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At least 469 records · Page 26
All sourcesNASA NTRSDOE OSTI
NASA NTRS1995

The effect of impact angle on craters formed by hypervelocity particles

The Space Power Institute (SPI) at Auburn University has conducted experiments on the effects of impact angle on crater morphology and impactor residue retention for hypervelocity impacts. Copper target plates were set at angles of 30 deg, 45 deg, 60 deg, and 75 deg from the particle flight path. For the 30 deg and 45 deg impacts, in the velocity regime greater than 8 km s(exp -1) the resultant craters are almost identical to normal incidence impacts. The only difference found was in the apparent distribution of particle residue within the crater, and further research is needed to verify this. The 60 deg and 75 deg impacts showed marked differences in crater symmetry, crater lip shape, and particle residue distribution in the same velocity regime. Impactor residue shock fractionation effects have been quantified in first-order. It is concluded that a combination of analysis techniques can yield further information on impact velocity, direction, and angle of incidence.

Hill, David C.↗
NASA NTRS1995

PARC3D calculations of the F/A-18A HARV inlet vortex generators

NASA Lewis Research Center is currently engaged in a research effort as a team member of the High Alpha Technology Program within the NASA agency. This program uses a specially-equipped F/A-18A aircraft called the High Alpha Research Vehicle (HARV), in an effort to improve the maneuverability of high performance military aircraft at low-subsonic-speed, high-angle-of-attack conditions. The overall objective of the NASA Lewis effort is to develop inlet analysis technology towards efficient airflow delivery to the engine during these maneuvers. One portion of this inlet analysis technology uses computational fluid dynamics to predict installed inlet performance. Most of the F/A-18A HARV geometry, which includes the ramp/splitter plate, side diverter and slot, inlet lip and upper diverter, and deflected leading-edge flap has been modeled. The empennage and rear fuselage have not. A pair of vortex generators located on the bottom wall of the inlet were not modeled initially. These vortex generators were installed to alleviate any flow separation that may be induced by the wheel well protrusion into the inlet wall. Calculations completed with the PARC3D code showed that the pressure recovery has been underpredicted and the flow distortion over-predicted. To improve the correlation of PARC3D predictions with flight and wind tunnel tests, the vortex generators were included in the grid geometry and the results are presented in this report. The grid totals 27 blocks or 1.3 million grid points for the half model, which includes the vortex generator grid blocks. Two flight cases were calculated, a high speed case with a Mach number of 0.8 and angle of attack of 3.4; and a low speed case with a Mach number of 0.43 and angle of attack of 32.2. The vortex generators have a significant effect on the inlet boundary layers at high speed, low angle of attack; and have no effect at low speed, high angle of attack.

Podleski, Steve D.↗
NASA NTRS1996

Experimental Investigation of the Performance of a Mach-2.7 Two-dimensional Bifurcated Duct Inlet with 30 Percent Internal Contraction

An experimental study was conducted to determine the performance of a two-dimensional, mixed-compression bifurcated duct inlet system designed for a free-stream Mach number of 2.7. Thirty percent of the supersonic area contraction occurred internally. A movable ramp was used to vary the contraction ratio for off-design operation. Boundary layer bleed regions were located on the cowl, centerbody, and sidewall surfaces. There were also provisions for vortex generators on the cowl and centerbody of the subsonic diffuser. Data were obtained over the Mach number range of 2.0 to 2.8 and at angles of yaw from 0 deg. to the maximum value prior to inlet un-start. The test at Mach 2.8 was to obtain data for an over- speed condition. The Reynolds number varied from 2.5 to 2.3 million/ft for Mach numbers above 2.5. At Mach numbers of 2.5 and lower, the Reynolds number was set at 2.5 million/ft. Bleed patterns, vortex generator patterns, and ramp position were varied, and three inlet configurations were selected for more extensive study. Two of these configurations had self-starting capability. The self-starting configuration that was developed produced 89 percent total pressure recovery at the compressor face station with 6.8 percent total bleed. The compressor face distortion was about 16 percent. Vortex generators were extremely effective in re-distributing flow but were not as effective in reducing distortion. Excellent flow symmetry was achieved between the separated halves of the inlet, and twin-duct instability was not observed. The ramp tip shock was steeper than expected. This caused the cowl lip shock to be reflected from the ramp instead of being cancelled at the shoulder. However, peak recovery at the throat was still obtained with the ramp near the design position.

