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At least 631 records · Page 35
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

On Characterization of Laminar to Turbulent Transitional Flow in the Truncated IHF Semi-Elliptical Nozzle

This paper reports computational analysis and experimental data to characterize the laminar to turbulent flow transition in a high-enthalpy panel test facility at NASA Ames Research Center. The recently developed facility configuration uses a truncated version of an existing semi-elliptical nozzle in the 60-MW Interaction Heating Facility, and it provides testing capability for panel test articles as large as 43.2 cm x 43.2 cm (17 in x 17 in). A water-cooled calibration plate, which is attached to the semi-elliptical nozzle bottom surface and instrumented with twenty Gardon-gage heat flux calorimeters and twelve pressure gages, is used to characterize the convective heating environment over the panel test articles. The present analysis comprises computational fluid dynamics simulations of the nonequilibrium flowfield in the facility nozzle and test box, and the flowfields over panel test articles, and comparisons with the measured test data. Both laminar and turbulent simulations are performed. Constant current studies are performed: the measured data are grouped at the same arc current levels with varying mass flow rates, and they are compared with corresponding computations. For all currents studied, these studies indicate that the semi-elliptical nozzle flow is transitional at relatively higher mass flow rates within the facility operating envelope.

Arc-jets↗
NASA NTRSDate not supplied

On Characterization of Laminar to Turbulent Transitional Flow in the Truncated IHF Semi-Elliptical Nozzle

This paper reports computational analysis and experimental data to characterize the laminar to turbulent flow transition in a high-enthalpy panel test facility at NASA Ames Research Center. The recently developed facility configuration uses a truncated version of an existing semi-elliptical nozzle in the 60-MW Interaction Heating Facility, and it provides testing capability for panel test articles as large as 43.2 cm x 43.2 cm (17 in x 17 in). A water-cooled calibration plate, which is attached to the semi-elliptical nozzle bottom surface and instrumented with twenty Gardon-gage heat flux calorimeters and twelve pressure gages, is used to characterize the convective heating environment over the panel test articles. The present analysis comprises computational fluid dynamics simulations of the nonequilibrium flowfield in the facility nozzle and test box, and the flowfields over panel test articles, and comparisons with the measured test data. Both laminar and turbulent simulations are performed. Constant current studies are conducted: the measured data are grouped at the same arc current levels with varying mass flow rates, and they are compared with corresponding computations. For all currents considered, these studies indicate that the semi-elliptical nozzle flow is transitional at relatively higher mass flow rates within the facility operating envelope. An approximate facility map for the transitional flow characterization is developed, showing laminar and transitional to turbulent flow regions.

Arc-jets↗
NASA NTRS1994

The use of multispectral thermal infrared image data to estimate the sulfur dioxide flux from volcanoes: A case study from Mount Etna, Sicily, July 29, 1986

We have found that image data acquired with NASA's airborne Thermal Infrared Multispectral Scanner (TIMS) can be used to make estimates of the SO2 content of volcanic plumes. TIMS image data are most applicable to the study of partially transparent SO2 plumes, such as those released during quiescent periods or nonexplosive eruptions. The estimation procedure is based on the LOWTRAN 7 radiative transfer code, which we use to model the radiance perceived by TIMS as it views the ground through an SO2 plume. The input to the procedure includes the altitudes of the aircraft and ground, the altitude and thickness of the SO2 plume, the emissivity of the ground, and altitude profiles of the atmospheric pressure, temperature, and relative humidity. We use the TIMS data to estimate both ground temperatures beneath a plume and SO2 concentrations within a plume. Applying our procedure to TIMS data acquired over Mount Etna, Sicily, on July 29, 1986, we estimate that the SO2 flux from the volcano was approximately 6700 t d(exp -1). The use of TIMS to study SO2 plumes represents a bridge between highly localized methods, such as correlation spectroscopy or direct sampling, and small-scale mapping techniques involving satellite instruments such as the Total Ozone Mapping Spectrometer or Microwave Limb Sounder. We require further airborne experiments to refine our estimation procedure. This refinement is a necessary preparation for the schedueled 1998 launch of the Advanced Spaceborne Thermal Emission and Reflectance Radiometer, which will allow large-scale multispectral thermal infrared image data to be collected over virtually any volcano on Earth at least once every 16 days.

