimuthal variation) at the exit of the module yields low NO{sub x} (5--10 ppm) at inlet conditions of 1 MPa ({approximately}10 atm) and temperatures as high as 616 K (650 F). The close proximity of adjacent modules and lower confinement in the liner most likely reduces the size of the recirculation zone associated with each module, thereby reducing the NO{sub x} formed therein. The CO emissions are probably lowered due to the reduced cool liner surface area per module resulting when several modules feed into the same liner.« less

Effects of inlet distortion on gas turbine combustion chamber exit temperature profiles

NASA Astrophysics Data System (ADS)

Maqsood, Omar Shahzada

Damage to a nozzle guide vane or blade, caused by non-uniform temperature distributions at the combustion chamber exit, is deleterious to turbine performance and can lead to expensive and time consuming overhaul and repair. A test rig was designed and constructed for the Allison 250-C20B combustion chamber to investigate the effects of inlet air distortion on the combustion chamber's exit temperature fields. The rig made use of the engine's diffuser tubes, combustion case, combustion liner, and first stage nozzle guide vane shield. Rig operating conditions simulated engine cruise conditions, matching the quasi-non-dimensional Mach number, equivalence ratio and Sauter mean diameter. The combustion chamber was tested with an even distribution of inlet air and a 4% difference in airflow at either side. An even distribution of inlet air to the combustion chamber did not create a uniform temperature profile and varying the inlet distribution of air exacerbated the profile's non-uniformity. The design of the combustion liner promoted the formation of an oval-shaped toroidal vortex inside the chamber, creating localized hot and cool sections separated by 90° that appeared in the exhaust. Uneven inlet air distributions skewed the oval vortex, increasing the temperature of the hot section nearest the side with the most mass flow rate and decreasing the temperature of the hot section on the opposite side. Keywords: Allison 250, Combustion, Dual-Entry, Exit Temperature Profile, Gas Turbine, Pattern Factor, Reverse Flow.

Convair XF-102 Model in the 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1953-08-21

A .10-scale model of Convair’s XF-102 in the 8- by 6-Foot Supersonic Wind Tunnel at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory for jet exit studies. The XF-102 was a prototype of the F-102 Delta Dagger. The F-102 served as an interceptor against long range bombers from the Soviet Union. The aircraft was powered by a Pratt and Whitney J57 turbojet. The first prototype crashed two weeks after is first flight on October 24, 1953, just months after this photograph. Engineers then incorporated the fixed-wing design to reduce drag at supersonic speeds. The production model F-102 became the first delta-wing supersonic aircraft in operation. The 8- by 6-Foot Supersonic Wind Tunnel is used to study propulsion systems, including inlets and exit nozzles, combustion fuel injectors, flame holders, exit nozzles, and controls on ramjet and turbojet engines. Flexible sidewalls alter the tunnel’s nozzle shape to vary the Mach number during operation. A seven-stage axial compressor, driven by three electric motors that yield a total of 87,000 horsepower, generates air speeds from Mach 0.36 to 2.0.

Spatial Correlation in the Ambient Core Noise Field of a Turbofan Engine

NASA Technical Reports Server (NTRS)

Miles, Jeffrey Hilton

2012-01-01

An acoustic transfer function relating combustion noise and turbine exit noise in the presence of enclosed ambient core noise is investigated using a dynamic system model and an acoustic system model for the particular turbofan engine studied and for a range of operating conditions. Measurements of cross-spectra magnitude and phase between the combustor and turbine exit and auto-spectra at the turbine exit and combustor are used to show the presence of indirect and direct combustion noise over the frequency range of 0 400 Hz. The procedure used evaluates the ratio of direct to indirect combustion noise. The procedure used also evaluates the post-combustion residence time in the combustor which is a factor in the formation of thermal NOx and soot in this region. These measurements are masked by the ambient core noise sound field in this frequency range which is observable since the transducers are situated within an acoustic wavelength of one another. An ambient core noise field model based on one and two dimensional spatial correlation functions is used to replicate the spatially correlated response of the pair of transducers. The spatial correlation function increases measured attenuation due to destructive interference and masks the true attenuation of the turbine.

Study of compressible flow through a rectangular-to-semiannular transition duct

NASA Technical Reports Server (NTRS)

Foster, Jeffry; Okiishi, Theodore H.; Wendt, Bruce J.; Reichert, Bruce A.

1995-01-01

Detailed flow field measurements are presented for compressible flow through a diffusing rectangular-to-semiannular transition duct. Comparisons are made with published computational results for flow through the duct. Three-dimensional velocity vectors and total pressures were measured at the exit plane of the diffuser model. The inlet flow was also measured. These measurements are made using calibrated five-hole probes. Surface oil flow visualization and surface static pressure data were also taken. The study was conducted with an inlet Mach number of 0.786. The diffuser Reynolds based on the inlet centerline velocity and the exit diameter of the diffuser was 3,200,000. Comparison of the measured data with previously published computational results are made. Data demonstrating the ability of vortex generators to reduce flow separation and circumferential distortion is also presented.

Development of base pressure similarity parameters for application to space shuttle launch vehicle power-on aerodynamic testing

NASA Technical Reports Server (NTRS)

Sulyma, P. R.; Penny, M. M.

1978-01-01

A base pressure data correlation study was conducted to define exhaust plume similarity parameters for use in Space Shuttle power-on launch vehicle aerodynamic test programs. Data correlations were performed for single bodies having, respectively, single and triple nozzle configurations and for a triple body configuration with single nozzles on each of the outside bodies. Base pressure similarity parameters were found to differ for the single nozzle and triple nozzle configurations. However, the correlation parameter for each was found to be a strong function of the nozzle exit momentum. Results of the data base evaluation are presented indicating an assessment of all data points. Analytical/experimental data comparisons were made for nozzle calibrations and correction factors derived, where indicated for use in nozzle exit plane data calculations.

High-temperature, high-pressure oxygen metering valve

NASA Technical Reports Server (NTRS)

Christianson, Rollin C. (Inventor); Lycou, Peter P. (Inventor); Daniel, James A. (Inventor)

1993-01-01

A control valve includes a body defining a central cavity arranged between a fluid inlet and outwardly-diverging first and second fluid outlets respectively disposed in a common transverse plane. A valve member is arranged in the cavity for rotation between first and second operating positions where a transverse fluid passage through the valve member alternatively communicates the fluid inlet with one or the other of the fluid outlets. To minimize fluid turbulence when the valve member is rotated to an alternate operating position, the fluid passage has a convergent entrance for maintaining the passage in permanent communication with the fluid inlet as well as an oblong exit opening with spaced side walls for enabling the exit opening to temporarily span the first and second fluid outlets as the valve member is turned between its respective operating positions.

Views of a Cal/OSHA Inspector.

PubMed

Oudiz, Jack

2009-01-01

Retiring CAL/OSHA Industrial Hygienist and Senior Safety Engineer Jack Oudiz offers his thoughts in the nature of a voluntary "exit interview" on his years working for the agency and its performance in its mission.

Effect of Exit-Slot Position and Opening on the Available Cooling Pressure for NACA Nose-Slot Cowlings

NASA Technical Reports Server (NTRS)

Stickle, George W; Naiman, Irven; Crigler, John L

1940-01-01

Report presents the results of an investigation of full-scale nose-slot cowlings conducted in the NACA 20-foot wind tunnel to furnish information on the pressure drop available for cooling. Engine conductances from 0 to 0.12 and exit-slot conductances from 0 to 0.30 were covered. Two basic nose shapes were tested to determine the effect of the radius of curvature of the nose contour; the nose shape with the smaller radius of curvature gave the higher pressure drop across the engine. The best axial location of the slot for low-speed operation was found to be in the region of maximum negative pressure for the basic shape for the particular operating condition. The effect of the pressure operating condition on the available cooling pressure is shown.

Faraday Probe Analysis, Part 2: Evaluation of Facility Effects on Ion Migration in a Hall Thruster Plume (Preprint)

DTIC Science & Technology

2010-02-24

A nested Faraday probe was designed and fabricated to assess facility effects in a systematic study of ion migration in a Hall thruster plume...Current density distributions were studied at 8, 12, 16, and 20 thruster diameters downstream of the Hall thruster exit plane with four probe configurations...measurements are a significant improvement for comparisons with numerical simulations and investigations of Hall thruster performance.

Quantum-shutter approach to tunneling time scales with wave packets

DOE Office of Scientific and Technical Information (OSTI.GOV)

Yamada, Norifumi; Garcia-Calderon, Gaston; Villavicencio, Jorge

2005-07-15

The quantum-shutter approach to tunneling time scales [G. Garcia-Calderon and A. Rubio, Phys. Rev. A 55, 3361 (1997)], which uses a cutoff plane wave as the initial condition, is extended to consider certain type of wave packet initial conditions. An analytical expression for the time-evolved wave function is derived. The time-domain resonance, the peaked structure of the probability density (as the function of time) at the exit of the barrier, originally found with the cutoff plane wave initial condition, is studied with the wave packet initial conditions. It is found that the time-domain resonance is not very sensitive to themore » width of the packet when the transmission process occurs in the tunneling regime.« less

Computation of inlet reference plane flow-field for a subscale free-jet forebody/inlet model and comparison to experimental data

NASA Astrophysics Data System (ADS)

McClure, M. D.; Sirbaugh, J. R.

1991-02-01

The computational fluid dynamics (CFD) computer code PARC3D was used to predict the inlet reference plane (IRP) flow field for a side-mounted inlet and forebody simulator in a free jet for five different flow conditions. The calculations were performed for free-jet conditions, mass flow rates, and inlet configurations that matched the free-jet test conditions. In addition, viscous terms were included in the main flow so that the viscous free-jet shear layers emanating from the free-jet nozzle exit were modeled. A measure of the predicted accuracy was determined as a function of free-stream Mach number, angle-of-attack, and sideslip angle.

Deepwater Ports Approach/Exit Hazard and Risk Assessment.

DTIC Science & Technology

1979-02-01

study results reflect a less structured view of the real world, as opposed to the neat technical framework of analysts and engineers . There may...offshore at which the 1. U.S. Army Corps of Engineers , Waterborne Commerce of the United States, (Washington , D.C., 1969-1976), Volume 2. 2. This...multivessel casual ties. The number of tanker port calls was extracted from U.S. V Waterborne Commerce data published by the U.S. Army Corps of Engineers

A Brighter Choice for Safety

NASA Technical Reports Server (NTRS)

2003-01-01

Emergency exit signs can be lifesavers, but only if they remain visible when people need them. All too often, power losses or poor visibility can render the signs ineffective. Luna Technologies International, Inc., of Kent, Washington, is shining new light on this safety issue. The company s LUNAplast(trademark) product line illuminates without the need for electricity, maintenance, or a power connection. LUNAplast, which benefited from tests conducted at Johnson Space Center, is so successful that NASA engineers selected it for the emergency exit pathway indicators on the International Space Station (ISS).

Computer modeling of fan-exit-splitter spacing effects on F100 response to distortion

NASA Technical Reports Server (NTRS)

Shaw, M.; Murdoch, R. W.

1982-01-01

The distortion response of the F100(3) engine was effected by the fan exit splitter configuration. The sensitivity for a proximate splitter fan is calculated to be slightly greater than a remote splitter configuration with identical airfoils. Predicted response was based upon a multiple segment parallel compressor Model modified to include a bypass ratio representation that effects the performance characteristics of the last rotor and intermediate case struts. The predicted distortion response required an accurate definition of row pre- and post-stall undistorted operation.

Recent radial turbine research at the NASA Lewis Research Center

NASA Technical Reports Server (NTRS)

Rohlik, H. E.; Kofskey, M. G.

1971-01-01

The high efficiencies of small radial turbines led to their application in space power systems and numerous APU and shaft power engines. Experimental and analytical work associated with these systems included examination of blade-shroud clearance, blade loading, and exit diffuser design. Results indicate high efficiency over a wide range of specific speed and also insensitivity to clearance and blade loading in the radial part of the rotor. The exit diffuser investigation indicated that a conventional conical outer wall may not provide the velocity variation consistent with minimum overall diffuser loss.

33 CFR 145.10 - Locations and number of fire extinguishers required.

Code of Federal Regulations, 2011 CFR

2011-07-01

... vicinity of exits, either inside or outside of spaces. machinery spaces Gas-fired boilers B-II (CO2 or dry chemical) 2 required. Gas-fired boilers B-V 1 required. 1 Oil-fired boilers B-II 2 required. Oil-fired boilers B-V 2 required. 1 Internal combustion or gas turbine engines B-II 1 for each engine. 2 Electric...

33 CFR 145.10 - Locations and number of fire extinguishers required.

Code of Federal Regulations, 2012 CFR

2012-07-01

... vicinity of exits, either inside or outside of spaces. machinery spaces Gas-fired boilers B-II (CO2 or dry chemical) 2 required. Gas-fired boilers B-V 1 required. 1 Oil-fired boilers B-II 2 required. Oil-fired boilers B-V 2 required. 1 Internal combustion or gas turbine engines B-II 1 for each engine. 2 Electric...

33 CFR 145.10 - Locations and number of fire extinguishers required.

Code of Federal Regulations, 2013 CFR

2013-07-01

... vicinity of exits, either inside or outside of spaces. machinery spaces Gas-fired boilers B-II (CO2 or dry chemical) 2 required. Gas-fired boilers B-V 1 required. 1 Oil-fired boilers B-II 2 required. Oil-fired boilers B-V 2 required. 1 Internal combustion or gas turbine engines B-II 1 for each engine. 2 Electric...

33 CFR 145.10 - Locations and number of fire extinguishers required.

Code of Federal Regulations, 2014 CFR

2014-07-01

... vicinity of exits, either inside or outside of spaces. machinery spaces Gas-fired boilers B-II (CO2 or dry chemical) 2 required. Gas-fired boilers B-V 1 required. 1 Oil-fired boilers B-II 2 required. Oil-fired boilers B-V 2 required. 1 Internal combustion or gas turbine engines B-II 1 for each engine. 2 Electric...

33 CFR 145.10 - Locations and number of fire extinguishers required.

Code of Federal Regulations, 2010 CFR

2010-07-01

... vicinity of exits, either inside or outside of spaces. machinery spaces Gas-fired boilers B-II (CO2 or dry chemical) 2 required. Gas-fired boilers B-V 1 required. 1 Oil-fired boilers B-II 2 required. Oil-fired boilers B-V 2 required. 1 Internal combustion or gas turbine engines B-II 1 for each engine. 2 Electric...

Thrust Augmentation Measurements for a Pulse Detonation Engine Driven Ejector

NASA Technical Reports Server (NTRS)

Pal, S.; Santoro, Robert J.; Shehadeh, R.; Saretto, S.; Lee, S.-Y.

2005-01-01

Thrust augmentation results of an ongoing study of pulse detonation engine driven ejectors are presented and discussed. The experiments were conducted using a pulse detonation engine (PDE) setup with various ejector configurations. The PDE used in these experiments utilizes ethylene (C2H4) as the fuel, and an equi-molar mixture of oxygen and nitrogen as the oxidizer at an equivalence ratio of one. High fidelity thrust measurements were made using an integrated spring damper system. The baseline thrust of the PDE engine was first measured and agrees with experimental and modeling results found in the literature. Thrust augmentation measurements were then made for constant diameter ejectors. The parameter space for the study included ejector length, PDE tube exit to ejector tube inlet overlap distance, and straight versus rounded ejector inlets. The relationship between the thrust augmentation results and various physical phenomena is described. To further understand the flow dynamics, shadow graph images of the exiting shock wave front from the PDE were also made. For the studied parameter space, the results showed a maximum augmentation of 40%. Further increase in augmentation is possible if the geometry of the ejector is tailored, a topic currently studied by numerous groups in the field.

Aerosols and Particulates Workshop Sampling Procedures and Venues Working Group Summary

NASA Technical Reports Server (NTRS)

Pachlhofer, Peter; Howard, Robert

1999-01-01

The Sampling Procedures and Venues Workgroup discussed the potential venues available and issues associated with obtaining measurements. Some of the issues included Incoming Air Quality, Sampling Locations, Probes and Sample Systems. The following is a summary of the discussion of the issues and venues. The influence of inlet air to the measurement of exhaust species, especially trace chemical species, must be considered. Analysis procedures for current engine exhaust emissions regulatory measurements require adjustments for air inlet humidity. As a matter of course in scientific investigations, it is recommended that "background" measurements for any species, particulate or chemical, be performed during inlet air flow before initiation of combustion, if possible, and during the engine test period as feasible and practical. For current regulatory measurements, this would be equivalent to setting the "zero" level for conventional gas analyzers. As a minimum, it is recommended that measurements of the humidity and particulates in the incoming air be taken at the start and end of each test run. Additional measurement points taken during the run are desirable if they can be practically obtained. It was felt that the presence of trace gases in the incoming air is not a significant problem. However, investigators should consider the ambient levels and influences of local air pollution for species of interest. Desired measurement locations depend upon the investigation requirements. A complete investigation of phenomenology of particulate formation and growth requires measurements at a number of locations both within the engine and in the exhaust field downstream of the nozzle exit plane. Desirable locations for both extractive and in situ measurements include: (1) Combustion Zone (Multiple axial locations); (2) Combustor Exit (Multiple radial locations for annular combustors); (3) Turbine Stage (Inlet and exit of the stage); (4) Exit Nozzle (Multiple axial locations downstream of the nozzle). Actual locations with potential for extractive or non-intrusive measurements depend upon the test article and test configuration. Committee members expressed the importance of making investigators aware of various ports that could allow access to various stages of the existing engines. Port locations are engine si)ecific and might allow extractive sampling or innovative hybrid optical-probe access. The turbine stage region was one the most desirable locations for obtaining samples and might be accessed through boroscope ports available in some engine designs. Discussions of probes and sampling systems quickly identified issues dependent on particular measurement quantities. With general consensus, the group recommends SAE procedures for measurements and data analyses of currently regulated exhaust species (CO2, CO, THC, NO(x),) using conventional gas sampling techniques. Special procedures following sound scientific practices must be developed as required for species and/or measurement conditions not covered by SAE standards. Several issues arose concerning short lived radicals and highly reactive species. For conventional sampling, there are concerns of perturbing the sample during extraction, line losses, line-wall reactions, and chemical reactions during the sample transport to the analyzers. Sample lines coated with quartz.or other materials should be investigated for minimization of such effects. The group advocates the development of innovative probe techniques and non-intrusive optical techniques for measurement of short lived radicals and highly reactive species that cannot be sampled accurately otherwise. Two innovative probe concepts were discussed. One concept uses specially designed probes to transfer optical beams to and from a region of flow inaccessible by traditional ports or windows. The probe can perturb the flow field but must have a negligible impact on the region to be optically sampled. Such probes are referred to as hybrid probes and are under development at AEDC for measurement in the high pressure, high temperature of a combustor under development for power generation. The other concept consists of coupling an instrument directly to the probe. The probe would isolate a representative sample stream, freeze chemical reactions and direct the sample into the analyzer portion of the probe. Thus, the measurement would be performed in situ without sample line losses due either to reactions or binding at the wall surfaces. This concept was used to develop a fast, in situ, time-of-flight mass spectrometer measurement system for temporal quantification of NO in the IMPULSE facility at AEDC. Additional work is required in this area to determine the best probe and sampling technique for each species measurement requirement identified by the Trace Chemistry Working Group. A partial list of Venues was used as a baseline for discussion. Additional venues were added to the list and the list was broken out into the following categories: (1)Engines (a) Sea Level Test Stands (b) Altitude Chambers; (2) Annular Combustor Test Stands, (3) Sector Flametube Test Stands, (4) Fundamentals Rigs/Experiments.

Electro-optical detector for use in a wide mass range mass spectrometer

NASA Technical Reports Server (NTRS)

Giffin, Charles E. (Inventor)

1976-01-01

An electro-optical detector is disclosed for use in a wide mass range mass spectrometer (MS), in the latter the focal plane is at or very near the exit end of the magnetic analyzer, so that a strong magnetic field of the order of 1000G or more is present at the focal plane location. The novel detector includes a microchannel electron multiplier array (MCA) which is positioned at the focal plane to convert ion beams which are focused by the MS at the focal plane into corresponding electron beams which are then accelerated to form visual images on a conductive phosphored surface. These visual images are then converted into images on the target of a vidicon camera or the like for electronic processing. Due to the strong magnetic field at the focal plane, in one embodiment of the invention, the MCA with front and back parallel ends is placed so that its front end forms an angle of not less than several degrees, preferably on the order of 10.degree.-20.degree., with respect to the focal plane, with the center line of the front end preferably located in the focal plane. In another embodiment the MCA is wedge-shaped, with its back end at an angle of about 10.degree.-20.degree. with respect to the front end. In this embodiment the MCA is placed so that its front end is located at the focal plane.

Saturn Apollo Program

NASA Image and Video Library

1968-01-01

A complete F-1 engine assembly is shown in this photograph. Designed and developed by Rocketdye under the direction of the Marshall Space Flight Center, the engine measured 19-feet tall by 12.5 feet at the nozzle exit, and each engine produced a 1,500,000-pound thrust using liquid oxygen and kerosene as the propellant. A cluster of five F-1 engines was mounted on the Saturn V S-IC (first) stage and burned 15 tons of liquid oxygen and kerosene each second to produce 7,500,000 pounds of thrust.

Phase 2: HGM air flow tests in support of HEX vane investigation

NASA Technical Reports Server (NTRS)

Cox, G. B., Jr.; Steele, L. L.; Eisenhart, D. W.

1993-01-01

Following the start of SSME certification testing for the Pratt and Whitney Alternate Turbopump Development (ATD) High Pressure Oxidizer Turbopump (HPOTP), cracking of the leading edge of the inner HEX vane was experienced. The HEX vane, at the inlet of the oxidizer bowl in the Hot Gas Manifold (HGM), accepts the HPOTP turbine discharge flow and turns it toward the Gaseous Oxidizer Heat Exchanger (GOX HEX) coil. The cracking consistently initiated over a specific circumferential region of the hex vane, with other circumferential locations appearing with increased run time. Since cracking had not to date been seen with the baseline HPOTP, a fluid-structural interaction involving the ATD HPOTP turbine exit flowfield and the HEX inner vane was suspected. As part of NASA contract NAS8-36801, Pratt and Whitney conducted air flow tests of the ATD HPOTP turbine turnaround duct flowpath in the MSFC Phase 2 HGM air flow model. These tests included HEX vane strain gages and additional fluctuating pressure gages in the turnaround duct and HEX vane flowpath area. Three-dimensional flow probe measurements at two stations downstream of the turbine simulator exit plane were also made. Modifications to the HPOTP turbine simulator investigated the effects on turbine exit flow profile and velocity components, with the objective of reproducing flow conditions calculated for the actual ATD HPOTP hardware. Testing was done at the MSFC SSME Dynamic Fluid Air Flow (Dual-Leg) Facility, at air supply pressures between 50 and 250 psia. Combinations of turbine exit Mach number and pressure level were run to investigate the effect of flow regime. Information presented includes: (1) Descriptions of turbine simulator modifications to produce the desired flow environment; (2) Types and locations for instrumentation added to the flow model for improved diagnostic capability; (3) Evaluation of the effect of changes to the turbine simulator flowpath on the turbine exit flow environment; and (4) Comparison of the experimental turbine exit flow environment to the environment calculated for the ATD HPOTP.

The Role of Astro-Geodetic in Precise Guidance of Long Tunnels

NASA Astrophysics Data System (ADS)

Mirghasempour, M.; Jafari, A. Y.

2015-12-01

One of prime aspects of surveying projects is guidance of paths of a long tunnel from different directions and finally ending all paths in a specific place. This kind of underground surveying, because of particular condition, has some different points in relation to the ground surveying, including Improper geometry in underground transverse, low precise measurement in direction and length due to condition such as refraction, distinct gravity between underground point and corresponding point on the ground (both value and direction of gravity) and etc. To solve this problems, astro-geodetic that is part of geodesy science, can help surveying engineers. In this article, the role of astronomy is defined in two subjects: 1- Azimuth determination of directions from entrance and exit nets of tunnel and also calibration of gyro-theodolite to use them in Underground transvers: By astronomical methods, azimuth of directions can be determine with an accuracy of 0.5 arcsecond, whereas, nowadays, no gyroscope can measure the azimuth in this accuracy; For instance, accuracy of the most precise gyroscope (Gyromat 5000) is 1.2 cm over a distance of one kilometre (2.4 arcsecond). Furthermore, the calibration methods that will be mention in this article, have significance effects on underground transverse. 2- Height relation between entrance point and exit point is problematic and time consuming; For example, in a 3 km long tunnel ( in Arak- Khoram Abad freeway), to relate entrance point to exit point, it is necessary to perform levelling about 90 km. Other example of this boring and time consuming levelling is in Kerman tunnel. This tunnel is 36 km length, but to transfer the entrance point height to exit point, 150 km levelling is needed. According to this paper, The solution for this difficulty is application of astro-geodetic and determination of vertical deflection by digital zenith camera system TZK2-D. These two elements make possible to define geoid profile in terms of tunnel azimuth in entrance and exit of tunnel; So by doing this, surveying engineers are able to transfer entrance point height to exit point of tunnels in easiest way.

A novel method for patient exit and entrance dose prediction based on water equivalent path length measured with an amorphous silicon electronic portal imaging device.

PubMed

Kavuma, Awusi; Glegg, Martin; Metwaly, Mohamed; Currie, Garry; Elliott, Alex

2010-01-21

In vivo dosimetry is one of the quality assurance tools used in radiotherapy to monitor the dose delivered to the patient. Electronic portal imaging device (EPID) images for a set of solid water phantoms of varying thicknesses were acquired and the data fitted onto a quadratic equation, which relates the reduction in photon beam intensity to the attenuation coefficient and material thickness at a reference condition. The quadratic model is used to convert the measured grey scale value into water equivalent path length (EPL) at each pixel for any material imaged by the detector. For any other non-reference conditions, scatter, field size and MU variation effects on the image were corrected by relative measurements using an ionization chamber and an EPID. The 2D EPL is linked to the percentage exit dose table, for different thicknesses and field sizes, thereby converting the plane pixel values at each point into a 2D dose map. The off-axis ratio is corrected using envelope and boundary profiles generated from the treatment planning system (TPS). The method requires field size, monitor unit and source-to-surface distance (SSD) as clinical input parameters to predict the exit dose, which is then used to determine the entrance dose. The measured pixel dose maps were compared with calculated doses from TPS for both entrance and exit depth of phantom. The gamma index at 3% dose difference (DD) and 3 mm distance to agreement (DTA) resulted in an average of 97% passing for the square fields of 5, 10, 15 and 20 cm. The exit dose EPID dose distributions predicted by the algorithm were in better agreement with TPS-calculated doses than phantom entrance dose distributions.

A novel method for patient exit and entrance dose prediction based on water equivalent path length measured with an amorphous silicon electronic portal imaging device

NASA Astrophysics Data System (ADS)

Kavuma, Awusi; Glegg, Martin; Metwaly, Mohamed; Currie, Garry; Elliott, Alex

2010-01-01

In vivo dosimetry is one of the quality assurance tools used in radiotherapy to monitor the dose delivered to the patient. Electronic portal imaging device (EPID) images for a set of solid water phantoms of varying thicknesses were acquired and the data fitted onto a quadratic equation, which relates the reduction in photon beam intensity to the attenuation coefficient and material thickness at a reference condition. The quadratic model is used to convert the measured grey scale value into water equivalent path length (EPL) at each pixel for any material imaged by the detector. For any other non-reference conditions, scatter, field size and MU variation effects on the image were corrected by relative measurements using an ionization chamber and an EPID. The 2D EPL is linked to the percentage exit dose table, for different thicknesses and field sizes, thereby converting the plane pixel values at each point into a 2D dose map. The off-axis ratio is corrected using envelope and boundary profiles generated from the treatment planning system (TPS). The method requires field size, monitor unit and source-to-surface distance (SSD) as clinical input parameters to predict the exit dose, which is then used to determine the entrance dose. The measured pixel dose maps were compared with calculated doses from TPS for both entrance and exit depth of phantom. The gamma index at 3% dose difference (DD) and 3 mm distance to agreement (DTA) resulted in an average of 97% passing for the square fields of 5, 10, 15 and 20 cm. The exit dose EPID dose distributions predicted by the algorithm were in better agreement with TPS-calculated doses than phantom entrance dose distributions.

Langmuir Probe Measurements Within the Discharge Channel of the 20-kW NASA-300M and NASA-300MS Hall Thrusters

NASA Technical Reports Server (NTRS)

Shastry, Rohit; Huang, Wensheng; Haag, Thomas W.; Kamhawi, Hani

2013-01-01

NASA is presently developing a high-power, high-efficiency, long-lifetime Hall thruster for the Solar Electric Propulsion Technology Demonstration Mission. In support of this task, studies have been performed on the 20-kW NASA-300M Hall thruster to aid in the overall design process. The ability to incorporate magnetic shielding into a high-power Hall thruster was also investigated with the NASA- 300MS, a modified version of the NASA-300M. The inclusion of magnetic shielding would allow the thruster to push existing state-of-the-art technology in regards to service lifetime, one of the goals of the Technology Demonstration Mission. Langmuir probe measurements were taken within the discharge channels of both thrusters in order to characterize differences at higher power levels, as well as validate ongoing modeling efforts using the axisymmetric code Hall2De. Flush-mounted Langmuir probes were also used within the channel of the NASA-300MS to verify that magnetic shielding was successfully applied. Measurements taken from 300 V, 10 kW to 600 V, 20 kW have shown plasma potentials near anode potential and electron temperatures of 4 to 12 eV at the walls near the thruster exit plane of the NASA-300MS, verifying magnetic shielding and validating the design process at this power level. Channel centerline measurements on the NASA-300M from 300 V, 10 kW to 500 V, 20 kW show the electron temperature peak at approximately 0.1 to 0.2 channel lengths upstream of the exit plane, with magnitudes increasing with discharge voltage. The acceleration profiles appear to be centered about the exit plane with a width of approximately 0.3 to 0.4 channel lengths. Channel centerline measurements on the NASA-300MS were found to be more challenging due to additional probe heating. Ionization and acceleration zones appeared to move downstream on the NASA-300MS compared to the NASA-300M, as expected based on the shift in peak radial magnetic field. Additional measurements or alternative diagnostics will be needed to verify peak electron temperatures in the NASA-300MS and compare them with model predictions.

Near Discharge Cathode Assembly Plasma Potential Measurements in a 30-cm NSTAR Type Ion Engine During Beam Extraction

NASA Technical Reports Server (NTRS)

Herman, Daniel A.; Gallimore, Alec D.

2006-01-01

Floating emissive probe plasma potential data are presented over a two-dimensional array of locations in the near Discharge Cathode Assembly (DCA) region of a 30-cm diameter ring-cusp ion thruster. Discharge plasma data are presented with beam extraction at throttling conditions comparable to the NASA TH Levels 8, 12, and 15. The operating conditions of the Extended Life Test (ELT) of the Deep Space One (DS1) flight spare ion engine, where anomalous discharge keeper erosion occurred, were TH 8 and TH 12 consequently they are of specific interest in investigating discharge keeper erosion phenomena. The data do not validate the presence of a potential hill plasma structure downstream of the DCA, which has been proposed as a possible erosion mechanism. The data are comparable in magnitude to data taken by other researchers in ring-cusp electron-bombardment ion thrusters. The plasma potential structures are insensitive to thruster throttling level with a minimum as low as 14 V measured at the DCA exit plane and increasing gradually in the axial direction. A sharp increase in plasma potential to the bulk discharge value of 26 to 28 volts, roughly 10 mm radially from DCA centerline, was observed. Plasma potential measurements indicate a low-potential plume structure that is roughly 20 mm in diameter emanating from the discharge cathode that may be attributed to a free-standing plasma double layer.

An Analytical Study of Icing Similitude for Aircraft Engine Testing. Revision

DTIC Science & Technology

1987-02-01

MODELING GEOMETRIES Component Cowl Spinner Fan Blade Fan Stator Exit Vane Probe Approximating Geometry NACA 0012 Airfoil Sphere NACA 0012...DOT/FAA/CT·86/35 AEDC·TR·86·26 An Analytical Study of Icing Similitude for Aircraft Engine Testing c. Scott Bartlett Sverdrup Technology, Inc...8217~,feCa.ORI A n AnalYtical Study )f Icin~ Similitude for Aircraft Engine Tes t tu~ 12. PERSONAL AUTHOR/S) B a r t l e t t , C. Scot t , Sverdrup

Engine environmental effects on composite behavior

NASA Technical Reports Server (NTRS)

Chamis, C. C.; Smith, G. T.

1980-01-01

A series of programs were conducted to investigate and develop the application of composite materials to turbojet engines. A significant part of that effort was directed to establishing the impact resistance and defect growth chracteristics of composite materials over the wide range of environmental conditions found in commercial turbojet engine operations. Both analytical and empirical efforts were involved. The experimental programs and the analytical methodology development as well as an evaluation program for the use of composite materials as fan exit guide vanes are summarized.

Tuned intake air system for a rotary engine

DOE Office of Scientific and Technical Information (OSTI.GOV)

Corbett, W.D.

This patent describes a rotary internal combustion engine for an outboard board motor. It comprises a plenum chamber attached to the rear of the engine; and the plenum chamber including an inner wall attached to the exhaust manifold; an inlet conduit connecting the cooling air exit passage and the inlet air opening; an outlet conduit connecting the outlet air opening and the combustion air inlet; and the outlet conduit terminating in a combustion air outlet in the inner wall of the plenum chamber.

Pioneer 10 and 11 radio occultations by Jupiter. [atmospheric temperature structure

NASA Technical Reports Server (NTRS)

Kliore, A. J.; Woiceshyn, P. M.; Hubbard, W. B.

1977-01-01

Results on the temperature structure of the Jovian atmosphere are reviewed which were obtained by applying an integral inversion technique combined with a model for the planet's shape based on gravity data to Pioneer 10 and 11 radio-occultation data. The technique applied to obtain temperature profiles from the Pioneer data consisted of defining a center of refraction based on a computation of the radius of curvature in the plane of refraction and the normal direction to the equipotential surface at the closest approach point of a ray. Observations performed during the Pioneer 10 entry and exit and the Pioneer 11 exit are analyzed, sources of uncertainty are identified, and representative pressure-temperature profiles are presented which clearly show a temperature inversion between 10 and 100 mb. Effects of zonal winds on the reliability of radio-occultation temperature profiles are briefly discussed.

Investigation of an Anomaly Observed in Impedance Eduction Techniques

NASA Technical Reports Server (NTRS)

Watson, W. R.; Jones, M. G.; Parrott, T. L.

2008-01-01

An intensive investigation into the cause of anomalous behavior commonly observed in impedance eduction techniques is performed. The investigation consists of grid refinement studies, detailed evaluation of results at and near anti-resonance frequencies, comparisons of different model results with synthesized and measured data, assessment or optimization techniques, and evaluation or boundary condition effects. Results show that the root cause of the anomalous behavior is the sensitivity of the educed impedance to small errors in the measured termination resistance at frequencies near anti-resonance or cut-on of a higher-order mode. Evidence is presented to show that the common usage of an anechoic, plane wave termination boundary condition in ducts where the "true" termination is reflective may act as a trigger for these anomalies. Replacing the exit impedance boundary condition by an exit pressure condition is shown to reduce the anomalous results.

Computation of Three-Dimensional Compressible Flow From a Rectangular Nozzle with Delta Tabs

NASA Technical Reports Server (NTRS)

Reddy, D. R.; Steffen, C. J., Jr.; Zaman, K. B. M. Q.

1999-01-01

A three-dimensional viscous flow analysis is performed using a time-marching Reynolds-averaged Navier-Stokes code for a 3:1 rectangular nozzle with two delta tabs located at the nozz1e exit plane to enhance mixing. Two flow configurations, a subsonic jet case and a supersonic jet case using the same rate configuration, which were previously studied experimentally, are computed and compared with the experimental data. The experimental data include streamwise velocity and vorticity distributions for the subsonic case, and Mach number distributions for the supersonic case, at various axial locations downstream of the nozzle exit. The computational results show very good agreement with the experimental data. In addition, the effect of compressibility on vorticity dynamics is examined by comparing the vorticity contours of the subsonic jet case with those of the supersonic jet case which were not measured in the experiment.

Turbojet-exhaust-nozzle secondary-airflow pumping as an exit control of an inlet-stability bypass system for a Mach 2.5 axisymmetric mixed-compression inlet. [Lewis 10- by 10-ft. supersonic wind tunnel test

NASA Technical Reports Server (NTRS)

Sanders, B. W.

1980-01-01

The throat of a Mach 2.5 inlet that was attached to a turbojet engine was fitted with large, porous bleed areas to provide a stability bypass system that would allow a large, stable airflow range. Exhaust-nozzle, secondary-airflow pumping was used as the exit control for the stability bypass airflow. Propulsion system response and stability bypass performance were obtained for several transient airflow disturbances, both internal and external. Internal airflow disturbances included reductions in overboard bypass airflow, power lever angle, and primary-nozzle area, as well as compressor stall. Nozzle secondary pumping as a stability bypass exit control can provide the inlet with a large stability margin with no adverse effects on propulsion system performance.

Internal Acoustics Measurements of a Full Scale Advanced Ducted Propulsor Demonstrator

NASA Technical Reports Server (NTRS)

Santa Maria, O. L.; Soderman, P. T.; Horne, W. C.; Jones, M. G.; Bock, L. A.

1995-01-01

Acoustics measurements of a Pratt & Whitney full-scale ADP (Advanced Ducted Propulsor), an ultrahigh by-pass ratio engine, were conducted in the NASA Ames 40- by 80-Foot Wind Tunnel. This paper presents data from measurements taken from sensors on a fan exit guide vane in the ADP. Data from two sensors, one at mid-span and the other at the tip of the fan exit guide vane, are presented. At the blade passage frequency (BPF), the levels observed at the various engine and wind speeds were higher at the mid-span sensor than the tip sensor. The coherence between these internal sensors and external microphones were calculated and plotted as a function of angle (angles ranged from 5 degrees to 160 degrees) relative to the ADP longitudinal axis. At the highest engine and wind speeds, the coherence between the tip sensor and the external microphones was observed to decrease at higher multiples of the BPF. These results suggest that the rotor-stator interaction tones are stronger in the mid-span region than at the tip.

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

The F-1 engine was developed and built by Rocketdyne under the direction of the Marshall Space Flight Center. It measured 19 feet tall by 12.5 feet at the nozzle exit, and produced a 1,500,000-pound thrust using liquid oxygen and kerosene as the propellant. The image shows an F-1 engine being test fired at the Test Stand 1-C at the Edwards Air Force Base in California.

An investigation of the effects of the high maximum-thickness-to-chord ratio on the performance of nozzle guide vanes in a transonic planar cascade

NASA Astrophysics Data System (ADS)

Radmard, Rama

1993-03-01

The performance of turbine airfoils is usually predicted by empirical correlations, which however are inadequate for the case of airfoils with maximum thickness to chord ratio (MTCR) higher than 25 percent. Studies were conducted to create a data base from which the performance of turbine airfoils with a MTCR higher than 25 percent could be predicted. A planar cascade consisting of four airfoils was constructed to allow the investigation of the effect of the MTCR on the airfoil performance. Three airfoil sets with MTCR of 15.2 percent (baseline), 26.6 percent, and 48.2 percent were used. Measurements included surface Mach number distributions for the baseline airfoil, total pressure loss coefficients, and deviation angles for isentropic exit Mach numbers of 0.7 (design), 0.9, and 1.1. The effect of varying the inlet boundary layer thickness and free-stream turbulence level was also examined. The results showed that the 26.6 percent airfoil produced lower losses as predicted by the Kacker and Okapuu (1982) correlation. The introduction of turbulence produced a significant redistribution of losses in the exit plane. The secondary loss decreased as the leading edge diameter was increased. Except for the baseline blade where high under-turning in exit flow angle was observed, the airfoils showed a decrease in over-turning with increasing exit Mach number, as predicted by Ainley and Mathieson (1951).