Wasserbauer, J. F.↗
NASA NTRS1996

Pulse Detonation Engine Air Induction System Analysis

A preliminary mixed-compression inlet design concept for potential pulse-detonation engine (PDE) powered supersonic aircraft was defined and analyzed. The objectives of this research were to conceptually design and integrate an inlet/PDE propulsion system into a supersonic aircraft, perform time-dependent CFD analysis of the inlet flowfield, and to estimate the installed PDE cycle performance. The study was baselined to a NASA Mach 5 Waverider study vehicle in which the baseline over/under turboramjet engines were replaced with a single flowpath PDE propulsion system. As much commonality as possible was maintained with the baseline configuration, including the engine location and forebody lines. Modifications were made to the inlet system's external ramp angles and a rotating cowl lip was incorporated to improve off-design inlet operability and performance. Engines were sized to match the baseline vehicle study's ascent trajectory thrust requirement at Mach 1.2. The majority of this study was focused on a flight Mach number of 3.0. The time-dependent Navier Stokes CFD analyses of a two-dimensional approximation of the inlet was conducted for the Mach 3.0 condition. The Lockheed Martin Tactical Aircraft Systems-developed FALCON CFD code with a two equation 'k-1' turbulence model was used. The downstream PDE was simulated by an array of four sonic nozzles in which the flow areas were rapidly varied in various opening/closing combinations. Results of the CFD study indicated that the inlet design concept operated successfully at the Mach 3.0 condition, satisfying mass capture, total pressure recovery, and operability requirements. Time-dependent analysis indicated that pressure and expansion waves from the simulated valve perturbations did not effect the inlet's operability or performance.

Pegg, R. J.↗
NASA NTRS1995

Computational Investigation of Tangential Slot Blowing on a Generic Chined Forebody

The effect of tangential slot blowing on the flowfield about a generic chined forebody at high angles of attack is investigated numerically using solutions of the thin-layer, Reynolds-averaged, Navier-Stokes equations. The effects of jet mass now ratios, angle of attack, and blowing slot location in the axial and circumferential directions are studied. The computed results compare well with available wind-tunnel experimental data. Computational results show that for a given mass now rate, the yawing moments generated by slot blowing increase as the body angle of attack increases. It is observed that greater changes in the yawing moments are produced by a slot located closest to the lip of the nose. Also, computational solutions show that inboard blowing across the top surface is more effective at generating yawing moments than blowing outboard from the bottom surface.

Agosta-Greenman, Roxana M.↗
NASA NTRS1995

Nozzle Thrust Optimization While Reducing Jet Noise

A Bluebell nozzle design concept is proposed for jet noise reduction with minimal thrust loss or even thrust augmentation. A Bluebell nozzle has a sinusoidal lip line edge (chevrons) and a sinusoidal cross section shape with linear amplitude increasing downstream in the divergent nozzle part (corrugations). The experimental tests of several Bluebell nozzle designs have shown nose reduction relative to a convergent-divergent round nozzle with design exhaust number M(e) = 1.5. The best design provides an acoustic benefit near 4dB with about 1 percent thrust augmentation. For subsonic flow ((M(e)= 0.6)), the tests indicated that the present method for design of Bluebell nozzles gives less acoustic benefit and in most cases jet noise increased. The proposed designs incorporate analytical theory and 2D and 3D numerical simulations. Full Navier-Stokes and Euler solvers were utilized. Boundary layer effects were used. Several different designs were accounted for in the Euler applications.

Seiner, J. M.↗
NASA NTRS1995

Numerical Analysis of Base Flowfield for a Four-Engine Clustered Nozzle Configuration

Excessive base heating has been a problem for many launch vehicles. For certain designs such as the direct dump of turbine exhaust inside and at the lip of the nozzle, the potential burning of the turbine exhaust in the base region can be of great concern. Accurate prediction of the base environment at altitudes is therefore very important during the vehicle design phase. Otherwise, undesirable consequences may occur. In this study, the turbulent base flowfield of a cold flow experimental investigation for a four-engine clustered nozzle was numerically benchmarked using a pressure-based computational fluid dynamics (CFD) method. This is a necessary step before the benchmarking of hot flow and combustion flow tests can be considered. Since the medium was unheated air, reasonable prediction of the base pressure distribution at high altitude was the main goal. Several physical phenomena pertaining to the multiengine clustered nozzle base flow physics were deduced from the analysis.