Realmuto, Vincent J.↗
NASA NTRS1996

Pressure Oscillations in Shuttle Solid Rocket Motors

All Shuttle Solid Rocket Motors (SRM's) exhibit low amplitude longitudinal pressure oscillations during motor burn. Although the oscillations have no known deleterious effect on motor ballistics, the acoustic pressure variations cause thrust oscillations that might affect Shuttle systems or components. The acoustic mode of greatest interest is the first or fundamental mode which, in the SRM, has a nominal frequency of 14-Hz. Oscillations in the SRM are believed to be caused by coupling between large scale vortices and the acoustic modes of the motor chamber. The vortices are thought to be created in the region of the motor segment interfaces and are inherent in the design of the motor. In such a situation the usual approach is to measure the oscillations and assess their impact on any sensitive components through tests and analysis. Questionable components can be altered to survive the vibration environment. As motor firings occur, oscillations are monitored to determine whether there are changes in the nature of the oscillations. Since the first static test, SRM's have been equipped with instrumentation especially designed to acquire chamber pressure oscillation data. Data from the first SRM static tests were used to establish predicted upper bounds for the maximum amplitudes in the latter half of burn. Those bounds have been used as a basis for worst-case simulation scenarios by specialists in structural dynamics at NASA and Rockwell International and to provide a basis for evaluating data from individual motors which were tested subsequent to the original SRM's. This paper updates the upper bounds prediction the High Performance Motors (HPM) by including data from all static tests performed to date including both original SRM's and post Challenger SRM's or Reusable Solid Rocket Motors (RSRM) in which the joint design was changed. All together, this study examines 27 SRM motors, 16 HPM motors and 11 RSRM motors. Predicted upper bounds will be made for both the first and second longitudinal modes. The first mode upper bounds will be compared to the original seven standard rocket motors (STD). The results indicate that, although the upper bounds have increased, they are still within acceptable bounds.

Blomshield, Fred S.↗
NASA NTRS1998

Model Engine Performance Measurement From Force Balance Instrumentation

A large scale model representative of a low-noise, high bypass ratio turbofan engine was tested for acoustics and performance in the NASA Lewis 9- by 15-Foot Low-Speed Wind Tunnel. This test was part of NASA's continuing Advanced Subsonic Technology Noise Reduction Program. The low tip speed fan, nacelle, and an un-powered core passage (with core inlet guide vanes) were simulated. The fan blades and hub are mounted on a rotating thrust and torque balance. The nacelle, bypass duct stators, and core passage are attached to a six component force balance. The two balance forces, when corrected for internal pressure tares, measure the total thrust-minus-drag of the engine simulator. Corrected for scaling and other effects, it is basically the same force that the engine supports would feel, operating at similar conditions. A control volume is shown and discussed, identifying the various force components of the engine simulator thrust and definitions of net thrust. Several wind tunnel runs with nearly the same hardware installed are compared, to identify the repeatability of the measured thrust-minus-drag. Other wind tunnel runs, with hardware changes that affected fan performance, are compared to the baseline configuration, and the thrust and torque effects are shown. Finally, a thrust comparison between the force balance and nozzle gross thrust methods is shown, and both yield very similar results.

Jeracki, Robert J.↗
NASA NTRS1962

NASA Scout ST-1 Flight-Test Results and Analyses, Launch Operations, and Test Vehicle Description