Combustion response to acoustic perturbation in liquid rocket engines

NASA Astrophysics Data System (ADS)

Ghafourian, Akbar

An experimental study of the effect of acoustic perturbations on combustion behavior of a model liquid propellant rocket engine has been carried out. A pair of compression drivers were used to excite transverse and longitudinal acoustic fields at strengths of up to 156.6 dB and 159.5 dB respectively in the combustion chamber of the experimental rocket engine. Propellant simulants were injected into the combustion chamber through a single element shear coaxial injector. Water and air were used in cold flow studies and ethanol and oxygen-enriched air were used as fuel and oxidizer in reacting hot flow studies. In cold flow studies an imposed transverse acoustic field had a more pronounced effect on the spray pattern than a longitudinal acoustic fields. A transverse acoustic field widened the spray by as much as 33 percent and the plane of impingement of the spray with chamber walls moved up closer to the injection plane. The behavior was strongly influenced by the gas phase velocity but was less sensitive to changes in the liquid phase velocity. In reacting hot flow studies the effects of changes in equivalence ratio, excitation amplitude, excitation frequency, liquid and gas phase velocity and chamber pressure on the response of the injector to imposed high frequency transverse acoustic excitation were measured. Reducing the equivalence ratio from 7.4 to 3.8 increased the chamber pressure response to the imposed excitation at 3000 Hz. Increasing the excitation amplitude from 147 dB to 155.6 dB at 3000 Hz increased the chamber pressure response to the excitation. In the frequency range of 1240 Hz to 3220 Hz, an excitation frequency of 3000 Hz resulted in the largest response of the chamber pressure indicating the importance of fluid dynamic coupling. Increasing the liquid phase velocity from 9.2 m/sec to 22.7 m/sec, did not change the amplitude of the chamber pressure response to excitation. This implied the importance of local equivalence ratio and not the overall equivalence ratio on chamber pressure response to excitation. Increasing the chamber pressure from 1.5 atm to 3.1 atm and gas phase velocity from 93.2 m/sec to 105.1 m/sec significantly increased the chamber pressure response to acoustic excitation. This emphasized the significance of the gas phase density and velocity. Measurements of the free radical C2 emission zone and Schlieren images indicated that transverse acoustic excitation moved the combustion zone closer to the injection plane and longitudinal acoustic excitation widened the combustion zone. The histogram of these images indicates that the area over which combustion takes place in the chamber increases under imposed acoustic excitation. This implied that more propellants combust prior to exiting from the exhaust nozzle under unsteady conditions.

Cold Aero Performance of a Two-Dimensional Mixer Ejector Nozzle

NASA Technical Reports Server (NTRS)

Balan, C.

2005-01-01

Since 1986, NASA and the U.S. aerospace industry have been assessing the economic viability and environmental acceptability of a second-generation supersonic civil transport, or High Speed Civil Transport (HSCT). Environmental acceptability in terms of airport community noise and economic viability are critical elements in this endeavor. Development of a propulsion system that satisfies strict airport noise regulations (FAR36 Stage III levels), at acceptable performance and weight, is critical to the success of any HSCT program. Two-dimensional mixer-ejector (2DME) exhaust systems are one approach in achieving this goal. In support of HSCT development, GEAE (GE Aircraft Engines), under contract to the NASA Glenn Research Center, conducted this test program at the NASA Langley 16 ft transonic wind tunnel to evaluate the cold aerodynamic performance aspects of the 2DME exhaust system concept. The effects of SAR (SAR, suppressor area ratio, = mixed-flow area/primary nozzle throat area), MAR (MAR = overall exhaust system exit/mixing-plane area), flap length, CER (suppressor chute expansion ratio), chute alignment, and free stream Mach number were investigated on a 1/11th cold aerodynamic scale model of a 2DME exhaust system.

Investigation of Primary Mirror Segment's Residual Errors for the Thirty Meter Telescope

NASA Technical Reports Server (NTRS)

Seo, Byoung-Joon; Nissly, Carl; Angeli, George; MacMynowski, Doug; Sigrist, Norbert; Troy, Mitchell; Williams, Eric

2009-01-01

The primary mirror segment aberrations after shape corrections with warping harness have been identified as the single largest error term in the Thirty Meter Telescope (TMT) image quality error budget. In order to better understand the likely errors and how they will impact the telescope performance we have performed detailed simulations. We first generated unwarped primary mirror segment surface shapes that met TMT specifications. Then we used the predicted warping harness influence functions and a Shack-Hartmann wavefront sensor model to determine estimates for the 492 corrected segment surfaces that make up the TMT primary mirror. Surface and control parameters, as well as the number of subapertures were varied to explore the parameter space. The corrected segment shapes were then passed to an optical TMT model built using the Jet Propulsion Laboratory (JPL) developed Modeling and Analysis for Controlled Optical Systems (MACOS) ray-trace simulator. The generated exit pupil wavefront error maps provided RMS wavefront error and image-plane characteristics like the Normalized Point Source Sensitivity (PSSN). The results have been used to optimize the segment shape correction and wavefront sensor designs as well as provide input to the TMT systems engineering error budgets.

ISS Expedition 41 Return to Ellington Field

NASA Image and Video Library

2015-03-12

Footage of ISS Expedition 41 Flight Engineer Reid Wiseman's return to Ellington Field. Includes footage of the aircraft landing at Ellington Field; Wiseman as he exits the aircraft and is greeted by family, friends and NASA Officials.

Expedition 9 Landing

NASA Image and Video Library

2004-10-24

Expedition 9 Flight Engineer Michael Fincke, lower left, greets his wife and children for the first time in six months after exiting the Gargarin Cosmonaut Training Center's airplane in Star City, Russia, Sunday October 24, 2004. Photo Credit: (NASA/Bill Ingalls)

Cold-air performance of a 15.41-cm-tip-diameter axial-flow power turbine with variable-area stator designed for a 75-kW automotive gas turbine engine

NASA Technical Reports Server (NTRS)

Mclallin, K. L.; Kofskey, M. G.; Wong, R. Y.

1982-01-01

An experimental evaluation of the aerodynamic performance of the axial flow, variable area stator power turbine stage for the Department of Energy upgraded automotive gas turbine engine was conducted in cold air. The interstage transition duct, the variable area stator, the rotor, and the exit diffuser were included in the evaluation of the turbine stage. The measured total blading efficiency was 0.096 less than the design value of 0.85. Large radial gradients in flow conditions were found at the exit of the interstage duct that adversely affected power turbine performance. Although power turbine efficiency was less than design, the turbine operating line corresponding to the steady state road load power curve was within 0.02 of the maximum available stage efficiency at any given speed.

Three-Dimensional Flow Field Measurements in a Transonic Turbine Cascade

NASA Technical Reports Server (NTRS)

Giel, P. W.; Thurman, D. R.; Lopez, I.; Boyle, R. J.; VanFossen, G. J.; Jett, T. A.; Camperchioli, W. P.; La, H.

1996-01-01

Three-dimensional flow field measurements are presented for a large scale transonic turbine blade cascade. Flow field total pressures and pitch and yaw flow angles were measured at an inlet Reynolds number of 1.0 x 10(exp 6) and at an isentropic exit Mach number of 1.3 in a low turbulence environment. Flow field data was obtained on five pitchwise/spanwise measurement planes, two upstream and three downstream of the cascade, each covering three blade pitches. Three-hole boundary layer probes and five-hole pitch/yaw probes were used to obtain data at over 1200 locations in each of the measurement planes. Blade and endwall static pressures were also measured at an inlet Reynolds number of 0.5 x 10(exp 6) and at an isentropic exit Mach number of 1.0. Tests were conducted in a linear cascade at the NASA Lewis Transonic Turbine Blade Cascade Facility. The test article was a turbine rotor with 136 deg of turning and an axial chord of 12.7 cm. The flow field in the cascade is highly three-dimensional as a result of thick boundary layers at the test section inlet and because of the high degree of flow turning. The large scale allowed for very detailed measurements of both flow field and surface phenomena. The intent of the work is to provide benchmark quality data for CFD code and model verification.

The effect of non-zero radial velocity on the impulse and circulation of starting jets

NASA Astrophysics Data System (ADS)

Krieg, Michael; Mohseni, Kamran

2011-11-01

Vortex ring formation dynamics are generally studied using two basic types of vortex generators. Piston cylinder vortex generators eject fluid through a long tube which ensures a purely axial jet; whereas, vortex ring generators which expel fluid through a flat plate with a circular orifice produce 2-D jets (non-zero radial velocity). At the nozzle exit plane of the orifice type vortex generator the radial component of velocity is linearly proportional to the radial distance from the axis of symmetry, reaching a maximum at the edge of the orifice with a magnitude around 10 % of the piston velocity (the ratio of the volume flux and the nozzle area). As the jet advances downstream the radial velocity quickly dissipates, and becomes purely axial less than a diameter away from the nozzle exit plane. The radial velocity gradient in the axial direction plays a key role in the rate at which circulation and impulse are ejected from the vortex generator. Though the radial component of velocity is small compared to the axial velocity, it has a significant effect on both the circulation and impulse of the starting jet because of this gradient. The extent of circulation and impulse enhancement is investigated through experimental DPIV data showing that the orifice device produces nearly double both circulation and energy (with identical piston velocity and stroke ratios).

Design method of freeform light distribution lens for LED automotive headlamp based on DMD

NASA Astrophysics Data System (ADS)

Ma, Jianshe; Huang, Jianwei; Su, Ping; Cui, Yao

2018-01-01

We propose a new method to design freeform light distribution lens for light-emitting diode (LED) automotive headlamp based on digital micro mirror device (DMD). With the Parallel optical path architecture, the exit pupil of the illuminating system is set in infinity. Thus the principal incident rays of micro lens in DMD is parallel. DMD is made of high speed digital optical reflection array, the function of distribution lens is to distribute the emergent parallel rays from DMD and get a lighting pattern that fully comply with the national regulation GB 25991-2010.We use DLP 4500 to design the light distribution lens, mesh the target plane regulated by the national regulation GB 25991-2010 and correlate the mesh grids with the active mirror array of DLP4500. With the mapping relations and the refraction law, we can build the mathematics model and get the parameters of freeform light distribution lens. Then we import its parameter into the three-dimensional (3D) software CATIA to construct its 3D model. The ray tracing results using Tracepro demonstrate that the Illumination value of target plane is easily adjustable and fully comply with the requirement of the national regulation GB 25991-2010 by adjusting the exit brightness value of DMD. The theoretical optical efficiencies of the light distribution lens designed using this method could be up to 92% without any other auxiliary lens.

Spatial correlation in the ambient core noise field of a turbofan engine.

PubMed

Miles, Jeffrey Hilton

2012-06-01

An acoustic transfer function relating combustion noise and turbine exit noise in the presence of enclosed ambient core noise is investigated using a dynamic system model and an acoustic system model for the particular turbofan engine studied and for a range of operating conditions. Measurements of cross-spectra magnitude and phase between the combustor and turbine exit and auto-spectra at the turbine exit and combustor are used to show the presence of indirect and direct combustion noise over the frequency range of 0-400 Hz. The procedure used evaluates the ratio of direct to indirect combustion noise. The procedure used also evaluates the post-combustion residence time in the combustor which is a factor in the formation of thermal NO(x) and soot in this region. These measurements are masked by the ambient core noise sound field in this frequency range which is observable since the transducers are situated within an acoustic wavelength of one another. An ambient core noise field model based on one and two dimensional spatial correlation functions is used to replicate the spatially correlated response of the pair of transducers. The spatial correlation function increases measured attenuation due to destructive interference and masks the true attenuation of the turbine.

Investigation at Mach Numbers 2.98 and 2.18 of Axially Symmetric Free-jet Diffusion with a Ram-jet Engine

NASA Technical Reports Server (NTRS)

Hunczak, Henry R

1952-01-01

An investigation was conducted to determine the effectiveness of a free-jet diffuser in reducing the over-all pressure ratios required to operate a free jet with a large air-breathing engine as a test vehicle. Efficient operation of the free jet was determined with and without the considerations required for producing suitable engine-inlet flow conditions. A minimum operating pressure ration of 5.5 was attained with a ratio of nozzle-exit to engine-inlet area of 1.85. Operation of the free jet with unstable engine-inlet flow (buzz) is also included.

La réflexion des ondes électromagnétiques du milieu de gyrotropie anisotrope, avec un axe d'anisotropie et un angle d'incidence arbitraires

NASA Astrophysics Data System (ADS)

Tchalyk, A.

1994-01-01

The interaction of plane electromagnetic wave with half-infinit gyrotropic media for different directions of propagation of incident microwave radiation and various directions of gyrotropic axis are investigated. The equations for calculation of wave's parameters exited in media and reflected from media-surface are obtained. On étudie l'interaction d'une onde électromagnétique plane avec un milieu gyrotrope à demi-infini dans le cas d'une direction arbitraire de propagation de la radiation électromagnétique incidente et de l'axe de gyrotropie. On a obtenu des expressions permettant de calculer les caractéristiques des ondes excitées dans et par le milieu.

Guidance and control strategies for aerospace vehicles

NASA Technical Reports Server (NTRS)

Hibey, Joseph L.; Naidu, Desineni S.

1990-01-01

The first part of the report concerns broadly the summary of the work done in the areas of singular perturbations and time scales (SPaTS), aerobraking technology, guidance and aerocruise. The synergistic plane change problem connected with orbital transfer employing aeroassist technology, is addressed. The mission involves transfer from high Earth orbit to low Earth orbit with plane change being performed within the atmosphere. The complete mission consists of a deorbit phase, atmospheric phase, and finally reorbit phase. The atmospheric maneuver is composed of an entry mode, a cruise mode, and finally an exit mode. During the cruise mode, constant altitude and velocity are maintained by means of bank angle control with constant thrust or thrust control with constant bank angle. Comparisons between these two control strategies bring out some interesting features.

Anti-air pollution & energy conservation system for automobiles using leaded or unleaded gasoline, diesel or alternate fuel

DOEpatents

Bose, Ranendra K.

2002-06-04

Exhaust gases from an internal combustion engine operating with leaded or unleaded gasoline or diesel or natural gas, are used for energizing a high-speed gas turbine. The convoluting gas discharge causes a first separation stage by stratifying of heavier and lighter exhaust gas components that exit from the turbine in opposite directions, the heavier components having a second stratifying separation in a vortex tube to separate combustible pollutants from non-combustible components. The non-combustible components exit a vortex tube open end to atmosphere. The lighter combustible, pollutants effected in the first separation are bubbled through a sodium hydroxide solution for dissolving the nitric oxide, formaldehyde impurities in this gas stream before being piped to the engine air intake for re-combustion, thereby reducing the engine's exhaust pollution and improving its fuel economy. The combustible, heavier pollutants from the second separation stage are piped to air filter assemblies. This gas stream convoluting at a high-speed through the top stator-vanes of the air filters, centrifugally separates the coalescent water, aldehydes, nitrogen dioxides, sulfates, sulfur, lead particles which collect at the bottom of the bowl, wherein it is periodically released to the roadway. Whereas, the heavier hydrocarbon, carbon particles are piped through the air filter's porous element to the engine air intake for re-combustion, further reducing the engine's exhaust pollution and improving its fuel economy.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Reininger, Ruben; Dhesi, Sarnjeet

The main requirement of the Nanoscience Beamline at Diamond is to deliver the highest possible flux at the sample position of a PEEM with a resolving power of about 5000 in the energy range 80-2000 eV. The source of the beamline is a couple of APPLE II helical undulators in tandem that can also be used separately to allow for faster switching of the circular polarization. Based on its versatility, a collimated plane grating monochromator using sagittally focusing elements was chosen to cover the required energy range with three gratings. The operation of this monochromator requires a collimated beam incidentmore » on the grating along the dispersion direction. This can be achieved either with a toroid, focusing with its major radius along the non-dispersive direction at the exit slit, or with a sagittal cylinder. The former option uses a sagittal cylinder after the grating to focus the collimated beam at the exit slit. In the latter case, a toroid after the grating is used to focus in both directions at the exit slit. The advantage of the toroid downstream the grating is the higher horizontal demagnification. This configuration fulfills the Nanoscience Beamline's required resolving power but cannot be used to achieve very high resolution due to the astigmatic coma aberration of the toroidal mirror. The focusing at the sample position is performed with a KB pair of plane elliptical mirrors. Assuming achievable values for the errors on all the optical surfaces, the expected spots FWHW in the horizontal and vertical directions are 10 {mu}m and 3 {mu}m, respectively. The calculated photon flux at this spot at 5000 resolving power is >1012 photons/sec between 80 and 1600 eV for linearly polarized light and between 106 and 1200 eV for circularly polarized light. The beamline is expected to be operational in January 2007.« less

Laser propagation through full-scale, high-gain MagLIF gas pipes using the NIF

NASA Astrophysics Data System (ADS)

Pollock, Bradley; Sefkow, Adam; Goyon, Clement; Strozzi, David; Khan, Shahab; Rosen, Mordy; Campbell, Mike; Logan, Grant; Peterson, Kyle; Moody, John

2016-10-01

The first relevant measurements of laser propagation through surrogate high-gain MagLIF gas pipe targets at full scale have been performed at the NIF, using 30 kJ of laser drive from one quad in a 10 ns pulse at an intensity of 2e14 W/cm2. The unmagnetized pipe is filled with 1 atm of 99%/1% neopentane/Ar, and uses an entrance window of 0.75 um polyimide and an exit window of 0.3 um of Ta backed with 5 um of polyimide. Side-on x-ray emission from the plasma is imaged through the 100 um-thick epoxy wall onto a framing camera at four times during the drive, and is in excellent agreement with pre-shot HYDRA radiation-hydrodynamics modeling. X-ray emission from the Ta exit plane is imaged onto a streak camera to determine the timing and intensity of the laser burning through the pipe, and the Ar emission from the center of the pipe is spectrally- and temporally-resolved to determine the plasma electron temperature. Backscatter is measured throughout the laser drive, and is found to be of significance only when the laser reaches the Ta exit plane and produces SBS. These first results in unmagnetized surrogate gas fills are encouraging since they demonstrate sufficient laser energy absorption and low LPI losses within high-density long-scale-length plasmas for proposed high-gain MagLIF target designs. We will discuss plans to magnetize targets filled with high-density DT gas in future experiments. This work performed under the auspices of the U.S. Department of Energy by Lawrence Livermore National Laboratory under Contract DE-AC52-07NA27344. Sandia is a multiprogram laboratory operated by Sandia Corporation, a Lockheed Martin Company, for the National Nuclear Security Administration under Contract No. DE-AC04-94AL85000.

An Experimental and Numerical Investigation of Endwall Aerodynamics and Heat Transfer in a Gas Turbine Nozzle Guide Vane with Slot Film Cooling

NASA Astrophysics Data System (ADS)

Alqefl, Mahmood Hasan

In many regions of the high-pressure gas turbine, film cooling flows are used to protect the turbine components from the combustor exit hot gases. Endwalls are challenging to cool because of the complex system of secondary flows that disturb surface film coolant coverage. The secondary flow vortices wash the film coolant from the surface into the mainstream significantly decreasing cooling effectiveness. In addition to being effected by secondary flow structures, film cooling flow can also affect these structures by virtue of their momentum exchange. In addition, many studies in the literature have shown that endwall contouring affects the strength of passage secondary flows. Therefore, to develop better endwall cooling schemes, a good understanding of passage aerodynamics and heat transfer as affected by interactions of film cooling flows with secondary flows is required. This experimental and computational study presents results from a linear, stationary, two-passage cascade representing the first stage nozzle guide vane of a high-pressure gas turbine with an axisymmetrically contoured endwall. The sources of film cooling flows are upstream combustor liner coolant and endwall slot film coolant injected immediately upstream of the cascade passage inlet. The operating conditions simulate combustor exit flow features, with a high Reynolds number of 390,000 and approach flow turbulence intensity of 11% with an integral length scale of 21% of the chord length. Measurements are performed with varying slot film cooling mass flow to mainstream flow rate ratios (MFR). Aerodynamic effects are documented with five-hole probe measurements at the exit plane. Heat transfer is documented through recovery temperature measurements with a thermocouple. General secondary flow features are observed. Total pressure loss measurements show that varying the slot film cooling MFR has some effects on passage loss. Velocity vectors and vorticity distributions show a very thin, yet intense, cross-pitch flow on the contoured endwall side. Endwall adiabatic effectiveness values and coolant distribution thermal fields show minimal effects of varying slot film coolant MFR. This suggests the dominant effects of combustor liner coolant. show dominant effects of combustor liner coolant on cooling the endwall. A coolant vorticity correlation presenting the advective mixing of the coolant due to secondary flow vorticity at the exit plane is also discussed.

Feasibility evaluation of the monolithic braided ablative nozzle

NASA Astrophysics Data System (ADS)

Director, Mark N.; McPherson, Douglass J., Sr.

1992-02-01

The feasibility of the monolithic braided ablative nozzle was evaluated as part of an independent research and development (IR&D) program complementary to the National Aeronautics and Space Administration/Marshall Space Flight Center (NASA/MSFC) Low-Cost, High-Reliability Case, Insulation and Nozzle for Large Solid Rocket Motors (LOCCIN) Program. The monolithic braided ablative nozzle is a new concept that utilizes a continuous, ablative, monolithic flame surface that extends from the nozzle entrance, through the throat, to the exit plane. The flame surface is fabricated using a Through-the-Thickness braided carbon-fiber preform, which is impregnated with a phenolic or phenolic-like resin. During operation, the braided-carbon fiber/resin material ablates, leaving the structural backside at temperatures which are sufficiently low to preclude the need for any additional insulative materials. The monolithic braided nozzle derives its potential for low life cycle cost through the use of automated processing, one-component fabrication, low material scrap, low process scrap, inexpensive raw materials, and simplified case attachment. It also has the potential for high reliability because its construction prevents delamination, has no nozzle bondlines or leak paths along the flame surface, is amenable to simplified analysis, and is readily inspectable. In addition, the braided construction has inherent toughness and is damage-tolerant. Two static-firing tests were conducted using subscale, 1.8 - 2.0-inch throat diameter, hardware. Tests were approximately 15 seconds in duration, using a conventional 18 percent aluminum/ammonium perchlorate propellant. The first of these tests evaluated the braided ablative as an integral backside insulator and exit cone; the second test evaluated the monolithic braided ablative as an integral entrance/throat/exit cone nozzle. Both tests met their objectives. Radial ablation rates at the throat were as predicted, approximately 0.017 in/sec; these rates are comparable to those for tapewrapped carbon phenolic materials. The maximum temperature rise on the outside surface occurred one inch from the nozzle exit plane and was less than 50 F at the end of the test. Further development for this concept is scheduled as part of phase 2 on the NASA/MSFC LOCCIN Program. During this effort, the nozzle materials, architecture, and processing will be optimized and tested in nozzles with 3- and 10-inch diameter throats. Further, a design and manufacturing plan for a full-scale, 20-inch-diameter throat, nozzle will be developed.

Influence of the cooling degree upon performances of internal combustion engine

NASA Astrophysics Data System (ADS)

Grǎdinariu, Andrei Cristian; Mihai, Ioan

2016-12-01

Up to present, air cooling systems still raise several unsolved problems due to conditions imposed by the environment in terms of temperature and pollution levels. The present paper investigates the impact of the engine cooling degree upon its performances, as important specific power is desired for as low as possible fuel consumption. A technical solution advanced by the authors[1], consists of constructing a bi-flux compressor, which can enhance the engine's performances. The bi-flux axial compressor accomplishes two major functions, that is it cools down the engine and it also turbocharges it. The present paper investigates the temperature changes corresponding to the fresh load, during the use of a bi-flux axial compressor. This compressor is economically simple, compact, and offers an optimal response at low rotational speeds of the engine, when two compression steps are used. The influence of the relative coefficient of air temperature drop upon working agent temperature at the intercooler exit is also investigated in the present work. The variation of the thermal load coefficient by report to the working agent temperature is also investigated during engine cooling. The variation of the average combustion temperature is analyzed in correlation to the thermal load coefficient and the temperatures of the working fluid at its exit from the cooling system. An exergetic analysis was conducted upon the influence of the cooling degree on the motor fluid and the gases resulted from the combustion process.

Real-time simulation of an F110/STOVL turbofan engine

NASA Technical Reports Server (NTRS)

Drummond, Colin K.; Ouzts, Peter J.

1989-01-01

A traditional F110-type turbofan engine model was extended to include a ventral nozzle and two thrust-augmenting ejectors for Short Take-Off Vertical Landing (STOVL) aircraft applications. Development of the real-time F110/STOVL simulation required special attention to the modeling approach to component performance maps, the low pressure turbine exit mixing region, and the tailpipe dynamic approximation. Simulation validation derives by comparing output from the ADSIM simulation with the output for a validated F110/STOVL General Electric Aircraft Engines FORTRAN deck. General Electric substantiated basic engine component characteristics through factory testing and full scale ejector data.

An Integrated Optimal Estimation Approach to Spitzer Space Telescope Focal Plane Survey

NASA Technical Reports Server (NTRS)

Bayard, David S.; Kang, Bryan H.; Brugarolas, Paul B.; Boussalis, D.

2004-01-01

This paper discusses an accurate and efficient method for focal plane survey that was used for the Spitzer Space Telescope. The approach is based on using a high-order 37-state Instrument Pointing Frame (IPF) Kalman filter that combines both engineering parameters and science parameters into a single filter formulation. In this approach, engineering parameters such as pointing alignments, thermomechanical drift and gyro drifts are estimated along with science parameters such as plate scales and optical distortions. This integrated approach has many advantages compared to estimating the engineering and science parameters separately. The resulting focal plane survey approach is applicable to a diverse range of science instruments such as imaging cameras, spectroscopy slits, and scanning-type arrays alike. The paper will summarize results from applying the IPF Kalman Filter to calibrating the Spitzer Space Telescope focal plane, containing the MIPS, IRAC, and the IRS science Instrument arrays.

Corps of Engineers Hydraulic Design Criteria. Volume I

DTIC Science & Technology

1977-01-01

DESIGN CRITERIA CLASSIFICATION INDEX S000-GENERAL 000 Physical Constants 001 Fluid Properties 010 Open Channel Flow 020 Free Overflow 030 Pressure Flow...Dissipation 113 Erosion below Spillways 120 Chute Spillways 121 Approach Channel 122 Ogee Crests 123 Spillay Chutes S ii124 Spillway Stilling Basins 125...Spillvay Exit Channel Revised 5-59 .. . j1.I.i edCORPS OF ENGINEERS HYDRAULIC DESIGN CRITERIA CLASSIFICATION INDEX (Continued) %. IO0-SPILLWAYS

Effects of injection nozzle exit width on rotating detonation engine

NASA Astrophysics Data System (ADS)

Sun, Jian; Zhou, Jin; Liu, Shijie; Lin, Zhiyong; Cai, Jianhua

2017-11-01

A series of numerical simulations of RDE modeling real injection nozzles with different exit widths are performed in this paper. The effects of nozzle exit width on chamber inlet state, plenum flowfield and detonation propagation are analyzed. The results are compared with that using an ideal injection model. Although the ideal injection model is a good approximation method to model RDE inlet, the two-dimensional effects of real nozzles are ignored in the ideal injection model so that some complicated phenomena such as the reflected waves caused by the nozzle walls and the reversed flow into the nozzles can not be modeled accurately. Additionally, the ideal injection model overpredicts the block ratio. In all the cases that stabilize at one-wave mode, the block ratio increases as the nozzle exit width gets smaller. The dual-wave mode case also has a relatively high block ratio. A pressure oscillation in the plenum with the same main frequency with the rotating detonation wave is observed. A parameter σ is applied to describe the non-uniformity in the plenum. σ increases as the nozzle exit width gets larger. Under some condition, the heat release on the interface of fresh premixed gas layer and detonation products can be strong enough to induce a new detonation wave. A spontaneous mode-transition process is observed for the smallest exit width case. Due to the detonation products existing in the premixed gas layer before the detonation wave, the detonation wave will propagate through reactants and products alternately, and therefore its strength will vary with time, especially near the chamber inlet. This tendency gets weaker as the injection nozzle exit width increases.

Design, durability and low cost processing technology for composite fan exit guide vanes

NASA Technical Reports Server (NTRS)

Blecherman, S. S.

1979-01-01

A lightweight composite fan exit guide vane for high bypass ratio gas turbine engine application was investigated. Eight candidate material/design combinations were evaluated by NASTRAN finite element analyses. A total of four combinations were selected for further analytical evaluation, part fabrication by two ventors, and fatigue test in dry and wet condition. A core and shell vane design was chosen in which the unidirectional graphite core fiber was the same for all candidates. The shell material, fiber orientation, and ply configuration were varied. Material tests were performed on raw material and composite specimens to establish specification requirements. Pre-test and post-test microstructural examination and nondestructive analyses were conducted to determine the effect of material variations on fatigue durability and failure mode. Relevant data were acquired with respect to design analysis, materials properties, inspection standards, improved durability, weight benefits, and part price of the composite fan exit guide vane.

Assessment of suturing in the vertical plane shows the efficacy of the multi-degree-of-freedom needle driver for neonatal laparoscopy.

PubMed

Takazawa, Shinya; Ishimaru, Tetsuya; Fujii, Masahiro; Harada, Kanako; Sugita, Naohiko; Mitsuishi, Mamoru; Iwanaka, Tadashi

2013-11-01

We have developed a thin needle driver with multiple degrees-of-freedom (DOFs) for neonatal laparoscopic surgery. The tip of this needle driver has three DOFs for grasp, deflection and rotation. Our aim was to evaluate the performance of the multi-DOF needle driver in vertical plane suturing. Six pediatric surgeons performed four directional suturing tasks in the vertical plane using the multi-DOF needle driver and a conventional one. Assessed parameters were the accuracy of insertion and exit, the depth of suture, the inclination angle of the needle and the force applied on the model. In left and right direction sutures, the inclination angle of the needle with the multi-DOF needle driver was significantly smaller than that with the conventional one (p = 0.014, 0.042, respectively). In left and right direction sutures, the force for pulling the model with the multi-DOF needle driver was smaller than that with the conventional one (p = 0.036, 0.010, respectively). This study showed that multi-directional suturing on a vertical plane using the multi-DOF needle driver had better needle trajectories and was less invasive as compared to a conventional needle driver.

Hover and Wind-Tunnel Testing of Shrouded Rotors for Improved Micro Air Vehicle Design

DTIC Science & Technology

2008-01-01

and the shroud surface pressure distributions. The uniformity of the wake was improved by the presence of the shrouds and by decreasing the blade tip...213 3.35 Effect of blade tip clearance on shrouded-rotor exit-plane wake profiles215 3.36 Effects of changing blade tip clearance on induced...Wright [139] developed a vortex wake model for heavily loaded ducted fans, in which the “inner vortex sheets [shed from the blades ] move at a different

High-Obliquity Impact of a Compact Penetrator on a Thin Plate: Penetrator Splitting and Adiabatic Shear

DTIC Science & Technology

1998-01-01

nonideal penetrator on a thin plate at high obliquities. These computations simulated two series of experiments at velocities of 1.5 km/ s and 4.1 km/ s ...3 2. Combined Effects of Obliquity, 0, and Rotation, 4, on Debris Cloud Evolution at 4.1 km/ s and 26 p s ; Impact Velocity Vector Lies in x-z Plane...7 3. Time History of the Penetrator Mass Fraction Exiting the Bottom of the Target at 4.1 km / s

Acoustic radiation from lined, unflanged ducts: Acoustic source distribution program

NASA Technical Reports Server (NTRS)

Beckemeyer, R. J.; Sawdy, D. T.

1971-01-01

An acoustic radiation analysis was developed to predict the far-field characteristics of fan noise radiated from an acoustically lined unflanged duct. This analysis is comprised of three modular digital computer programs which together provide a capability of accounting for the impedance mismatch at the duct exit plane. Admissible duct configurations include circular or annular, with or without an extended centerbody. This variation in duct configurations provides a capability of modeling inlet and fan duct noise radiation. The computer programs are described in detail.

Expedition 9 Landing

NASA Image and Video Library

2004-10-24

Expedition 9 Commander Gennady Padalka, center, and Expedition 9 Flight Engineer Michael Fincke, background, exit the Gargarin Cosmonaut Training Center's airplane to an awaiting crowd of officials and family members after their flight from Kustanay, Kazakhstan, Sunday, October 24, 2004. Photo Credit: (NASA/Bill Ingalls)

The impact of the fuel chemical composition on volatile organic compounds emitted by an in-service aircraft gas turbine engine

NASA Astrophysics Data System (ADS)

Setyan, A.; Kuo, Y. Y.; Brem, B.; Durdina, L.; Gerecke, A. C.; Heeb, N. V.; Haag, R.; Wang, J.

2017-12-01

Aircraft emissions received increased attention recently because of the steady growth of aviation transport in the last decades. Aircraft engines substantially contribute to emissions of particulate matter and gaseous pollutants in the upper and lower troposphere. Among all the pollutants emitted by aircrafts, volatile organic compounds (VOCs) are particularly important because they are mainly emitted at ground level, posing a serious health risk for people living or working near airports. A series of measurements was performed at the aircraft engine testing facility of SR Technics (Zürich airport, Switzerland). Exhausts from an in-service turbofan engine were sampled at the engine exit plane by a multi-point sampling probe. A wide range of instruments was connected to the common sampling line to determine physico-chemical characteristics of non-volatile particulate matter and gaseous pollutants. Conventional Jet A-1 fuel was used as the base fuel, and measurements were performed with the base fuel doped with two different mixtures of aromatic compounds (Solvesso 150 and naphthalene-depleted Solvesso 150) and an alternative fuel (hydro-processed esters and fatty acids [HEFA] jet fuel). During this presentation, we will show results obtained for VOCs. These compounds were sampled with 3 different adsorbing cartridges, and analyzed by thermal desorption gas chromatography/mass spectrometry (TD-GC/MS, for Tenax TA and Carboxen 569) and by ultra-performance liquid chromatography/ mass spectrometry (UPLC/MS, for DNPH). The total VOC concentration was also measured with a flame ionization detector (FID). In addition, fuel samples were also analyzed by GC/MS, and their chemical compositions were compared to the VOCs emitted via engine exhaust. Total VOCs concentrations were highest at ground idle (>200 ppm C at 4-7% thrust), and substantially lower at high thrust (<3 ppm C during take-off, 100% thrust). Fuel samples were dominated by alkanes, whereas VOCs emitted by the aircraft engine were mainly constituted of alkanes, oxygenated compounds, and aromatics. More than 50 % of the compounds identified in the exhaust were not present in the fuel, and thus were formed during combustion. The impact of the fuel doping with aromatics and the alternative fuel on VOCs emitted by the engine will also be discussed.

Langley Mach 4 scramjet test facility

NASA Technical Reports Server (NTRS)

Andrews, E. H., Jr.; Torrence, M. G.; Anderson, G. Y.; Northam, G. B.; Mackley, E. A.

1985-01-01

An engine test facility was constructed at the NASA Langley Research Center in support of a supersonic combustion ramjet (scramjet) technology development program. Hydrogen combustion in air with oxygen replenishment provides simulated air at Mach 4 flight velocity, pressure, and true total temperature for an altitude range from 57,000 to 86,000 feet. A facility nozzle with a 13 in square exit produces a Mach 3.5 free jet flow for engine propulsion tests. The facility is described and calibration results are presented which demonstrate the suitability of the test flow for conducting scramjet engine research.

Fuel injector nozzle for an internal combustion engine

DOEpatents

Cavanagh, Mark S [Bloomington, IL; Urven, Jr., Roger L.; Lawrence, Keith E [Peoria, IL

2011-03-22

A direct injection fuel injector includes a nozzle tip having a plurality of passages allowing fluid communication between an inner nozzle tip surface portion and an outer nozzle tip surface portion and directly into a combustion chamber of an internal combustion engine. A first group of the passages have inner surface apertures located substantially in a first common plane. A second group of the passages have inner surface apertures located substantially in at least a second common plane substantially parallel to the first common plane. The second group has more passages than the first group.

Fuel Injector Nozzle For An Internal Combustion Engine

DOEpatents

Cavanagh, Mark S.; Urven, Jr.; Roger L.; Lawrence, Keith E.

2006-04-25

A direct injection fuel injector includes a nozzle tip having a plurality of passages allowing fluid communication between an inner nozzle tip surface portion and an outer nozzle tip surface portion and directly into a combustion chamber of an internal combustion engine. A first group of the passages have inner surface apertures located substantially in a first common plane. A second group of the passages have inner surface apertures located substantially in at least a second common plane substantially parallel to the first common plane. The second group has more passages than the first group.

Fuel injector nozzle for an internal combustion engine

DOEpatents

Cavanagh, Mark S.; Urven, Jr., Roger L.; Lawrence, Keith E.

2007-11-06

A direct injection fuel injector includes a nozzle tip having a plurality of passages allowing fluid communication between an inner nozzle tip surface portion and an outer nozzle tip surface portion and directly into a combustion chamber of an internal combustion engine. A first group of the passages have inner surface apertures located substantially in a first common plane. A second group of the passages have inner surface apertures located substantially in at least a second common plane substantially parallel to the first common plane. The second group has more passages than the first group.

Fuel injector nozzle for an internal combustion engine

DOEpatents

Cavanagh, Mark S.; Urven, Jr., Roger L.; Lawrence, Keith E.

2008-11-04

A direct injection fuel injector includes a nozzle tip having a plurality of passages allowing fluid communication between an inner nozzle tip surface portion and an outer nozzle tip surface portion and directly into a combustion chamber of an internal combustion engine. A first group of the passages have inner surface apertures located substantially in a first common plane. A second group of the passages have inner surface apertures located substantially in at least a second common plane substantially parallel to the first common plane. The second group has more passages than the first group.

Orbital transfer rocket engine technology: Advanced engine study

NASA Technical Reports Server (NTRS)

Hayden, Warren R.

1992-01-01

An advanced LOX/LH2 engine study for the use of NASA and vehicle prime contractors in developing concepts for manned missions to the Moon, Mars, and Phobos is documented. Parametric design data was obtained at five engine thrusts from 7.5K lbf to 50K lbf. Also, a separate task evaluated engine throttling over a 20:1 range and operation at a mixture ratio of 12 plus or minus 1 versus the 6 plus or minus 1 nominal. Cost data was also generated for DDT&E, first unit production, and factors in other life cycle costs. The major limitation of the study was lack of contact with vehicle prime contractors to resolve the issues in vehicle/engine interfaces. The baseline Aerojet dual propellant expander cycle was shown capable of meeting all performance requirements with an expected long operational life due to the high thermal margins. The basic engine design readily accommodated the 20:1 throttling requirement and operation up to a mixture ratio of 10 without change. By using platinum for baffled injector construction the increased thermal margin allowed operation up to mixture ratio 13. An initial engine modeling with an Aerojet transient simulation code (named MLETS) indicates stable engine operation with the baseline control system. A throttle ratio of 4 to 5 seconds from 10 percent to 100 percent thrust is also predicted. Performance predictions are 483.1 sec at 7.5K lbf, 487.3 sec at 20K lbf, and 485.2 sec at 50K lbf with a mixture ratio of 6 and an area ratio of 1200. Engine envelopes varied from 120 in. length/53 in. exit diameter at 7.5K lbf to 305 in. length/136 in. exit diameter at 50 K lbf. Packaging will be an important consideration. Continued work is recommended to include more vehicle prime contractor/engine contractor joint assessment of the interface issues.

Experimental Simulation of Turbine-Exhaust Oxygen Recovery

NASA Technical Reports Server (NTRS)

Clark, Jim A.; Branch, Ryan W.

2004-01-01

In some liquid-propellant rocket engines, the liquid-oxygen boost pump is driven by a turbine that is powered by high-pressure gaseous oxygen. Once it exits the turbine, this gaseous oxygen can be salvaged by injecting it into the subcooled liquid oxygen exiting the boost pump. If the main LOX pump is to function correctly under these circumstances, complete condensation of the gaseous oxygen must quickly follow its injection into the boost-pump discharge. The current investigation uses steam and water in a simple rig that allows the condensation process to be visualized and quantified. This paper offers dimensionless-parameter correlations of the data and trends observed.

A study on evacuation time from lecture halls in Faculty of Engineering, Universiti Putra Malaysia

NASA Astrophysics Data System (ADS)

Othman, W. N. A. W.; Tohir, M. Z. M.

2018-04-01

An evacuation situation in any building involves many risks. The geometry of building and high potential of occupant load may affect the efficiency of evacuation process. Although fire safety rules and regulations exist, they remain insufficient to guarantee the safety of all building occupants and do not prevent the dramatic events to be repeated. The main objective of this project is to investigate the relationship between the movement time, travel speed and occupant density during a series of evacuation drills specifically for lecture halls. Generally, this study emphasizes on the movement of crowd within a limited space and includes the aspects of human behaviour. A series of trial evacuations were conducted in selected lecture halls at Faculty of Engineering, Universiti Putra Malaysia with the aim of collecting actual data for numerical analysis. The numerical data obtained during trial evacuations were used to determine the evacuation time, crowd movement and behaviour during evacuation process particularly for lecture halls. The evacuation time and number of occupants exiting from each exit were recorded. Video camera was used to record and observe the movement behaviour of occupants during evacuations. EvacuatioNZ was used to simulate the trials evacuations of DK 5 and the results predicted were compared with experimental data. EvacuatioNZ was also used to predict the evacuation time and the flow of occupants exiting from each door for DK 4 and DK 8.