The first of a series of flight tests for the development of the four-stage, solid-propellant Scout vehicle was conducted at the NASA Wallops Station under the direction of the Langley Research Center. Vehicle designation for the test was NASA Scout ST-1. Performance characteristics of the vehicle and components were recorded during a high-altitude probe mission. Flight-simulation studies are presented and show that the accuracy of the guidance system during the flight was within control-system design specifications. The control system functioned normally during the flight with the exception of an overpowering of the reaction-control roll jets near burnout of the third-stage rocket motor. The resulting roll displacement of the vehicle is shown to have caused the monitor tracking radar which had been erroneously tracking a radar beacon in the vehicle on a side lobe to reorient to the major lobe of the receiving antenna. This tracking switch falsely indicated a violent turning maneuver on the monitor plot board and resulted in a hold-fire decision for the fourth-stage rocket motor. Although data for the final thrusting and coast phase of the flight were not obtained, the majority of the test objectives were achieved. In-flight thrust misalignment angles for the second- and third-stage rocket motors derived from control-system error data and for the first-stage motor determined from flight-simulation studies are presented. All rocket-motor thrust misalignment angles were well within the tolerances used for control-system design. Rocket-motor flight performance is presented, and velocity increments attained from the first three stages substantiated the predicted nominal performance. Operation of the rocket motors was satisfactory with the exception of high-level vibrations which were encountered during third-stage motor burning. Rolling moments which overpowered the reaction-control jets are also attributed to the burning characteristics of the third-stage motor. A discussion of the premature loss of the third-stage heat shield is given and shows that the heat-shield latching mechanism failed from pressure loads as the vehicle entered the transonic speed range. Although venting was provided to relieve the high negative pressures known to exist on the heat shield at these speeds, a field modification of the wiring tunnel had the same effect as opening the inside of the heat shield to ambient pressures. Consequently, the latching mechanism failed from pressure loads which were of about the same magnitude as the latching-mechanism yield loads. Skin temperatures were recorded at several locations on the vehicle and were generally in good agreement with theoretical values. Aerodynamic heating presented no problem during the flight since the maximum temperatures recorded during the flight were only about half the design values because of the high-launch-angle trajectory. Environmental vibrations recorded in the vicinity of the guidance package showed that no significant continuous amplitude levels above the general instrumentation noise level were present during first- and second-stage burning. Large vibration amplitudes were recorded during third-stage burning which coincided with the large roll disturbance experienced by the vehicle near burnout of the third-stage motor.

Scout Launch Vehicle↗
NASA NTRS1990

A model for the Space Shuttle Main Engine High Pressure Oxidizer Turbopump shaft seal system

A model of the High Pressure Oxidizer Turbopump (HPOTP) shaft seal system on the Space Shuttle Main Engine (SSME) is described. The model predicts the fluid properties and flow rates throughout this system for a number of conditions simulating failed seals. The results agree well with qualitative expectations and redline values but cannot be verified with actual data due to the lack thereof. The results indicate that each failure mode results in a unique distribution of properties throughout the seal system and can therefore be individually identified given the proper instrumentation. Furthermore, the detection process can be built on the principle of qualitative reasoning without the use of exact fluid property values. A simplified implementation of the model which does not include the slinger/labyrinth seal combination has been developed and will be useful for inclusion in a real-time diagnostic system.

Paxson, Daniel E.↗
NASA NTRS2001

Central Elemental Abundance Ratios In the Perseus Cluster: Resonant Scattering or SN Ia Enrichment?

We have determined abundance ratios in the core of the Perseus Cluster for several elements. These ratios indicate a central dominance of Type 1a supernova (SN Ia) ejects similar to that found for A496, A2199 and A3571. Simultaneous analysis of ASCA spectra from SIS1, GIS2, and GIS3 shows that the ratio of Ni to Fe abundances is approx. 3.4 +/- 1.1 times solar within the central 4'. This ratio is consistent with (and more precise than) that observed in other clusters whose central regions are dominated by SN Ia ejecta. Such a large Ni overabundance is predicted by "convective deflagration" explosion models for SNe Ia such as W7 but is inconsistent with delayed detonation models. We note that with current instrumentation the Ni K(alpha) line is confused with Fe K(beta) and that the Ni overabundance we observe has been interpreted by others as an anomalously large ratio of Fe K(beta) to Fe K(alpha) caused by resonant scattering in the Fe K(alpha) line. We argue that a central enhancement of SN Ia ejecta and hence a high ratio of Ni to Fe abundances are naturally explained by scenarios that include the generation of chemical gradients by suppressed SN Ia winds or ram pressure stripping of cluster galaxies. It is not necessary to suppose that the intracluster gas is optically thick to resonant scattering of the Fe K(alpha) line.

Dupke, Renato A.↗
NASA NTRS2012

Real-Time, Non-Intrusive Detection of Liquid Nitrogen in Liquid Oxygen at High Pressure and High Flow

An integrated fiber-optic Raman sensor has been designed for real-time, nonintrusive detection of liquid nitrogen in liquid oxygen (LOX) at high pressures and high flow rates in order to monitor the quality of LOX used during rocket engine ground testing. The integrated sensor employs a high-power (3-W) Melles Griot diode-pumped, solid-state (DPSS), frequency-doubled Nd:YAG 532- nm laser; a modified Raman probe that has built-in Raman signal filter optics; two high-resolution spectrometers; and photomultiplier tubes (PMTs) with selected bandpass filters to collect both N2 and O2 Raman signals. The PMT detection units are interfaced with National Instruments Lab- VIEW for fast data acquisition. Studies of sensor performance with different detection systems (i.e., spectrometer and PMT) were carried out. The concentration ratio of N2 and O2 can be inferred by comparing the intensities of the N2 and O2 Raman signals. The final system was fabricated to measure N2 and O2 gas mixtures as well as mixtures of liquid N2 and LOX

Singh, Jagdish P.↗
NASA NTRS2019

Ice-Crystal Icing Accretion Studies at the NASA Propulsion Systems Laboratory

This paper describes an ice-crystal icing experiment conducted at the NASA Propulsion System Laboratory during June 2018. This test produced ice shape data on an airfoil for different test conditions similar to those inside the compressor region of a turbo-fan jet engine. Mixed-phase icing conditions were generated by partially freezing out a water spray using the relative humidity of flow as the primary parameter to control freeze-out. The paper presents the ice shape data and associated conditions which include pressure, velocity, temperature, humidity, total water content, melt ratio, and particle size distribution. The test featured a new instrument traversing system which allowed surveys of the flow and cloud. The purpose of this work was to provide experimental ice shape data and associated conditions to help develop and validate ice-crystal icing accretion models. The results support previous experimental observations of a minimum melt-ratio threshold for accretion to occur as well as the existence of a plateau region where the icing severity is high for a range of melt ratios. However, a maximum limit for melt ratio, which is suggested in the ice crystal icing literature, was not observed perhaps complicated by the potential for some supercooling of the water at these conditions.

Engine Icing↗
NASA NTRSDate not supplied

A Case Study on the Challenges and Opportunities for the Deployment of PHM Capabilities in Existing Engineering Systems

The field of Prognostics and Health Management (PHM) of engineering systems has experienced considerable growth over the last decade. From benefits associated with faster and more powerful hardware in the form of wireless sensors, edge devices, and general computing capabilities (GPU’s and cloud computing), to development of powerful algorithms for anomaly detection and remaining useful life (RUL) estimation, the number of engineering systems featuring advanced diagnostics and prognostics capabilities continues to grow at an increasingly faster pace. However, the deployment of PHM capabilities as part of the upgrade of existing engineering systems presents multiple challenges to the PHM practitioner charged with retrofitting such systems. Issues include a lack of specific instrumentation needed to capture the signals of interest; insufficient data and sampling rates required for fault detection and diagnosis, and for detection of failure/degradation indicators; and difficulties in the identification of a system’s nominal behavior as a result of age induced degradation. Today’s PHM practitioner must be able to quickly identify and assess these types of issues to effectively evaluate and select the optimal PHM strategies required to achieve the desired results. This paper presents results from the preliminary evaluation of the High-Pressure Gas Facility (HPGF) infrastructure at NASA’s Stennis Space Center in Hancock County, Mississippi. This evaluation is part of a feasibility study conducted prior to the deployment of prognostics and diagnostics capabilities in the pumps skids of the liquid nitrogen (LN2) system of the HPGF.

Condition Based Maintenance↗
NASA NTRS1976

Ozone spectroscopy with a CO2 waveguide laser

Ozone spectral measurements were conducted with a waveguide CO2 laser, which was continuously tunable over a + or - 450-MHz region, about several lines in the P-branch of the band near the 9.5-micron wavelength region. A passively stabilized low-pressure CO2 laser was also used to provide reference CO2 frequencies. The frequencies of several ozone lines were determined in the nu-3 band, which are near CO2 laser line frequencies in the 9.5-micron wavelength region. The absolute measurement accuracy is estimated to be + or - 6 MHz, and the measurements indicate that the ozone line frequencies listed in currently available spectral compilations are 0.003-0.006 per cm too high. The size of this error is small by conventional IR standards, but extremely important when CO2 laser instruments are used to monitor stratospheric ozone, where line widths in this spectral region are typically 0.001/cm. Empirically determined absorption curves must be constructed in order to calibrate laser differential absorption instrument properly when low-altitude ozone measurements are conducted.