CARS Temperature and Species Concentration Measurements in a Supersonic Combustor with Normal Injection

NASA Technical Reports Server (NTRS)

Tedder, S. A.; OByrne, S.; Danehy, P. M.; Cutler, A. D.

2005-01-01

The dual-pump coherent anti-Stokes Raman spectroscopy (CARS) method was used to measure temperature and the absolute mole fractions of N2, O2 and H2 in a supersonic combustor. Experiments were conducted in the NASA Langley Direct-Connect Supersonic Combustion Test Facility. CARS measurements were performed at the facility nozzle exit and at three planes downstream of fuel injection. Processing the CARS measurements produced maps of the mean temperature, as well as quantitative N2 and O2 and qualitative H2 mean mole fraction fields at each plane. The CARS measurements were also used to compute correlations between fluctuations of the different simultaneously measured parameters. Comparisons were made between this 90 degree angle fuel injection case and a 30 degree fuel injection case previously presented at the 2004 Reno AIAA Meeting.

Experimental Investigation of the Near-Wall Region in the NASA HiVHAc EDU2 Hall Thruster

NASA Technical Reports Server (NTRS)

Shastry, Rohit; Kamhawi, Hani; Huang, Wensheng; Haag, Thomas W.

2015-01-01

The HiVHAc propulsion system is currently being developed to support Discovery-class NASA science missions. Presently, the thruster meets the required operational lifetime by utilizing a novel discharge channel replacement mechanism. As a risk reduction activity, an alternative approach is being investigated that modifies the existing magnetic circuit to shift the ion acceleration zone further downstream such that the magnetic components are not exposed to direct ion impingement during the thruster's lifetime while maintaining adequate thruster performance and stability. To measure the change in plasma properties between the original magnetic circuit configuration and the modified, "advanced" configuration, six Langmuir probes were flush-mounted within each channel wall near the thruster exit plane. Plasma potential and electron temperature were measured for both configurations across a wide range of discharge voltages and powers. Measurements indicate that the upstream edge of the acceleration zone shifted downstream by as much as 0.104 channel lengths, depending on operating condition. The upstream edge of the acceleration zone also appears to be more insensitive to operating condition in the advanced configuration, remaining between 0.136 and 0.178 channel lengths upstream of the thruster exit plane. Facility effects studies performed on the original configuration indicate that the plasma and acceleration zone recede further upstream into the channel with increasing facility pressure. These results will be used to inform further modifications to the magnetic circuit that will provide maximum protection of the magnetic components without significant changes to thruster performance and stability.

Enhanced microlithography using coated objectives and image duplication

NASA Astrophysics Data System (ADS)

Erdelyi, Miklos; Bor, Zsolt; Szabo, Gabor; Tittel, Frank K.

1998-06-01

Two processes were investigated theoretically using both a scalar wave optics model and a microlithography simulation tool (Solid-C). The first method introduces a phase- transmission filter into the exit pupil plane. The results of both the scalar optics calculation (aerial image) and the Solid-C simulation (resist image) show that the final image profile is optimum, when the exit pupil plane filter is divided into two zones with the inner zone having a phase retardation of (pi) rad with respect to the outer one and the ratio of the radii of the zones is 0.3. Using this optimized filter for the fabrication of isolated contact holes, the focus-exposure process window increases significantly, and the depth of focus (DOF) can be enhanced by a factor of 1.5 to 2. The second technique enhances the DOF of the aerial image by means of a birefringent plate inserted between the projection lens and the wafer. As the shift in focus introduced by the plate strongly depends on the refractive index, two focal points will appear when using a birefringent plate instead of an isotropic plate: the first one is created by the ordinary, and the second one is created by the extraordinary ray. The distance between these images can be controlled by the thickness of the plate. The results of the calculations show that application of a thin but strongly birefringent material is a better candidate than using a slightly birefringent but thick plate, since aberrations proportional to the thickness can cause undesirable effects.

Linearly polarized photoluminescence of anisotropically strained c-plane GaN layers on stripe-shaped cavity-engineered sapphire substrate

NASA Astrophysics Data System (ADS)

Kim, Jongmyeong; Moon, Daeyoung; Lee, Seungmin; Lee, Donghyun; Yang, Duyoung; Jang, Jeonghwan; Park, Yongjo; Yoon, Euijoon

2018-05-01

Anisotropic in-plane strain and resultant linearly polarized photoluminescence (PL) of c-plane GaN layers were realized by using a stripe-shaped cavity-engineered sapphire substrate (SCES). High resolution X-ray reciprocal space mapping measurements revealed that the GaN layers on the SCES were under significant anisotropic in-plane strain of -0.0140% and -0.1351% along the directions perpendicular and parallel to the stripe pattern, respectively. The anisotropic in-plane strain in the GaN layers was attributed to the anisotropic strain relaxation due to the anisotropic arrangement of cavity-incorporated membranes. Linearly polarized PL behavior such as the observed angle-dependent shift in PL peak position and intensity comparable with the calculated value based on k.p perturbation theory. It was found that the polarized PL behavior was attributed to the modification of valence band structures induced by anisotropic in-plane strain in the GaN layers on the SCES.

Numerical Investigation of Cavity-Vane Interactions within the Ultra Combat Combustor

DTIC Science & Technology

2006-03-01

nozzle guide vane and the turbine blades are highly dependent on the temperature distribution of the combustor exit. 20 PatternFactor = T4max − T4avg...Procedure for the Calculation of Gaseous Emissions from Aircraft Turbine Engines ”. Society of Automotive Engineers , June 1996. 5. Bernard, Peter S. and...Whipkey. “Locked Vortex Afterbodies”. Journal of Aircraft , Volume 16, No. 5, May 1979. 17. Liu, Feng and William Sirignano. “ Turbojet and Turbofan

Hypersonic trajectory control of aerospace plane with integrated SCRAMJET engine

NASA Astrophysics Data System (ADS)

Yonemoto, Koichi

The aerospace plane is an airbreathing 'propulsion configured' vehicle having proper forebody contour for inflow pre-compression to the inlet and afterbody that operates as an external expansion nozzle. Since the whole lower side of the body acts as important compression and expansion elements for the airbreathing engine, the flight attitude influences its performance such as specific impulse and thrust coefficient considerably. The stability of ascent trajectory controlling dynamic pressure or heat-input rate is analyzed considering the performance change due to attitude fluctuation. The performance of scramjet engine, a typical hypersonic airbreathing engine, is estimated by a rapid prediction methodology of the combustor proposed by Ikawa.

Aircraft Particle Emissions eXperiment (APEX)

NASA Technical Reports Server (NTRS)

Wey, C. C.; Anderson, B. E.; Hudgins, C.; Wey, C.; Li-Jones, X.; Winstead, E.; Thornhill, L. K.; Lobo, P.; Hagen, D.; Whitefield, P.

2006-01-01

APEX systematically investigated the gas-phase and particle emissions from a CFM56-2C1 engine on NASA's DC-8 aircraft as functions of engine power, fuel composition, and exhaust plumage. Emissions parameters were measured at 11 engine power, settings, ranging from idle to maximum thrust, in samples collected at 1, 10, and 30 m downstream of the exhaust plane as the aircraft burned three fuels to stress relevant chemistry. Gas-phase emission indices measured at 1 m were in good agreement with the ICAO data and predictions provided by GEAE empirical modeling tools. Soot particles emitted by the engine exhibited a log-normal size distribution peaked between 15 and 40 nm, depending on engine power. Samples collected 30 m downstream of the engine exhaust plane exhibited a prominent nucleation mode.

Effects of high combustion chamber pressure on rocket noise environment

NASA Technical Reports Server (NTRS)

Pao, S. P.

1972-01-01

The acoustical environment for a high combustion chamber pressure engine was examined in detail, using both conventional and advanced theoretical analysis. The influence of elevated chamber pressure on the rocket noise environment was established, based on increase in exit velocity and flame temperature, and changes in basic engine dimensions. Compared to large rocket engines, the overall sound power level is found to be 1.5 dB higher, if the thrust is the same. The peak Strouhal number shifted about one octave lower to a value near 0.01. Data on apparent sound source location and directivity patterns are also presented.

Scale model performance test investigation of exhaust system mixers for an Energy Efficient Engine /E3/ propulsion system

NASA Technical Reports Server (NTRS)

Kuchar, A. P.; Chamberlin, R.

1980-01-01

A scale model performance test was conducted as part of the NASA Energy Efficient Engine (E3) Program, to investigate the geometric variables that influence the aerodynamic design of exhaust system mixers for high-bypass, mixed-flow engines. Mixer configuration variables included lobe number, penetration and perimeter, as well as several cutback mixer geometries. Mixing effectiveness and mixer pressure loss were determined using measured thrust and nozzle exit total pressure and temperature surveys. Results provide a data base to aid the analysis and design development of the E3 mixed-flow exhaust system.

Preliminary Study on Acoustic Detection of Faults Experienced by a High-Bypass Turbofan Engine

NASA Technical Reports Server (NTRS)

Boyle, Devin K.

2014-01-01

The vehicle integrated propulsion research (VIPR) effort conducted by NASA and several partners provided an unparalleled opportunity to test a relatively low TRL concept regarding the use of far field acoustics to identify faults occurring in a high bypass turbofan engine. Though VIPR Phase II ground based aircraft installed engine testing wherein a multitude of research sensors and methods were evaluated, an array of acoustic microphones was used to determine the viability of such an array to detect failures occurring in a commercially representative high bypass turbofan engine. The failures introduced during VIPR testing included commanding the engine's low pressure compressor (LPC) exit and high pressure compressor (HPC) 14th stage bleed values abruptly to their failsafe positions during steady state

Dispersion of turbojet engine exhaust in flight

NASA Technical Reports Server (NTRS)

Holdeman, J. D.

1973-01-01

The dispersion of the exhaust of turbojet engines into the atmosphere is estimated by using a model developed for the mixing of a round jet with a parallel flow. The analysis is appropriate for determining the spread and dilution of the jet exhaust from the engine exit until it is entrained in the aircraft trailing vortices. Chemical reactions are not expected to be important and are not included in the flow model. Calculations of the dispersion of the exhaust plumes of three aircraft turbojet engines with and without afterburning at typical flight conditions are presented. Calculated average concentrations for the exhaust plume from a single engine jet fighter are shown to be in good agreement with measurements made in the aircraft wake during flight.

Prospects for Nonlinear Laser Diagnostics in the Jet Noise Laboratory

NASA Technical Reports Server (NTRS)

Herring, Gregory C.; Hart, Roger C.; Fletcher, mark T.; Balla, R. Jeffrey; Henderson, Brenda S.

2007-01-01

Two experiments were conducted to test whether optical methods, which rely on laser beam coherence, would be viable for off-body flow measurement in high-density, compressible-flow wind tunnels. These tests measured the effects of large, unsteady density gradients on laser diagnostics like laser-induced thermal acoustics (LITA). The first test was performed in the Low Speed Aeroacoustics Wind Tunnel (LSAWT) of NASA Langley Research Center's Jet Noise Laboratory (JNL). This flow facility consists of a dual-stream jet engine simulator (with electric heat and propane burners) exhausting into a simulated flight stream, reaching Mach numbers up to 0.32. A laser beam transited the LSAWT flow field and was imaged with a high-speed gated camera to measure beam steering and transverse mode distortion. A second, independent test was performed on a smaller laboratory jet (Mach number < 1.2 and mass flow rate < 0.1 kg/sec). In this test, time-averaged LITA velocimetry and thermometry were performed at the jet exit plane, where the effect of unsteady density gradients is observed on the LITA signal. Both experiments show that LITA (and other diagnostics relying on beam overlap or coherence) faces significant hurdles in the high-density, compressible, and turbulent flow environments similar to those of the JNL.

Multispectral imaging of a space shuttle primary reaction control system firing

NASA Astrophysics Data System (ADS)

Rall, David L. A.; Kofsky, Irving L.; Viereck, Rodney A.; Pike, Charles P.

1996-11-01

A series of three-second firings of Space Shuttle Orbiter's 870-lbf Primary Reaction Control System thruster motors were photographed from the crew cabin with an intensified video camera. The spectral imager sequentially recorded 4 ms exposures at 30 Hz in six 20 to 30 nm FWHM channels centered from 400 to 800 nm, chosen specifically to study bi- propellant (monomethyl hydrazine fuel/nitrogen dioxide oxidizer) thruster exhaust chemistry. The species producing the visible radiance were earlier identified as CN, CH, C2, NO2, and HNO; the electronic bands originating from the same excited states of CN (B-X) and CH (A-X) extend into the near UV. Images of the vacuum core viewing within a few degrees of perpendicular to the first several meters from the exit plane were analyzed to relate the spatial distribution of exhaust product species and afterburning chemistry to a flowfield-kinetics model. Profiles of radiance transverse to the exhaust symmetry-axis show substantial limb brightening in all six channels, indicating that the distribution of the radiating species corresponds to a `zone'-type model of liquid-fuel film-cooled engine performance. Profiles of band radiance along the axis indicate the production and quenching of excited species as the exhaust gas adiabatically expands and cools.

Data Quality Assurance for Supersonic Jet Noise Measurements

NASA Technical Reports Server (NTRS)

Brown, Clifford A.; Henderson, Brenda S.; Bridges, James E.

2010-01-01

The noise created by a supersonic aircraft is a primary concern in the design of future high-speed planes. The jet noise reduction technologies required on these aircraft will be developed using scale-models mounted to experimental jet rigs designed to simulate the exhaust gases from a full-scale jet engine. The jet noise data collected in these experiments must accurately predict the noise levels produced by the full-scale hardware in order to be a useful development tool. A methodology has been adopted at the NASA Glenn Research Center s Aero-Acoustic Propulsion Laboratory to insure the quality of the supersonic jet noise data acquired from the facility s High Flow Jet Exit Rig so that it can be used to develop future nozzle technologies that reduce supersonic jet noise. The methodology relies on mitigating extraneous noise sources, examining the impact of measurement location on the acoustic results, and investigating the facility independence of the measurements. The methodology is documented here as a basis for validating future improvements and its limitations are noted so that they do not affect the data analysis. Maintaining a high quality jet noise laboratory is an ongoing process. By carefully examining the data produced and continually following this methodology, data quality can be maintained and improved over time.

Design and Fabrication of Oxygen/RP-2 Multi-Element Oxidizer-Rich Staged Combustion Thrust Chamber Injectors

NASA Technical Reports Server (NTRS)

Garcia, C. P.; Medina, C. R.; Protz, C. S.; Kenny, R. J.; Kelly, G. W.; Casiano, M. J.; Hulka, J. R.; Richardson, B. R.

2016-01-01

As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center was contracted to assemble and hot-fire test a multi-element integrated test article demonstrating combustion characteristics of an oxygen/hydrocarbon propellant oxidizer-rich staged-combustion engine thrust chamber. Such a test article simulates flow through the main injectors of oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines, or future U.S.-built engine systems such as the Aerojet-Rocketdyne AR-1 engine or the Hydrocarbon Boost program demonstration engine. On the current project, several configurations of new main injectors were considered for the thrust chamber assembly of the integrated test article. All the injector elements were of the gas-centered swirl coaxial type, similar to those used on the Russian oxidizer-rich staged-combustion rocket engines. In such elements, oxidizer-rich combustion products from the preburner/turbine exhaust flow through a straight tube, and fuel exiting from the combustion chamber and nozzle regenerative cooling circuits is injected near the exit of the oxidizer tube through tangentially oriented orifices that impart a swirl motion such that the fuel flows along the wall of the oxidizer tube in a thin film. In some elements there is an orifice at the inlet to the oxidizer tube, and in some elements there is a sleeve or "shield" inside the oxidizer tube where the fuel enters. In the current project, several variations of element geometries were created, including element size (i.e., number of elements or pattern density), the distance from the exit of the sleeve to the injector face, the width of the gap between the oxidizer tube inner wall and the outer wall of the sleeve, and excluding the sleeve entirely. This paper discusses the design rationale for each of these element variations, including hydraulic, structural, thermal, combustion performance, and combustion stability considerations. This paper also discusses the fabrication and assembly of the injector components, including the injector body/interpropellant plate, the additive manufactured GRCop-84 faceplate, and the pieces that make up the injector elements including the oxidizer tube, an inlet to the oxidizer tube, and a facenut that includes the fuel tangential inlets and forms the initial recessed volume where oxidizer and fuel first interact. Hot-fire test results of these main injector designs in an integrated test article that includes an oxidizer-rich preburner are described in companion papers at this JANNAF meeting.

Experimental and Numerical Examination of a Hall Thruster Plume (Preprint)

DTIC Science & Technology

2007-07-31

Hall thruster has been characterized through measurements from various plasma electrostatic probes. Ion current flux, plasma potential, plasma density, and electron temperatures were measured from the near-field plume to 60 cm downstream of the exit plane. These experimentally derived measurements were compared to numerical simulations run with the plasma plume code DRACO. A major goal of this study was to determine the fidelity of the DRACO numerical simulation. The effect of background pressure on the thruster plume was also examined using ion current flux measurements

Hall Thruster With an External Acceleration Zone

DTIC Science & Technology

2005-09-14

Hall Thruster in a high vacuum environment. The ionized propellant velocities were measured using laser induced fluorescence of the excited state xenon ionic transition at 834.7 nm. Ion velocities were interrogated from the channel exit plane to a distance 30 mm from it. Both axial and cross-field (along the electron Hall current direction) velocities were measured. The results presented here, combined with those of previous work, highlight the high sensitivity of electron mobility inside and outside the channel, depending on the background gas density, type of wall

Ion Velocity Measurements in a Linear Hall Thruster (Postprint)

DTIC Science & Technology

2005-06-14

Hall Thruster in a high vacuum environment. The ionized propellant velocities were measured using laser induced fluorescence of the excited state xenon ionic transition at 834.7 nm. Ion velocities were interrogated from the channel exit plane to a distance 30 mm from it. Both axial and cross-field (along the electron Hall current direction) velocities were measured. The results presented here, combined with those of previous work, highlight the high sensitivity of electron mobility inside and outside the channel, depending on the background gas density, type of wall

KSC-08pd0658

NASA Image and Video Library

2008-03-08

KENNEDY SPACE CENTER, FLA. -- The crew of space shuttle Endeavour's STS-123 mission arrive at NASA Kennedy Space Center's Shuttle Landing Facility for launch at 2:28 a.m. EDT on March 11. Exiting the plane are Pilot Gregory H. Johnson, followed by Mission Specialists Rick Linnehan and Takao Doi of the Japan Aerospace Exploration Agency. On this mission to the International Space Station, Endeavour and its crew will deliver the first section of the Japan Aerospace Exploration Agency's Kibo laboratory and the Canadian Space Agency's two-armed robotic system, Dextre. Photo credit: NASA/Kim Shiflett

Real time Faraday spectrometer

DOEpatents

Smith, Jr., Tommy E.; Struve, Kenneth W.; Colella, Nicholas J.

1991-01-01

This invention uses a dipole magnet to bend the path of a charged particle beam. As the deflected particles exit the magnet, they are spatially dispersed in the bend-plane of the magnet according to their respective momenta and pass to a plurality of chambers having Faraday probes positioned therein. Both the current and energy distribution of the particles is then determined by the non-intersecting Faraday probes located along the chambers. The Faraday probes are magnetically isolated from each other by thin metal walls of the chambers, effectively providing real time current-versus-energy particle measurements.

Design study of an advanced gas generator. [which can be ignited during start-up period of turbine engines

NASA Technical Reports Server (NTRS)

Kim, S.; Trinh, H. P.

1993-01-01

A gas generator which can be ignited reliably during the initial start-up period and offers fairly uniform gas temperature at the exit was studied numerically. Various sizes and shapes of the mixing enhancement devices and their positions were examined to evaluate the uniformity of the exit gas temperature and the change of internal pressure drop incurred by introducing the mixing enhancement devices. By introducing a turbulence ring and a splash plate with an appropriate size and position, it was possible to obtain fairly uniform gas temperature distributions and a maximum gas temperature that is within the design limit temperature of 1600 R at the generator exit. However, with the geometry studied, the pressure drop across the generator was great, approximately 1150 psi, to satisfy the assigned design limit temperature. If the design limit temperature is increased to 1650 R, the pressure drop across the generator could be lowered by as much as 350 psi.

Optimal trajectories for an aerospace plane. Part 1: Formulation, results, and analysis

NASA Technical Reports Server (NTRS)

Miele, Angelo; Lee, W. Y.; Wu, G. D.

1990-01-01

The optimization of the trajectories of an aerospace plane is discussed. This is a hypervelocity vehicle capable of achieving orbital speed, while taking off horizontally. The vehicle is propelled by four types of engines: turbojet engines for flight at subsonic speeds/low supersonic speeds; ramjet engines for flight at moderate supersonic speeds/low hypersonic speeds; scramjet engines for flight at hypersonic speeds; and rocket engines for flight at near-orbital speeds. A single-stage-to-orbit (SSTO) configuration is considered, and the transition from low supersonic speeds to orbital speeds is studied under the following assumptions: the turbojet portion of the trajectory has been completed; the aerospace plane is controlled via the angle of attack and the power setting; the aerodynamic model is the generic hypersonic aerodynamics model example (GHAME). Concerning the engine model, three options are considered: (EM1), a ramjet/scramjet combination in which the scramjet specific impulse tends to a nearly-constant value at large Mach numbers; (EM2), a ramjet/scramjet combination in which the scramjet specific impulse decreases monotonically at large Mach numbers; and (EM3), a ramjet/scramjet/rocket combination in which, owing to stagnation temperature limitations, the scramjet operates only at M approx. less than 15; at higher Mach numbers, the scramjet is shut off and the aerospace plane is driven only by the rocket engines. Under the above assumptions, four optimization problems are solved using the sequential gradient-restoration algorithm for optimal control problems: (P1) minimization of the weight of fuel consumed; (P2) minimization of the peak dynamic pressure; (P3) minimization of the peak heating rate; and (P4) minimization of the peak tangential acceleration.

Design of a Mach-3 Nozzle for TBCC Testing in the NASA LaRC 8-ft High Temperature Tunnel

NASA Technical Reports Server (NTRS)

Gaffney, Richard L., Jr.; Norris, Andrew T.

2008-01-01

A new nozzle is being constructed for the NASA Langley Research Center 8-Foot High Temperature Tunnel. The axisymmetric nozzle was designed with a Mach-3 exit flow for testing Turbine-Based Combined-Cycle engines at a Mach number in the vicinity of the transition from turbojet to ramjet operation. The nozzle contour was designed using the NASA Langley IMOCND computer program which solves the potential equation using the classical method of characteristics. To include viscous effects, the design procedure iterated the MOC contour generation with CFD Navier-Stokes calculations, adjusting MOC input parameters until target nozzle-exit conditions were achieved in the Navier-Stokes calculations. The design process was complicated by a requirement to use the final 29.5 inches of an existing 54.5-inch exit-diameter Mach-5 nozzle contour. This was accomplished by generating a Mach-3 contour that matched the radius of the Mach-5 contour at the match point and using a 3rd order polynomial to create a smooth transition between the two contours. During the final evaluation of the design it was realized that the throat diameter is more than half that of the upstream mixing chamber. This led to the concern that large vortical structures generated in the mixer would persist downstream, affecting nozzle-exit flow. This concern was addressed by analyzing the results of three-dimensional, viscous, numerical simulations of the entire flowfield, from the exit of the facility combustor to the nozzle exit. An analysis of the solution indicated that large scale structures do not pass through the throat and that both the total temperature and species (CO2) are well mixed in the mixer, providing uniform flow to the nozzle and subsequently the test cabin.

TF34 convertible engine control system design

NASA Technical Reports Server (NTRS)

Gilmore, D. R., Jr.

1984-01-01

The characteristics of the TF34 convertible engine, capable of producing shaft power, thrust, or a combination of both, is investigated with respect to the control system design, development, bench testing, and the anticipated transient response during engine testing at NASA. The modifications to the prototype standard TF34-GE-400 turbofan, made primarily in the fan section, consist of the variable inlet guide vanes and variable exit guide vanes. The control system was designed using classical frequency domain techniques and was based on the anticipated convertible/VTOL airframe requirements. The engine has been run in the fan mode and in the shaft mode, exhibiting a response of 0.14 second to a 5-percent thrust change.

The influence of the stagnation zone on the fluid dynamics at the nozzle exit of a confined and submerged impinging jet

NASA Astrophysics Data System (ADS)

Jeffers, Nicholas; Stafford, Jason; Conway, Ciaran; Punch, Jeff; Walsh, Edmond

2016-02-01

Low profile impinging jets provide a means to achieve high heat transfer coefficients while occupying a small quantity of space. Consequently, they are found in many engineering applications such as electronics cooling, annealing of metals, food processing, and others. This paper investigates the influence of the stagnation zone fluid dynamics on the nozzle exit flow condition of a low profile, submerged, and confined impinging water jet. The jet was geometrically constrained to a round, 16-mm diameter, square-edged nozzle at a jet exit to target surface spacing ( H/ D) that varied between 0.25 < {{ H}{/}{ D}} < 8.75. The influence of turbulent flow regimes is the main focus of this paper; however, laminar flow data are also presented between 1350 < Re < 17{,}300. A custom measurement facility was designed and commissioned to utilise particle image velocimetry in order to quantitatively measure the fluid dynamics both before and after the jet exits its nozzle. The velocity profiles are normalised with the mean velocity across the nozzle exit, and turbulence statistics are also presented. The primary objective of this paper is to present accurate flow profiles across the nozzle exit of an impinging jet confined to a low H/ D, with a view to guide the boundary conditions chosen for numerical simulations confined to similar constraints. The results revealed in this paper suggest that the fluid dynamics in the stagnation zone strongly influences the nozzle exit velocity profile at confinement heights between 0 < {{ H}{/}{ D}} < 1. This is of particular relevance with regard to the choice of inlet boundary conditions in numerical models, and it was found that it is necessary to model a jet tube length {{ L}{/}{ D}} > 0.5—where D is the inner diameter of the jet—in order to minimise modelling uncertainty.

Electrostatic dispersion lenses and ion beam dispersion methods

DOEpatents

Dahl, David A [Idaho Falls, ID; Appelhans, Anthony D [Idaho Falls, ID

2010-12-28

An EDL includes a case surface and at least one electrode surface. The EDL is configured to receive through the EDL a plurality of ion beams, to generate an electrostatic field between the one electrode surface and either the case surface or another electrode surface, and to increase the separation between the beams using the field. Other than an optional mid-plane intended to contain trajectories of the beams, the electrode surface or surfaces do not exhibit a plane of symmetry through which any beam received through the EDL must pass. In addition or in the alternative, the one electrode surface and either the case surface or the other electrode surface have geometries configured to shape the field to exhibit a less abrupt entrance and/or exit field transition in comparison to another electrostatic field shaped by two nested, one-quarter section, right cylindrical electrode surfaces with a constant gap width.

Experimental and computational investigation of the NASA Low-Speed Centrifugal Compressor flow field

NASA Technical Reports Server (NTRS)

Hathaway, M. D.; Chriss, R. M.; Wood, J. R.; Strazisar, A. J.

1992-01-01

An experimental and computational investigation of the NASA Low-Speed Centrifugal Compressor (LSCC) flow field has been conducted using laser anemometry and Dawes' 3D viscous code. The experimental configuration consists of a backswept impeller followed by a vaneless diffuser. Measurements of the three-dimensional velocity field were acquired at several measurement planes through the compressor. The measurements describe both the throughflow and secondary velocity field along each measurement plane. In several cases the measurements provide details of the flow within the blade boundary layers. Insight into the complex flow physics within centrifugal compressors is provided by the computational analysis, and assessment of the CFD predictions is provided by comparison with the measurements. Five-hole probe and hot-wire surveys at the inlet and exit to the rotor as well as surface flow visualization along the impeller blade surfaces provide independent confirmation of the laser measurement technique.

Fuel-optimal trajectories of aeroassisted orbital transfer with plane change

NASA Technical Reports Server (NTRS)

Naidu, Desineni Subbaramaiah; Hibey, Joseph L.

1989-01-01

The problem of minimization of fuel consumption during the atmospheric portion of an aeroassisted, orbital transfer with plane change is addressed. The complete mission has required three characteristic velocities, a deorbit impulse at high earth orbit (HEO), a boost impulse at the atmospheric exit, and a reorbit impulse at low earth orbit (LEO). A performance index has been formulated as the sum of these three impulses. Application of optimal control principles has led to a nonlinear, two-point, boundary value problem which was solved by using a multiple shooting algorithm. The strategy for the atmospheric portion of the minimum-fuel transfer is to start initially with the maximum positive lift in order to recover from the downward plunge, and then to fly with a gradually decreasing lift such that the vehicle skips out of the atmosphere with a flight path angle near zero degrees.

WATER PUMP HOUSE, TRA619. VIEW OF PUMP HOUSE UNDER CONSTRUCTION. ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

WATER PUMP HOUSE, TRA-619. VIEW OF PUMP HOUSE UNDER CONSTRUCTION. CAMERA IS ON WATER TOWER AND FACES NORTHWEST. TWO RESERVOIR TANKS ALREADY ARE COMPLETED. NOTE EXCAVATIONS FOR PIPE LINES EXITING FROM BELOW GROUND ON SOUTH SIDE OF PUMP HOUSE. BUILDING AT LOWER RIGHT IS ELECTRICAL CONTROL BUILDING, TRA-623. SWITCHYARD IS IN LOWER RIGHT CORNER OF VIEW. INL NEGATIVE NO. 2753. Unknown Photographer, ca. 6/1951 - Idaho National Engineering Laboratory, Test Reactor Area, Materials & Engineering Test Reactors, Scoville, Butte County, ID

Expedition 37 Soyuz Landing Preparation

NASA Image and Video Library

2013-11-10

A member of Russian search and rescue exits a helicopter moments after landing at Zhezkazgan airport in Kazakhstan, Sunday, Nov. 10, 2013, a day ahead of the scheduled landing of the Soyuz TMA-09M spacecraft with Expedition 37 Commander Fyodor Yurchikhin of the Russian Federal Space Agency (Roscosmos), Flight Engineer Karen Nyberg of NASA and Flight Engineer Luca Parmitano of the European Space Agency. Yurchikhin, Nyberg and Parmitano are returning to Earth after five and a half months on the International Space Station. Photo Credit: (NASA/Carla Cioffi)

The Aerothermodynamics of Aircraft Gas Turbine Engines

DTIC Science & Technology

1978-07-01

engine will deteriorate. 1.6.2 Experimental Testing It is easy to fall int9 the organiza- tional trap of four isolated groups . One group does the... Quasi -Dne-Dimensional Fluid Flows The First Law for a F1mdng System-- The Control Volume • . • The Channel Flow Equations Stagnation Properties...exit to control volume (Eq • 2. 14 . 2) CHAPTER TWO THERMODYNAMICS AND QUASI -ONE-DUlENSIONAL FLUID FLO’’{S 2.0 INTRODUCTION This chapter "ill be

Exhaust gas bypass valve control for thermoelectric generator

DOEpatents

Reynolds, Michael G; Yang, Jihui; Meisner, Greogry P.; Stabler, Francis R.; De Bock, Hendrik Pieter Jacobus; Anderson, Todd Alan

2012-09-04

A method of controlling engine exhaust flow through at least one of an exhaust bypass and a thermoelectric device via a bypass valve is provided. The method includes: determining a mass flow of exhaust exiting an engine; determining a desired exhaust pressure based on the mass flow of exhaust; comparing the desired exhaust pressure to a determined exhaust pressure; and determining a bypass valve control value based on the comparing, wherein the bypass valve control value is used to control the bypass valve.

HOT CELL BUILDING, TRA632, INTERIOR. CONTEXTUAL VIEW OF HOT CELL ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

HOT CELL BUILDING, TRA-632, INTERIOR. CONTEXTUAL VIEW OF HOT CELL NO. 2 FROM STAIRWAY ALONG NORTH WALL. OBSERVATION WINDOW ALONG WEST SIDE BENEATH "CELL 2" SIGN. DOORWAY IN LEFT OF VIEW LEADS TO CELL 1 WORK AREA OR TO EXIT OUTDOORS TO NORTH. RADIATION DETECTION MONITOR TO RIGHT OF DOOR. CAMERA FACING SOUTHWEST. INL NEGATIVE NO. HD46-28-3. Mike Crane, Photographer, 2/2005 - Idaho National Engineering Laboratory, Test Reactor Area, Materials & Engineering Test Reactors, Scoville, Butte County, ID

DEVELOPMENT OF A SUPERSONIC TRANSPORT AIRCRAFT ENGINE - PHASE II-A.

DTIC Science & Technology

JET TRANSPORT PLANES, *SUPERSONIC AIRCRAFT ) (U) TURBOJET ENGINES , PERFORMANCE( ENGINEERING ), TURBOFAN ENGINES , AFTERBURNING, SPECIFICATIONS...COMPRESSORS, GEOMETRY, TURBOJET INLETS, COMBUSTION, TEST EQUIPMENT, TURBINE BLADES , HEAT TRANSFER, AIRFOILS , CASCADE STRUCTURES, EVAPOTRANSPIRATION, PLUG NOZZLES, ANECHOIC CHAMBERS, BEARINGS, SEALS, DESIGN, FATIGUE(MECHANICS)

Particle clustering within a two-phase turbulent pipe jet

NASA Astrophysics Data System (ADS)

Lau, Timothy; Nathan, Graham

2016-11-01

A comprehensive study of the influence of Stokes number on the instantaneous distributions of particles within a well-characterised, two-phase, turbulent pipe jet in a weak co-flow was performed. The experiments utilised particles with a narrow size distribution, resulting in a truly mono-disperse particle-laden jet. The jet Reynolds number, based on the pipe diameter, was in the range 10000 <= ReD <= 40000 , while the exit Stokes number was in the range 0 . 3 <= SkD <= 22 . 4 . The particle mass loading was fixed at ϕ = 0 . 4 , resulting in a flow that was in the two-way coupling regime. Instantaneous particle distributions within a two-dimensional sheet was measured using planar nephelometry while particle clusters were identified and subsequently characterised using an in-house developed technique. The results show that particle clustering is significantly influenced by the exit Stokes number. Particle clustering was found to be significant for 0 . 3 <= SkD <= 5 . 6 , with the degree of clustering increasing as SkD is decreased. The clusters, which typically appeared as filament-like structures with high aspect ratio oriented at oblique angles to the flow, were measured right from the exit plane, suggesting that they were generated inside the pipe. The authors acknowledge the financial contributions by the Australian Research Council (Grant No. DP120102961) and the Australian Renewable Energy Agency (Grant No. USO034).

Three-dimensional displacement measurement of image point by point-diffraction interferometry

NASA Astrophysics Data System (ADS)

He, Xiao; Chen, Lingfeng; Meng, Xiaojie; Yu, Lei

2018-01-01

This paper presents a method for measuring the three-dimensional (3-D) displacement of an image point based on point-diffraction interferometry. An object Point-light-source (PLS) interferes with a fixed PLS and its interferograms are captured by an exit pupil. When the image point of the object PLS is slightly shifted to a new position, the wavefront of the image PLS changes. And its interferograms also change. Processing these figures (captured before and after the movement), the wavefront difference of the image PLS can be obtained and it contains the information of three-dimensional (3-D) displacement of the image PLS. However, the information of its three-dimensional (3-D) displacement cannot be calculated until the distance between the image PLS and the exit pupil is calibrated. Therefore, we use a plane-parallel-plate with a known refractive index and thickness to determine this distance, which is based on the Snell's law for small angle of incidence. Thus, since the distance between the exit pupil and the image PLS is a known quantity, the 3-D displacement of the image PLS can be simultaneously calculated through two interference measurements. Preliminary experimental results indicate that its relative error is below 0.3%. With the ability to accurately locate an image point (whatever it is real or virtual), a fiber point-light-source can act as the reticle by itself in optical measurement.

Wind tunnel test results of a 1/8-scale fan-in-wing model

NASA Technical Reports Server (NTRS)

Wilson, John C.; Gentry, Garl L.; Gorton, Susan A.

1996-01-01

A 1/8-scale model of a fan-in-wing concept considered for development by Grumman Aerospace Corporation for the U.S. Army was tested in the Langley 14- by 22-Foot Subsonic Tunnel. Hover testing, which included height above a pressure-instrumented ground plane, angle of pitch, and angle of roll for a range of fan thrust, was conducted in a model preparation area near the tunnel. The air loads and surface pressures on the model were measured for several configurations in the model preparation area and in the tunnel. The major hover configuration change was varying the angles of the vanes attached to the exit of the fans for producing propulsive force. As the model height above the ground was decreased, there was a significant variation of thrust-removed normal force with constant fan speed. The greatest variation was generally for the height-to-fan exit diameter ratio of less than 2.5; the variation was reduced by deflecting fan exit flow outboard with the vanes. In the tunnel angles of pitch and sideslip, height above the tunnel floor, and wind speed were varied for a range of fan thrust and different vane angle configurations. Other configuration features such as flap deflections and tail incidence were evaluated as well. Though the V-tail empennage provided an increase in static longitudinal stability, the total model configuration remained unstable.

Radial and circumferential flow surveys at the inlet and exit of the Space Shuttle Main Engine High Pressure Fuel Turbine Model

NASA Technical Reports Server (NTRS)

Hudson, S. T.; Bordelon, W. J., Jr.; Smith, A. W.; Ramachandran, N.

1995-01-01

The main objective of this test was to obtain detailed radial and circumferential flow surveys at the inlet and exit of the SSME High Pressure Fuel Turbine model using three-hole cobra probes, hot-film probes, and a laser velocimeter. The test was designed to meet several objectives. First, the techniques for making laser velocimeter, hot-film probe, and cobra probe measurements in turbine flows were developed and demonstrated. The ability to use the cobra probes to obtain static pressure and, therefore, velocity had to be verified; insertion techniques had to be established for the fragile hot-film probes; and a seeding method had to be established for the laser velocimetry. Once the measurement techniques were established, turbine inlet and exit velocity profiles, temperature profiles, pressure profiles, turbulence intensities, and boundary layer thicknesses were measured at the turbine design point. The blockage effect due to the model inlet and exit total pressure and total temperature rakes on the turbine performance was also studied. A small range of off-design points were run to obtain the profiles and to verify the rake blockage effects off-design. Finally, a range of different Reynolds numbers were run to study the effect of Reynolds number on the various measurements.

Design study of an air pump and integral lift engine ALF-504 using the Lycoming 502 core

NASA Technical Reports Server (NTRS)

Rauch, D.

1972-01-01

Design studies were conducted for an integral lift fan engine utilizing the Lycoming 502 fan core with the final MQT power turbine. The fan is designed for a 12.5 bypass ratio and 1.25:1 pressure ratio, and provides supercharging for the core. Maximum sea level static thrust is 8370 pounds with a specific fuel consumption of 0.302 lb/hr-lb. The dry engine weight without starter is 1419 pounds including full-length duct and sound-attenuating rings. The engine envelope including duct treatment but not localized accessory protrusion is 53.25 inches in diameter and 59.2 inches long from exhaust nozzle exit to fan inlet flange. Detailed analyses include fan aerodynamics, fan and reduction gear mechanical design, fan dynamic analysis, engine noise analysis, engine performance, and weight analysis.

Coupled-Flow Simulation of HP-LP Turbines Has Resulted in Significant Fuel Savings

NASA Technical Reports Server (NTRS)

Veres, Joseph P.