Menzies, R. T.↗
NASA NTRS1983

Flutter and steady/unsteady aerodynamic characteristics of supercritical and conventional transport wings

This paper presents the technical details and results of a high-speed wind-tunnel test program of an aeroelastic cantilevered transport type wing with two pylon-mounted engines. The tests were conducted in the NASA-Langley 16-foot Transonic Dynamic tunnel (TDT) during December 1981. Flutter of identical planforms, mass properties, and stiffness. The test parameters included different values of model stiffness and wing loading at various angles of attack. The models were instrumented at span-wise wing stations to determine bending and torsion deflections and vertical accelerations. At two model wing stations, pressure transducers were distributed along the chord to record static and unsteady oscillatory pressures during the approach to and onset of flutter. This paper presents the test program with results of the flutter characteristics and selected steady and unsteady aerodynamic data for both airfoils at different angles of attack for various Mach numbers and dynamic pressures.

Grosser, W. F.↗
NASA NTRS1993

In situ detection of tropospheric OH, HO2, NO2, and NO by laser-induced fluorescence in detection chambers at reduced pressures

This report is a brief summary of the status of work on the grant entitled 'In situ detection of tropospheric OH, HO2, NO2, and NO by laser induced fluorescence in detection chambers at low pressures.' The basic instrument characteristics have been established, and have been reported in a manuscript, included as an appendix to this report, that has been accepted by the Journal of Geophysical Research. Currently, two efforts are under way. First, instrument tests and calibrations are continuing. These efforts include field measurements and an informal inter comparison in Colorado last August and September. Second, new technologies in lasers and detectors are being implemented to make the instrument smaller, lighter, and more energy efficient. Such instrument modifications are essential for measurements from aircraft, high scaffolding in forests, and ships.

Brune, William H.↗
NASA NTRS2003

Gas Scintillation Proportional Counters for High-Energy X-ray Astronomy

A focal plane array of high-pressure gas scintillation proportional counters (GSPC) for a balloon-borne hard-x-ray telescope is under development at the Marshall Space Flight Center. These detectors have an active area of approx. 20 sq cm, and are filled with a high pressure (10(exp 6) Pa) xenon-helium mixture. Imaging is via crossed-grid position-sensitive phototubes sensitive in the UV region. The performance of the GSPC is well matched to that of the telescopes x-ray optics which have response to 75 keV and a focal spot size of approx. 500 microns. The detector s energy resolution, 4% FWHM at 60 keV, is adequate for resolving the broad spectral lines of astrophysical importance and for accurate continuum measurements. Full details of the instrument and its performance will be provided.

Gubarev, Mikhail↗
NASA NTRS2004

Microphysical Characterization of Ice in Hurricane Erin for Wideband Passive Microwave Comparisons

In order to better understand the characteristics of frozen hydrometeors in hurricane systems, computed brightness temperatures were compared with observations of Hurricane Erin from CAMEX-4 radiometers on the ER-2 aircraft. These wideband observations ranged from 10.7 to 183 plus or minus 10 GHz. In order to initialize the cloud profiles used in the radiative transfer calculations, radar reflectivities from an ER-2 instrument were converted to mass contents at specified height levels. Dropsonde data from the ER-2 and modeled Hurricane Andrew data were used to generate profiles of temperature, pressure, and relative humidity and to provide surface wind speed and ocean temperature. Four different ice (size and density) parameterizations were used to compute brightness temperature values. The comparisons between the calculated and observed brightness temperatures show that in general the Grenfell and Warren particle size distribution does the best job of matching the observations for most of the high frequencies. These high frequencies (greater than or equal to 85 GHz) are particularly sensitive to frozen particles. The Sehkon and Srivastava particle size distribution produced brightness temperatures that were too cold in most cases. On the other hand, the fluffy snow particle parameterizations with and without a melting layer produced brightness temperatures that were much warmer than the observations at these high frequencies. All low frequency calculations track the observations well except in the eye of Hurricane Erin where the surface wind speed and/or near surface temperatures were not properly adjusted to warm the brightness temperature calculations to values closer to the observations. Overall, this work shows that selecting a proper ice particle parameterization can be difficult and additional in situ measurements are needed to define appropriate parameterizations.