2001-01-01

Our objective was to create a high-fidelity Navier-Stokes computer simulation of the flow through the turbines of a modern high-bypass-ratio turbofan engine. The simulation would have to capture the aerodynamic interactions between closely coupled high- and low-pressure turbines. A computer simulation of the flow in the GE90 turbofan engine's high-pressure (HP) and low-pressure (LP) turbines was created at GE Aircraft Engines under contract with the NASA Glenn Research Center. The three-dimensional steady-state computer simulation was performed using Glenn's average-passage approach named APNASA. The areas upstream and downstream of each blade row mutually interact with each other during engine operation. The embedded blade row operating conditions are modeled since the average passage equations in APNASA actively include the effects of the adjacent blade rows. The turbine airfoils, platforms, and casing are actively cooled by compressor bleed air. Hot gas leaks around the tips of rotors through labyrinth seals. The flow exiting the high work HP turbines is partially transonic and, therefore, has a strong shock system in the transition region. The simulation was done using 121 processors of a Silicon Graphics Origin 2000 (NAS 02K) cluster at the NASA Ames Research Center, with a parallel efficiency of 87 percent in 15 hr. The typical average-passage analysis mesh size per blade row was 280 by 45 by 55, or approx.700,000 grid points. The total number of blade rows was 18 for a combined HP and LP turbine system including the struts in the transition duct and exit guide vane, which contain 12.6 million grid points. Design cycle turnaround time requirements ran typically from 24 to 48 hr of wall clock time. The number of iterations for convergence was 10,000 at 8.03x10(exp -5) sec/iteration/grid point (NAS O2K). Parallel processing by up to 40 processors is required to meet the design cycle time constraints. This is the first-ever flow simulation of an HP and LP turbine. In addition, it includes the struts in the transition duct and exit guide vanes.

Energy efficient engine: Turbine intermediate case and low-pressure turbine component test hardware detailed design report

NASA Technical Reports Server (NTRS)

Leach, K.; Thulin, R. D.; Howe, D. C.

1982-01-01

A four stage, low pressure turbine component has been designed to power the fan and low pressure compressor system in the Energy Efficient Engine. Designs for a turbine intermediate case and an exit guide vane assembly also have been established. The components incorporate numerous technology features to enhance efficiency, durability, and performance retention. These designs reflect a positive step towards improving engine fuel efficiency on a component level. The aerodynamic and thermal/mechanical designs of the intermediate case and low pressure turbine components are presented and described. An overview of the predicted performance of the various component designs is given.

Preliminary Results From a Heavily Instrumented Engine Ice Crystal Icing Test in a Ground Based Altitude Test Facility

NASA Technical Reports Server (NTRS)

Flegel, Ashlie B.; Oliver, Michael J.

2016-01-01

Preliminary results from the Heavily Instrumented ALF503R-5 Engine test conducted in the NASA Glenn Research Center Propulsion Systems Laboratory will be discussed. The effects of ice crystal icing on a full scale engine is examined and documented. This model engine, serial number LF01, was used during the inaugural icing test in the PSL facility. The reduction of thrust (rollback) events experienced by this engine in flight were replicated in the facility. Limited instrumentation was used to detect icing. Metal temperature on the exit guide vanes and outer shroud and the load measurement were the only indicators of ice formation. The current study features a similar engine, serial number LF11, which is instrumented to characterize the cloud entering the engine, detect characterize ice accretion, and visualize the ice accretion in the region of interest.

Evaluation of innovative rocket engines for single-stage earth-to-orbit vehicles

NASA Astrophysics Data System (ADS)

Manski, Detlef; Martin, James A.

1988-07-01

Computer models of rocket engines and single-stage-to-orbit vehicles that were developed by the authors at DFVLR and NASA have been combined. The resulting code consists of engine mass, performance, trajectory and vehicle sizing models. The engine mass model includes equations for each subsystem and describes their dependences on various propulsion parameters. The engine performance model consists of multidimensional sets of theoretical propulsion properties and a complete thermodynamic analysis of the engine cycle. The vehicle analyses include an optimized trajectory analysis, mass estimation, and vehicle sizing. A vertical-takeoff, horizontal-landing, single-stage, winged, manned, fully reusable vehicle with a payload capability of 13.6 Mg (30,000 lb) to low earth orbit was selected. Hydrogen, methane, propane, and dual-fuel engines were studied with staged-combustion, gas-generator, dual bell, and the dual-expander cycles. Mixture ratio, chamber pressure, nozzle exit pressure liftoff acceleration, and dual fuel propulsive parameters were optimized.

Evaluation of innovative rocket engines for single-stage earth-to-orbit vehicles

NASA Technical Reports Server (NTRS)

Manski, Detlef; Martin, James A.

1988-01-01

Computer models of rocket engines and single-stage-to-orbit vehicles that were developed by the authors at DFVLR and NASA have been combined. The resulting code consists of engine mass, performance, trajectory and vehicle sizing models. The engine mass model includes equations for each subsystem and describes their dependences on various propulsion parameters. The engine performance model consists of multidimensional sets of theoretical propulsion properties and a complete thermodynamic analysis of the engine cycle. The vehicle analyses include an optimized trajectory analysis, mass estimation, and vehicle sizing. A vertical-takeoff, horizontal-landing, single-stage, winged, manned, fully reusable vehicle with a payload capability of 13.6 Mg (30,000 lb) to low earth orbit was selected. Hydrogen, methane, propane, and dual-fuel engines were studied with staged-combustion, gas-generator, dual bell, and the dual-expander cycles. Mixture ratio, chamber pressure, nozzle exit pressure liftoff acceleration, and dual fuel propulsive parameters were optimized.

National Aerospace Plane Engine Seals: High Temperature Seal Performance Evaluation

NASA Technical Reports Server (NTRS)

Steinetz, Bruce M.

1991-01-01

The key to the successful development of the single stage to orbit National Aerospace Plane (NASP) is the successful development of combined cycle ramjet/scramjet engines that can propel the vehicle to 17,000 mph to reach low Earth orbit. To achieve engine performance over this speed range, movable engine panels are used to tailor engine flow that require low leakage, high temperature seals around their perimeter. NASA-Lewis is developing a family of new high temperature seals to form effective barriers against leakage of extremely hot (greater than 2000 F), high pressure (up to 100 psi) flow path gases containing hydrogen and oxygen. Preventing backside leakage of these explosive gas mixtures is paramount in preventing the potential loss of the engine or the entire vehicle. Seal technology development accomplishments are described in the three main areas of concept development, test, and evaluation and analytical development.

Wave Engine Technology Development

DTIC Science & Technology

1984-01-01

were the usual minor but time consuming problems of converting a program to run on a new computer with a new operating system and Fortran compiler...Exit Port. - - I _ _- i - - ~ = _ _ o71 - .. (I 00 kfC ) C: 4 03 \\. ft~ d) Ix- 3:- 0r i lzz 𔃾 14- Wave Field 81 and the associated port printouts are

34. Photographic copy of photograph (ca. 1962, original print in ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

34. Photographic copy of photograph (ca. 1962, original print in possession of Army Corps of Engineers, Ft. Belvoir, Virginia) Photographer unknown. View of launch control facility under construction, security gate at left - Ellsworth Air Force Base, Delta Flight, 10 mile radius around Exit 127 off Interstate 90, Interior, Jackson County, SD

New Research on the Cowling and Cooling of Radial Engines

NASA Technical Reports Server (NTRS)

Molloy, Richard C.; Brewster, James H., III

1943-01-01

An extensive series of wind-tunnel tests on a half-scale conventional, nacelle model were made by the United Aircraft Corporation to determine and correlate the effects of many variables on cooling air flow and nacelle drag. The primary investigation was concerned with the reaction of these factors to varying conditions ahead of, across, and behind the engine. In the light of this investigation, common misconceptions and factors which are frequently overlooked in the cooling and cowling of radial engines are considered in some detail. Data are presented to support certain design recommendations and conclusions which should lead toward the improvement of present engine installations. Several charts are included to facilitate the estimation of cooling drag, available cooling pressure, and cowl exit area.

Pumping Performance or RBCC Engine under Sea Level Static Condition

NASA Astrophysics Data System (ADS)

Kouchi, Toshinori; Tomioka, Sadatake; Kanda, Takeshi

Numerical simulations were conducted to predict the ejector pumping performance of a rocket-ramjet combined-cycle engine under a take-off condition. The numerical simulations revealed that the suction airflow was chocked at the exit of the engine throat when the ejector rocket was driven by cold N2 gas at the chamber pressure of 3MPa. When the ejector-driving gas was changed from cold N2 gas to hot combustion gas, the suction performance decreased remarkably. Mach contours in the engine revealed that the rocket plume constricted when the driving gas was the hot combustion gas. The change of the area of the stream tube area seemed to induce the pressure rise in the duct and decreasing in the pumping performance.

Hybrid bandgap engineering for super-hetero-epitaxial semiconductor materials, and products thereof

NASA Technical Reports Server (NTRS)

Park, Yeonjoon (Inventor); Choi, Sang H. (Inventor); King, Glen C. (Inventor); Elliott, James R. (Inventor)

2012-01-01

"Super-hetero-epitaxial" combinations comprise epitaxial growth of one material on a different material with different crystal structure. Compatible crystal structures may be identified using a "Tri-Unity" system. New bandgap engineering diagrams are provided for each class of combination, based on determination of hybrid lattice constants for the constituent materials in accordance with lattice-matching equations. Using known bandgap figures for previously tested materials, new materials with lattice constants that match desired substrates and have the desired bandgap properties may be formulated by reference to the diagrams and lattice matching equations. In one embodiment, this analysis makes it possible to formulate new super-hetero-epitaxial semiconductor systems, such as systems based on group IV alloys on c-plane LaF.sub.3; group IV alloys on c-plane langasite; Group III-V alloys on c-plane langasite; and group II-VI alloys on c-plane sapphire.

Application of the scalar and vector potentials to the aerodynamics of jets

NASA Technical Reports Server (NTRS)

Russell, H. L.; Skifstad, J. G.

1973-01-01

The applicability of a method based on the Stokes potentials (vector and scalar potentials) to computations associated with the aerodynamics of jets was examined. The aerodynamic field near the nozzle could be represented and that the influence of a nonuniform velocity profile at the nozzle exit plane could be determined. Also computations were made for an axisymmetric jet exhausting into a quiescient atmosphere. The velocity at the axis of the jet, and the location of the half-velocity points along the jet yield accurate aerodynamic field computations. Inconsistencies among the different theoretical characterizations of jet flowfields are shown.

Analysis of SRB reentry acoustic environments. [aeroacoustic spectra determined from wind tunnel tests

NASA Technical Reports Server (NTRS)

Coffin, T.; Dandridge, R. E.; Haddock, U. W.

1979-01-01

Space shuttle solid rocket booster reentry aeroacoustic environments were estimated. Particular emphasis was given to the aft skirt/exit plane region for the Mach number regime 0.6 = or greater than M infinity = or less than 3.5. The analysis is based on the evaluation of wind tunnel model results in conjunction with Monte Carlo simulation of trajectory parameters. The experimental approach is described as well as the evaluation process utilized. Predicted environments are presented in terms of one-third octave band spectra representing space averaged values for critical regions on the solid rocket booster.

ExB Measurements of a 200 W Xenon Hall Thruster (Preprint)

DTIC Science & Technology

2007-08-28

Hall thruster Busek BHT-200 plume were measured using an ExB probe under a variety of thruster operating conditions and background pressures. The thruster was operated at several operating conditions by varying the anode potential of the thruster from 200 V to 325 V in 25 V increments. Measurements of the ion species fractions were made 90 from thruster centerline 60 cm downstream of the exit plane. At reduced discharge voltages, the species fractions of multiply-charged xenon ions were lower, while at increased discharge voltages, Xe+2 and Xe+3 showed an increase in their

Turbofan forced mixer-nozzle internal flowfield. Volume 1: A benchmark experimental study

NASA Technical Reports Server (NTRS)

Paterson, R. W.

1982-01-01

An experimental investigation of the flow field within a model turbofan forced mixer nozzle is described. Velocity and thermodynamic state variable data for use in assessing the accuracy and assisting the further development of computational procedures for predicting the flow field within mixer nozzles are provided. Velocity and temperature data suggested that the nozzle mixing process was dominated by circulations (secondary flows) of a length scale on the order the lobe dimensions which were associated with strong radial velocities observed near the lobe exit plane. The 'benchmark' model mixer experiment conducted for code assessment purposes is discussed.

Experimental performance of three design factors for ventral nozzles for SSTOVL aircraft

NASA Technical Reports Server (NTRS)

Esker, Barbara S.; Perusek, Gail P.

1992-01-01

An experimental study of three variations of a ventral nozzle system for supersonic short-takeoff and vertical-landing (SSTOVL) aircraft was performed at the NASA LeRC Powered Lift Facility. These test results include the effects of an annular duct flow into the ventral duct, a blocked tailpipe, and a short ventral duct length. An analytical study was also performed on the short ventral duct configuration using the PARC3D computational dynamics code. Data presented include pressure losses, thrust and flow performance, internal flow visualization, and pressure distributions at the exit plane of the ventral nozzle.

Aerodynamic Effects of Turbulence Intensity on a Variable-Speed Power-Turbine Blade with Large Incidence and Reynolds Number Variations

NASA Technical Reports Server (NTRS)

Flegel, Ashlie Brynn; Giel, Paul W.; Welch, Gerard E.

2014-01-01

The effects of inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The high turbulence study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Downstream total pressure and exit angle data were acquired for ten incidence angles ranging from +15.8 to 51.0. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.12105 to 2.12106 and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 0.25 - 0.4 for the low Tu tests and 8- 15 for the high Tu study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitchyaw probe located in a survey plane 7 axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the turbulence grid imposes.

A Design Study of Onboard Navigation and Guidance During Aerocapture at Mars. M.S. Thesis

NASA Technical Reports Server (NTRS)

Fuhry, Douglas Paul

1988-01-01

The navigation and guidance of a high lift-to-drag ratio sample return vehicle during aerocapture at Mars are investigated. Emphasis is placed on integrated systems design, with guidance algorithm synthesis and analysis based on vehicle state and atmospheric density uncertainty estimates provided by the navigation system. The latter utilizes a Kalman filter for state vector estimation, with useful update information obtained through radar altimeter measurements and density altitude measurements based on IMU-measured drag acceleration. A three-phase guidance algorithm, featuring constant bank numeric predictor/corrector atmospheric capture and exit phases and an extended constant altitude cruise phase, is developed to provide controlled capture and depletion of orbital energy, orbital plane control, and exit apoapsis control. Integrated navigation and guidance systems performance are analyzed using a four degree-of-freedom computer simulation. The simulation environment includes an atmospheric density model with spatially correlated perturbations to provide realistic variations over the vehicle trajectory. Navigation filter initial conditions for the analysis are based on planetary approach optical navigation results. Results from a selection of test cases are presented to give insight into systems performance.

Analytical and Experimental Study of Flow Through an Axial Turbine Stage with a Nonuniform Inlet Radial Temperature Profile

NASA Technical Reports Server (NTRS)

Schwab, J. R.; Stabe, R. G.; Whitney, W. J.

1983-01-01

Results are presented for a typical nonuniform inlet radial temperature profile through an advanced single-stage axial turbine and compared with the results obtained for a uniform profile. Gas temperature rises of 40 K to 95 K are predicted at the hub and tip corners at the trailing edges of the pressure surfaces in both the stator and rotor due to convection of hot fluid from the mean by the secondary flow. The inlet temperature profile is shown to be mixed out at the rotor exit survey plane (2.3 axial chords downstream of the rotor trailing edge) in both the analysis and the experiment. The experimental rotor exit angle profile for the nonuniform inlet temperature profile indicates underturning at the tip caused by increased clearance. Severe underturning also occurs at the mean, both with and without the nonuniform inlet temperature profile. The inviscid rotational flow code used in the analysis fails to predict the underturning at the mean, which may be caused by viscous effects.

Analytical and experimental study of flow through an axial turbine stage with a nonuniform inlet radial temperature profile

NASA Technical Reports Server (NTRS)

Schwab, J. R.; Stabe, R. G.; Whitney, W. J.

1983-01-01

Results are presented for a typical nonuniform inlet radial temperature profile through an advanced single-stage axial turbine and compared with the results obtained for a uniform profile. Gas temperature rises of 40 K to 95 K are predicted at the hub and tip corners at the trailing edges of the pressure surfaces in both the stator and rotor due to convection of hot fluid from the mean by the secondary flow. The inlet temperature profile is shown to be mixed out at the rotor exit survey plane (2.3 axial chords downstream of the rotor trailing edge) in both the analysis and the experiment. The experimental rotor exit angle profile for the nonuniform inlet temperature profile indicates underturning at the tip caused by increased clearance. Severe underturning also occurs at the mean, both with and without the nonuniform inlet temperature profile. The inviscid rotational flow code used in the analysis fails to predict the underturning at the mean, which may be caused by viscous effects. Previously announced in STAR as N83-27958

Integrated Design Engineering Analysis (IDEA) Environment - Aerodynamics, Aerothermodynamics, and Thermal Protection System Integration Module

NASA Technical Reports Server (NTRS)

Kamhawi, Hilmi N.

2011-01-01

This report documents the work performed during from March 2010 October 2011. The Integrated Design and Engineering Analysis (IDEA) environment is a collaborative environment based on an object-oriented, multidisciplinary, distributed environment using the Adaptive Modeling Language (AML) as the underlying framework. This report will focus on describing the work done in the area of extending the aerodynamics, and aerothermodynamics module using S/HABP, CBAERO, PREMIN and LANMIN. It will also detail the work done integrating EXITS as the TPS sizing tool.

Foreign body in scrotum following a boat engine blast accident.

PubMed

Mante, S D; Yeboah, E D; Adusei, B; Edusa, S

2013-03-01

Male genital injuries, demand prompt management to prevent long-term sexual and psychological damage. Injuries to the scrotum and contents may produce impaired fertility.We report our experience in diagnosing and managing a case of a foreign body in the scrotum following a boat engine blast accident. This case report highlights the need for a good history and thorough general examination to establish the mechanism of injury in order to distinguish between an embedded penetrating projectile injury and an injury with an exit wound. Prompt surgical exploration with hematoma evacuation limits complications.

Engineering: Defining and differentiating its unique culture

NASA Astrophysics Data System (ADS)

Pilotte, Mary K.

The world of work for engineering professionals is changing. At a rapid pace, experienced engineers are exiting the workforce due to retirement of the Baby Boomer generation, while at the same time the problems facing engineers are increasingly complex and frequently global in nature. For firms to protect their knowledge assets, they must ensure that acquired understandings are shared among their engineering work groups. Engineering teaching and learning in the workplace (i.e., knowledge sharing), is a social activity that resides in a social context governed by the professional engineering culture. This quantitative study uses Hofstede's Organizational Cultural Values Model (Hofstede, Neuijen, Ohayv, & Sanders, 1990) to examine dimensions of engineering culture in the workplace, producing a central tendency profile of engineering's cultural practices. Further, it explores through hypotheses if demographic differentiators, including birth generation, gender, race, industry sector of employment, and engineering discipline, play roles in forming engineering cultural practices. Results both corroborate aspects of Hofstede's model and assert new understandings relative to factors influencing dimensions of engineering practice. Outcomes are discussed in terms of their potential impact on industrial knowledge sharing and formation of beneficial engineering cultures.

Investigation of conjugate circular arcs in rocket nozzle contour design

NASA Astrophysics Data System (ADS)

Schomberg, K.; Olsen, J.; Neely, A.; Doig, G.

2018-05-01

The use of conjugate circular arcs in rocket nozzle contour design has been investigated by numerically comparing three existing sub-scale nozzles to a range of equivalent arc-based contour designs. Three performance measures were considered when comparing nozzle designs: thrust coefficient, nozzle exit wall pressure, and a transition between flow separation regimes during the engine start-up phase. In each case, an equivalent arc-based contour produced an increase in the thrust coefficient and exit wall pressure of up to 0.4 and 40% respectively, in addition to suppressing the transition between a free and restricted shock separation regime. A general approach to arc-based nozzle contour design has also been presented to outline a rapid and repeatable process for generating sub-scale arc-based contours with an exit Mach number of 3.8-5.4 and a length between 60 and 100% of a 15° conical nozzle. The findings suggest that conjugate circular arcs may represent a viable approach for producing sub-scale rocket nozzle contours, and that a further investigation is warranted between arc-based and existing full-scale rocket nozzles.

Heat exchange assembly

DOEpatents

Lowenstein, Andrew; Sibilia, Marc; Miller, Jeffrey; Tonon, Thomas S.

2004-06-08

A heat exchange assembly comprises a plurality of plates disposed in a spaced-apart arrangement, each of the plurality of plates includes a plurality of passages extending internally from a first end to a second end for directing flow of a heat transfer fluid in a first plane, a plurality of first end-piece members equaling the number of plates and a plurality of second end-piece members also equaling the number of plates, each of the first and second end-piece members including a recessed region adapted to fluidly connect and couple with the first and second ends of the plate, respectively, and further adapted to be affixed to respective adjacent first and second end-piece members in a stacked formation, and each of the first and second end-piece members further including at least one cavity for enabling entry of the heat transfer fluid into the plate, exit of the heat transfer fluid from the plate, or 180.degree. turning of the fluid within the plate to create a serpentine-like fluid flow path between points of entry and exit of the fluid, and at least two fluid conduits extending through the stacked plurality of first and second end-piece members for providing first fluid connections between the parallel fluid entry points of adjacent plates and a fluid supply inlet, and second fluid connections between the parallel fluid exit points of adjacent plates and a fluid discharge outlet so that the heat transfer fluid travels in parallel paths through each respective plate.

Hear Exchange Assembly

DOEpatents

Lowenstein, Andrew; Sibilia, Marc; Miller, Jeffrey; Tonon, Thomas S.

2003-05-27

A heat exchange assembly comprises a plurality of plates disposed in a spaced-apart arrangement, each of the plurality of plates includes a plurality of passages extending internally from a first end to a second end for directing flow of a heat transfer fluid in a first plane, a plurality of first end-piece members equaling the number of plates and a plurality of second end-piece members also equaling the number of plates, each of the first and second end-piece members including a recessed region adapted to fluidly connect and couple with the first and second ends of the plate, respectively, and further adapted to be affixed to respective adjacent first and second end-piece members in a stacked formation, and each of the first and second end-piece members further including at least one cavity for enabling entry of the heat transfer fluid into the plate, exit of the heat transfer fluid from the plate, or 180.degree. turning of the fluid within the plate to create a serpentine-like fluid flow path between points of entry and exit of the fluid, and at least two fluid conduits extending through the stacked plurality of first and second end-piece members for providing first fluid connections between the parallel fluid entry points of adjacent plates and a fluid supply inlet, and second fluid connections between the parallel fluid exit points of adjacent plates and a fluid discharge outlet so that the heat transfer fluid travels in parallel paths through each respective plate.

Bubble formation dynamics in a planar co-flow configuration: Influence of geometric and operating characteristics

NASA Astrophysics Data System (ADS)

Gutiérrez-Montes, Cándido; Bolaños-Jiménez, Rocío; Martínez-Bazán, Carlos; Sevilla, Alejandro

2014-11-01

An experimental and numerical study has been performed to explore the influence of different geometric features and operating conditions on the dynamics of a water-air-water planar co-flow. Specifically, regarding the nozzle used, the inner-to-outer thickness ratio of the air injector, β = Hi/Ho, the water-to-air thickness ratio, h = Hw/Ho, and the shape of the injector tip, have been described. As for the operating conditions, the water exit velocity profile under constant flow rate and constant air feeding pressure has been assessed. The results show that the jetting-bubbling transition is promoted for increasing values of β, decreasing values of h, rounded injector tip, and for uniform water exit velocity profiles. As for the bubble formation frequency, it increases with increasing values of β, decreasing values of h, rounded injector and parabolic-shaped water exit profiles. Furthermore, the bubble formation frequency has been shown to be lower under constant air feeding pressure conditions than at constant gas flow rate conditions. Finally, the effectiveness of a time-variable air feeding stream has been numerically studied, determining the forcing receptivity space in the amplitude-frequency plane. Experimental results corroborate the effectiveness of this control technique. Work supported by Spanish MINECO, Junta de Andalucía, European Funds and UJA under Projects DPI2011-28356-C03-02, DPI2011-28356-C03-03, P11-TEP7495 and UJA2013/08/05.

Microscopic information processing and communication in crowd dynamics

NASA Astrophysics Data System (ADS)

Henein, Colin Marc; White, Tony

2010-11-01

Due, perhaps, to the historical division of crowd dynamics research into psychological and engineering approaches, microscopic crowd models have tended toward modelling simple interchangeable particles with an emphasis on the simulation of physical factors. Despite the fact that people have complex (non-panic) behaviours in crowd disasters, important human factors in crowd dynamics such as information discovery and processing, changing goals and communication have not yet been well integrated at the microscopic level. We use our Microscopic Human Factors methodology to fuse a microscopic simulation of these human factors with a popular microscopic crowd model. By tightly integrating human factors with the existing model we can study the effects on the physical domain (movement, force and crowd safety) when human behaviour (information processing and communication) is introduced. In a large-room egress scenario with ample exits, information discovery and processing yields a crowd of non-interchangeable individuals who, despite close proximity, have different goals due to their different beliefs. This crowd heterogeneity leads to complex inter-particle interactions such as jamming transitions in open space; at high crowd energies, we found a freezing by heating effect (reminiscent of the disaster at Central Lenin Stadium in 1982) in which a barrier formation of naïve individuals trying to reach blocked exits prevented knowledgeable ones from exiting. Communication, when introduced, reduced this barrier formation, increasing both exit rates and crowd safety.

Experimental Measurements of the Effects of Photo-chemical Oxidation on Aerosol Emissions in Aircraft Exhaust

NASA Astrophysics Data System (ADS)

Miracolo, M. A.; Presto, A. A.; Hennigan, C. J.; Nguyen, N.; Ranjan, M.; Reeder, A.; Lipsky, E.; Donahue, N. M.; Robinson, A. L.

2009-12-01

Many military and commercial airfields are located in non-attainment areas for particulate matter (PM2.5), but the contribution of emissions from in-use aircraft to local and regional PM2.5 concentrations is uncertain. In collaboration with the Pennsylvania Air National Guard 171st Air Refueling Wing, the Carnegie Mellon University (CMU) Mobile Laboratory was deployed to measure fresh and aged emissions from a CFM56-2B1 gas-turbine engine mounted on a KC-135 Stratotanker airframe. The CFM-56 family of engine powers many different types of military and civilian aircraft, including the Boeing 737 and several Airbus models. It is one of the most widely deployed models of engines in the world. The goal of this work was to measure the gas-particle partitioning of the fresh emissions at atmospherically relevant conditions and to investigate the effect of atmospheric oxidation on aerosol loadings as the emissions age. Emissions were sampled from an inlet installed one meter downstream of the engine exit plane and transferred into a portable smog chamber via a heated inlet line. Separate experiments were conducted at different engine loads ranging from ground idle to take-off rated thrust. During each experiment, some diluted exhaust was added to the chamber and the volatility of the fresh emissions was then characterized using a thermodenuder. After this characterization, the chamber was exposed to either ambient sunlight or UV lights to initiate photochemical oxidation, which produced secondary aerosol and ozone. A suite of gas and particle-phase instrumentation was used to characterize the evolution of the gas and particle-phase emissions, including an aerosol mass spectrometer (AMS) to measure particle size and composition distributions. Fresh emissions of fine particles varied with engine load with peak emission factors at low and high loads. At high engine loads, the fresh emissions were dominated by black carbon; at low loads volatile organic carbon emissions were dominant. At low loads, photo-oxidation increased aerosol loadings in the chamber by a factor of fifty. We attribute this substantial secondary organic aerosol (SOA) production to oxidation of low-volatility organic vapors emitted under low loads. At higher loads, we see more modest secondary aerosol production from both organics and inorganics. Therefore secondary aerosol production can substantially exceed the direct aerosol emissions from aircraft. The results underscore the dramatic effects that photo-oxidation has on aerosol emissions from aircraft.

Modeling of Highly Instrumented Honeywell Turbofan Engine Tested with Ice Crystal Ingestion in the NASA Propulsion System Laboratory

NASA Technical Reports Server (NTRS)

Veres, Joseph P.; Jorgenson, Philip C. E.; Jones, Scott M.

2016-01-01

The Propulsion Systems Laboratory (PSL), an altitude test facility at NASA Glenn Research Center, has been used to test a highly instrumented turbine engine at simulated altitude operating conditions. This is a continuation of the PSL testing that successfully duplicated the icing events that were experienced in a previous engine (serial LF01) during flight through ice crystal clouds, which was the first turbofan engine tested in PSL. This second model of the ALF502R-5A serial number LF11 is a highly instrumented version of the previous engine. The PSL facility provides a continuous cloud of ice crystals with controlled characteristics of size and concentration, which are ingested by the engine during operation at simulated altitudes. Several of the previous operating points tested in the LF01 engine were duplicated to confirm repeatability in LF11. The instrumentation included video cameras to visually illustrate the accretion of ice in the low pressure compressor (LPC) exit guide vane region in order to confirm the ice accretion, which was suspected during the testing of the LF01. Traditional instrumentation included static pressure taps in the low pressure compressor inner and outer flow path walls, as well as total pressure and temperature rakes in the low pressure compressor region. The test data was utilized to determine the losses and blockages due to accretion in the exit guide vane region of the LPC. Multiple data points were analyzed with the Honeywell Customer Deck. A full engine roll back point was modeled with the Numerical Propulsion System Simulation (NPSS) code. The mean line compressor flow analysis code with ice crystal modeling was utilized to estimate the parameters that indicate the risk of accretion, as well as to estimate the degree of blockage and losses caused by accretion during a full engine roll back point. The analysis provided additional validation of the icing risk parameters within the LPC, as well as the creation of models for estimating the rates of blockage growth and losses.

iSTEM: The Aerospace Engineering Challenge

ERIC Educational Resources Information Center

English, Lyn D.; King, Donna T.; Hudson, Peter; Dawes, Les

2014-01-01

The authors developed The Paper Plane Challenge as one of a three-part response to The Aerospace Engineering Challenge. The Aerospace Engineering Challenge was the second of three multi-part activities that they had developed with the teachers during the year. Their aim was to introduce students to the exciting world of engineering, where they…

Design, fabrication and test of the RL10 derivative II chamber/primary nozzle

NASA Technical Reports Server (NTRS)

Marable, R. W.

1989-01-01

The design, fabrication and test of the RL10-II chamber/primary nozzle was accomplished as part of the RL10 Product Improvement Program (PIP). The overall goal of the RL10 PIP was to gain the knowledge and experience necessary to develop new cryogenic upper stage engines to fulfill future NASA requirements. The goal would be reached by producing an RL10 engine designed to be reusable, operate at several thrust levels, and have increased performance. The goals for the chamber/primary nozzle task were: (1) to design a reusable assembly capable of operation at increased mixture ratio and low thrust; (2) to fabricate three assemblies using new or updated techniques where possible; and (3) to test one assembly to verify the design and construction. The design and fabrication phases produced an assembly having improved features such as single piece reinforcing band segments (i.e., Mae West segments) and relocated tube exit braze joints (i.e., hooked tube exit). In addition, a computer program was developed to design the chamber tubes to meet both performance and heat transfer requirements. The test phase showed the specific impulse of the test bed engine system to be as predicted. These results, along with the heat transfer data obtained, sufficiently proved the overall design of the RL10-II recontoured and shortened chamber/primary nozzle assembly.

Asymptotic problems for stochastic partial differential equations

NASA Astrophysics Data System (ADS)

Salins, Michael

Stochastic partial differential equations (SPDEs) can be used to model systems in a wide variety of fields including physics, chemistry, and engineering. The main SPDEs of interest in this dissertation are the semilinear stochastic wave equations which model the movement of a material with constant mass density that is exposed to both determinstic and random forcing. Cerrai and Freidlin have shown that on fixed time intervals, as the mass density of the material approaches zero, the solutions of the stochastic wave equation converge uniformly to the solutions of a stochastic heat equation, in probability. This is called the Smoluchowski-Kramers approximation. In Chapter 2, we investigate some of the multi-scale behaviors that these wave equations exhibit. In particular, we show that the Freidlin-Wentzell exit place and exit time asymptotics for the stochastic wave equation in the small noise regime can be approximated by the exit place and exit time asymptotics for the stochastic heat equation. We prove that the exit time and exit place asymptotics are characterized by quantities called quasipotentials and we prove that the quasipotentials converge. We then investigate the special case where the equation has a gradient structure and show that we can explicitly solve for the quasipotentials, and that the quasipotentials for the heat equation and wave equation are equal. In Chapter 3, we study the Smoluchowski-Kramers approximation in the case where the material is electrically charged and exposed to a magnetic field. Interestingly, if the system is frictionless, then the Smoluchowski-Kramers approximation does not hold. We prove that the Smoluchowski-Kramers approximation is valid for systems exposed to both a magnetic field and friction. Notably, we prove that the solutions to the second-order equations converge to the solutions of the first-order equation in an Lp sense. This strengthens previous results where convergence was proved in probability.

Turbofan Engine Core Compartment Vent Aerodynamic Configuration Development Methodology

NASA Technical Reports Server (NTRS)

Hebert, Leonard J.

2006-01-01

This paper presents an overview of the design methodology used in the development of the aerodynamic configuration of the nacelle core compartment vent for a typical Boeing commercial airplane together with design challenges for future design efforts. Core compartment vents exhaust engine subsystem flows from the space contained between the engine case and the nacelle of an airplane propulsion system. These subsystem flows typically consist of precooler, oil cooler, turbine case cooling, compartment cooling and nacelle leakage air. The design of core compartment vents is challenging due to stringent design requirements, mass flow sensitivity of the system to small changes in vent exit pressure ratio, and the need to maximize overall exhaust system performance at cruise conditions.

Atomization characteristics of swirl injector sprays

NASA Technical Reports Server (NTRS)

Feikema, Douglas A.

1996-01-01

Stable combustion within rocket engines is a continuing concern for designers of rocket engine systems. The swirl-coaxial injector has demonstrated effectiveness in achieving atomization and mixing, and therefore stable combustion. Swirl-coaxial injector technology is being deployed in the American RL1OA rocket design and Russian engine systems already make wide spread use of this technology. The present requirement for swirl injector research is derived from NASA's current Reusable Launch Vehicle (RLV) technology program. This report describes some of the background and literature on this topic including drop size measurements, comparison with theoretical predictions, the effect of surface tension on the atomization process, and surface wave characteristics of liquid film at the exit of the injector.

Scale Model Experiments on Sound Propagation From a Mach 2.5 Cold Nitrogen Jet Flowing Through a Rigid-Walled Duct With a J-Deflector

NASA Technical Reports Server (NTRS)

Kandula, Max; Vu, Bruce

2003-01-01

The Launch Systems Testbed (LST) represents the evolution of vibroacoustics research and development work performed at NASA John F. Kennedy Space Center (KSC) over the last 15 years. The LST is located at the Launch Equipment Test Facility (LETF) in the KSC industrial complex. The LETF is operated by Sierra Lobo, Inc., as a member of University-Affiliated Technology Development Contract (USTDC) to KSC Spaceport and Engineering and Technology Directorate (YA), with ASRC Aerospace Corporation as a the prime contractor. Trajectory Simulation Mechanism (TSM) is a major component of the LST, developed specifically to simulate nonstationary acoustic loads on launch pad structures, vehicles, and payloads. TSM enhances the capabilities within LST for simulating launch environments of future vehicles. The scaled launch environments will be used to predict the full-scale launch environment via an appropriate scaling procedure. Air Force Research Laboratory (AFRL) has tasked NASA KSC to perform a basic technology test program in support of developing a low-cost clean pad (incorporating passive mitigation techniques) for future launch vehicles. The overall goal of the program is to develop innovative launch exhaust management systems, which effectively reduce launch acoustic environment with innovative duct designs, while eliminating traditional sound suppression water systems. Passive techniques, such as nontraditional duct geometries, resonators, and diffusers, etc., will be investigated. The overall goals are to advance innovative concepts for a clean pad while developing ideas to reduce transmitted sound via investigation and modeling of jet exhaust acoustic and flow field characteristics. The series of tests outlined in this report represent baseline tests and are geared towards defining the acoustic load environment on the TSM pad for open and closed duct configurations. This report summarizes the cold jet acoustic testing for Mach 2.5 supersonic nitrogen jet issuing from a nozzle with 1-inch exit diameter. Acoustic data, including spectral sound power and Overall Sound Pressure Level (OASPL), are obtained both for a free jet and with the jet flowing through a rigid-walled duct with a J-deflector. The relative performance of closed duct and open duct is evaluated. The results show that the closed duct is superior to the partially open duct, and results in about 3-decibel (dB) noise reduction (near the duct axis) relative to the free jet. The location of the nozzle exit plane (NEP) relative to the duct inlet plane (DIP) has a significant effect on the acoustic field. The results suggest that the location of NEP at 10 inches above the DIP results in reduced acoustic loads relative to 5 inches above the duct inlet and 1 inch into the duct inlet.

Dr. von Braun Standing by Five F-1 Engines

NASA Technical Reports Server (NTRS)

2004-01-01

A pioneer of America's space program, Dr. von Braun stands by the five F-1 engines of the Saturn V launch vehicle. This Saturn V vehicle is an actual test vehicle which has been displayed at the U.S. Space Rocket Center in Huntsville, Alabama. Designed and developed by Rocketdyne under the direction of the Marshall Space Flight Center, a cluster of five F-1 engines was mounted on the Saturn V S-IC (first) stage. The engines measured 19-feet tall by 12.5-feet at the nozzle exit and burned 15 tons of liquid oxygen and kerosene each second to produce 7,500,000 pounds of thrust. The S-IC stage is the first stage, or booster, of a 364-foot long rocket that ultimately took astronauts to the Moon.

Experimental studies of thermal preparation of internal combustion engine

NASA Astrophysics Data System (ADS)

Karnaukhov, N. N.; Merdanov, Sh M.; V, Konev V.; Borodin, D. M.

2018-05-01

In conditions of autonomous functioning of road construction machines, it becomes necessary to use its internal sources. This can be done by using a heat recovery system of an internal combustion engine (ICE). For this purpose, it is proposed to use heat accumulators that accumulate heat of the internal combustion engine during the operation of the machine. Experimental studies have been carried out to evaluate the efficiency of using the proposed pre-start thermal preparation system, which combines a regular system based on liquid diesel fuel heaters and an ICE heat recovery system. As a result, the stages of operation of the preheating thermal preparation system, mathematical models and the dependence of the temperature change of the antifreeze at the exit from the internal combustion engine on the warm-up time are determined.

Luminescent Measurement Systems for the Investigation of a Scramjet Inlet-Isolator

PubMed Central

Idris, Azam Che; Saad, Mohd Rashdan; Zare-Behtash, Hossein; Kontis, Konstantinos

2014-01-01

Scramjets have become a main focus of study for many researchers, due to their application as propulsive devices in hypersonic flight. This entails a detailed understanding of the fluid mechanics involved to be able to design and operate these engines with maximum efficiency even at their off-design conditions. It is the objective of the present cold-flow investigation to study and analyse experimentally the mechanics of the fluid structures encountered within a generic scramjet inlet at M = 5. Traditionally, researchers have to rely on stream-thrust analysis, which requires the complex setup of a mass flow meter, a force balance and a heat transducer in order to measure inlet-isolator performance. Alternatively, the pitot rake could be positioned at inlet-isolator exit plane, but this method is intrusive to the flow, and the number of pitot tubes is limited by the model size constraint. Thus, this urgent need for a better flow diagnostics method is addressed in this paper. Pressure-sensitive paint (PSP) has been applied to investigate the flow characteristics on the compression ramp, isolator surface and isolator sidewall. Numerous shock-shock interactions, corner and shoulder separation regions, as well as shock trains were captured by the luminescent system. The performance of the scramjet inlet-isolator has been shown to improve when operated in a modest angle of attack. PMID:24721773

Combustor Computations for CO2-Neutral Aviation

NASA Technical Reports Server (NTRS)

Hendricks, Robert C.; Brankovic, Andreja; Ryder, Robert C.; Huber, Marcia

2011-01-01

Knowing the pure component C(sub p)(sup 0) or mixture C(sub p) (sup 0) as computed by a flexible code such as NIST-STRAPP or McBride-Gordon, one can, within reasonable accuracy, determine the thermophysical properties necessary to predict the combustion characteristics when there are no tabulated or computed data for those fluid mixtures 3or limited results for lower temperatures. (Note: C(sub p) (sup 0) is molar heat capacity at constant pressure.) The method can be used in the determination of synthetic and biological fuels and blends using the NIST code to compute the C(sub p) (sup 0) of the mixture. In this work, the values of the heat capacity were set at zero pressure, which provided the basis for integration to determine the required combustor properties from the injector to the combustor exit plane. The McBride-Gordon code was used to determine the heat capacity at zero pressure over a wide range of temperatures (room to 6,000 K). The selected fluids were Jet-A, 224TMP (octane), and C12. It was found that each heat capacity loci were form-similar. It was then determined that the results [near 400 to 3,000 K] could be represented to within acceptable engineering accuracy with the simplified equation C(sub p) (sup 0) = A/T + B, where A and B are fluid-dependent constants and T is temperature (K).