Skofronick-Jackson, G.↗
NASA NTRS2009

Compact Reconnaissance Imaging Spectrometer Observations of Water Vapor and Carbon Monoxide

The Compact Reconnaissance Imaging Spectrometer for Mars (CRISM) onboard the Mars Reconnaissance Orbiter (MRO) spacecraft began taking observations in September 2006 and has now collected more than a full Martian year of data. Retrievals performed using the near-infrared spectra obtained by CRISM are used to characterize the seasonal and spatial variation of the column abundance of water vapor and the column-averaged mixing ratio of carbon monoxide. CRISM retrievals show nominal behavior in water vapor during northern hemisphere spring and summer with maximum abundance reaching 50 precipitable micrometers. Water vapor abundance during the southern hemisphere spring and summer appears significantly reduced compared to observations by other instruments taken during previous years. The CRISM retrievals show the seasonally and globally averaged carbon monoxide mixing ratio to be 700 ppm, but with strong seasonal variations at high latitudes. The summertime near-polar carbon monoxide mixing ratio falls to 200 ppm in the south and 400 ppm in the north as carbon dioxide sublimates from the seasonal polar ice caps and dilutes noncondensable species including carbon monoxide. At low latitudes, the carbon monoxide mixing ratio varies in response to the mean seasonal cycle of surface pressure.

Smith, Michael D.↗
NASA NTRS2000

Test Cases for a Clipped Delta Wing with Pitching and Trailing-Edge Control Surface Oscillations

Steady and unsteady measured pressures for a Clipped Delta Wing (CDW) undergoing pitching oscillations and trailing-edge control surface oscillations have been presented . From the several hundred compiled data points, 22 static cases, 12 pitching-oscillation cases, and 12 control-surface-oscillation cases have been proposed for Computational Test Cases to illustrate the trends with Mach number, reduced frequency, and angle of attack. The planform for this wing was derived by simplifying the planform of a proposed design for a supersonic transport which is described as the Boeing 2707-300. The strake was deleted, the resulting planform was approximated by a trapezoid with an unswept trailing edge, and the twist and camber were removed. In order to facilitate pressure instrumentation, the thickness was increased to 6 percent from the typical 2.5 to 3 percent for the supersonic transport. The airfoil is thus a symmetrical circular arc section with t/c = 0.06. A wing of similar planform but with a thinner airfoil of t/c = 0.03 was used in the flutter investigations, and the buffet and stall flutter investigation . Flutter results are also reported both for the 3 per cent thick simplified wing and for a more complex SST model. One of the consequences of the increased thickness of the clipped delta wing is that transonic effects are enhanced for Mach numbers near one. They are significantly stronger than would be the case for the thinner wing. Also, with the combination of high leading edge sweep of 50.5, and the sharp leading edge, a leading edge vortex forms on the wing at relatively low angles of attack, on the order of three degrees. The Appendix discusses some of the vortex flow effects. In addition, a shock develops over the aft portion of the wing at transonic speeds such that at some angles of attack, there is both a leading edge vortex and a shock wave on the wing. Such cases are a computational challenge. Some previous applications of this data set have been for the evaluation of an aerodynamic panel method and for evaluation of a Navier-Stokes capability. Linear theory and panel method results are also presented, which demonstrated the need for inclusion of transonic effects. Flutter calculations for the related wing with t/c=O.O3 are given. In this report several Test Cases are selected to illustrate trends for a variety of different conditions with emphasis on transonic flow effects. An overview of the model and tests are given, and the standard formulary for these data is listed. For each type of data, a sample table and a sample plot of the measured pressures are presented. A complete tabulation and plotting of the Test Cases is given. Only the static pressures and the 1st harmonic real and imaginary parts of the pressures are available. All of the data for the test are included in a microfiche document in the original report and are available in electronic file form. The Test Cases are also available as separate electronic files.

Bennett, Robert M.↗