Application of Optimization Techniques to Design of Unconventional Rocket Nozzle Configurations

NASA Technical Reports Server (NTRS)

Follett, W.; Ketchum, A.; Darian, A.; Hsu, Y.

1996-01-01

Several current rocket engine concepts such as the bell-annular tri-propellant engine, and the linear aerospike being proposed for the X-33 require unconventional three dimensional rocket nozzles which must conform to rectangular or sector shaped envelopes to meet integration constraints. These types of nozzles exist outside the current experience database, therefore, the application of efficient design methods for these propulsion concepts is critical to the success of launch vehicle programs. The objective of this work is to optimize several different nozzle configurations, including two- and three-dimensional geometries. Methodology includes coupling computational fluid dynamic (CFD) analysis to genetic algorithms and Taguchi methods as well as implementation of a streamline tracing technique. Results of applications are shown for several geometeries including: three dimensional thruster nozzles with round or super elliptic throats and rectangualar exits, two- and three-dimensional thrusters installed within a bell nozzle, and three dimensional thrusters with round throats and sector shaped exits. Due to the novel designs considered for this study, there is little experience which can be used to guide the effort and limit the design space. With a nearly infinite parameter space to explore, simple parametric design studies cannot possibly search the entire design space within the time frame required to impact the design cycle. For this reason, robust and efficient optimization methods are required to explore and exploit the design space to achieve high performance engine designs. Five case studies which examine the application of various techniques in the engineering environment are presented in this paper.

Aerodynamic Investigation of Incidence Angle Effects in a Large Scale Transonic Turbine Cascade

NASA Technical Reports Server (NTRS)

McVetta, Ashlie B.; Giel, Paul W.; Welch, Gerard E.

2013-01-01

Aerodynamic measurements showing the effects of large incidence angle variations on an HPT turbine blade set are presented. Measurements were made in NASA's Transonic Turbine Blade Cascade Facility which has been used in previous studies to acquire detailed aerodynamic and heat transfer measurements for CFD code validation. The current study supports the development of variable-speed power turbine (VSPT) speed-change technology for the NASA Large Civil Tilt Rotor (LCTR) vehicle. In order to maintain acceptable main rotor propulsive efficiency, the VSPT operates over a nearly 50 percent speed range from takeoff to altitude cruise. This results in 50deg or more variations in VSPT blade incidence angles. The cascade facility has the ability to operate over a wide range of Reynolds numbers and Mach numbers, but had to be modified in order to accommodate the negative incidence angle variation required by the LCTR VSPT operation. Using existing blade geometry with previously acquired aerodynamic data, the tunnel was re-baselined and the new incidence angle range was exercised. Midspan exit total pressure and flow angle measurements were obtained at seven inlet flow angles. For each inlet angle, data were obtained at five flow conditions with inlet Reynolds numbers varying from 6.83×10(exp 5) to 0.85×10(exp 5) and two isentropic exit Mach numbers of 0.74 and 0.34. The midspan flowfield measurements were acquired using a three-hole pneumatic probe located in a survey plane 8.6 percent axial chord downstream of the blade trailing edge plane and covering three blade passages. Blade and endwall static pressure distributions were also acquired for each flow condition.

Aerodynamic Investigation of Incidence Angle Effects in a Large Scale Transonic Turbine Cascade. Revision 1

NASA Technical Reports Server (NTRS)

McVetta, Ashlie B.; Giel, Paul W.; Welch, Gerard E.

2014-01-01

Aerodynamic measurements showing the effects of large incidence angle variations on an HPT turbine blade set are presented. Measurements were made in NASA's Transonic Turbine Blade Cascade Facility which has been used in previous studies to acquire detailed aerodynamic and heat transfer measurements for CFD code validation. The current study supports the development of variable-speed power turbine (VSPT) speed-change technology for the NASA Large Civil Tilt Rotor (LCTR) vehicle. In order to maintain acceptable main rotor propulsive efficiency, the VSPT operates over a nearly 50 percent speed range from takeoff to altitude cruise. This results in 50 deg or more variations in VSPT blade incidence angles. The cascade facility has the ability to operate over a wide range of Reynolds numbers and Mach numbers, but had to be modified in order to accommodate the negative incidence angle variation required by the LCTR VSPT operation. Using existing blade geometry with previously acquired aerodynamic data, the tunnel was re-baselined and the new incidence angle range was exercised. Midspan exit total pressure and flow angle measurements were obtained at seven inlet flow angles. For each inlet angle, data were obtained at five flow conditions with inlet Reynolds numbers varying from 6.83×10 (exp 5) to 0.85×10(exp 5) and two isentropic exit Mach numbers of 0.74 and 0.34. The midspan flowfield measurements were acquired using a three-hole pneumatic probe located in a survey plane 8.6 percent axial chord downstream of the blade trailing edge plane and covering three blade passages. Blade and endwall static pressure distributions were also acquired for each flow condition.

Aerodynamic Investigation of Incidence Angle Effects in a Large Scale Transonic Turbine Cascade

NASA Technical Reports Server (NTRS)

McVetta, Ashlie B.; Giel, Paul W.; Welch, Gerard E.

2012-01-01

Aerodynamic measurements showing the effects of large incidence angle variations on an HPT turbine blade set are presented. Measurements were made in NASA's Transonic Turbine Blade Cascade Facility which has been used in previous studies to acquire detailed aerodynamic and heat transfer measurements for CFD code validation. The current study supports the development of variable-speed power turbine (VSPT) speed-change technology for the NASA Large Civil Tilt Rotor (LCTR) vehicle. In order to maintain acceptable main rotor propulsive efficiency, the VSPT operates over a nearly 50% speed range from takeoff to altitude cruise. This results in 50 degrees or more variations in VSPT blade incidence angles. The cascade facility has the ability to operate over a wide range of Reynolds numbers and Mach numbers, but had to be modified in order to accommodate the negative incidence angle variation required by the LCTR VSPT operation. Using existing blade geometry with previously acquired aerodynamic data, the tunnel was re-baselined and the new incidence angle range was exercised. Midspan exit total pressure and flow angle measurements were obtained at seven inlet flow angles. For each inlet angle, data were obtained at five flow conditions with inlet Reynolds numbers varying from 6.83 × 10(exp 5) to 0.85 ×10(exp 5) and two isentropic exit Mach numbers of 0.74 and 0.34. The midspan flowfield measurements were acquired using a three-hole pneumatic probe located in a survey plane 8.6% axial chord downstream of the blade trailing edge plane and covering three blade passages. Blade and endwall static pressure distributions were also acquired for each flow condition

Partially integrated exhaust manifold

DOEpatents

Hayman, Alan W; Baker, Rodney E

2015-01-20

A partially integrated manifold assembly is disclosed which improves performance, reduces cost and provides efficient packaging of engine components. The partially integrated manifold assembly includes a first leg extending from a first port and terminating at a mounting flange for an exhaust gas control valve. Multiple additional legs (depending on the total number of cylinders) are integrally formed with the cylinder head assembly and extend from the ports of the associated cylinder and terminate at an exit port flange. These additional legs are longer than the first leg such that the exit port flange is spaced apart from the mounting flange. This configuration provides increased packaging space adjacent the first leg for any valving that may be required to control the direction and destination of exhaust flow in recirculation to an EGR valve or downstream to a catalytic converter.

Turbine exhaust diffuser with region of reduced flow area and outer boundary gas flow

DOEpatents

Orosa, John

2014-03-11

An exhaust diffuser system and method for a turbine engine. The outer boundary may include a region in which the outer boundary extends radially inwardly toward the hub structure and may direct at least a portion of an exhaust flow in the diffuser toward the hub structure. At least one gas jet is provided including a jet exit located on the outer boundary. The jet exit may discharge a flow of gas downstream substantially parallel to an inner surface of the outer boundary to direct a portion of the exhaust flow in the diffuser toward the outer boundary to effect a radially outward flow of at least a portion of the exhaust gas flow toward the outer boundary to balance an aerodynamic load between the outer and inner boundaries.

Performance of a small annular turbojet combustor designed for low cost

NASA Technical Reports Server (NTRS)

Fear, J. S.

1972-01-01

Performance investigations were conducted on a combustor utilizing several cost-reducing innovations and designed for use in a low-cost 4448-N thrust turbojet engine for commercial light aircraft. Low-cost features included simple, air-atomizing fuel injectors; combustor liners of perforated sheet; and the use of inexpensive type 304 stainless-steel material. Combustion efficiencies at the cruise and sea-level-takeoff design points were approximately 97 and 98 percent, respectively. The combustor isothermal pressure loss was 6.3 percent at the cruise-condition diffuser inlet Mach number of 0.34. The combustor exit temperature pattern factor was less than 0.24 at both the cruise and sea-level-takeoff design points. The combustor exit average radial temperature profiles at all conditions were in very good agreement with the design profile.

Design of a Mach-15 Total-Enthalpy Nozzle With Non-uniform Inflow Using Rotational MOC

NASA Technical Reports Server (NTRS)

Gaffney, Richard L., Jr.

2004-01-01

A new computer program to design nozzles with non-uniform inflow has been developed using the rotational method of characteristics (MOC). This program has been used to design a nozzle for the NASA's HYPULSE shock-expansion tunnel for use in scramjet engine tests at a Mach-15 flight-enthalpy condition. The nozzle has an area ratio of 9.5:1 that expands the inflow from Mach 6 along the centerline to Mach 8.7. Although the density and Mach number vary radially at the exit due to the non-uniformities of the inflow, the MOC procedure produces exit flow that is parallel and has uniform static pressure. The design has been verified with CFD which compares favorably with the MOC solution.

An improved numerical procedure for the parametric optimization of three dimensional scramjet nozzles. [supersonic combustion ramjet engines - computer programs

NASA Technical Reports Server (NTRS)

Dash, S.; Delguidice, P. D.

1975-01-01

A parametric numerical procedure permitting the rapid determination of the performance of a class of scramjet nozzle configurations is presented. The geometric complexity of these configurations ruled out attempts to employ conventional nozzle design procedures. The numerical program developed permitted the parametric variation of cowl length, turning angles on the cowl and vehicle undersurface and lateral expansion, and was subject to fixed constraints such as the vehicle length and nozzle exit height. The program required uniform initial conditions at the burner exit station and yielded the location of all predominant wave zones, accounting for lateral expansion effects. In addition, the program yielded the detailed pressure distribution on the cowl, vehicle undersurface and fences, if any, and calculated the nozzle thrust, lift and pitching moments.

Particle model of full-size ITER-relevant negative ion source.

PubMed

Taccogna, F; Minelli, P; Ippolito, N

2016-02-01

This work represents the first attempt to model the full-size ITER-relevant negative ion source including the expansion, extraction, and part of the acceleration regions keeping the mesh size fine enough to resolve every single aperture. The model consists of a 2.5D particle-in-cell Monte Carlo collision representation of the plane perpendicular to the filter field lines. Magnetic filter and electron deflection field have been included and a negative ion current density of j(H(-)) = 660 A/m(2) from the plasma grid (PG) is used as parameter for the neutral conversion. The driver is not yet included and a fixed ambipolar flux is emitted from the driver exit plane. Results show the strong asymmetry along the PG driven by the electron Hall (E × B and diamagnetic) drift perpendicular to the filter field. Such asymmetry creates an important dis-homogeneity in the electron current extracted from the different apertures. A steady state is not yet reached after 15 μs.

Construction of mathematical models the parachute jumper with change position acrobatic

NASA Astrophysics Data System (ADS)

Asmianto, Asmianto; Hariyanto, Hariyanto; Herisman, Iis

2018-03-01

Construction of mathematical models the movement of a parachutist during the air using newton’s II law is ΣF = ma. Position parachutist after exiting the plane immediately unfurled his body so as to create a large air resistance. The presence of air drag resulting movement indirectly parachutist moves down vertically downwards but also shifting toward horizontal and form a parabolic trajectory. Parachutist speed getting down increased until eventually the parachutist reaches terminal velocity it’s the position where the air drag is equal to the gravitational force (gravity) jumpers. In this paper is assumed to be parachutist with 91.6 kg mass (including equipment) jumping from a plane at an altitude of 3.000 meters and reach a height of parachutist ± 1000 meters with velocity ± 57 m/s. So the parachutist have to be clever in taking account of everything, because if just a little too late can dangerous the safety of the parachutist.

Experimental and computational investigation of the NASA low-speed centrifugal compressor flow field

NASA Technical Reports Server (NTRS)

Hathaway, Michael D.; Chriss, Randall M.; Wood, Jerry R.; Strazisar, Anthony J.

1993-01-01

An experimental and computational investigation of the NASA Lewis Research Center's low-speed centrifugal compressor (LSCC) flow field was conducted using laser anemometry and Dawes' three-dimensional viscous code. The experimental configuration consisted of a backswept impeller followed by a vaneless diffuser. Measurements of the three-dimensional velocity field were acquired at several measurement planes through the compressor. The measurements describe both the throughflow and secondary velocity field along each measurement plane. In several cases the measurements provide details of the flow within the blade boundary layers. Insight into the complex flow physics within centrifugal compressors is provided by the computational fluid dynamics analysis (CFD), and assessment of the CFD predictions is provided by comparison with the measurements. Five-hole probe and hot-wire surveys at the inlet and exit to the impeller as well as surface flow visualization along the impeller blade surfaces provided independent confirmation of the laser measurement technique. The results clearly document the development of the throughflow velocity wake that is characteristic of unshrouded centrifugal compressors.

Engine component improvement: Performance improvement, JT9D-7 3.8 AR fan

NASA Technical Reports Server (NTRS)

Gaffin, W. O.

1980-01-01

A redesigned, fuel efficient fan for the JT9D-7 engine was tested. Tests were conducted to determine the effect of the 3.8 AR fan on performance, stability, operational characteristics, and noise of the JT9D-7 engine relative to the current 4.6 AR Bill-of-Material fan. The 3.8 AR fan provides increased fan efficiency due to a more advanced blade airfoil with increased chord, eliminating one part span shroud and reducing the number of fan blades and fan exit guide vanes. Engine testing at simulated cruise conditions demonstrated the predicted 1.3 percent improvement in specific fuel consumption with the redesigned 3.8 AR fan. Flight testing and sea level stand engine testing demonstrated exhaust gas temperature margins, fan and low pressure compressor stability, operational suitability, and noise levels comparable to the Bill-of-Material fan.

Perspectives On Dilution Jet Mixing

NASA Technical Reports Server (NTRS)

Holdeman, J. D.; Srinivasan, R.

1990-01-01

NASA recently completed program of measurements and modeling of mixing of transverse jets with ducted crossflow, motivated by need to design or tailor temperature pattern at combustor exit in gas turbine engines. Objectives of program to identify dominant physical mechanisms governing mixing, extend empirical models to provide near-term predictive capability, and compare numerical code calculations with data to guide future analysis improvement efforts.

Exchange inlet optimization by genetic algorithm for improved RBCC performance

NASA Astrophysics Data System (ADS)

Chorkawy, G.; Etele, J.

2017-09-01

A genetic algorithm based on real parameter representation using a variable selection pressure and variable probability of mutation is used to optimize an annular air breathing rocket inlet called the Exchange Inlet. A rapid and accurate design method which provides estimates for air breathing, mixing, and isentropic flow performance is used as the engine of the optimization routine. Comparison to detailed numerical simulations show that the design method yields desired exit Mach numbers to within approximately 1% over 75% of the annular exit area and predicts entrained air massflows to between 1% and 9% of numerically simulated values depending on the flight condition. Optimum designs are shown to be obtained within approximately 8000 fitness function evaluations in a search space on the order of 106. The method is also shown to be able to identify beneficial values for particular alleles when they exist while showing the ability to handle cases where physical and aphysical designs co-exist at particular values of a subset of alleles within a gene. For an air breathing engine based on a hydrogen fuelled rocket an exchange inlet is designed which yields a predicted air entrainment ratio within 95% of the theoretical maximum.

Optimal Area Profiles for Ideal Single Nozzle Air-Breathing Pulse Detonation Engines

NASA Technical Reports Server (NTRS)

Paxson, Daniel E.

2003-01-01

The effects of cross-sectional area variation on idealized Pulse Detonation Engine performance are examined numerically. A quasi-one-dimensional, reacting, numerical code is used as the kernel of an algorithm that iteratively determines the correct sequencing of inlet air, inlet fuel, detonation initiation, and cycle time to achieve a limit cycle with specified fuel fraction, and volumetric purge fraction. The algorithm is exercised on a tube with a cross sectional area profile containing two degrees of freedom: overall exit-to-inlet area ratio, and the distance along the tube at which continuous transition from inlet to exit area begins. These two parameters are varied over three flight conditions (defined by inlet total temperature, inlet total pressure and ambient static pressure) and the performance is compared to a straight tube. It is shown that compared to straight tubes, increases of 20 to 35 percent in specific impulse and specific thrust are obtained with tubes of relatively modest area change. The iterative algorithm is described, and its limitations are noted and discussed. Optimized results are presented showing performance measurements, wave diagrams, and area profiles. Suggestions for future investigation are also discussed.

Merits of full flow vs. conventional staged combustion cycles for reusable launch vehicle propulsion

NASA Astrophysics Data System (ADS)

Peery, Steven D.; Parsley, Randy C.

1996-03-01

This paper provides a comparison between full-flow and conventional staged combustion thermodynamic O2/H2 rocket engine cycles for Reusable Launch Vehicle, RLV, single-stage-to-orbit applications. The impact of the cycle thermodynamics, component configuration, and component operating parameters on engine performance and weight for the two approaches is presented. Both cycles were modeled with equivalent technology turbomachinery and chamber/nozzle RLV life requirements. The first order impact of cycle selection, pump exit pressure, and turbine temperature on the empty weight of an SSTO Reusable Launch Vehicle is presented.

Preparing America for Deep Space Exploration Episode 11: Committed to Exploration

NASA Image and Video Library

2015-12-09

Engineers around the country are making progress developing NASA’s Space Launch System, Orion spacecraft and the ground systems at Kennedy Space Center in Florida needed to send astronauts on missions to deep space destinations. Between July and September, progress continued as pieces of Orion’s crew module and the SLS core stage tanks were welded together at NASA’s Michoud Assembly Facility in New Orleans, modifications were made to the mobile launcher at Kennedy, astronauts tested techniques for exiting Orion after a mission, and an RS-25 engine was tested at Stennis Space Center in Mississippi.

Noise-free recovery of optodigital encrypted and multiplexed images.

PubMed

Henao, Rodrigo; Rueda, Edgar; Barrera, John F; Torroba, Roberto

2010-02-01

We present a method that allows storing multiple encrypted data using digital holography and a joint transform correlator architecture with a controllable angle reference wave. In this method, the information is multiplexed by using a key and a different reference wave angle for each object. In the recovering process, the use of different reference wave angles prevents noise produced by the nonrecovered objects from being superimposed on the recovered object; moreover, the position of the recovered object in the exit plane can be fully controlled. We present the theoretical analysis and the experimental results that show the potential and applicability of the method.

FIBER AND INTEGRATED OPTICS: Compact fiber-optic compressor of ultrashort pulses

NASA Astrophysics Data System (ADS)

Nikitin, S. P.; Onishchukov, G. I.; Fomichev, A. A.

1992-02-01

A theoretical design of a universal compact fiber-optic compressor based on a monochromator with a spherical mirror in the plane of its exit slit was considered. Ultrashort pulses emitted by an actively mode-locked YAG:Nd3+ laser, whose spectrum was broadened in a fiber-optic waveguide, were compressed experimentally to 2.7 ns. A universal compact compressor was developed: it produced 4-ns pulses with an average radiation power of about 1 W. The dimensions of this compressor were several times smaller than those of a traditional scheme using a diffraction grating to compress pulses having an initial duration of about 100 ns.

Laser anemometer measurements in an annular cascade of core turbine vanes and comparison with theory

NASA Technical Reports Server (NTRS)

Goldman, L. J.; Seashultz, R. G.

1982-01-01

Laser measurements were made in an annular cascade of stator vanes operating at an exit critical velocity ratio of 0.78. Velocity and flow angles in the blade to blade plane were obtained at every 10 percent of axial chord within the passage and at 1/2 axial chord downstream of the vanes for radial positions near the hub, mean and tip. Results are presented in both plot and tabulated form and are compared with calculations from an inviscid, quasi three dimensional computer program. The experimental measurements generally agreed well with these theoretical calculations, an indication of the usefulness of this analytic approach.

Experimental demonstration of free-space optical vortex transmutation with polygonal lenses.

PubMed

Gao, Nan; Xie, Changqing

2012-08-01

Vortex transmutation was predicted to take place when vortices interact with systems possessing discrete rotational symmetries of finite order [Phys. Rev. Lett.95, 123901 (2005)]. Here we report what is believed to be the first experimental demonstration of vortex transmutation. We show that in free space, by simply inserting polygonal lenses into the optical path, the central vorticity of a coaxially incident optical vortex can be changed following the modular transmutation rule. We generate the wavefront at the exit face of the lenses with computer generated holograms and measure the output vorticity using the interference patterns at the focal plane. The results agree well with theoretical predictions.

MPD work at MIT

NASA Technical Reports Server (NTRS)

Martinez-Sanchez, Manuel

1991-01-01

MPD work at MIT is presented in the form of the view-graphs. The following subject areas are covered: the MIT program, its goals, achievements, and roadblocks; quasi one-dimensional modeling; two-dimensional modeling - transport effects and Hall effect; microscopic instabilities in MPD flows and modified two stream instability; electrothermal stability theory; separation of onset and anode depletion; exit plane spectroscopic measurements; phenomena of onset as performance limiter; explanations of onset; geometry effects on onset; onset at full ionization and its consequences; relationship to anode depletion; summary on self-field MPD; applied field MPD - the logical growth path; the case for AF; the challenges of AF MPD; and recommendations.

Near-Field Diffraction Imaging from Multiple Detection Planes

NASA Astrophysics Data System (ADS)

Loetgering, L.; Golembusch, M.; Hammoud, R.; Wilhein, T.

2017-06-01

We present diffraction imaging results obtained from multiple near-field diffraction constraints. An iterative phase retrieval algorithm was implemented that uses data redundancy achieved by measuring near-field diffraction intensities at various sample-detector distances. The procedure allows for reconstructing the exit surface wave of a sample within a multiple constraint satisfaction framework neither making use of a priori knowledge as enforced in coherent diffraction imaging (CDI) nor exact scanning grid knowledge as required in ptychography. We also investigate the potential of the presented technique to deal with polychromatic radiation as important for potential application in diffraction imaging by means of tabletop EUV and X-ray sources.

Earth Observing-1 Advanced Imager Flight Performance Assessment: Investigating Dark Current Stability Over One-Half Orbit Period during the First 60 Days

NASA Technical Reports Server (NTRS)

Mendenhall, J. A.

2001-01-01

The stability of the EO-1 Advanced Land Imager dark current levels over the period of one-half orbit is investigated. A series of two-second dark current collections, over the course of 40 minutes, was performed during the first sixty days the instrument was in orbit. Analysis of this data indicates only two dark current reference periods, obtained entering and exiting eclipse, are required to remove ALI dark current offsets for 99.9% of the focal plane to within 1.5 digital numbers for any observation on the solar illuminated portion of the orbit.

Stage separation study of Nike-Black Brant V Sounding Rocket System

NASA Technical Reports Server (NTRS)

Ferragut, N. J.

1976-01-01

A new Sounding Rocket System has been developed. It consists of a Nike Booster and a Black Brant V Sustainer with slanted fins which extend beyond its nozzle exit plane. A cursory look was taken at different factors which must be considered when studying a passive separation system. That is, one separation system without mechanical constraints in the axial direction and which will allow separation due to drag differential accelerations between the Booster and the Sustainer. The equations of motion were derived for rigid body motions and exact solutions were obtained. The analysis developed could be applied to any other staging problem of a Sounding Rocket System.

Pulsed plasma thruster contamination studies

NASA Technical Reports Server (NTRS)

Rudolph, L. K.; Jones, R. M.

1979-01-01

The exhaust plume of the one millipound pulsed plasma thruster has a measurable backflow upstream of the nozzle exit plane which may deposit on and degrade the performance of exposed spacecraft surfaces. High speed photographs and Faraday cup measurements suggest that this backflow is predominantly an electrically neutral, relatively low energy vapor. Articulated collimator quartz crystal microbalance measurements of this backflow were made for a thruster with a radically modified nozzle and a flat plate backflow shield, to determine the backflow sensitivity to nozzle design changes. The results are compared with the original nozzle backflow and show a measurable reduction in the backflow directly upstream of the shield.

The TORMOZ Gene Encodes a Nucleolar Protein Required for Regulated Division Planes and Embryo Development in Arabidopsis[W

PubMed Central

Griffith, Megan E.; Mayer, Ulrike; Capron, Arnaud; Ngo, Quy A.; Surendrarao, Anandkumar; McClinton, Regina; Jürgens, Gerd; Sundaresan, Venkatesan

2007-01-01

Embryogenesis in Arabidopsis thaliana is marked by a predictable sequence of oriented cell divisions, which precede cell fate determination. We show that mutation of the TORMOZ (TOZ) gene yields embryos with aberrant cell division planes and arrested embryos that appear not to have established normal patterning. The defects in toz mutants differ from previously described mutations that affect embryonic cell division patterns. Longitudinal division planes of the proembryo are frequently replaced by transverse divisions and less frequently by oblique divisions, while divisions of the suspensor cells, which divide only transversely, appear generally unaffected. Expression patterns of selected embryo patterning genes are altered in the mutant embryos, implying that the positional cues required for their proper expression are perturbed by the misoriented divisions. The TOZ gene encodes a nucleolar protein containing WD repeats. Putative TOZ orthologs exist in other eukaryotes including Saccharomyces cerevisiae, where the protein is predicted to function in 18S rRNA biogenesis. We find that disruption of the Sp TOZ gene results in cell division defects in Schizosaccharomyces pombe. Previous studies in yeast and animal cells have identified nucleolar proteins that regulate the exit from M phase and cytokinesis, including factors involved in pre-rRNA processing. Our study suggests that in plant cells, nucleolar functions might interact with the processes of regulated cell divisions and influence the selection of longitudinal division planes during embryogenesis. PMID:17616738

Wernher von Braun

NASA Image and Video Library

2004-04-15

A pioneer of America's space program, Dr. von Braun stands by the five F-1 engines of the Saturn V launch vehicle. This Saturn V vehicle is an actual test vehicle which has been displayed at the U.S. Space Rocket Center in Huntsville, Alabama. Designed and developed by Rocketdyne under the direction of the Marshall Space Flight Center, a cluster of five F-1 engines was mounted on the Saturn V S-IC (first) stage. The engines measured 19-feet tall by 12.5-feet at the nozzle exit and burned 15 tons of liquid oxygen and kerosene each second to produce 7,500,000 pounds of thrust. The S-IC stage is the first stage, or booster, of a 364-foot long rocket that ultimately took astronauts to the Moon.

Experimental Evaluation of a Low Emissions High Performance Duct Burner for Variable Cycle Engines (VCE)

NASA Technical Reports Server (NTRS)

Lohmann, R. P.; Mador, R. J.

1979-01-01

An evaluation was conducted with a three stage Vorbix duct burner to determine the performance and emissions characteristics of the concept and to refine the configuration to provide acceptable durability and operational characteristics for its use in the variable cycle engine (VCE) testbed program. The tests were conducted at representative takeoff, transonic climb, and supersonic cruise inlet conditions for the VSCE-502B study engine. The test stand, the emissions sampling and analysis equipment, and the supporting flow visualization rigs are described. The performance parameters including the fuel-air ratio, the combustion efficiency/exit temperature, thrust efficiency, and gaseous emissions calculations are defined. The test procedures are reviewed and the results are discussed.

Pulse Detonation Rocket Engine Research at NASA Marshall

NASA Technical Reports Server (NTRS)

Morris, Christopher I.

2003-01-01

This viewgraph representation provides an overview of research being conducted on Pulse Detonation Rocket Engines (PDRE) by the Propulsion Research Center (PRC) at the Marshall Space Flight Center. PDREs have a theoretical thermodynamic advantage over Steady-State Rocket Engines (SSREs) although unsteady blowdown processes complicate effective use of this advantage in practice; PRE is engaged in a fundamental study of PDRE gas dynamics to improve understanding of performance issues. Topics covered include: simplified PDRE cycle, comparison of PDRE and SSRE performance, numerical modeling of quasi 1-D rocket flows, time-accurate thrust calculations, finite-rate chemistry effects in nozzles, effect of F-R chemistry on specific impulse, effect of F-R chemistry on exit species mole fractions and PDRE performance optimization studies.

Experimental Clean Combustor Program (ECCP), phase 3. [commercial aircraft turbofan engine tests with double annular combustor

NASA Technical Reports Server (NTRS)

Gleason, C. C.; Bahr, D. W.

1979-01-01

A double annular advanced technology combustor with low pollutant emission levels was evaluated in a series of CF6-50 engine tests. Engine lightoff was readily obtained and no difficulties were encountered with combustor staging. Engine acceleration and deceleration were smooth, responsive and essentially the same as those obtainable with the CF6-50 combustor. The emission reductions obtained in carbon monoxide, hydrocarbons, and nitrogen oxide levels were 55, 95, and 30 percent, respectively, at an idle power setting of 3.3 percent of takeoff power on an EPA parameter basis. Acceptable smoke levels were also obtained. The exit temperature distribution of the combustor was found to be its major performance deficiency. In all other important combustion system performance aspects, the combustor was found to be generally satisfactory.

Convoluted nozzle design for the RL10 derivative 2B engine

NASA Technical Reports Server (NTRS)

1985-01-01

The convoluted nozzle is a conventional refractory metal nozzle extension that is formed with a portion of the nozzle convoluted to show the extendible nozzle within the length of the rocket engine. The convoluted nozzle (CN) was deployed by a system of four gas driven actuators. For spacecraft applications the optimum CN may be self-deployed by internal pressure retained, during deployment, by a jettisonable exit closure. The convoluted nozzle is included in a study of extendible nozzles for the RL10 Engine Derivative 2B for use in an early orbit transfer vehicle (OTV). Four extendible nozzle configurations for the RL10-2B engine were evaluated. Three configurations of the two position nozzle were studied including a hydrogen dump cooled metal nozzle and radiation cooled nozzles of refractory metal and carbon/carbon composite construction respectively.

Flow in out-of-plane double S-bends

NASA Technical Reports Server (NTRS)

Schmidt, M. C.; Whitelaw, J. H.; Yianneskis, M.

1987-01-01

An experimental investigation of developing flows through a combination of out-of-plane S-bend ducts was conducted to gain insight into the redirection of flow in geometries similar to those encountered in practical aircraft wing-root intake ducts. The present double S-bend was fabricated by placing previously investigated S-ducts and S-diffusers in series and with perpendicular planes of symmetry. Laser-Doppler anemometry was employed to measure the three components of mean velocity, the corresponding rms quantities, and Reynolds stresses in the rectangular cross-section ducts. Due to limited optical access, only two mean and rms velocity components were resolved in the circular cross-section ducts. The velocity measurements were complemented by wall static pressure measurements. The data indicates that the flows at the exit are complex and asymmetric. Secondary flows generated by the pressure field in the first S-duct are complemented or counteracted by the secondary flows produced by the area expansion and the curvature of the S-diffuser. The results indicate the dominance of the inlet conditions and geometry upon the development of secondary flows and demonstrate that the flows are predominantly pressure-controlled. The pressure distribution caused by the duct geometry determines the direction and magnitude of the bulk flow while the turbulence dictates the mixing characteristics and profiles in the near wall region.

Conceptual study of space plane powered by hypersonic airbreathing propulsion system

NASA Astrophysics Data System (ADS)

Maita, Masataka; Ohkami, Yoshiaki; Yamanaka, Tatsuo; Mori, Takashige

1990-10-01

The paper describes the investigations of aerospace plane concept, conducted by the National Aerospace Laboratory (NAL) of Japan, with particular attention given to a concept which integrates a scram/liquid air cycle engine (LACE) hypersonic propulsion system fueling with slush hydrogen. The key requirements in achieving the space plane using scram/LACE propulsion system are described along with the mission requirements and the vehicle characteristics. Typical outputs of SSTO analysis are presented.

A simple analytical method to estimate all exit parameters of a cross-flow air dehumidifier using liquid desiccant.

PubMed

Bassuoni, M M

2014-03-01

The dehumidifier is a key component in liquid desiccant air-conditioning systems. Analytical solutions have more advantages than numerical solutions in studying the dehumidifier performance parameters. This paper presents the performance results of exit parameters from an analytical model of an adiabatic cross-flow liquid desiccant air dehumidifier. Calcium chloride is used as desiccant material in this investigation. A program performing the analytical solution is developed using the engineering equation solver software. Good accuracy has been found between analytical solution and reliable experimental results with a maximum deviation of +6.63% and -5.65% in the moisture removal rate. The method developed here can be used in the quick prediction of the dehumidifier performance. The exit parameters from the dehumidifier are evaluated under the effects of variables such as air temperature and humidity, desiccant temperature and concentration, and air to desiccant flow rates. The results show that hot humid air and desiccant concentration have the greatest impact on the performance of the dehumidifier. The moisture removal rate is decreased with increasing both air inlet temperature and desiccant temperature while increases with increasing air to solution mass ratio, inlet desiccant concentration, and inlet air humidity ratio.

Alternate Propulsion Subsystem Concepts Tripropellant Comparison Study

NASA Technical Reports Server (NTRS)

Levack, Daniel

1995-01-01

A study was conducted under MSFC contract NAS8-39210 to compare tripropellant and bipropellant engine configurations for the SSTO mission. The objective was to produce an 'apples-to-apples' comparison to isolate the effects of design implementation, designing company, year of design, or technologies included from the basic tripropellant/bipropellant comparison. Consequently, identical technologies were included (e.g., jet pumps) and the same design groundrules and practices were used. Engine power cycles were examined as were turbomachinery/preburner arrangements for each cycle. The bipropellant approach and two tripropellant approaches were separately optimized in terms of operating parameters: exit pressures, mixture ratios, thrust splits, etc. This briefing presents the results of the study including engine weights for both tripropellant and bipropellant engines; dry vehicle weight performance for a range of engine chamber pressures; discusses the basis for the results; examines vehicle performance due to engine cycles and the margin characteristics of various cycles; and identifies technologies with significant payoffs for this application.

Preliminary design of propulsion system for V/STOL research and technology aircraft

NASA Technical Reports Server (NTRS)

1977-01-01

The V/STOL Research and Technology Aircraft (RTA)propulsion system design effort is limited to components of the lift/cruise engines, turboshaft engine modifications, lift fan assembly, and propulsion system performance generation. The uninstalled total net thrust with all engines and fans operating at intermediate power was 37,114 pounds. Uninstalled system total net thrust was 27,102 pounds when one lift/cruise is inoperative. Components have lives above the 500 hours of the RTA duty cycle. The L/C engine used in a fixed nacelle has the cross shaft forward of the reduction gear whereas the cross shaft is aft of the reduction gear in a tilt nacelle L/C engine. The lift/cruise gearbox contains components and technologies from other DDA engines. The rotor has a 62-inch diameter and contains 22 composite blades that have a hub/tip ratio of 0.454. The blade pitch change mechanism contains hydraulic and mechanical redundancy. The lift fan assembly is completely self-contained including oil cooling in 10 exit vanes.

Optical diagnostics integrated with laser spark delivery system

DOEpatents

Yalin, Azer [Fort Collins, CO; Willson, Bryan [Fort Collins, CO; Defoort, Morgan [Fort Collins, CO; Joshi, Sachin [Fort Collins, CO; Reynolds, Adam [Fort Collins, CO

2008-09-02

A spark delivery system for generating a spark using a laser beam is provided, and includes a laser light source and a laser delivery assembly. The laser delivery assembly includes a hollow fiber and a launch assembly comprising launch focusing optics to input the laser beam in the hollow fiber. The laser delivery assembly further includes exit focusing optics that demagnify an exit beam of laser light from the hollow fiber, thereby increasing the intensity of the laser beam and creating a spark. Other embodiments use a fiber laser to generate a spark. Embodiments of the present invention may be used to create a spark in an engine. Yet other embodiments include collecting light from the spark or a flame resulting from the spark and conveying the light for diagnostics. Methods of using the spark delivery systems and diagnostic systems are provided.

Fiber laser coupled optical spark delivery system

DOEpatents

Yalin, Azer [Fort Collins, CO; Willson, Bryan [Fort Collins, CO; Defoort, Morgan [Fort Collins, CO; Joshi, Sachin [Fort Collins, CO; Reynolds, Adam [Fort Collins, CO

2008-03-04

A spark delivery system for generating a spark using a laser beam is provided, and includes a laser light source and a laser delivery assembly. The laser delivery assembly includes a hollow fiber and a launch assembly comprising launch focusing optics to input the laser beam in the hollow fiber. The laser delivery assembly further includes exit focusing optics that demagnify an exit beam of laser light from the hollow fiber, thereby increasing the intensity of the laser beam and creating a spark. Other embodiments use a fiber laser to generate a spark. Embodiments of the present invention may be used to create a spark in an engine. Yet other embodiments include collecting light from the spark or a flame resulting from the spark and conveying the light for diagnostics. Methods of using the spark delivery systems and diagnostic systems are provided.

Experimental results for a two-dimensional supersonic inlet used as a thrust deflecting nozzle

NASA Technical Reports Server (NTRS)

Johns, Albert L.; Burstadt, Paul L.

1984-01-01

Nearly all supersonic V/STOL aircraft concepts are dependent on the thrust deflecting capability of a nozzle. In one unique concept, referred to as the reverse flow dual fan, not only is there a thrust deflecting nozzle for the fan and core engine exit flow, but because of the way the propulsion system operates during vertical takeoff and landing, the supersonic inlet is also used as a thrust deflecting nozzle. This paper presents results of an experimental study to evaluate the performance of a supersonic inlet used as a thrust deflecting nozzle for this reverse flow dual fan concept. Results are presented in terms of nozzle thrust coefficient and thrust vector angle for a number of inlet/nozzle configurations. Flow visualization and nozzle exit flow survey results are also shown.

Computational Analyses in Support of Sub-scale Diffuser Testing for the A-3 Facility. Part 1; Steady Predictions

NASA Technical Reports Server (NTRS)

Allgood, Daniel C.; Graham, Jason S.; Ahuja, Vineet; Hosangadi, Ashvin

2008-01-01

Simulation technology can play an important role in rocket engine test facility design and development by assessing risks, providing analysis of dynamic pressure and thermal loads, identifying failure modes and predicting anomalous behavior of critical systems. Advanced numerical tools assume greater significance in supporting testing and design of high altitude testing facilities and plume induced testing environments of high thrust engines because of the greater inter-dependence and synergy in the functioning of the different sub-systems. This is especially true for facilities such as the proposed A-3 facility at NASA SSC because of a challenging operating envelope linked to variable throttle conditions at relatively low chamber pressures. Facility designs in this case will require a complex network of diffuser ducts, steam ejector trains, fast operating valves, cooling water systems and flow diverters that need to be characterized for steady state performance. In this paper, we will demonstrate with the use of CFD analyses s advanced capability to evaluate supersonic diffuser and steam ejector performance in a sub-scale A-3 facility at NASA Stennis Space Center (SSC) where extensive testing was performed. Furthermore, the focus in this paper relates to modeling of critical sub-systems and components used in facilities such as the A-3 facility. The work here will address deficiencies in empirical models and current CFD analyses that are used for design of supersonic diffusers/turning vanes/ejectors as well as analyses for confined plumes and venting processes. The primary areas that will be addressed are: (1) supersonic diffuser performance including analyses of thermal loads (2) accurate shock capturing in the diffuser duct; (3) effect of turning duct on the performance of the facility (4) prediction of mass flow rates and performance classification for steam ejectors (5) comparisons with test data from sub-scale diffuser testing and assessment of confidence levels in CFD based flowpath modeling of the facility. The analyses tools used here expand on the multi-element unstructured CFD which has been tailored and validated for impingement dynamics of dry plumes, complex valve/feed systems, and high pressure propellant delivery systems used in engine and component test stands at NASA SSC. The analyses performed in the evaluation of the sub-scale diffuser facility explored several important factors that influence modeling and understanding of facility operation such as (a) importance of modeling the facility with Real Gas approximation, (b) approximating the cluster of steam ejector nozzles as a single annular nozzle, (c) existence of mixed subsonic/supersonic flow downstream of the turning duct, and (d) inadequacy of two-equation turbulence models in predicting the correct pressurization in the turning duct and expansion of the second stage steam ejectors. The procedure used for modeling the facility was as follows: (i) The engine, test cell and first stage ejectors were simulated with an axisymmetric approximation (ii) the turning duct, second stage ejectors and the piping downstream of the second stage ejectors were analyzed with a three-dimensional simulation utilizing a half-plane symmetry approximation. The solution i.e. primitive variables such as pressure, velocity components, temperature and turbulence quantities were passed from the first computational domain and specified as a supersonic boundary condition for the second simulation. (iii) The third domain comprised of the exit diffuser and the region in the vicinity of the facility (primary included to get the correct shock structure at the exit of the facility and entrainment characteristics). The first set of simulations comprising the engine, test cell and first stage ejectors was carried out both as a turbulent real gas calculation as well as a turbulent perfect gas calculation. A comparison for the two cases (Real Turbulent and Perfect gas turbulent) of the Ma Number distribution and temperature distributions are shown in Figures 1 and 2 respectively. The Mach Number distribution shows small yet distinct differences between the two cases such as locations of shocks/shock reflections and a slightly different impingement point on the wall of the diffuser from the expansion at the exit of the nozzle. Similarly the temperature distribution indicates different flow recirculation patterns in the test cell. Both cases capture all the essential flow phenomena such as the shock-boundary layer interaction, plume expansion, expansion of the first stage ejectors, mixing between the engine plume and the first stage ejector flow and pressurization due to the first stage ejectors. The final paper will discuss thermal loads on the walls of the diffuser and cooling mechanisms investigated.

The relation between knee flexion angle and anterior cruciate ligament femoral tunnel characteristics: a cadaveric study comparing a standard and a far anteromedial portal.

PubMed

Alavekios, Damon; Peterson, Alexander; Patton, John; McGarry, Michelle H; Lee, Thay Q

2014-11-01

The purpose of this study was to compare the anterior cruciate ligament (ACL) femoral tunnel characteristics between 2 common arthroscopic portals used for ACL reconstruction, a standard anteromedial portal and a far anteromedial portal. Seven cadaveric knees were used. A 1.25-mm Kirschner wire was drilled through the center of the ACL femoral footprint and through the distal femur from the standard anteromedial and far anteromedial portals at knee flexion angles of 100°, 120°, and 140°. No formal tunnels were drilled. Each tunnel exit point was marked with a colored pin. After all tunnels were created, the specimens were digitized with a MicroScribe device (Revware, Raleigh, NC) to measure the tunnel length; distance to the posterior femoral cortical wall (posterior cortical margin); and tunnel orientation in the sagittal, coronal, and axial planes. The standard anteromedial portal resulted in a longer tunnel length, a less horizontal tunnel in the coronal plane, and a greater posterior cortical margin compared with the far anteromedial portal at all knee flexion angles. For both portal locations, the tunnel length and posterior cortical margin increased, and the tunnel position became more horizontal in the coronal plane, more anterior in the sagittal plane, and less horizontal in the transverse plane as knee flexion increased. Portal position affects femoral tunnel characteristics, with results favoring the more laterally positioned standard anteromedial portal at all flexion angles. Increasing the knee flexion angle leads to a longer femoral tunnel length and posterior femoral cortical margin with either portal position. Understanding how portal positioning and knee flexion angle affect femoral tunnel orientation and characteristics may lead to improved surgical outcomes after ACL reconstruction. Published by Elsevier Inc.

4. COMPLETE X15 VEHICLE TEST STAND, DETAIL OF THRUST MOUNTING ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

4. COMPLETE X-15 VEHICLE TEST STAND, DETAIL OF THRUST MOUNTING STRUCTURE AT ENGINE END OF PLANE. - Edwards Air Force Base, X-15 Engine Test Complex, Rocket Engine & Complete X-15 Vehicle Test Stands, Rogers Dry Lake, east of runway between North Base & South Base, Boron, Kern County, CA

Studies on an aerial propellant transfer space plane (APTSP)

NASA Astrophysics Data System (ADS)

Jayan, N.; Biju Kumar, K. S.; Gupta, Anish Kumar; Kashyap, Akhilesh Kumar; Venkatraman, Kartik; Mathew, Joseph; Mukunda, H. S.

2004-04-01

This paper presents a study of a fully reusable earth-to-orbit launch vehicle concept with horizontal take-off and landing, employing a turbojet engine for low speed, and a rocket for high-speed acceleration and space operations. This concept uses existing technology to the maximum possible extent, thereby reducing development time, cost and effort. It uses the experience in aerial filling of military aircrafts for propellant filling at an altitude of 13 km at a flight speed of M=0.85. Aerial filling of propellant reduces the take-off weight significantly thereby minimizing the structural weight of the vehicle. The vehicle takes off horizontally and uses turbojet engines till the end of the propellant filling operation. The rocket engines provide thrust for the next phase till the injection of a satellite at LEO. A sensitivity analysis of the mission with respect to rocket engine specific impulse and overall vehicle structural factor is also presented in this paper. A conceptual design of space plane with a payload capability of 10 ton to LEO is carried out. The study shows that the realization of an aerial propellant transfer space plane is possible with limited development of new technology thus reducing the demands on the finances required for achieving the objectives.

The Exit Gradient As a Measure of Groundwater Dependency of Watershed Ecosystem Services

NASA Astrophysics Data System (ADS)

Faulkner, B. R.; Canfield, T. J.; Justin, G. F.

2014-12-01

Flux of groundwater to surface water is often of great interest for the determination of the groundwater dependency of ecosystem services, such as maintenance of wetlands and of baseflow as a contributor to stream channel storage. It is difficult to measure. Most methods are based on coarse mass balance estimates or seepage meters. One drawback of these methods is they are not entirely spatially explicit. The exit gradient is commonly used in engineering studies of hydraulic structures affected by groundwater flow. It can be simply defined in the groundwater modeling context as the ratio of the difference between the computed head and the land surface elevation, for each computational cell, to the thickness of the cell, as it varies in space. When combined with calibrated groundwater flow models, it also has the potential to be useful in watershed scale evaluations of groundwater dependency in a spatially explicit way. We have taken advantage of calibrated models for the Calapooia watershed, Oregon, to map exit gradients for the watershed. Streams in the Calapooia are hydraulically well connected with groundwater. Not surprisingly, we found large variations in exit gradients between wet and dry season model runs, supporting the notion of stream expansion, as observed in the field, which may have a substantial influence on water quality. We have mapped the exit gradients in the wet and dry seasons, and compared them to regions which have been mapped in wetland surveys. Those classified as Palustrine types fell largest in the area of contribution from groundwater. In many cases, substantially high exit gradients, even on average, did not correspond to mapped wetland areas, yet nutrient retention ecosystem services may still be playing a role in these areas. The results also reinforce the notion of the importance of baseflow to maintenance of stream flow, even in the dry summer season in this Temperate/Mediterranean climate. Exit gradient mapping is a simple, yet potentially very useful and underutilized tool for measuring groundwater dependency in watershed scale ecosystem services studies, and could potentially be used to predict effects due to groundwater stresses resulting from water withdrawals. This is an abstract of a proposed presentation and does not necessarily reflect EPA policy.

Hypersonic research engine project. Phase 2: Aerothermodynamic integration model development

NASA Technical Reports Server (NTRS)

Jilly, L. F. (Editor)

1970-01-01

The analytical effort was directed towards (1) completing the design of the combustor exit instrumentation assembly, (2) analyzing the coolant flow distribution of the cowl leading edge tip section, (3) determining effects of purge gas pressure on AIM performance analysis, and (4) analyzing heat transfer and associated stress problems related to the cowl leading edge tip section and the nozzle shroud assembly for test conditions.

Aerodynamic Effects of High Turbulence Intensity on a Variable-Speed Power-Turbine Blade with Large Incidence and Reynolds Number Variations

NASA Technical Reports Server (NTRS)

Flegel, Ashlie B.; Giel, Paul W.; Welch, Gerard E.

2014-01-01

The effects of high inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. These results are compared to previous measurements made in a low turbulence environment. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The current study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Assessing the effects of turbulence at these large incidence and Reynolds number variations complements the existing database. Downstream total pressure and exit angle data were acquired for 10 incidence angles ranging from +15.8deg to -51.0deg. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.12×10(exp 5) to 2.12×10(exp 6) and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 8 to 15 percent for the current study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitch/yaw probe located in a survey plane 7 percent axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the turbulence grid imposes.

Aerodynamic Effects of High Turbulence Intensity on a Variable-Speed Power-Turbine Blade With Large Incidence and Reynolds Number Variations

NASA Technical Reports Server (NTRS)

Flegel, Ashlie B.; Giel, Paul W.; Welch, Gerard E.

2014-01-01

The effects of high inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. These results are compared to previous measurements made in a low turbulence environment. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The current study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Assessing the effects of turbulence at these large incidence and Reynolds number variations complements the existing database. Downstream total pressure and exit angle data were acquired for 10 incidence angles ranging from +15.8deg to -51.0deg. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.12×10(exp 5) to 2.12×10(exp 6) and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 8 to 15 percent for the current study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitch/yaw probe located in a survey plane 7 percent axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the turbulence grid imposes.

Radiation of sound from unflanged cylindrical ducts

NASA Technical Reports Server (NTRS)

Hartharan, S. L.; Bayliss, A.

1983-01-01

Calculations of sound radiated from unflanged cylindrical ducts are presented. The numerical simulation models the problem of an aero-engine inlet. The time dependent linearized Euler equations are solved from a state of rest until a harmonic solution is attained. A fourth order accurate finite difference scheme is used and solutions are obtained from a fully vectorized Cyber-203 computer program. Cases of both plane waves and spin modes are treated. Spin modes model the sound generated by a turbofan engine. Boundary conditions for both plane waves and spin modes are treated. Solutions obtained are compared with experiments conducted at NASA Langley Research Center.

Multiresonant Composite Optical Nanoantennas by Out-of-plane Plasmonic Engineering.

PubMed

Song, Junyeob; Zhou, Wei

2018-06-27

Optical nanoantennas can concentrate light and enhance light-matter interactions in subwavelength domain, which is useful for photodetection, light emission, optical biosensing, and spectroscopy. However, conventional optical nanoantennas operating at a single wavelength band are not suitable for multiband applications. Here, we propose and exploit an out-of-plane plasmonic engineering strategy to design and create composite optical nanoantennas that can support multiple nanolocalized modes at different resonant wavelengths. These multiresonant composite nanoantennas are composed of vertically stacked building blocks of metal-insulator-metal loop nanoantennas. Studies of multiresonant composite nanoantennas demonstrate that the number of supported modes depends on the number of vertically stacked building blocks and the resonant wavelengths of individual modes are tunable by controlling the out-of-plane geometries of their building blocks. In addition, numerical studies show that the resonant wavelengths of individual modes in composite nanoantennas can deviate from the optical response of building blocks due to hybridization of magnetic modes in neighboring building blocks. Using Au nanohole arrays as deposition masks to fabricate arrays of multilayered composite nanoantennas, we experimentally demonstrate their multiresonant optical properties in good agreement with theory predictions. These studies show that out-of-plane engineered multiresonant composite nanoantennas can provide new opportunities for fundamental nanophotonics research and practical applications involving optical multiband operations, such as multiphoton process, broadband solar energy conversion, and wavelength-multiplexed optical system.

System and method for optical fiber based image acquisition suitable for use in turbine engines

DOEpatents

Baleine, Erwan; A V, Varun; Zombo, Paul J.; Varghese, Zubin

2017-05-16

A system and a method for image acquisition suitable for use in a turbine engine are disclosed. Light received from a field of view in an object plane is projected onto an image plane through an optical modulation device and is transferred through an image conduit to a sensor array. The sensor array generates a set of sampled image signals in a sensing basis based on light received from the image conduit. Finally, the sampled image signals are transformed from the sensing basis to a representation basis and a set of estimated image signals are generated therefrom. The estimated image signals are used for reconstructing an image and/or a motion-video of a region of interest within a turbine engine.

A 3-D Navier-Stokes CFD study of turbojet/ramjet nozzle plume interactions at Mach 3.0 and comparison with data

NASA Technical Reports Server (NTRS)

Chang, Ing; Hunter, Louis G.

1995-01-01

Advanced airbreathing propulsion systems used in Mach 4-6 mission scenarios, usually consist of a single integrated turboramjet or as in this study, a turbojet housed in an upper bay with a separate ramjet housed in a lower bay. As the engines transition from turbojet to ramjet, there is an operational envelope where both engines operate simultaneously. One nozzle concept under consideration has a common nozzle, where the plumes from the turbojet and ramjet interact with one another as they expand to ambient conditions. In this paper, the two plumes interact at the end of a common 2-D cowl, when they both reach an approximate Mach 3.0 condition and then jointly expand to Mach 3.6 at the common nozzle exit plane. At this condition, the turbojet engine operated at a higher NPR than the ramjet, where the turbojet overpowers the ramjet plume, deflecting it approximately 12 degrees downward and in turn the turbojet plume is deflected 6 degrees upward. In the process, shocks were formed at the deflections and a shear layer formed at the confluence of the two jets. This particular case was experimentally tested and the data used to compare with the PARC3D code with k-kl two equation turbulence model. The 2-D and 3-D centerline CFD solutions are in good agreement, but as the CFD solutions approach the outer sidewall, a slight variance occurs. The outer wall boundary layers are thin and do not present much of an interaction, however, where the confluence interaction shocks interact with the thin boundary layer on the outer wall, strong vortices run down each shock causing substantial disturbances in the boundary layer. These disturbances amplify somewhat as they propagate downstream axially from the confluence point. The nozzle coefficient (CFG) is reduced 1/2 percent as a result of this sidewall interaction, from 0.9850 to 0.9807. This three-dimensional reduction is in better agreement with the experimental value of 0.9790.

Simultaneous Determination of Thermal Conductivity and Thermal Diffusivity of Food and Agricultural Materials Using a Transient Plane-Source Method

USDA-ARS?s Scientific Manuscript database

Thermal conductivity and thermal diffusivity are two important physical properties essential for designing any food engineering processes. Recently a new transient plane-source method was developed to measure a variety of materials, but its application in foods has not been documented. Therefore, ...

Purdue Plane Structures Analyzer II : a computerized wood engineering system

Treesearch

S. K. Suddarth; R. W. Wolfe

1984-01-01

The Purdue Plane Structures Analyzer (PPSA) is a computer program developed specifically for the analysis of wood structures. It uses recognized analysis procedures, in conjunction with recommendations of the 1982 National Design Specification for Wood Construction, to determine stresses and deflections of wood trusses and frames. The program offers several options for...

Optimization of a Turboprop UAV for Maximum Loiter and Specific Power Using Genetic Algorithm

NASA Astrophysics Data System (ADS)

Dinc, Ali

2016-09-01

In this study, a genuine code was developed for optimization of selected parameters of a turboprop engine for an unmanned aerial vehicle (UAV) by employing elitist genetic algorithm. First, preliminary sizing of a UAV and its turboprop engine was done, by the code in a given mission profile. Secondly, single and multi-objective optimization were done for selected engine parameters to maximize loiter duration of UAV or specific power of engine or both. In single objective optimization, as first case, UAV loiter time was improved with an increase of 17.5% from baseline in given boundaries or constraints of compressor pressure ratio and burner exit temperature. In second case, specific power was enhanced by 12.3% from baseline. In multi-objective optimization case, where previous two objectives are considered together, loiter time and specific power were increased by 14.2% and 9.7% from baseline respectively, for the same constraints.

Numerical Modeling of Pulse Detonation Rocket Engine Gasdynamics and Performance

NASA Technical Reports Server (NTRS)

2003-01-01

This paper presents viewgraphs on the numerical modeling of pulse detonation rocket engines (PDRE), with an emphasis on the Gasdynamics and performance analysis of these engines. The topics include: 1) Performance Analysis of PDREs; 2) Simplified PDRE Cycle; 3) Comparison of PDRE and Steady-State Rocket Engines (SSRE) Performance; 4) Numerical Modeling of Quasi 1-D Rocket Flows; 5) Specific PDRE Geometries Studied; 6) Time-Accurate Thrust Calculations; 7) PDRE Performance (Geometries A B C and D); 8) PDRE Blowdown Gasdynamics (Geom. A B C and D); 9) PDRE Geometry Performance Comparison; 10) PDRE Blowdown Time (Geom. A B C and D); 11) Specific SSRE Geometry Studied; 12) Effect of F-R Chemistry on SSRE Performance; 13) PDRE/SSRE Performance Comparison; 14) PDRE Performance Study; 15) Grid Resolution Study; and 16) Effect of F-R Chemistry on SSRE Exit Species Mole Fractions.

Experimental clean combustor program: Diesel no. 2 fuel addendum, phase 3

NASA Technical Reports Server (NTRS)

Gleason, C. C.; Bahr, D. W.

1979-01-01

A CF6-50 engine equipped with an advanced, low emission, double annular combustor was operated 4.8 hours with No. 2 diesel fuel. Fourteen steady-state operating conditions ranging from idle to full power were investigated. Engine/combustor performance and exhaust emissions were obtained and compared to JF-5 fueled test results. With one exception, fuel effects were very small and in agreement with previously obtained combustor test rig results. At high power operating condition, the two fuels produced virtually the same peak metal temperatures and exhaust emission levels. At low power operating conditions, where only the pilot stage was fueled, smoke levels tended to be significantly higher with No. 2 diesel fuel. Additional development of this combustor concept is needed in the areas of exit temperature distribution, engine fuel control, and exhaust emission levels before it can be considered for production engine use.

Use of cooling air heat exchangers as replacements for hot section strategic materials

NASA Technical Reports Server (NTRS)

Gauntner, J. W.

1983-01-01

Because of financial and political constraints, strategic aerospace materials required for the hot section of future engines might be in short supply. As an alternative to these strategic materials, this study examines the use of a cooling air heat exchanger in combination with less advanced hot section materials. Cycle calculations are presented for future turbofan systems with overall pressure ratios to 65, bypass ratios near 13, and combustor exit temperatures to 3260 R. These calculations quantify the effect on TSFC of using a decreased materials technology in a turbofan system. The calculations show that the cooling air heat exchanger enables the feasibility of these engines.

Application of a Computer Model to Various Specifications of Fuel Injection System for DI Diesel Engines

NASA Astrophysics Data System (ADS)

Yamanishi, Manabu

A combined experimental and computational investigation was performed in order to evaluate the effects of various design parameters of an in-line injection pump on the nozzle exit characteristics for DI diesel engines. Measurements of the pump chamber pressure and the delivery valve lift were included for validation by using specially designed transducers installed inside the pump. The results confirm that the simulation model is capable of predicting the pump operation for all the different designs investigated pump operating conditions. Following the successful validation of this model, parametric studies were performed which allow for improved fuel injection system design.

A reflection mechanism for aft fan tone noise from turbofan engines

NASA Astrophysics Data System (ADS)

Topol, D. A.; Holhubner, S. C.; Mathews, D. C.

1987-10-01

A fan tone noise mechanism is proposed which results from reflections from the fan of forward propagating rotor wake/fan exit guide vane interaction tone noise. These fan noise tones are often more dominant out of the rear than out of the front of an engine. To simulate this effect a simple qualitative prediction model was formulated and a scaled model test program was conducted. Results from each of these investigations are compared with each other and with full-scale engine data. These comparisons substantiate the potential importance of this mechanism. Further support is provided by mode measurement data from full-scale testing. This study concluded that for certain vane/blade ratios and tip Mach numbers the contribution of the reflection noise mechanism is significant.

Physically elastic analysis of a cylindrical ring as a unit cell of a complete composite under applied stress in the complex plane using cubic polynomials

NASA Astrophysics Data System (ADS)

Monfared, Vahid

2018-03-01

Elastic analysis is analytically presented to predict the behaviors of the stress and displacement components in the cylindrical ring as a unit cell of a complete composite under applied stress in the complex plane using cubic polynomials. This analysis is based on the complex computation of the stress functions in the complex plane and polar coordinates. Also, suitable boundary conditions are considered and assumed to analyze along with the equilibrium equations and bi-harmonic equation. This method has some important applications in many fields of engineering such as mechanical, civil and material engineering generally. One of the applications of this research work is in composite design and designing the cylindrical devices under various loadings. Finally, it is founded that the convergence and accuracy of the results are suitable and acceptable through comparing the results.

Kinematic and Kinetic Evaluation of High Speed Backward Running

DTIC Science & Technology

1999-06-30

Designed using Perform Pro , WHS/DIOR, Oct 94 KINEMATIC AND KINETIC EVALUATION OF HIGH SPEED BACKWARD RUNNING by ALAN WAYNE ARATA A DISSERTATION...Project Manager, Engineering Division, Kelly Air Force Base, Texas, 1983-86 AWARDS AND HONORS: All-American, 50yd Freestyle , 1979 Winner, Rocky...redirection #include <stdlib.h> // for exit #include <iomanip.h> // for set precision #include <string.h> // for string copy const int NUMPOINTS

Combustor Simulation

NASA Technical Reports Server (NTRS)

Norris, Andrew

2003-01-01

The goal was to perform 3D simulation of GE90 combustor, as part of full turbofan engine simulation. Requirements of high fidelity as well as fast turn-around time require massively parallel code. National Combustion Code (NCC) was chosen for this task as supports up to 999 processors and includes state-of-the-art combustion models. Also required is ability to take inlet conditions from compressor code and give exit conditions to turbine code.

A practical approach to the development of aircraft GTE's noise suppression system on the base of fiber optic sensors

NASA Astrophysics Data System (ADS)

Vinogradov, Vasiliy Yu.; Morozov, Oleg G.; Morozov, Gennady A.; Sakhabutdinov, Airat Zh.; Nureev, Ilnur I.; Kuznetsov, Artem A.; Faskhutdinov, Lenar M.; Sarvarova, Lutsia M.

2017-04-01

In this paper, we consider a number of different methods that form the modern approach to the development of aircraft GTE's noise suppression systems at service conditions. The herein-presented efficient noise suppression system on the base of fiber optic sensors makes it possible to reduce pulsations at the exhaust nozzle exit and noise levels at the engine outlet section.

Airfoil for a gas turbine engine

DOEpatents

Liang, George [Palm City, FL

2011-05-24

An airfoil is provided for a turbine of a gas turbine engine. The airfoil comprises: an outer structure comprising a first wall including a leading edge, a trailing edge, a pressure side, and a suction side; an inner structure comprising a second wall spaced from the first wall and at least one intermediate wall; and structure extending between the first and second walls so as to define first and second gaps between the first and second walls. The second wall and the at least one intermediate wall define at least one pressure side supply cavity and at least one suction side supply cavity. The second wall may include at least one first opening near the leading edge of the first wall. The first opening may extend from the at least one pressure side supply cavity to the first gap. The second wall may further comprise at least one second opening near the trailing edge of the outer structure. The second opening may extend from the at least one suction side supply cavity to the second gap. The first wall may comprise at least one first exit opening extending from the first gap through the pressure side of the first wall and at least one second exit opening extending from the second gap through the suction side of the second wall.

A Study on the Performance of the Split Reaction Water Turbine with Guide Ribs

NASA Astrophysics Data System (ADS)

Allen, Deuel H.; Villanueva, Eliseo P.

2015-09-01

The development of technologies that make use of renewable energy is of great significance presently. A new kind of turbine called Split Reaction Water Turbine (SRWT) using PVC pipes as material is a major contribution towards harnessing the energy potentials of small stream low head water resources. SRWTs of diameter to height ratio (D/H = 110 cm/160 cm) were tested at the MSU-IIT College of Engineering Fluid Engineering Laboratory. Data on volumetric flow and pressure head at the turbine inlet of the SRWT were recorded using National Instrument Data Processing System using LabView software. In later experiments, guide ribs were installed at the vane of the exit nozzles in order to determine the difference in the performance of the ribbed and the non-ribbed SRWT. Simulations of the running SRWT were conducted using SOLIDWORKS software. Results of the simulations aided in the thorough analyses of the data from the experimental runs. A comparison of data from the ribbed and non-ribbed SRWT shows that guide ribs were effective in directing the momentum of the exiting water to improve the speed of rotation. In this study, the increase in the speed of the Split Reaction Water Turbine was as much as 46%.

KENNEDY SPACE CENTER, FLA. - Standing inside Discovery’s payload bay, Carol Scott (right), lead orbiter engineer, talks about her job as part of a special feature for the KSC Web. With his back to the camera is Bill Kallus, Media manager in the KSC Web Studio. Behind Scott can be seen the open hatch of the airlock, which provides support functions such as airlock depressurization and repressurization, extravehicular activity equipment recharge, liquid-cooled garment water cooling, EVA equipment checkout, donning and communications. The outer hatch isolates the airlock from the unpressurized payload bay when closed and permits the EVA crew members to exit from the airlock to the payload bay when open.

NASA Image and Video Library

2004-01-22

KENNEDY SPACE CENTER, FLA. - Standing inside Discovery’s payload bay, Carol Scott (right), lead orbiter engineer, talks about her job as part of a special feature for the KSC Web. With his back to the camera is Bill Kallus, Media manager in the KSC Web Studio. Behind Scott can be seen the open hatch of the airlock, which provides support functions such as airlock depressurization and repressurization, extravehicular activity equipment recharge, liquid-cooled garment water cooling, EVA equipment checkout, donning and communications. The outer hatch isolates the airlock from the unpressurized payload bay when closed and permits the EVA crew members to exit from the airlock to the payload bay when open.

An experimental study of the natural noise in the Transit-M hypersonic wind tunnel

NASA Astrophysics Data System (ADS)

Gromyko, Yu. V.; Polivanov, P. A.; Sidorenko, A. A.; Buntin, D. A.; Maslov, A. A.

2013-12-01

Tests were performed at Mach number M = 6. Observations of flow pulsations were performed with the help of high-frequency measurements of pressure and heat fluxes and with the help of hot-wire anemometry. Measurements were carried out in a broad range of Re1 numbers at various orientations and locations of sensors with respect to the nozzle exit plane. It is shown that the Transit-M noise varies weakly over the longitudinal coordinate, whereas on moving along the radial coordinate, some increase of the fluctuation level is observed due to approaching the mixing layer. Also, the distribution of axial asymmetry of the noise level was obtained.

A comparison of arcjet plume properties to model predictions

NASA Technical Reports Server (NTRS)

Cappelli, M. A.; Liebeskind, J. G.; Hanson, R. K.; Butler, G. W.; King, D. Q.

1993-01-01

This paper describes an experimental study of the plasma plume properties of a 1 kW class hydrogen arcjet thruster and the comparison of measured temperature and velocity field to model predictions. The experiments are based on laser-induced fluorescence excitation of the Balmer-alpha transition. The model is based on a single-fluid magnetohydrodynamic description of the flow originally developed to predict arcjet thruster performance. Excellent agreement between model predictions and experimental velocity is found, despite the complex nature of the flow. Measured and predicted exit plane temperatures are in disagreement by as much as 2000K over a range of operating conditions. The possible sources for this discrepancy are discussed.

Mode transition induced by the magnetic field gradient in Hall thrusters

NASA Astrophysics Data System (ADS)

Han, Liang; Wei, Liqiu; Yu, Daren

2016-09-01

A mode transition phenomenon was found in Hall thrusters, which was induced by the increase of the magnetic field gradient. In the transition process, we observed experimentally that there have been obvious changes in the oscillation, the mean value of the discharge current, the thrust, the anode efficiency, and the plume pattern. The shifting and compression of the high magnetic field causes the electron density in the discharge channel to decrease and the ionization zone to move towards the exit plane. This also corresponds to a low atom density in the discharge channel, resulting in a loss of stability of the ionization at a high magnetic field gradient, which presents the transition of the discharge mode.

Aerodynamic Performance of a Compact, High Work-Factor Centrifugal Compressor at the Stage and Subcomponent Level

NASA Technical Reports Server (NTRS)

Braunscheidel, Edward P.; Welch, Gerard E.; Skoch, Gary J.; Medic, Gorazd; Sharma, Om P.

2014-01-01

The measured aerodynamic performance of a compact, high work factor, single-stage centrifugal compressor, comprising an impeller, diffuser, 90-bend, and exit guide vane (EGV), is reported. Performance levels are based on steady-state total-pressure and total-temperature rake and angularity-probe data acquired at key machine rating planes during recent testing at NASA Glenn Research Center. Aerodynamic performance at the stage level are reported for operation between 70 to 105 of design corrected speed, with subcomponent (impeller, diffuser, and exitguide-vane) detailed flow field measurements presented and discussed at the 100 design-speed condition. Individual component losses from measurements are compared with pre-test predictions on a limited basis.

The effect of initial flow nonuniformity on second-stage fuel injection and combustion in a supersonic duct. [supersonic combustion ramjet engine

NASA Technical Reports Server (NTRS)

Russin, W. R.

1975-01-01

The effects of flow nonuniformity on second-stage hydrogen fuel injection and combustion in supersonic flow were evaluated. The first case, second-stage fuel injection into a uniform duct flow, produced data indicating that fuel mixing is considerably slower than estimates based on an empirical mixing correlation. The second-case, two-stage fuel injection (or second-stage fuel injection into a nonuniform duct flow), produced a large interaction between stages with extensive flow separation. For this case the measured wall pressure, heat transfer, and amount of reaction at the duct exit were significantly greater than estimates based on the mixing correlation. Substantially more second-stage fuel burned in the second case than in the first case. Overall effects of unmixedness/chemical kinetics were found not to be significant at the exit for stoichiometric fuel injection.

Fiber coupled optical spark delivery system

DOEpatents

Yalin, Azer; Willson, Bryan; Defoort, Morgan

2008-08-12

A spark delivery system for generating a spark using a laser beam is provided, the spark delivery system including a laser light source and a laser delivery assembly. The laser delivery assembly includes a hollow fiber and a launch assembly comprising launch focusing optics to input the laser beam in the hollow fiber. In addition, the laser delivery assembly includes exit focusing optics that demagnify an exit beam of laser light from the hollow fiber, thereby increasing the intensity of the laser beam and creating a spark. In accordance with embodiments of the present invention, the assembly may be used to create a spark in a combustion engine. In accordance with other embodiments of the present invention, a method of using the spark delivery system is provided. In addition, a method of choosing an appropriate fiber for creating a spark using a laser beam is also presented.

Experimental evaluation of a cooled radial-inflow turbine

NASA Technical Reports Server (NTRS)

Tirres, Lizet; Dicicco, L. D.; Nowlin, Brent C.

1993-01-01

Two 14.4 inch tip diameter rotors were installed and tested in the Small Engines Component Turbine Facility (SECTF) at NASA Lewis Research Center. The rotors, a solid and a cooled version of a radial-inflow turbine, were tested with a 15 vane stat or over a set of rotational speeds ranging from 80 to 120 percent design speed (17,500 to 21,500 rpm). The total-to-total stage pressure ratios ranged from 2.5 to 5.5. The data obtained at the equivalent conditions using the solid version of the rotor are presented with the cooled rotor data. A Reynolds number of 381,000 was maintained for both rotors, whose stages had a design mass flow of 4.0 lbm/sec, a design work level of 59.61 Btu/lbm, and a design efficiency of 87 percent. The results include mass flow data, turbine torque, turbine exit flow angles, stage efficiency, and rotor inlet and exit surveys.

Experimental Evaluation of a Cooled Radial-inflow Turbine

NASA Technical Reports Server (NTRS)

Tirres, Lizet; Dicicco, L. Danielle; Nowlin, Brent C.

1993-01-01

Two 14.4 inch tip diameter rotors were installed and tested in the Small Engines Component Turbine Facility (SECTF) at NASA Lewis Research Center. The rotors, a solid and a cooled version of a radial-inflow turbine, were tested with a 15 vane stat or over a set of rotational speeds ranging from 80 to 120 percent design speed (17,500 to 21,500 rpm). The total-to-total stage pressure ratios ranged from 2.5 to 5.5. The data obtained at the equivalent conditions using the solid version of the rotor are presented with the cooled rotor data. A Reynolds number of 381,000 was maintained for both rotors, whose stages had a design mass flow of 4.0 Ibm/sec, a design work level of 59.61 Btu/lbm, and a design efficiency of 87 percent. The results include mass flow data, turbine torque, turbine exit flow angles, stage efficiency, and rotor inlet and exit surveys.

Theoretical investigation of operation modes of MHD generators for energy-bypass engines

NASA Astrophysics Data System (ADS)

Tang, Jingfeng; Li, Nan; Yu, Daren

2014-12-01

A MHD generator with different arrangements of electromagnetic fields will lead the generator working in three modes. A quasi-one-dimensional approximation is used for the model of the MHD generator to analyze the inner mechanism of operation modes. For the MHD generator with a uniform constant magnetic field, a specific critical electric field E cr is required to decelerate a supersonic entrance flow into a subsonic exit flow. Otherwise, the generator works in a steady mode with a larger electric field than E cr in which a steady supersonic flow is provided at the exit, or the generator works in a choked mode with a smaller electric field than E cr in which the supersonic entrance flow is choked in the channel. The detailed flow field characteristics in different operation modes are discussed, demonstrating the relationship of operation modes with electromagnetic fields.

Experimental evaluation of expendable supersonic nozzle concepts

NASA Technical Reports Server (NTRS)

Baker, V.; Kwon, O.; Vittal, B.; Berrier, B.; Re, R.

1990-01-01

Exhaust nozzles for expendable supersonic turbojet engine missile propulsion systems are required to be simple, short and compact, in addition to having good broad-range thrust-minus-drag performance. A series of convergent-divergent nozzle scale model configurations were designed and wind tunnel tested for a wide range of free stream Mach numbers and nozzle pressure ratios. The models included fixed geometry and simple variable exit area concepts. The experimental and analytical results show that the fixed geometry configurations tested have inferior off-design thrust-minus-drag performance in the transonic Mach range. A simple variable exit area configuration called the Axi-Quad nozzle, combining features of both axisymmetric and two-dimensional convergent-divergent nozzles, performed well over a broad range of operating conditions. Analytical predictions of the flow pattern as well as overall performance of the nozzles, using a fully viscous, compressible CFD code, compared very well with the test data.

Turbine exhaust diffuser with a gas jet producing a coanda effect flow control

DOEpatents

Orosa, John; Montgomery, Matthew

2014-02-11

An exhaust diffuser system and method for a turbine engine includes an inner boundary and an outer boundary with a flow path defined therebetween. The inner boundary is defined at least in part by a hub structure that has an upstream end and a downstream end. The outer boundary may include a region in which the outer boundary extends radially inward toward the hub structure and may direct at least a portion of an exhaust flow in the diffuser toward the hub structure. The hub structure includes at least one jet exit located on the hub structure adjacent to the upstream end of the tail cone. The jet exit discharges a flow of gas substantially tangential to an outer surface of the tail cone to produce a Coanda effect and direct a portion of the exhaust flow in the diffuser toward the inner boundary.

Illustrating Thermodynamic Concepts Using a Hero's Engine

NASA Astrophysics Data System (ADS)

Muiño, Pedro L.; Hodgson, James R.

2000-05-01

A modified Hero's engine is used to illustrate concepts of thermodynamics and engineering design suitable for introductory chemistry courses and more advanced physical chemistry courses. The engine is a boiler made of Pyrex with two off-center nozzles. Upon boiling, the vapor exits the nozzles, creating two opposite, off-center forces that result in a circular motion by the engine around the vertical axis. The engine is suspended from a horizontal bar by means of two parallel threads. The rotation of the engine results in the twisting of the threads, with two important effects: the engine is raised vertically, and potential energy is stored in the coiling of the threads. When the engine is raised, it is removed from the heating source. This stops the boiling. The stored potential energy is then released into kinetic energy; that is, the threads uncoil, and the engine rotates in the opposite direction. This lowers the engine into the flame, so the water resumes boiling and the engine can be raised again. This cycle continues until all the liquid water is vaporized. This demonstration is suitable to illustrate concepts like gas expansion, gas cooling through expansion (Joule-Thompson experiment), conversion of heat to work, interconversion between kinetic energy and potential energy, and feedback mechanisms.

Fine-scale features in the far-field of a turbulent jet

NASA Astrophysics Data System (ADS)

Buxton, Oliver; Ganapathisubramani, Bharathram

2008-11-01

The structure of a fully turbulent axisymmetric jet, at Reynolds number based on jet exit conditions of 5000, is investigated with cinematographic (1 kHz) stereoscopic PIV in a plane normal to the jet axis. Taylor's hypothesis is employed to calculate all three velocity gradients in the axial direction. The technique's resolution allows all terms of the velocity gradient tensor, hence strain rate tensor and kinetic energy dissipation, to be computed at each point within the plane. The data reveals that the vorticity field is dominated by high enstrophy tube-like structures. Conversely, the dissipation field appears to consist of sheet-like structures. Several criteria for isolating these strongly swirling vortical structures from the background turbulence were employed. One such technique involves isolating points in which the velocity gradient tensor has a real and a pair of complex conjugate eigenvectors. Once identified, the alignment of the various structures with relation to the vorticity vector and the real velocity gradient tensor eigenvector is investigated. The effect of the strain field on the geometry of the structures is also examined.

Solution of the three-dimensional Helmholtz equation with nonlocal boundary conditions

NASA Technical Reports Server (NTRS)

Hodge, Steve L.; Zorumski, William E.; Watson, Willie R.

1995-01-01

The Helmholtz equation is solved within a three-dimensional rectangular duct with a nonlocal radiation boundary condition at the duct exit plane. This condition accurately models the acoustic admittance at an arbitrarily-located computational boundary plane. A linear system of equations is constructed with second-order central differences for the Helmholtz operator and second-order backward differences for both local admittance conditions and the gradient term in the nonlocal radiation boundary condition. The resulting matrix equation is large, sparse, and non-Hermitian. The size and structure of the matrix makes direct solution techniques impractical; as a result, a nonstationary iterative technique is used for its solution. The theory behind the nonstationary technique is reviewed, and numerical results are presented for radiation from both a point source and a planar acoustic source. The solutions with the nonlocal boundary conditions are invariant to the location of the computational boundary, and the same nonlocal conditions are valid for all solutions. The nonlocal conditions thus provide a means of minimizing the size of three-dimensional computational domains.

In-cylinder flows of a motored four-stroke engine with flat-crown and slightly concave-crown pistons

DOE Office of Scientific and Technical Information (OSTI.GOV)

Huang, R.F.; Yang, H.S.; Yeh, C.-N.

2008-04-15

The temporal and spatial evolution processes of the in-cylinder flow structures and turbulence intensities in the symmetry and offset planes of a motored four-valve, four-stroke engine during the intake and compression strokes are diagnosed by using a particle image velocimeter. Two pistons of different crown shapes (flat-crown and slightly concave-crown pistons) are studied. The inception, establishment, and evolution of the tumbling vortical flow structures during the intake and compression strokes are clearly depicted. Quantitative strengths of the rotating vortical flow motions are presented by a dimensionless parameter, the tumble ratio, which can represent the mean angular velocity of the vorticesmore » in the target plane. The turbulence intensity of the in-cylinder flow is also calculated by using the measured time-varying velocity data. The results show that the flat-crown piston induces higher bulk-averaged tumble ratio and turbulence intensity than the slightly concave-crown piston does because the tumble ratio and turbulence generated by the flat-crown piston in the offset planes during the compression stroke are particularly large. The engine with the flat-crown piston also presents larger torque and power outputs and lower hydrocarbon emission than that with the slightly concave-crown piston. This might be caused by the enhanced combustion in the engine cylinder due to the stronger tumble ratio and turbulence intensity. (author)« less

Strain engineered barium strontium titanate for tunable thin film resonators

DOE Office of Scientific and Technical Information (OSTI.GOV)

Khassaf, H.; Khakpash, N.; Sun, F.

2014-05-19

Piezoelectric properties of epitaxial (001) barium strontium titanate (BST) films are computed as functions of composition, misfit strain, and temperature using a non-linear thermodynamic model. Results show that through adjusting in-plane strains, a highly adaptive rhombohedral ferroelectric phase can be stabilized at room temperature with outstanding piezoelectric response exceeding those of lead based piezoceramics. Furthermore, by adjusting the composition and the in-plane misfit, an electrically tunable piezoelectric response can be obtained in the paraelectric state. These findings indicate that strain engineered BST films can be utilized in the development of electrically tunable and switchable surface and bulk acoustic wave resonators.

Joint US/Russia TU-144 Engine Ground Tests

NASA Technical Reports Server (NTRS)

Acosta, Waldo A.; Balser, Jeffrey S.; McCartney, Timothy P.; Richter, Charles A.; Woike, Mark R.

1997-01-01

Two engine research experiments were recently completed in Moscow, Russia using an engine from the Tu-144 supersonic transport airplane. This was a joint project between the United States and Russia. Personnel from the NASA Lewis Research Center, General Electric Aircraft Engines, Pratt & Whitney, the Tupolev Design Bureau, and EBP Aircraft LTD worked together as a team to overcome the many technical and cultural challenges. The objective was to obtain large scale inlet data that could be used in the development of a supersonic inlet system for a future High Speed Civil Transport (HSCT). The-first experiment studied the impact of typical inlet structures that have trailing edges in close proximity to the inlet/engine interface plane on the flow characteristics at that plane. The inlet structure simulated the subsonic diffuser of a supersonic inlet using a bifurcated splitter design. The centerbody maximum diameter was designed to permit choking and slightly supercritical operation. The second experiment measured the reflective characteristics of the engine face to incoming perturbations of pressure amplitude. The basic test rig from the first experiment was used with a longer spacer equipped with fast actuated doors. All the objectives set forth at the beginning of the project were met.

Numerical Investigation of the Interaction between Mainstream and Tip Shroud Leakage Flow in a 2-Stage Low Pressure Turbine

NASA Astrophysics Data System (ADS)

Jia, Wei; Liu, Huoxing

2014-06-01

The pressing demand for future advanced gas turbine requires to identify the losses in a turbine and to understand the physical mechanisms producing them. In low pressure turbines with shrouded blades, a large portion of these losses is generated by tip shroud leakage flow and associated interaction. For this reason, shroud leakage losses are generally grouped into the losses of leakage flow itself and the losses caused by the interaction between leakage flow and mainstream. In order to evaluate the influence of shroud leakage flow and related losses on turbine performance, computational investigations for a 2-stage low pressure turbine is presented and discussed in this paper. Three dimensional steady multistage calculations using mixing plane approach were performed including detailed tip shroud geometry. Results showed that turbines with shrouded blades have an obvious advantage over unshrouded ones in terms of aerodynamic performance. A loss mechanism breakdown analysis demonstrated that the leakage loss is the main contributor in the first stage while mixing loss dominates in the second stage. Due to the blade-to-blade pressure gradient, both inlet and exit cavity present non-uniform leakage injection and extraction. The flow in the exit cavity is filled with cavity vortex, leakage jet attached to the cavity wall and recirculation zone induced by main flow ingestion. Furthermore, radial gap and exit cavity size of tip shroud have a major effect on the yaw angle near the tip region in the main flow. Therefore, a full calculation of shroud leakage flow is necessary in turbine performance analysis and the shroud geometric features need to be considered during turbine design process.

Approximate heating analysis for the windward-symmetry plane of Shuttle-like bodies at large angle of attack

NASA Technical Reports Server (NTRS)

Zoby, E. V.

1981-01-01

An engineering method has been developed for computing the windward-symmetry plane convective heat-transfer rates on Shuttle-like vehicles at large angles of attack. The engineering code includes an approximate inviscid flowfield technique, laminar and turbulent heating-rate expressions, an approximation to account for the variable-entropy effects on the surface heating and the concept of an equivalent axisymmetric body to model the windward-ray flowfields of Shuttle-like vehicles at angles of attack from 25 to 45 degrees. The engineering method is validated by comparing computed heating results with corresponding experimental data measured on Shuttle and advanced transportation models over a wide range of flow conditions and angles of attack from 25 to 40 degrees and also with results of existing prediction techniques. The comparisons are in good agreement.

Multifocal planes head-mounted displays.

PubMed

Rolland, J P; Krueger, M W; Goon, A

2000-07-01

Stereoscopic head-mounted displays (HMD's) provide an effective capability to create dynamic virtual environments. For a user of such environments, virtual objects would be displayed ideally at the appropriate distances, and natural concordant accommodation and convergence would be provided. Under such image display conditions, the user perceives these objects as if they were objects in a real environment. Current HMD technology requires convergent eye movements. However, it is currently limited by fixed visual accommodation, which is inconsistent with real-world vision. A prototype multiplanar volumetric projection display based on a stack of laminated planes was built for medical visualization as discussed in a paper presented at a 1999 Advanced Research Projects Agency workshop (Sullivan, Advanced Research Projects Agency, Arlington, Va., 1999). We show how such technology can be engineered to create a set of virtual planes appropriately configured in visual space to suppress conflicts of convergence and accommodation in HMD's. Although some scanning mechanism could be employed to create a set of desirable planes from a two-dimensional conventional display, multiplanar technology accomplishes such function with no moving parts. Based on optical principles and human vision, we present a comprehensive investigation of the engineering specification of multiplanar technology for integration in HMD's. Using selected human visual acuity and stereoacuity criteria, we show that the display requires at most 27 equally spaced planes, which is within the capability of current research and development display devices, located within a maximal 26-mm-wide stack. We further show that the necessary in-plane resolution is of the order of 5 microm.

Effect of tube processing methods on the texture and grain boundary characteristics of 14YWT nanostructured ferritic alloys

DOE Office of Scientific and Technical Information (OSTI.GOV)

Aydogan, E.; Pal, S.; Anderoglu, O.

In this paper, texture and microstructure of tubes and plates fabricated from a nanostructured ferritic alloy (14YWT), produced either by spray forming followed by hydrostatic extrusion (Process I) or hot extrusion and cross-rolling a plate followed by hydrostatic tube extrusion (Process II) have been characterized in terms of their effects on texture and grain boundary character. Hydrostatic extrusion results in a combination of plane strain and shear deformations which generate low intensity α- and γ-fiber components of {001}<110> and {111}<110> together with a weak ζ-fiber component of {011}<211> and {011}<011>. In contrast, multi-step plane strain deformation by hot extrusion andmore » cross-rolling of the plate leads to a strong texture component of {001}<110> together with a weaker {111}<112> component. Although the total strains are similar, shear dominated deformation leads to much lower texture indexes compared to plane strain deformations. Further, the texture intensity decreases after hydrostatic extrusion of the alloy plate formed by plane strain deformation, due to a lower number of activated slip systems during shear dominated deformation. Finally and notably, hot extruded and cross-rolled plate subjected to plane strain deformation to ~50% engineering strain creates only a modest population of low angle grain boundaries, compared to the much larger population observed following the combination of plane strain and shear deformation of ~44% engineering strain resulting from subsequent hydrostatic extrusion.« less

Effect of tube processing methods on the texture and grain boundary characteristics of 14YWT nanostructured ferritic alloys

DOE PAGES

Aydogan, E.; Pal, S.; Anderoglu, O.; ...

2016-03-08

In this paper, texture and microstructure of tubes and plates fabricated from a nanostructured ferritic alloy (14YWT), produced either by spray forming followed by hydrostatic extrusion (Process I) or hot extrusion and cross-rolling a plate followed by hydrostatic tube extrusion (Process II) have been characterized in terms of their effects on texture and grain boundary character. Hydrostatic extrusion results in a combination of plane strain and shear deformations which generate low intensity α- and γ-fiber components of {001}<110> and {111}<110> together with a weak ζ-fiber component of {011}<211> and {011}<011>. In contrast, multi-step plane strain deformation by hot extrusion andmore » cross-rolling of the plate leads to a strong texture component of {001}<110> together with a weaker {111}<112> component. Although the total strains are similar, shear dominated deformation leads to much lower texture indexes compared to plane strain deformations. Further, the texture intensity decreases after hydrostatic extrusion of the alloy plate formed by plane strain deformation, due to a lower number of activated slip systems during shear dominated deformation. Finally and notably, hot extruded and cross-rolled plate subjected to plane strain deformation to ~50% engineering strain creates only a modest population of low angle grain boundaries, compared to the much larger population observed following the combination of plane strain and shear deformation of ~44% engineering strain resulting from subsequent hydrostatic extrusion.« less

Noise Reduction with Lobed Mixers: Nozzle-Length and Free-Jet Speed Effects

NASA Technical Reports Server (NTRS)

Mengle, Vinod G.; Dalton, William N.; Bridges, James C.; Boyd, Kathy C.

1997-01-01

Acoustic test results are presented for 1/4th-scaled nozzles with internal lobed mixers used for reduction of subsonic jet noise of turbofan engines with bypass ratio above 5 and jet speeds up to 830 ft/s. One coaxial and three forced lobe mixers were tested with variations in lobe penetration, cut-outs in lobe-sidewall, lobe number and nozzle-length. Measured exit flow profiles and thrusts are used to assist the inferences from acoustic data. It is observed that lobed mixers reduce the low-frequency noise due to more uniformly mixed exit flow; but they may also increase the high-frequency noise at peak perceived noise (PNL) angle and angles upstream of it due to enhanced mixing inside the nozzle. Cut-outs and low lobe penetration reduce the annoying portion of the spectrum but lead to less uniform exit flow. Due to the dominance of internal duct noise in unscalloped, high-penetration mixers their noise is not reduced as much with increase in free-jet speed as that of coaxial or cut-out lobed mixers. The latter two mixers also show no change in PNL over the wide range of nozzle-lengths tested because most of their noise sources are outside the nozzle; whereas, the former show an increase in noise with decrease in nozzle-length.

A simple analytical method to estimate all exit parameters of a cross-flow air dehumidifier using liquid desiccant

PubMed Central

Bassuoni, M.M.

2013-01-01

The dehumidifier is a key component in liquid desiccant air-conditioning systems. Analytical solutions have more advantages than numerical solutions in studying the dehumidifier performance parameters. This paper presents the performance results of exit parameters from an analytical model of an adiabatic cross-flow liquid desiccant air dehumidifier. Calcium chloride is used as desiccant material in this investigation. A program performing the analytical solution is developed using the engineering equation solver software. Good accuracy has been found between analytical solution and reliable experimental results with a maximum deviation of +6.63% and −5.65% in the moisture removal rate. The method developed here can be used in the quick prediction of the dehumidifier performance. The exit parameters from the dehumidifier are evaluated under the effects of variables such as air temperature and humidity, desiccant temperature and concentration, and air to desiccant flow rates. The results show that hot humid air and desiccant concentration have the greatest impact on the performance of the dehumidifier. The moisture removal rate is decreased with increasing both air inlet temperature and desiccant temperature while increases with increasing air to solution mass ratio, inlet desiccant concentration, and inlet air humidity ratio. PMID:25685485

Design of Astrometric Mission (JASMINE) by Applying Model Driven System Engineering

NASA Astrophysics Data System (ADS)

Yamada, Y.; Miyashita, H.; Nakamura, H.; Suenaga, K.; Kamiyoshi, S.; Tsuiki, A.

2010-12-01

We are planning space astrometric satellite mission named JASMINE. The target accuracy of parallaxes in JASMINE observation is 10 micro arc second, which corresponds to 1 nm scale on the focal plane. It is very hard to measure the 1 nm scale deformation of focal plane. Eventually, we need to add the deformation to the observation equations when estimating stellar astrometric parameters, which requires considering many factors such as instrument models and observation data analysis. In this situation, because the observation equations become more complex, we may reduce the stability of the hardware, nevertheless, we require more samplings due to the lack of rigidity of each estimation. This mission imposes a number of trades-offs in the engineering choices and then decide the optimal design from a number of candidates. In order to efficiently support such decisions, we apply Model Driven Systems Engineering (MDSE), which improves the efficiency of the engineering by revealing and formalizing requirements, specifications, and designs to find a good balance among various trade-offs.

Force of resistance to pipeline pulling in plane and volumetrically curved wells

NASA Astrophysics Data System (ADS)

Toropov, V. S.; Toropov, S. Yu; Toropov, E. S.

2018-05-01

A method has been developed for calculating the component of the pulling force of a pipeline, arising from the well curvature in one or several planes, with the assumption that the pipeline is ballasted by filling with water or otherwise until zero buoyancy in the drilling mud is reached. This paper shows that when calculating this force, one can neglect the effect of sections with zero curvature. In the other case, if buoyancy of the pipeline is other than zero, the resistance force in the curvilinear sections should be calculated taking into account the difference between the normal components of the buoyancy force and weight. In the paper, it is proved that without taking into account resistance forces from the viscosity of the drilling mud, if buoyancy of the pipeline is zero, the total resistance force is independent of the length of the pipe and is determined by the angle equal to the sum of the entry angle and the exit angle of the pipeline to the day surface. For the case of the well curvature in several planes, it is proposed to perform the calculation of such volumetrically curved well by the central angle of the well profile. Analytical dependences are obtained that allow calculating the pulling force for well profiles with a variable curvature radius, i.e. at different angles of deviation between the drill pipes along the well profile.

Exhaust-stack nozzle area and shape for individual cylinder exhaust-gas jet-propulsion system

NASA Technical Reports Server (NTRS)

Pinkel, Benjamin; Turner, Richard; Voss, Fred; Humble, Leroy V

1943-01-01

This report presents the results of an investigation conducted on the effect of exhaust-stack nozzle area, shape, and length on engine power, jet thrust, and gain in net thrust (engine propeller plus jet). Single-cylinder engine data were obtained using three straight stacks 25, 44, and 108 inches in length; an S-shaped stack, a 90 degree bend, a 180 degree bend, and a short straight stack having a closed branch faired into it. Each stack was fitted with nozzles varying in exit area from 0.91 square inch to the unrestricted area of the stack of 4.20 square inches. The engine was generally operated over a range of engine speeds from 1300 to 2100 r.p.m, inlet-manifold pressures from 22 to 30 inches of mercury absolute, and a fuel-air ratio of 0.08. The loss in engine power, the jet thrust, and the gain in net thrust are correlated in terms of several simple parameters. An example is given for determining the optimum nozzle area and the overall net thrust.

High-speed monodisperse droplet generation by ultrasonically controlled micro-jet breakup

NASA Astrophysics Data System (ADS)

Frommhold, Philipp Erhard; Lippert, Alexander; Holsteyns, Frank Ludwig; Mettin, Robert

2014-04-01

A liquid jet that is ejected from a nozzle into air will disintegrate into drops via the well-known Plateau-Rayleigh instability within a certain range of Ohnesorge and Reynolds numbers. With the focus on the micrometer scale, we investigate the control of this process by superimposing a suitable ultrasonic signal, which causes the jet to break up into a very precise train of monodisperse droplets. The jet leaves a pressurized container of liquid via a small orifice of about 20 μm diameter. The break-up process and the emerging droplets are recorded via high-speed imaging. An extended parameter study of exit speed and ultrasonic frequency is carried out for deionized water to evaluate the jet's state and the subsequent generation of monodisperse droplets. Maximum exit velocities obtained reach almost 120 m s-1, and frequencies have been applied up to 1.8 MHz. Functionality of the method is confirmed for five additional liquids for moderate jet velocities 38 m s-1. For the uncontrolled jet disintegration, the drop size spectra revealed broad distributions and downstream drop growth by collision, while the acoustic control generated monodisperse droplets with a standard deviation less than 0.5 %. By adjustment of the acoustic excitation frequency, drop diameters could be tuned continuously from about 30 to 50 μm for all exit speeds. Good agreement to former experiments and theoretical approaches is found for the relation of overpressure and jet exit speed, and for the observed stability regions of monodisperse droplet generation in the parameter plane of jet speed and acoustic excitation frequency. Fitting of two free parameters of the general theory to the liquids and nozzles used is found to yield an even higher precision. Furthermore, the high-velocity instability limit of regular jet breakup described by von Ohnesorge has been superseded by more than a factor of two without entering the wind-induced instability regime, and monodisperse droplet generation was always achievable. Thus, the reliable and robust realization of tunable high-speed monodisperse micro-droplet trains is demonstrated. Some implication for applications is discussed.

CFD Study of Turbo-Ramjet Interactions in Hypersonic Airbreathing Propulsion System

NASA Technical Reports Server (NTRS)

Chang, Ing; Hunter, Louis G.

1996-01-01

Advanced airbreathing propulsion systems used in Mach 4-6 mission scenarios, usually involve turbo-ramjet configurations. As the engines transition from turbojet to ramjet, there is an operational envelope where both engines operate simultaneously. In the first phase of our study, an over/under nozzle configuration was analyzed. The two plumes from the turbojet and ramjet interact at the end of a common 2-D cowl, where they both reach an approximate Mach 3.0 condition and then jointly expand to Mach 3.6 at the common nozzle exit plane. For the problem analyzed, the turbojet engine operates at a higher nozzle pressure ratio than the ramjet, causes the turbojet plume overpowers the ramjet plume, deflecting it approximately 12 degrees downward and in turn the turbojet plume is deflected 6 degrees upward. In the process, shocks were formed at the deflections and a shear layer formed at the confluence of the two jets. This particular case was experimentally tested and the data were used to compare with a computational fluid dynamics (CFD) study using the PARC2D code. The CFD results were in good agreement with both static pressure distributions on the cowl separator and on nozzle walls. The thrust coefficients were also in reasonable agreement. In addition, inviscid relationships were developed around the confluence point, where the two exhaust jets meet, and these results compared favorably with the CFD results. In the second phase of our study, a 3-D CFD solution was generated to compare with the 2-D solution. The major difference between the 2-D and 3-D solutions was the interaction of the shock waves, generated by the plume interactions, on the sidewall. When a shock wave interacts with a sidewall and sidewall boundary layer, it is called a glancing shock sidewall interaction. These interactions entrain boundary layer flow down the shockline into a vortical flow pattern. The 3-D plots show the streamlines being entrained down the shockline. The pressure of the flow also decreases slightly as the sidewall is approached. Other difference between the 2-D and 3-D solutions were a lowering of the nozzle thrust coefficient value from 0.9850 (2-D) to 0.9807 (3-D), where the experimental value was 0.9790. In the third phase of our study, a different turbo-ramjet configuration was analyzed. The confluence of a supersonic turbojet and a subsonic ramjet in the turbine based combined-cycle (TBCC) propulsion system was studied by a 2-D CFD code. In the analysis, Mach 1.4 primary turbojet was mixed with the subsonic ramjet secondary flow in an ejector mode operation. Reasonable agreements were obtained with the supplied I-D TBCC solutions. For low downstream backpressure, the Fabri choke condition (Break-Point condition) was observed in the secondary flow within mixing zone. For sufficient high downstream backpressure, the Fabri choke no longer exist, the ramjet flow was reduced and the ejector flow became backpressure dependent. Highly non-uniform flow at ejector exit were observed, indicated that for smooth downstream combustion, the mixing of the two streams probably required some physical devices.

Effect of scapular external rotation on the axillary nerve during the arthroscopic Latarjet procedure: an anatomical investigation.

PubMed

Reinares, Felipe; Werthel, Jean-David; Moraiti, Constantina; Valenti, Philippe

2017-10-01

The first purpose of this study is to measure the distance between the axillary nerve and the exit point of K-wires placed retrograde through the glenoid in the setting of an arthroscopic Latarjet procedure. The second objective is to evaluate whether manual external rotation of the scapula alters that distance. In seven fresh-frozen specimens, two 2.0-mm K-wires were drilled through the glenoid using an arthroscopic Latarjet retrograde glenoid guide. These were drilled into the glenoid at the 7- and 8-o'clock positions (right shoulders) and at the 4- and 5-o'clock positions (left). K-wires were oriented parallel to the glenoid articular surface and perpendicular to the long superoinferior axis of the glenoid, 7 mm medial to the joint surface. Two independent evaluators measured the distances between the axillary nerve and the exit point of the K-wires in the horizontal plane (AKHS for the superior K-wire and AKHI for the inferior K-wire) and in the vertical plane (AKV). Measurements were taken with the scapula left free and were repeated with the scapula placed at 15° and 30° of external rotation. With the scapula left free, scapular external rotation was 34° ± 2.3°. In this position, the AKHS was 2.5 ± 1.6, 6.3 ± 1.2 mm at 15° of external rotation (ER) and 11.4 ± 1.4 mm at 30° ER. The AKHI distance was 0.37 ± 1.6, 3.4 ± 1.4 and 10.6 ± 2.1 mm, respectively, for the scapula left free, at 15° ER and 30° of ER. The AKV distances were, respectively, 0.12 ± 0.2, 4.9 ± 1.6 and 9.9 ± 1.7 mm. The increase in all distances was statistically significant (p < 0.001). Increasing scapular external rotation significantly increases the distance between the axillary nerve and the exit point of the K-wires, increasing the margin of safety during this procedure. Therefore, increased external rotation of the scapula could be an effective tool to decrease the risk of iatrogenic axillary nerve injury. Cadaveric study, Level V.

40 CFR 94.203 - Application for certification.

Code of Federal Regulations, 2010 CFR

2010-07-01

... Section 94.203 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR PROGRAMS... § 94.109(d) for Category 3 engines. Small-volume manufacturers may omit measurement and reporting of... application of the engine (e.g., used to propel planing vessels, use to propel vessels with variable-pitch...

Large-Area Permanent-Magnet ECR Plasma Source

NASA Technical Reports Server (NTRS)

Foster, John E.

2007-01-01

A 40-cm-diameter plasma device has been developed as a source of ions for material-processing and ion-thruster applications. Like the device described in the immediately preceding article, this device utilizes electron cyclotron resonance (ECR) excited by microwave power in a magnetic field to generate a plasma in an electrodeless (noncontact) manner and without need for an electrically insulating, microwave-transmissive window at the source. Hence, this device offers the same advantages of electrodeless, windowless design - low contamination and long operational life. The device generates a uniform, high-density plasma capable of sustaining uniform ion-current densities at its exit plane while operating at low pressure [<10(exp -4) torr (less than about 1.3 10(exp -2) Pa)] and input power <200 W at a frequency of 2.45 GHz. Though the prototype model operates at 2.45 GHz, operation at higher frequencies can be achieved by straightforward modification to the input microwave waveguide. Higher frequency operation may be desirable in those applications that require even higher background plasma densities. In the design of this ECR plasma source, there are no cumbersome, power-hungry electromagnets. The magnetic field in this device is generated by a permanent-magnet circuit that is optimized to generate resonance surfaces. The microwave power is injected on the centerline of the device. The resulting discharge plasma jumps into a "high mode" when the input power rises above 150 W. This mode is associated with elevated plasma density and high uniformity. The large area and uniformity of the plasma and the low operating pressure are well suited for such material-processing applications as etching and deposition on large silicon wafers. The high exit-plane ion-current density makes it possible to attain a high rate of etching or deposition. The plasma potential is <3 V low enough that there is little likelihood of sputtering, which, in plasma processing, is undesired because it is associated with erosion and contamination. The electron temperature is low and does not vary appreciably with power.

6th Annual CMMI Technology Conference and User Group

DTIC Science & Technology

2006-11-17

Operationally Oriented; Customer Focused Proven Approach – Level of Detail Beginner Decision Table (DT) is a tabular representation with tailoring options to...written to reflect the experience of the author Software Engineering led the process charge in the ’80s – Used Flowcharts – CASE tools – “data...Postpo ned PCR. Verification Steps • EPG configuration audits • EPG configuration status reports Flowcharts and Entry, Task, Verification and eXit

Computational Investigation of Combustion Instabilities in a Laboratory-Scale LDI Gas Turbine Engine

DTIC Science & Technology

2013-06-01

combustor by the insertion of a slotted inlet and an exit nozzle , whereas the reduced geometry is acoustically open. Table 2 Summary of Cases Considered... nozzle located at the right-end surface, an outlet condition is imposed by a characteristic back pressure condition. The fuel spray is injected at the...Computational Mesh visualized around the fuel nozzle and swirler III. Decomposition Methods For Combustion Dynamics Diagnostics To understand the

Advanced and Adaptable Military Propulsion

DTIC Science & Technology

2008-01-22

turbine. The accelerating flow in the turbine environment would mitigate this somewhat (i.e. (favorable) axial pressure gradient vs radial pressure...For instance in a low bypass ratio (0.8) turbofan engine operating at flight Mach numbers ranging from 0.85 to 2.5, the specific fuel consumption...ratio, T8/PT2 - Fan inlet axial Mach No, M2, capability - Compressor exit axial Mach No, M3 - Compressor pressure ratio, PT3/PT2 - Turbine nozzle area

Lattice strain effects on the optical properties of MoS2 nanosheets

PubMed Central

Yang, Lei; Cui, Xudong; Zhang, Jingyu; Wang, Kan; Shen, Meng; Zeng, Shuangshuang; Dayeh, Shadi A.; Feng, Liang; Xiang, Bin

2014-01-01

“Strain engineering” in functional materials has been widely explored to tailor the physical properties of electronic materials and improve their electrical and/or optical properties. Here, we exploit both in plane and out of plane uniaxial tensile strains in MoS2 to modulate its band gap and engineer its optical properties. We utilize X-ray diffraction and cross-sectional transmission electron microscopy to quantify the strains in the as-synthesized MoS2 nanosheets and apply measured shifts of Raman-active modes to confirm lattice strain modification of both the out-of-plane and in-plane phonon vibrations of the MoS2 nanosheets. The induced band gap evolution due to in-plane and out-of-plane tensile stresses is validated by photoluminescence (PL) measurements, promising a potential route for unprecedented manipulation of the physical, electrical and optical properties of MoS2. PMID:25008782

Experimental demonstration of in-plane negative-angle refraction with an array of silicon nanoposts.

PubMed

Wu, Aimin; Li, Hao; Du, Junjie; Ni, Xingjie; Ye, Ziliang; Wang, Yuan; Sheng, Zhen; Zou, Shichang; Gan, Fuwan; Zhang, Xiang; Wang, Xi

2015-03-11

Controlling an optical beam is fundamental in optics. Recently, unique manipulation of optical wavefronts has been successfully demonstrated by metasurfaces. However, these artificially engineered nanostructures have thus far been limited to operate on light beams propagating out-of-plane. The in-plane operation is critical for on-chip photonic applications. Here, we demonstrate an anomalous negative-angle refraction of a light beam propagating along the plane, by designing a thin dielectric array of silicon nanoposts. The circularly polarized dipoles induced by the high-permittivity nanoposts at the scattering resonance significantly shape the wavefront of the light beam and bend it anomalously. The unique capability of a thin line of the nanoposts for manipulating in-plane wavefronts makes the device extremely compact. The low loss all-dielectric structure is compatible with complementary metal-oxide semiconductor technologies, offering an effective solution for in-plane beam steering and routing for on-chip photonics.

Exhaust-Gas Pressure and Temperature Survey of F404-GE-400 Turbofan Engine

NASA Technical Reports Server (NTRS)

Walton, James T.; Burcham, Frank W., Jr.

1986-01-01

An exhaust-gas pressure and temperature survey of the General Electric F404-GE-400 turbofan engine was conducted in the altitude test facility of the NASA Lewis Propulsion System Laboratory. Traversals by a survey rake were made across the exhaust-nozzle exit to measure the pitot pressure and total temperature. Tests were performed at Mach 0.87 and a 24,000-ft altitude and at Mach 0.30 and a 30,000-ft altitude with various power settings from intermediate to maximum afterburning. Data yielded smooth pressure and temperature profiles with maximum jet temperatures approximately 1.4 in. inside the nozzle edge and maximum jet temperatures from 1 to 3 in. inside the edge. A low-pressure region located exactly at engine center was noted. The maximum temperature encountered was 3800 R.

Effects of Structural Flexibility on Aircraft-Engine Mounts

NASA Technical Reports Server (NTRS)

Phillips, W. H.

1986-01-01

Analysis extends technique for design of widely used type of vibration-isolating mounts for aircraft engines, in which rubber mounting pads located in plane behind center of gravity of enginepropeller combination. New analysis treats problem in statics. Results of simple approach useful in providing equations for design of vibrationisolating mounts. Equations applicable in usual situation in which engine-mount structure itself relatively light and placed between large mass of engine and other heavy components of airplane.

High-Temperature Rocket Engine

NASA Technical Reports Server (NTRS)

Schneider, Steven J.; Rosenberg, Sanders D.; Chazen, Melvin L.

1994-01-01

Two rocket engines that operate at temperature of 2,500 K designed to provide thrust for station-keeping adjustments of geosynchronous satellites, for raising and lowering orbits, and for changing orbital planes. Also useful as final propulsion stages of launch vehicles delivering small satellites to low orbits around Earth. With further development, engines used on planetary exploration missions for orbital maneuvers. High-temperature technology of engines adaptable to gas-turbine combustors, ramjets, scramjets, and hot components of many energy-conversion systems.

Pullout of a Rigid Insert Adhesively Bonded to an Elastic Half Plane.

DTIC Science & Technology

1983-12-01

COMMAND UNITED STATES AIR FORCE C-= °84 02 13 071. C,, W % d 6 This document was prepared by the Department of Engineering Mechanics, USAF Academy Faculty...THOMAS E. KULLGREN, Lt Col, USAF Project Engineer /Scientist Professor and Acting Head, Department of Engineering Mechanics KENNETH E. SIEGETH Lt Col...Department of Engineering (Ifapphicable) Mechanics USAFA/DFEM 6c. ADDRESS (City. State and ZIP Code) 7b. ADDRESS (City, Slate and ZIP Code) USAF Academy

Use of cooling air heat exchangers as replacements for hot section strategic materials

NASA Technical Reports Server (NTRS)

Gauntner, J. W.

1983-01-01

Because of financial and political constraints, strategic aerospace materials required for the hot section of future engines might be in short supply. As an alternative to these strategic materials, this study examines the use of a cooling air heat exchanger in combination with less advanced hot section materials. Cycle calculations are presented for future turbofan systems with overall pressure ratios to 65, bypass ratios near 13, and combustor exit temperatures to 3260 R. These calculations quantify the effect on TSFC of using a decreased materials technology in a turbofan system. The calculations show that the cooling air heat exchanger enables the feasibility of these engines. Previously announced in STAR as N83-34946

Air/fuel supply system for use in a gas turbine engine

DOEpatents

Fox, Timothy A; Schilp, Reinhard; Gambacorta, Domenico

2014-06-17

A fuel injector for use in a gas turbine engine combustor assembly. The fuel injector includes a main body and a fuel supply structure. The main body has an inlet end and an outlet end and defines a longitudinal axis extending between the outlet and inlet ends. The main body comprises a plurality of air/fuel passages extending therethrough, each air/fuel passage including an inlet that receives air from a source of air and an outlet. The fuel supply structure communicates with and supplies fuel to the air/fuel passages for providing an air/fuel mixture within each air/fuel passage. The air/fuel mixtures exit the main body through respective air/fuel passage outlets.

Concept study of a hydrogen containment process during nuclear thermal engine ground testing

NASA Astrophysics Data System (ADS)

Wang, Ten-See; Stewart, Eric T.; Canabal, Francisco

A new hydrogen containment process was proposed for ground testing of a nuclear thermal engine. It utilizes two thermophysical steps to contain the hydrogen exhaust. First, the decomposition of hydrogen through oxygen-rich combustion at higher temperature; second, the recombination of remaining hydrogen with radicals at low temperature. This is achieved with two unit operations: an oxygen-rich burner and a tubular heat exchanger. A computational fluid dynamics methodology was used to analyze the entire process on a three-dimensional domain. The computed flammability at the exit of the heat exchanger was less than the lower flammability limit, confirming the hydrogen containment capability of the proposed process.

A combined Eulerian-Lagrangian two-phase flow analysis of SSME HPOTP nozzle plug trajectories. II - Results

NASA Technical Reports Server (NTRS)

Mcconnaughey, P. K.; Garcia, R.; Dejong, F. J.; Sabnis, J. S.; Pribik, D. A.

1989-01-01

An analysis of Space Shuttle Main Engine high-pressure oxygen turbopump nozzle plug trajectories has been performed, using a Lagrangian method to track nozzle plug particles expelled from a turbine through a high Reynolds number flow in a turnaround duct with turning vanes. Axisymmetric and parametric analyses reveal that if nozzle plugs exited the turbine they would probably impact the LOX heat exchanger with impact velocities which are significantly less than the penetration velocity. The finding that only slight to moderate damage will result from nozzle plug failure in flight is supported by the results of a hot-fire engine test with induced nozzle plug failures.

Low Dimensional Study of a Supersonic Multi-Stream Jet Flow

NASA Astrophysics Data System (ADS)

Tenney, Andrew; Berry, Matthew; Aycock-Rizzo, Halley; Glauser, Mark; Lewalle, Jacques

2017-11-01

In this study, the near field of a two stream supersonic jet flow is examined using low dimensional tools. The flow issues from a multi-stream nozzle as described in A near-field investigation of a supersonic, multi-stream jet: locating turbulence mechanisms through velocity and density measurements by Magstadt et al., with the bulk flow Mach number, M1, being 1.6, and the second stream Mach number, M2, reaching the sonic condition. The flow field is visualized using Particle Image Velocimetry (PIV), with frames captured at a rate of 4Hz. Time-resolved pressure measurements are made just aft of the nozzle exit, as well as in the far-field, 86.6 nozzle hydraulic diameters away from the exit plane. The methodologies used in the analysis of this flow include Proper Orthogonal Decomposition (POD), and the continuous wavelet transform. The results from this ``no deck'' case are then compared to those found in the study conducted by Berry et al. From this comparison, we draw conclusions about the effects of the presence of an aft deck on the low dimensional flow description, and near field spectral content. Supported by AFOSR Grant FA9550-15-1-0435, and AFRL, through an SBIR Grant with Spectral Energies, LLC.

Mixing noise reduction for rectangular supersonic jets by nozzle shaping and induced screech mixing

NASA Technical Reports Server (NTRS)

Rice, Edward J.; Raman, Ganesh

1993-01-01

Two methods of mixing noise modification were studied for supersonic jets flowing from rectangular nozzles with an aspect ratio of about five and a small dimension of about 1.4 cm. The first involves nozzle geometry variation using either single (unsymmetrical) or double bevelled (symmetrical) thirty degree cutbacks of the nozzle exit. Both converging (C) and converging-diverging (C-D) versions were tested. The double bevelled C-D nozzle produced a jet mixing noise reduction of about 4 dB compared to a standard rectangular C-D nozzle. In addition all bevelled nozzles produced an upstream shift in peak mixing noise which is conducive to improved attenuation when the nozzle is used in an acoustically treated duct. A large increase in high frequency noise also occurred near the plane of the nozzle exit. Because of near normal incidence, this noise can be easily attenuated with wall treatment. The second approach uses paddles inserted on the edge of the two sides of the jet to induce screech and greatly enhance the jet mixing. Although screech and mixing noise levels are increased, the enhanced mixing moves the source locations upstream and may make an enclosed system more amenable to noise reduction using wall acoustic treatment.

An Acoustical Comparison of Sub-Scale and Full-Scale Far-Field Measurements for the Reusable Solid Rocket Motor

NASA Technical Reports Server (NTRS)

Haynes, Jared; Kenny, R. Jeremy

2010-01-01

Recently, members of the Marshall Space Flight Center (MSFC) Fluid Dynamics Branch and Wyle Labs measured far-field acoustic data during a series of three Reusable Solid Rocket Motor (RSRM) horizontal static tests conducted in Promontory, Utah. The test motors included the Technical Evaluation Motor 13 (TEM-13), Flight Verification Motor 2 (FVM-2), and the Flight Simulation Motor 15 (FSM-15). Similar far-field data were collected during horizontal static tests of sub-scale solid rocket motors at MSFC. Far-field acoustical measurements were taken at multiple angles within a circular array centered about the nozzle exit plane, each positioned at a radial distance of 80 nozzle-exit-diameters from the nozzle. This type of measurement configuration is useful for calculating rocket noise characteristics such as those outlined in the NASA SP-8072 "Acoustic Loads Generated by the Propulsion System." Acoustical scaling comparisons are made between the test motors, with particular interest in the Overall Sound Power, Acoustic Efficiency, Non-dimensional Relative Sound Power Spectrum, and Directivity. Since most empirical data in the NASA SP-8072 methodology is derived from small rockets, this investigation provides an opportunity to check the data collapse between a sub-scale and full-scale rocket motor.

An ablative pulsed plasma thruster with a segmented anode

NASA Astrophysics Data System (ADS)

Zhang, Zhe; Ren, Junxue; Tang, Haibin; Ling, William Yeong Liang; York, Thomas M.

2018-01-01

An ablative pulsed plasma thruster (APPT) design with a ‘segmented anode’ is proposed in this paper. We aim to examine the effect that this asymmetric electrode configuration (a normal cathode and a segmented anode) has on the performance of an APPT. The magnetic field of the discharge arc, plasma density in the exit plume, impulse bit, and thrust efficiency were studied using a magnetic probe, Langmuir probe, thrust stand, and mass bit measurements, respectively. When compared with conventional symmetric parallel electrodes, the segmented anode APPT shows an improvement in the impulse bit of up to 28%. The thrust efficiency is also improved by 49% (from 5.3% to 7.9% for conventional and segmented designs, respectively). Long-exposure broadband emission images of the discharge morphology show that compared with a normal anode, a segmented anode results in clear differences in the luminous discharge morphology and better collimation of the plasma. The magnetic probe data indicate that the segmented anode APPT exhibits a higher current density in the discharge arc. Furthermore, Langmuir probe data collected from the central exit plane show that the peak electron density is 75% higher than with conventional parallel electrodes. These results are believed to be fundamental to the physical mechanisms behind the increased impulse bit of an APPT with a segmented electrode.

A generalized one-dimensional computer code for turbomachinery cooling passage flow calculations

NASA Technical Reports Server (NTRS)

Kumar, Ganesh N.; Roelke, Richard J.; Meitner, Peter L.

1989-01-01

A generalized one-dimensional computer code for analyzing the flow and heat transfer in the turbomachinery cooling passages was developed. This code is capable of handling rotating cooling passages with turbulators, 180 degree turns, pin fins, finned passages, by-pass flows, tip cap impingement flows, and flow branching. The code is an extension of a one-dimensional code developed by P. Meitner. In the subject code, correlations for both heat transfer coefficient and pressure loss computations were developed to model each of the above mentioned type of coolant passages. The code has the capability of independently computing the friction factor and heat transfer coefficient on each side of a rectangular passage. Either the mass flow at the inlet to the channel or the exit plane pressure can be specified. For a specified inlet total temperature, inlet total pressure, and exit static pressure, the code computers the flow rates through the main branch and the subbranches, flow through tip cap for impingement cooling, in addition to computing the coolant pressure, temperature, and heat transfer coefficient distribution in each coolant flow branch. Predictions from the subject code for both nonrotating and rotating passages agree well with experimental data. The code was used to analyze the cooling passage of a research cooled radial rotor.

Simultaneous in-plane and out-of-plane displacement measurement based on a dual-camera imaging system and its application to inspection of large-scale space structures

NASA Astrophysics Data System (ADS)

Ri, Shien; Tsuda, Hiroshi; Yoshida, Takeshi; Umebayashi, Takashi; Sato, Akiyoshi; Sato, Eiichi

2015-07-01

Optical methods providing full-field deformation data have potentially enormous interest for mechanical engineers. In this study, an in-plane and out-of-plane displacement measurement method based on a dual-camera imaging system is proposed. The in-plane and out-of-plane displacements are determined simultaneously using two measured in-plane displacement data observed from two digital cameras at different view angles. The fundamental measurement principle and experimental results of accuracy confirmation are presented. In addition, we applied this method to the displacement measurement in a static loading and bending test of a solid rocket motor case (CFRP material; 2.2 m diameter and 2.3 m long) for an up-to-date Epsilon rocket developed by JAXA. The effectiveness and measurement accuracy is confirmed by comparing with conventional displacement sensor. This method could be useful to diagnose the reliability of large-scale space structures in the rocket development.

Fan Noise Source Diagnostic Test: Rotor Alone Aerodynamic Performance Results

NASA Technical Reports Server (NTRS)

Hughes, Christopher E.; Jeracki, Robert J.; Woodward, Richard P.; Miller, Christopher J.

2005-01-01

The aerodynamic performance of an isolated fan or rotor alone model was measured in the NASA Glenn Research Center 9- by 15- Foot Low Speed Wind Tunnel as part of the Fan Broadband Source Diagnostic Test conducted at NASA Glenn. The Source Diagnostic Test was conducted to identify the noise sources within a wind tunnel scale model of a turbofan engine and quantify their contribution to the overall system noise level. The fan was part of a 1/5th scale model representation of the bypass stage of a current technology turbofan engine. For the rotor alone testing, the fan and nacelle, including the inlet, external cowl, and fixed area fan exit nozzle, were modeled in the test hardware; the internal outlet guide vanes located behind the fan were removed. Without the outlet guide vanes, the velocity at the nozzle exit changes significantly, thereby affecting the fan performance. As part of the investigation, variations in the fan nozzle area were tested in order to match as closely as possible the rotor alone performance with the fan performance obtained with the outlet guide vanes installed. The fan operating performance was determined using fixed pressure/temperature combination rakes and the corrected weight flow. The performance results indicate that a suitable nozzle exit was achieved to be able to closely match the rotor alone and fan/outlet guide vane configuration performance on the sea level operating line. A small shift in the slope of the sea level operating line was measured, which resulted in a slightly higher rotor alone fan pressure ratio at take-off conditions, matched fan performance at cutback conditions, and a slightly lower rotor alone fan pressure ratio at approach conditions. However, the small differences in fan performance at all fan conditions were considered too small to affect the fan acoustic performance.

Dual-Pump CARS Development and Application to Supersonic Combustion

NASA Astrophysics Data System (ADS)

Magnotti, Gaetano

Successful design of hypersonic air-breathing engines requires new computational fluid dynamics (CFD) models for turbulence and turbulence-chemistry interaction in supersonic combustion. Unfortunately, not enough data are available to the modelers to develop and validate their codes, due to difficulties in taking measurements in such a harsh environment. Dual-pump coherent anti-Stokes Raman spectroscopy (CARS) is a non-intrusive, non-linear, laser-based technique that provides temporally and spatially resolved measurements of temperature and absolute mole fractions of N2, O2 and H2 in H2-air flames. A dual-pump CARS instrument has been developed to obtain measurements in supersonic combustion and generate databases for the CFD community. Issues that compromised previous attempts, such as beam steering and high irradiance perturbation effects, have been alleviated or avoided. Improvements in instrument precision and accuracy have been achieved. An axis-symmetric supersonic combusting coaxial jet facility has been developed to provide a simple, yet suitable flow to CFD modelers. The facility provides a central jet of hot "vitiated air" simulating the hot air entering the engine of a hypersonic vehicle flying at Mach numbers between 5 and 7. Three different silicon carbide nozzles, with exit Mach number 1, 1.6 and 2, are used to provide flows with the effects of varying compressibility. H2 co-flow is available in order to generate a supersonic combusting free jet. Dual-pump CARS measurements have been obtained for varying values of flight and exit Mach numbers at several locations. Approximately one million Dual-pump CARS single shots have been collected in the supersonic jet for varying values of flight and exit Mach numbers at several locations. Data have been acquired with a H2 co-flow (combustion case) or a N 2 co-flow (mixing case). Results are presented and the effects of the compressibility and of the heat release are discussed.

An astronomy camera for low background applications in the 1. 0 to 2. 5. mu. m spectral region

DOE Office of Scientific and Technical Information (OSTI.GOV)

Kaki, S.A.; Bailey, G.C.; Hagood, R.W.

1989-02-01

A short wavelength (1.0-2.5 ..mu..m) 128 x 128 focal plane array forms the heart of this low background astronomy camera system. The camera is designed to accept either a 128 x 128 HgCdTe array for the 1-2.5 ..mu..m spectral region or an InSb array for the 3-5 ..mu..m spectral region. A cryogenic folded optical system is utilized to control excess stray light along with a cold eight-position filter wheel for spectral filtering. The camera head and electronics will also accept a 256 x 256 focal plane. Engineering evaluation of the complete system is complete along with two engineering runs atmore » the JPL Table Mountain Observatory. System design, engineering performance, and sample imagery are presented in this paper.« less

Modeling the Deterioration of Engine and Low Pressure Compressor Performance During a Roll Back Event Due to Ice Accretion

NASA Technical Reports Server (NTRS)

Veres, Joseph P.; Jorgenson, Philip, C. E.; Jones, Scott M.

2014-01-01

The main focus of this study is to apply a computational tool for the flow analysis of the engine that has been tested with ice crystal ingestion in the Propulsion Systems Laboratory (PSL) of NASA Glenn Research Center. A data point was selected for analysis during which the engine experienced a full roll back event due to the ice accretion on the blades and flow path of the low pressure compressor. The computational tool consists of the Numerical Propulsion System Simulation (NPSS) engine system thermodynamic cycle code, and an Euler-based compressor flow analysis code, that has an ice particle melt estimation code with the capability of determining the rate of sublimation, melting, and evaporation through the compressor blade rows. Decreasing the performance characteristics of the low pressure compressor (LPC) within the NPSS cycle analysis resulted in matching the overall engine performance parameters measured during testing at data points in short time intervals through the progression of the roll back event. Detailed analysis of the fan-core and LPC with the compressor flow analysis code simulated the effects of ice accretion by increasing the aerodynamic blockage and pressure losses through the low pressure compressor until achieving a match with the NPSS cycle analysis results, at each scan. With the additional blockages and losses in the LPC, the compressor flow analysis code results were able to numerically reproduce the performance that was determined by the NPSS cycle analysis, which was in agreement with the PSL engine test data. The compressor flow analysis indicated that the blockage due to ice accretion in the LPC exit guide vane stators caused the exit guide vane (EGV) to be nearly choked, significantly reducing the air flow rate into the core. This caused the LPC to eventually be in stall due to increasing levels of diffusion in the rotors and high incidence angles in the inlet guide vane (IGV) and EGV stators. The flow analysis indicating compressor stall is substantiated by the video images of the IGV taken during the PSL test, which showed water on the surface of the IGV flowing upstream out of the engine, indicating flow reversal, which is characteristic of a stalled compressor.

Modeling an anode layer Hall thruster and its plume

NASA Astrophysics Data System (ADS)

Choi, Yongjun

This thesis consists of two parts: a study of the D55 Hall thruster channel using a hydrodynamic model; and particle simulations of plasma plume flow from the D55 Hall thruster. The first part of this thesis investigates the xenon plasma properties within the D55 thruster channel using a hydrodynamic model. The discharge voltage (V) and current (I) characteristic of the D55 Hall thruster are studied. The hydrodynamic model fails to accurately predict the V-I characteristics. This analysis shows that the model needs to be improved. Also, the hydrodynamic model is used to simulate the plasma flow within the D55 Hall thruster. This analysis is performed to investigate the plasma properties of the channel exit. It is found that the hydrodynamic model is very sensitive to initial conditions, and fails to simulate the complete domain of the D55 Hall thruster. However, the model successfully calculates the channel domain of the D55 Hall thruster. The results show that, at the thruster exit, the plasma density has a maximum value while the ion velocity has a minimum at the channel center. Also, the results show that the flow angle varies almost linearly across the exit plane and increases from the center to the walls. Finally, the hydrodynamic model results are used to estimate the plasma properties at the thruster nozzle exit. The second part of the thesis presents two dimensional axisymmetric simulations of xenon plasma plume flow fields from the D55 anode layer Hall thruster. A hybrid particle-fluid method is used for the simulations. The magnetic field near the Hall thruster exit is included in the calculation. The plasma properties obtained from the hydrodynamic model are used to determine boundary conditions for the simulations. In these simulations, the Boltzmann model and a detailed fluid model are used to compute the electron properties, the direct simulation Monte Carlo method models the collisions of heavy particles, and the Particle-In-Cell method models the transport of ions in an electric field. The accuracy of the simulation is assessed through comparison with various sets of measured data. It is found that a magnetic field significantly affects the profile of the plasma in the Detailed model. For instance, the plasma potential decreases more rapidly with distance from the thruster in the presence of a magnetic field. Results predicted by the Detailed model with the magnetic field are in better agreement with experimental data than those obtained with other models investigated.

A novel porous Ffowcs-Williams and Hawkings acoustic methodology for complex geometries

NASA Astrophysics Data System (ADS)

Nitzkorski, Zane Lloyd

Predictive noise calculations from high Reynolds number flows in complex engineering geometry are becoming a possibility with the high performance computing resources that have become available in recent years. Increasing the applicability and reliability of solution methodologies have been two key challenges toward this goal. This dissertation develops a porous Ffowcs-Williams and Hawkings methodology that uses a novel endcap methodology, and can be applied to unstructured grids. The use of unstructured grids allows complex geometry to be represented while porous formulation eliminates difficulties with the choice of acoustic Green's function. Specifically, this dissertation (1) proposes and examines a novel endcap procedure to account for spurious noise, (2) uses the proposed methodology to investigate noise production from a range of subcritical Reynolds number circular cylinders, and (3) investigates a trailing edge geometry for noise production and to illustrate the generality of the Green's function. Porous acoustic analogies need an endcap scheme in order to prevent spurious noise due to truncation errors. A dynamic end cap methodology is proposed to account for spurious contributions to the far--field sound within the context of the Ffowcs--Williams and Hawkings (FW--H) acoustic analogy. The quadrupole source terms are correlated over multiple planes to obtain a convection velocity which is then used to determine a corrective convective flux at the FW--H porous surface. The proposed approach is first demonstrated for a convecting potential vortex. The correlation is investigated by examining it pass through multiple exit planes. It is then evaluated by computing the sound emitted by flow over a circular cylinder at Reynolds number of 150 and compared to other endcap methods, such as Shur et al. [1]. Insensitivity to end plane location and spacing and the effect of the dynamic convection velocity are computed. Subcritical Reynolds number circular cylinder flows are investigated at Re = 3900, 10000 and 89000 in order to evaluate the method and investigate the physical sources of noise production. The Re = 3900 case was chosen due to its highly validated flow-field and to serve as a basis of comparison. The Re = 10000 cylinder is used to validate the noise production at turbulent Reynolds numbers against other simulations. Finally the Re = 89000 simulations are used to compare to experiment serving as a rigorous test of the methods predictive ability. The proposed approach demonstrates better performance than other commonly used approaches with the added benefit of computational efficiency and the ability to query independent volumes. This gives the added benefit of discovering how much noise production is directly associated with volumetric noise contributions. These capabilities allow for a thorough investigation of the sources of noise production and a means to evaluate proposed theories. A physical description of the source of sound for subcritical Reynolds number cylinders is established. A 45° beveled trailing edge configuration is investigated due to its relevance to hydrofoil and propeller noise. This configuration also allows for the evaluation of the assumption associated with the free-space Green's function since the half-plane Green's function can be used to represent the solution to the wave equation for this geometry. Similar results for directivity and amplitudes of the two formulations confirm the flexibility of the porous surface implementation. Good agreement with experiment is obtained. The effect of boundary layer thickness is investigated. The noise produced in the upper half plane is significantly decreased for the thinner boundary layer while the noise production in the lower half plane is only slightly decreased.

Fundamental and applied research on core engine/combustion noise of aircraft engines

NASA Technical Reports Server (NTRS)

Plett, E. G.; Leshner, M. D.; Summerfield, M.

1974-01-01

Some results of a study of the importance of geometrical features of the combustor to combustion roughness and resulting noise are presented. Comparison is made among a perforated can flame holder, a plane slotted flame holder and a plane slotted flame holder which introduces two counter swirling streams. The latter is found to permit the most stable, quiet combustion. Crosscorrelations between the time derivative of chamber pressure fluctuations and far field noise are found to be stronger than between the far field noise and the direct chamber pressure signal. Temperature fluctuations in the combustor nozzle are also found to have a reasonably strong crosscorrelation with far field sound.

PRELIMINARY DESIGN ANALYSIS OF AXIAL FLOW TURBINES

NASA Technical Reports Server (NTRS)

Glassman, A. J.

1994-01-01

A computer program has been developed for the preliminary design analysis of axial-flow turbines. Rapid approximate generalized procedures requiring minimum input are used to provide turbine overall geometry and performance adequate for screening studies. The computations are based on mean-diameter flow properties and a stage-average velocity diagram. Gas properties are assumed constant throughout the turbine. For any given turbine, all stages, except the first, are specified to have the same shape velocity diagram. The first stage differs only in the value of inlet flow angle. The velocity diagram shape depends upon the stage work factor value and the specified type of velocity diagram. Velocity diagrams can be specified as symmetrical, zero exit swirl, or impulse; or by inputting stage swirl split. Exit turning vanes can be included in the design. The 1991 update includes a generalized velocity diagram, a more flexible meanline path, a reheat model, a radial component of velocity, and a computation of free-vortex hub and tip velocity diagrams. Also, a loss-coefficient calibration was performed to provide recommended values for airbreathing engine turbines. Input design requirements include power or pressure ratio, mass flow rate, inlet temperature and pressure, and rotative speed. The design variables include inlet and exit diameters, stator angle or exit radius ratio, and number of stages. Gas properties are input as gas constant, specific heat ratio, and viscosity. The program output includes inlet and exit annulus dimensions, exit temperature and pressure, total and static efficiencies, flow angles, blading angles, and last stage absolute and relative Mach numbers. This program is written in FORTRAN 77 and can be ported to any computer with a standard FORTRAN compiler which supports NAMELIST. It was originally developed on an IBM 7000 series computer running VM and has been implemented on IBM PC computers and compatibles running MS-DOS under Lahey FORTRAN, and DEC VAX series computers running VMS. Format statements in the code may need to be rewritten depending on your FORTRAN compiler. The source code and sample data are available on a 5.25 inch 360K MS-DOS format diskette. This program was developed in 1972 and was last updated in 1991. IBM and IBM PC are registered trademarks of International Business Machines. MS-DOS is a registered trademark of Microsoft Corporation. DEC VAX, and VMS are trademarks of Digital Equipment Corporation.

Noise suppression due to annulus shaping of conventional coaxial nozzle

NASA Technical Reports Server (NTRS)

Vonglahn, U.; Goodykoontz, J.

1980-01-01

A method which shows that increasing the annulus width of a conventional coaxial nozzle with constant bypass velocity will lower the noise level is described. The method entails modifying a concentric coaxial nozzle to provide an eccentric outer stream annulus while maintaining approximately the same through flow as that for the original concentric bypass nozzle. Acoustical tests to determine the noise generating characteristics of the nozzle over a range of flow conditions are described. The tests involved sequentially analyzing the noise signals and digitally recording the 1/3 octave band sound pressure levels. The measurements were made in a plane passing through the minimum and maximum annulus width points, as well as at 90 degrees in this plane, by rotating the outer nozzle about its axis. Representative measured spectral data in the flyover plane for the concentric nozzle obtained at model scale are discussed. Representative spectra for several engine cycles are presented for both the eccentric and concentric nozzles at engine size.

Restoring the magnetism of ultrathin LaMn O3 films by surface symmetry engineering

NASA Astrophysics Data System (ADS)

Peng, J. J.; Song, C.; Li, F.; Gu, Y. D.; Wang, G. Y.; Pan, F.

2016-12-01

The frustration of magnetization and conductivity properties of ultrathin manganite is detrimental to their device performance, preventing their scaling down process. Here we demonstrate that the magnetism of ultrathin LaMn O3 films can be restored by a SrTi O3 capping layer, which engineers the surface from a symmetry breaking induced out-of-plane orbital occupancy to the recovered in-plane orbital occupancy. The stabilized in-plane orbital occupancy would strengthen the intralayer double exchange and thus recovers the robust magnetism. This method is proved to be effective for films as thin as 2 unit cells, greatly shrinking the critical thickness of 6 unit cells for ferromagnetic LaMn O3 as demonstrated previously [Wang et al., Science 349, 716 (2015), 10.1126/science.aaa5198]. The achievement made in this work opens up new perspectives to an active control of surface states and thereby tailors the surface functional properties of transition metal oxides.

Wrinkling instability in nanoparticle-supported graphene: implications for strain engineering

NASA Astrophysics Data System (ADS)

Cullen, William; Yamamoto, Mahito; Pierre-Louis, Olivier; Huang, Jia; Fuhrer, Michael; Einstein, Theodore

2013-03-01

We have carried out a systematic study of the wrinkling instability of graphene membranes supported on SiO2 substrates with randomly placed silica nanoparticles. At small nanoparticle density, monolayer graphene adheres to the substrate and is highly conformal over the nanoparticles. With increasing nanoparticle density, and decreasing nanoparticle separation to ~100 nm, graphene's elastic response dominates substrate adhesion, and elastic stretching energy is reduced by the formation of wrinkles which connect protrusions. Above a critical nanoparticle density, the wrinkles form a percolating network through the sample. As the graphene membrane is made thicker, delamination from the substrate is observed. Since the wrinkling instability acts to remove inhomogeneous in-plane elastic strains through out-of-plane buckling, our results can be used to place limits on the possible in-plane strain magnitudes that may be created in graphene to realized strain-engineered electronic structures.[2] Supported by the UMD NSF-MRSEC under Grant No. DMR 05-20471, the US ONR MURI and UMD CNAM.

Design assurance test of the Thiokol Te-M-521-5 apogee kick motor tested in the spin mode at simulated altitude conditions

NASA Technical Reports Server (NTRS)

Cimino, A. A.

1973-01-01

One Thiokol Chemical Corporation TE-M-521-5 solid-propellant apogee rocket motor was successfully fired at an average simulated altitude of about 108,000 ft while spinning at 46 rpm. The general program objectives were to verify compliance of motor performance with the manufacturer's specifications. Specific primary objectives were to determine vacuum ballistic performance of the motor after prefire vibration conditioning and temperature conditioning at 40F, altitude ignition characteristics, motor structural integrity, and motor temperature-time history during and after motor operation. Additional objectives were to measure the lateral (nonaxial) thrust component during motor operation and to measure radiation heat flux in the vicinity of the nozzle exit plane.

Depth resolved investigations of boron implanted silicon

NASA Astrophysics Data System (ADS)

Sztucki, M.; Metzger, T. H.; Milita, S.; Berberich, F.; Schell, N.; Rouvière, J. L.; Patel, J.

2003-01-01

We have studied the depth distribution and structure of defects in boron implanted silicon (0 0 1). Silicon wafers were implanted with a boron dose of 6×10 15 ions/cm -2 at 32 keV and went through different annealing treatments. Using diffuse X-ray scattering at grazing incidence and exit angles we are able to distinguish between different kinds of defects (point defect clusters and extrinsic stacking faults on {1 1 1} planes) and to determine their depth distribution as a function of the thermal budget. Cross-section transmission electron microscopy was used to gain complementary information. In addition we have determined the strain distribution caused by the boron implantation as a function of depth from rocking curve measurements.

Computational Analyses in Support of Sub-scale Diffuser Testing for the A-3 Facility. Part 1; Steady Predictions

NASA Technical Reports Server (NTRS)

Allgood, Daniel C.; Graham, Jason S.; Ahuja, Vineet; Hosangadi, Ashvin

2010-01-01

Simulation technology can play an important role in rocket engine test facility design and development by assessing risks, providing analysis of dynamic pressure and thermal loads, identifying failure modes and predicting anomalous behavior of critical systems. Advanced numerical tools assume greater significance in supporting testing and design of high altitude testing facilities and plume induced testing environments of high thrust engines because of the greater inter-dependence and synergy in the functioning of the different sub-systems. This is especially true for facilities such as the proposed A-3 facility at NASA SSC because of a challenging operating envelope linked to variable throttle conditions at relatively low chamber pressures. Facility designs in this case will require a complex network of diffuser ducts, steam ejector trains, fast operating valves, cooling water systems and flow diverters that need to be characterized for steady state performance. In this paper, we will demonstrate with the use of CFD analyses s advanced capability to evaluate supersonic diffuser and steam ejector performance in a sub-scale A-3 facility at NASA Stennis Space Center (SSC) where extensive testing was performed. Furthermore, the focus in this paper relates to modeling of critical sub-systems and components used in facilities such as the A-3 facility. The work here will address deficiencies in empirical models and current CFD analyses that are used for design of supersonic diffusers/turning vanes/ejectors as well as analyses for confined plumes and venting processes. The primary areas that will be addressed are: (1) supersonic diffuser performance including analyses of thermal loads (2) accurate shock capturing in the diffuser duct; (3) effect of turning duct on the performance of the facility (4) prediction of mass flow rates and performance classification for steam ejectors (5) comparisons with test data from sub-scale diffuser testing and assessment of confidence levels in CFD based flowpath modeling of the facility. The analyses tools used here expand on the multi-element unstructured CFD which has been tailored and validated for impingement dynamics of dry plumes, complex valve/feed systems, and high pressure propellant delivery systems used in engine and component test stands at NASA SSC. The analyses performed in the evaluation of the sub-scale diffuser facility explored several important factors that influence modeling and understanding of facility operation such as (a) importance of modeling the facility with Real Gas approximation, (b) approximating the cluster of steam ejector nozzles as a single annular nozzle, (c) existence of mixed subsonic/supersonic flow downstream of the turning duct, and (d) inadequacy of two-equation turbulence models in predicting the correct pressurization in the turning duct and expansion of the second stage steam ejectors. The procedure used for modeling the facility was as follows: (i) The engine, test cell and first stage ejectors were simulated with an axisymmetric approximation (ii) the turning duct, second stage ejectors and the piping downstream of the second stage ejectors were analyzed with a three-dimensional simulation utilizing a half-plane symmetry approximation. The solution i.e. primitive variables such as pressure, velocity components, temperature and turbulence quantities were passed from the first computational domain and specified as a supersonic boundary condition for the second simulation. (iii) The third domain comprised of the exit diffuser and the region in the vicinity of the facility (primary included to get the correct shock structure at the exit of the facility and entrainment characteristics). The first set of simulations comprising the engine, test cell and first stage ejectors was carried out both as a turbulent real gas calculation as well as a turbulent perfect gas calculation. A comparison for the two cases (Real Turbulent and Perfect gas turbulent) of the Ma Number distribution and temperature distributions are shown in Figures 1 and 2 respectively.

Measurement and Computation of Supersonic Flow in a Lobed Diffuser-Mixer for Trapped Vortex Combustors

NASA Technical Reports Server (NTRS)

Brankovic, Andreja; Ryder, Robert C., Jr.; Hendricks, Robert C.; Liu, Nan-Suey; Gallagher, John R.; Shouse, Dale T.; Roquemore, W. Melvyn; Cooper, Clayton S.; Burrus, David L.; Hendricks, John A.

2002-01-01

The trapped vortex combustor (TVC) pioneered by Air Force Research Laboratories (AFRL) is under consideration as an alternative to conventional gas turbine combustors. The TVC has demonstrated excellent operational characteristics such as high combustion efficiency, low NO(x) emissions, effective flame stabilization, excellent high-altitude relight capability, and operation in the lean-burn or rich burn-quick quench-lean burn (RQL) modes of combustion. It also has excellent potential for lowering the engine combustor weight. This performance at low to moderate combustor mach numbers has stimulated interest in its ability to operate at higher combustion mach number, and for aerospace, this implies potentially higher flight mach numbers. To this end, a lobed diffuser-mixer that enhances the fuel-air mixing in the TVC combustor core was designed and evaluated, with special attention paid to the potential shock system entering the combustor core. For the present investigation, the lobed diffuser-mixer combustor rig is in a full annular configuration featuring sixfold symmetry among the lobes, symmetry within each lobe, and plain parallel, symmetric incident flow. During hardware cold-flow testing, significant discrepancies were found between computed and measured values for the pitot-probe-averaged static pressure profiles at the lobe exit plane. Computational fluid dynamics (CFD) simulations were initiated to determine whether the static pressure probe was causing high local flow-field disturbances in the supersonic flow exiting the diffuser-mixer and whether shock wave impingement on the pitot probe tip, pressure ports, or surface was the cause of the discrepancies. Simulations were performed with and without the pitot probe present in the modeling. A comparison of static pressure profiles without the probe showed that static pressure was off by nearly a factor of 2 over much of the radial profile, even when taking into account potential axial displacement of the probe by up to 0.25 in. (0.64 cm). Including the pitot probe in the CFD modeling and data interpretation lead to good agreement between measurement and prediction. Graphical inspection of the results showed that the shock waves impinging on the probe surface were highly nonuniform, with static pressure varying circumferentially among the pressure ports by over 10 percent in some cases. As part of the measurement methodology, such measurements should be routinely supplemented with CFD analyses that include the pitot probe as part of the flow-path geometry.

Aeroacoustic Prediction Codes

NASA Technical Reports Server (NTRS)

Gliebe, P; Mani, R.; Shin, H.; Mitchell, B.; Ashford, G.; Salamah, S.; Connell, S.; Huff, Dennis (Technical Monitor)

2000-01-01

This report describes work performed on Contract NAS3-27720AoI 13 as part of the NASA Advanced Subsonic Transport (AST) Noise Reduction Technology effort. Computer codes were developed to provide quantitative prediction, design, and analysis capability for several aircraft engine noise sources. The objective was to provide improved, physics-based tools for exploration of noise-reduction concepts and understanding of experimental results. Methods and codes focused on fan broadband and 'buzz saw' noise and on low-emissions combustor noise and compliment work done by other contractors under the NASA AST program to develop methods and codes for fan harmonic tone noise and jet noise. The methods and codes developed and reported herein employ a wide range of approaches, from the strictly empirical to the completely computational, with some being semiempirical analytical, and/or analytical/computational. Emphasis was on capturing the essential physics while still considering method or code utility as a practical design and analysis tool for everyday engineering use. Codes and prediction models were developed for: (1) an improved empirical correlation model for fan rotor exit flow mean and turbulence properties, for use in predicting broadband noise generated by rotor exit flow turbulence interaction with downstream stator vanes: (2) fan broadband noise models for rotor and stator/turbulence interaction sources including 3D effects, noncompact-source effects. directivity modeling, and extensions to the rotor supersonic tip-speed regime; (3) fan multiple-pure-tone in-duct sound pressure prediction methodology based on computational fluid dynamics (CFD) analysis; and (4) low-emissions combustor prediction methodology and computer code based on CFD and actuator disk theory. In addition. the relative importance of dipole and quadrupole source mechanisms was studied using direct CFD source computation for a simple cascadeigust interaction problem, and an empirical combustor-noise correlation model was developed from engine acoustic test results. This work provided several insights on potential approaches to reducing aircraft engine noise. Code development is described in this report, and those insights are discussed.

PIV Measurements of Chevrons on F400-Series Tactical Aircraft Nozzle Model

NASA Technical Reports Server (NTRS)

Bridges, James; Wernet, Mark P.; Frate, Franco C.

2011-01-01

Reducing noise of tactical jet aircraft has taken on fresh urgency as core engine technologies allow higher specific-thrust engines and as society become more concerned for the health of its military workforce. Noise reduction on this application has lagged the commercial field as incentives for quieting military aircraft have not been as strong as in their civilian counterparts. And noise reduction strategies employed on civilian engines may not be directly applicable due to the differences in exhaust system architecture and mission. For instance, the noise reduction technology of chevrons, examined in this study, will need to be modified to take into account the special features of tactical aircraft nozzles. In practice, these nozzles have divergent slats that are tied to throttle position, and at take off the jet flow is highly overexpanded as the nozzle is optimized for cruise altitude rather than sea level. In simple oil flow visualization experiments conducted at the onset of the current test program flow barely stays attached at end of nozzle at takeoff conditions. This adds a new twist to the design of chevrons. Upon reaching the nozzle exit the flow shrinks inward radially, meaning that for a chevron to penetrate the flow it must extend much farther away from the baseline nozzle streamline. Another wrinkle is that with a variable divergence angle on the nozzle, the effective penetration will differ with throttle position and altitude. The final note of realism introduced in these experiments was to simulate the manner in which bypass flow is bled into the nozzle wall in real engines to cool the nozzle, which might cause very fat boundary layer at exit. These factors, along with several other issues specific to the application of chevrons to convergent-divergent nozzles have been explored with particle image velocimetry measurements and are presented in this paper.

Tests and analysis of a vented D thrust deflecting nozzle on a turbofan engine. [conducted at the outdoor aerodynamic research facility of the Ames Research Center

NASA Technical Reports Server (NTRS)

Roseberg, E. W.

1982-01-01

The objectives were to: obtain nozzle performance characteristics in and out of ground effects; demonstrate the compatibility of the nozzle with a turbofan engine; obtain pressure and temperature distributions on the surface of the D vented nozzle; and establish a correlation of the nozzle performance between small scale and large scale models. The test nozzle was a boilerplate model of the MCAIR D vented nozzle configured for operation with a General Electric YTF-34-F5 turbofan engine. The nozzle was configured to provide: a thrust vectoring range of 0 to 115 deg; a yaw vectoring range of 0 to 10 deg; variable nozzle area control; and variable spacing between the core exit and nozzle entrance station. Compatibility between the YTF-34-T5 turbofan engine and the D vented nozzle was demonstrated. Velocity coefficients of 0.96 and greater were obtained for 90 deg of thrust vectoring. The nozzle walls remained cool during all test conditions.

Laser anemometer measurements and computations for transonic flow conditions in an annular cascade of high turning core turbine vanes

NASA Technical Reports Server (NTRS)

Goldman, Louis J.

1993-01-01

An advanced laser anemometer (LA) was used to measure the axial and tangential velocity components in an annular cascade of turbine stator vanes operating at transonic flow conditions. The vanes tested were based on a previous redesign of the first-stage stator in a two-stage turbine for a high-bypass-ratio engine. The vanes produced 75 deg of flow turning. Tests were conducted on a 0.771-scale model of the engine-sized stator. The advanced LA fringe system employed an extremely small 50-micron diameter probe volume. Window correction optics were used to ensure that the laser beams did not uncross in passing through the curved optical access port. Experimental LA measurements of velocity and turbulence were obtained at the mean radius upstream of, within, and downstream of the stator vane row at an exit critical velocity ratio of 1.050 at the hub. Static pressures were also measured on the vane surface. The measurements are compared, where possible, with calculations from a three-dimensional inviscid flow analysis. Comparisons were also made with the results obtained previously when these same vanes were tested at the design exit critical velocity ratio of 0.896 at the hub. The data are presented in both graphical and tabulated form so that they can be readily compared against other turbomachinery computations.

An Engine Research Program Focused on Low Pressure Turbine Aerodynamic Performance

NASA Technical Reports Server (NTRS)

Castner, Raymond; Wyzykowski, John; Chiapetta, Santo; Adamczyk, John

2002-01-01

A comprehensive test program was performed in the Propulsion Systems Laboratory at the NASA Glenn Research Center, Cleveland Ohio using a highly instrumented Pratt and Whitney Canada PW 545 turbofan engine. A key objective of this program was the development of a high-altitude database on small, high-bypass ratio engine performance and operability. In particular, the program documents the impact of altitude (Reynolds Number) on the aero-performance of the low-pressure turbine (fan turbine). A second objective was to assess the ability of a state-of-the-art CFD code to predict the effect of Reynolds number on the efficiency of the low-pressure turbine. CFD simulation performed prior and after the engine tests will be presented and discussed. Key findings are the ability of a state-of-the art CFD code to accurately predict the impact of Reynolds Number on the efficiency and flow capacity of the low-pressure turbine. In addition the CFD simulations showed the turbulent intensity exiting the low-pressure turbine to be high (9%). The level is consistent with measurements taken within an engine.

Initial Flow Matching Results of MHD Energy Bypass on a Supersonic Turbojet Engine Using the Numerical Propulsion System Simulation (NPSS) Environment

NASA Technical Reports Server (NTRS)

Benyo, Theresa L.

2010-01-01

Preliminary flow matching has been demonstrated for a MHD energy bypass system on a supersonic turbojet engine. The Numerical Propulsion System Simulation (NPSS) environment was used to perform a thermodynamic cycle analysis to properly match the flows from an inlet to a MHD generator and from the exit of a supersonic turbojet to a MHD accelerator. Working with various operating conditions such as the enthalpy extraction ratio and isentropic efficiency of the MHD generator and MHD accelerator, interfacing studies were conducted between the pre-ionizers, the MHD generator, the turbojet engine, and the MHD accelerator. This paper briefly describes the NPSS environment used in this analysis and describes the NPSS analysis of a supersonic turbojet engine with a MHD generator/accelerator energy bypass system. Results from this study have shown that using MHD energy bypass in the flow path of a supersonic turbojet engine increases the useful Mach number operating range from 0 to 3.0 Mach (not using MHD) to an explored and desired range of 0 to 7.0 Mach.

Shape memory alloy actuation for a variable area fan nozzle

NASA Astrophysics Data System (ADS)

Rey, Nancy; Tillman, Gregory; Miller, Robin M.; Wynosky, Thomas; Larkin, Michael J.; Flamm, Jeffrey D.; Bangert, Linda S.

2001-06-01

The ability to control fan nozzle exit area is an enabling technology for next generation high-bypass-ratio turbofan engines. Performance benefits for such designs are estimated at up to 9% in thrust specific fuel consumption (TSFC) relative to current fixed-geometry engines. Conventionally actuated variable area fan nozzle (VAN) concepts tend to be heavy and complicated, with significant aircraft integration, reliability and packaging issues. The goal of this effort was to eliminate these undesirable features and formulate a design that meets or exceeds leakage, durability, reliability, maintenance and manufacturing cost goals. A Shape Memory Alloy (SMA) bundled cable actuator acting to move an array of flaps around the fan nozzle annulus is a concept that meets these requirements. The SMA bundled cable actuator developed by the United Technologies Corporation (Patents Pending) provides significant work output (greater than 2200 in-lb per flap, through the range of motion) in a compact package and minimizes system complexity. Results of a detailed design study indicate substantial engine performance, weight, and range benefits. The SMA- based actuation system is roughly two times lighter than a conventional mechanical system, with significant aircraft direct operating cost savings (2-3%) and range improvements (5-6%) relative to a fixed-geometry nozzle geared turbofan. A full-scale sector model of this VAN system was built and then tested at the Jet Exit Test Facility at NASA Langley to demonstrate the system's ability to achieve 20% area variation of the nozzle under full scale aerodynamic loads. The actuator exceeded requirements, achieving repeated actuation against full-scale loads representative of typical cruise as well as greater than worst-case (ultimate) aerodynamic conditions. Based on these encouraging results, work is continuing with the goal of a flight test on a C-17 transport aircraft.

Preliminary design of an alternate high-temperature turbine. A topical report for Phase II of the High-Temperature-Turbine Technology Program

DOE Office of Scientific and Technical Information (OSTI.GOV)

Strough, R.I.

The feasibility of designing a convectively air-cooled turbine to operate in the environment of a 3000/sup 0/F combustor exit temperature with maximum turbine airfoil metal temperatures held to 1500/sup 0/F was established. The United Technologies-Kraftwerk Union V84.3 gas turbine design was used as the basic configuration for the design of the 3000/sup 0/F turbine. Turbine cooling requirements were determined based on the use of the modified V84.3 type silo combustor with a pattern factor of 0.1. The convective air-cooling technology levels in terms of cooling effectiveness required to satisfy the airfoil cooling requirements were identified. Cooling schemes and fabrication technologiesmore » required are discussed. Turbine airfoil cooling technology levels required for the 3000/sup 0/F engine were selected. The performance of the 3000/sup 0/F convectively air-cooled gas turbine in simple and combined cycle was calculated. The 3000/sup 0/F gas turbine combined-cycle system provides an increase in power of 61% and a decrease in heat rate of 10% compared to a similar system with a combustor exit temperature of 2210/sup 0/F and the same airflow. The development of a successful 3000/sup 0/F convectively air-cooled turbine can be accomplished with a reasonable design and fabrication development effort on the cooled turbine airfoils. Use of the convectively air-cooled turbine provides the transfer of technology from extensive aircraft engines developed programs and operating experience to industrial gas turbines. It eliminates the requirement for large investments in alternate cooling techniques tailored specifically for industrial engines which offer no additional benefits.« less

Licencing and Training Reform in the Australian Aircraft Maintenance Industry

ERIC Educational Resources Information Center

Hampson, Ian; Fraser, Doug

2016-01-01

The training and licencing of aircraft maintenance engineers fulfils a crucial protective function since it is they who perform and supervise aircraft maintenance and certify that planes are safe afterwards. In Australia, prior to training reform, a trades-based system of aircraft maintenance engineer training existed in an orderly relation with…

In-Plane Heterostructures Enable Internal Stress Assisted Strain Engineering in 2D Materials.

PubMed

Liu, Feng; Wang, Tzu-Chiang; Tang, Qiheng

2018-04-01

Conventional methods to induce strain in 2D materials can hardly catch up with the sharp increase in requirements to design specific strain forms, such as the pseudomagnetic field proposed in graphene, funnel effect of excitons in MoS 2 , and also the inverse funnel effect reported in black phosphorus. Therefore, a long-standing challenge in 2D materials strain engineering is to find a feasible scheme that can be used to design given strain forms. In this article, combining the ability of experimentally synthetizing in-plane heterostructures and elegant Eshelby inclusion theory, the possibility of designing strain fields in 2D materials to manipulate physical properties, which is called internal stress assisted strain engineering, is theoretically demonstrated. Particularly, through changing the inclusion's size, the stress or strain gradient can be controlled precisely, which is never achieved. By taking advantage of it, the pseudomagnetic field as well as the funnel effect can be accurately designed, which opens an avenue to practical applications for strain engineering in 2D materials. © 2018 WILEY-VCH Verlag GmbH & Co. KGaA, Weinheim.

Enhanced optical alignment of a digital micro mirror device through Bayesian adaptive exploration

NASA Astrophysics Data System (ADS)

Wynne, Kevin B.; Knuth, Kevin H.; Petruccelli, Jonathan

2017-12-01

As the use of Digital Micro Mirror Devices (DMDs) becomes more prevalent in optics research, the ability to precisely locate the Fourier "footprint" of an image beam at the Fourier plane becomes a pressing need. In this approach, Bayesian adaptive exploration techniques were employed to characterize the size and position of the beam on a DMD located at the Fourier plane. It couples a Bayesian inference engine with an inquiry engine to implement the search. The inquiry engine explores the DMD by engaging mirrors and recording light intensity values based on the maximization of the expected information gain. Using the data collected from this exploration, the Bayesian inference engine updates the posterior probability describing the beam's characteristics. The process is iterated until the beam is located to within the desired precision. This methodology not only locates the center and radius of the beam with remarkable precision but accomplishes the task in far less time than a brute force search. The employed approach has applications to system alignment for both Fourier processing and coded aperture design.

Adsorption induced modification of in-plane magnetic anisotropy in epitaxial Co and Fe/Co films on Fe(110)

NASA Astrophysics Data System (ADS)

Ślezak, M.; Ślezak, T.; Matlak, K.; DróŻdŻ, P.; Korecki, J.

2018-05-01

A study of in-plane magnetic anisotropy (MA) in epitaxial bcc Co films and Fe/Co bilayers on a Fe(110) surface is reported. Surface MA of as-deposited Co films and Fe/Co bilayers strongly depends on the Co (dCo) and Fe (dFe) thickness. Adsorption of residual gases drastically modifies in-plane MA of both Co films and Fe/Co bilayers. We present two dimensional MA maps in the (dCo, dFe) space for both as grown and adsorption-modified films. Our results indicate how to precisely engineer in-plane MA that can be controlled by dCo, dFe and is sensitive to the residual gas adsorption.

Energy Efficient Engine Exhaust Mixer Model Technology

NASA Technical Reports Server (NTRS)

Kozlowski, H.; Larkin, M.

1981-01-01

An exhaust mixer test program was conducted to define the technology required for the Energy Efficient Engine Program. The model configurations of 1/10 scale were tested in two phases. A parametric study of mixer design options, the impact of residual low pressure turbine swirl, and integration of the mixer with the structural pylon of the nacelle were investigated. The improvement of the mixer itself was also studied. Nozzle performance characteristics were obtained along with exit profiles and oil smear photographs. The sensitivity of nozzle performance to tailpipe length, lobe number, mixer penetration, and mixer modifications like scalloping and cutbacks were established. Residual turbine swirl was found detrimental to exhaust system performance and the low pressure turbine system for Energy Efficient Engine was designed so that no swirl would enter the mixer. The impact of mixer/plug gap was also established, along with importance of scalloping, cutbacks, hoods, and plug angles on high penetration mixers.

Gas-Generator Augmented Expander Cycle Rocket Engine

NASA Technical Reports Server (NTRS)

Greene, William D. (Inventor)

2011-01-01

An augmented expander cycle rocket engine includes first and second turbopumps for respectively pumping fuel and oxidizer. A gas-generator receives a first portion of fuel output from the first turbopump and a first portion of oxidizer output from the second turbopump to ignite and discharge heated gas. A heat exchanger close-coupled to the gas-generator receives in a first conduit the discharged heated gas, and transfers heat to an adjacent second conduit carrying fuel exiting the cooling passages of a primary combustion chamber. Heat is transferred to the fuel passing through the cooling passages. The heated fuel enters the second conduit of the heat exchanger to absorb more heat from the first conduit, and then flows to drive a turbine of one or both of the turbopumps. The arrangement prevents the turbopumps exposure to combusted gas that could freeze in the turbomachinery and cause catastrophic failure upon attempted engine restart.