Wang, Ten-See↗
NASA NTRS1997

Response of High Subsonic Jet to Nonaxisymmetric Disturbances

A model of sound generated in a high subsonic (Mach 0.9) circular jet is solved numerically in cylindrical coordinates for nonaxisymmetric disturbances. The jet is excited by transient mass injection by a finite duration pulse via a modulated ring source. The nonaxisymmetric solution is computed for long times after the initial disturbance has exited the computational domain. The long time behavior of the jet is dominated by vorticity and pressure disturbances generated at the nozzle lip and growing as they convect down-stream in the jet. These disturbances generate sound as they propagate. The primary non-axisymmetric effect that we simulate is that of a flapping mode where regions of high and low pressure alternate on opposite sides of the jet. The predominant feature of this mode is the appearance of relatively large deviations of the pressure from the ambient pressure on opposite sides of the jet and the convection of these regions downstream. We illustrate flow field, near field and far field data. Important nonaxisymmetric characteristics of the near and flow field disturbances include roughly periodic pressure elevations and depressions at opposite values of the azimuthal angle psi. These correspond to pressure disturbances propagating in the axial direction. The azimuthal velocity exhibits a sinusoidal dependence on psi with similar roughly periodic disturbances. For every azimuthal angle psi, the jet radiation peaks about 30 deg. from the jet axis, however there is now a pronounced dependence of the far field radiation pattern on psi.

Bayliss, A.↗
NASA NTRS1997

Screech Tones of Supersonic Jets from Bevelled Rectangular Nozzles

It is known experimentally that an imperfectly expanded rectangular jet from a thin-lip convergent nozzle emits only a single dominant screech tone. The frequency of the screech tone decreases continuously with increase in jet Mach number. However, for a supersonic jet issued from a bevelled nozzle or a convergent-divergent nozzle with straight side walls, the shock cell structure and the screech frequency pattern are fairly complicated and have not been predicted before. In this paper, it is shown that the shock cell structures of these jets can be decomposed into waveguide modes of the jet flow. The screech frequencies are related to the higher-order waveguide modes following the weakest-link screech tone theory. The measured screech frequencies are found to compare well with the predicted screech frequency curves.

Tam, Christopher K. W.↗
NASA NTRS1996

Comparison of the Aeroacoustics of Two Small-Scale Supersonic Inlets

An aerodynamic and acoustic investigation was performed on two small-scale supersonic inlets to determine which inlet would be more suitable for a High Speed Civil Transport (HSCT) aircraft during approach and takeoff flight conditions. The comparison was made between an axisymmetric supersonic P inlet and a bifurcated two-dimensional supersonic inlet. The 1/14 scale model supersonic inlets were used in conjunction with a 4.1 in (10.4 cm) turbofan engine simulator. A bellmouth was utilized on each inlet to eliminate lip separation commonly associated with airplane engine inlets that are tested under static conditions. Steady state measurements of the aerodynamic flowfield and acoustic farfield were made in order to evaluate the aeroacoustic performance of the inlets. The aerodynamic results show the total pressure recovery of the two inlets to be nearly identical, 99% at the approach condition and 98% at the takeoff condition. At the approach fan speed (60% design speed), there was no appreciable difference in the acoustic performance of either inlet over the entire 0 deg to 110 deg farfield measurement sector. The inlet flow field results at the takeoff fan speed (88% design speed), show the average inlet throat Mach number for the P inlet (Mach 0.52) to be approximately 2 times that of the 2D inlet (Mach 0.26). The difference in the throat Mach number is a result of the smaller throughflow area of the P inlet. This reduced area resulted in a 'soft choking' of the P inlet which lowered the tone and overall sound pressure levels of the simulator in the forward sector by an average of 9 dB and 3 dB, respectively, when compared to the 2D inlet.

Ng, Wing↗
NASA NTRS1997

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

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.

Ponton, Michael K.↗
NASA NTRS1997

Survey of Aerothermodynamics Facilities Useful for the Design of Hypersonic Vehicles Using Air-Breathing Propulsion

This paper surveys the use of aerothermodynamic facilities which have been useful in the study of external flows and propulsion aspects of hypersonic, air-breathing vehicles. While the paper is not a survey of all facilities, it covers the utility of shock tunnels and conventional hypersonic blow-down facilities which have been used for hypersonic air-breather studies. The problems confronting researchers in the field of aerothermodynamics are outlined. Results from the T5 GALCIT tunnel for the shock-on lip problem are outlined. Experiments on combustors and short expansion nozzles using the semi-free jet method have been conducted in large shock tunnels. An example which employed the NASA Ames 16-Inch shock tunnel is outlined, and the philosophy of the test technique is described. Conventional blow-down hypersonic wind tunnels are quite useful in hypersonic air-breathing studies. Results from an expansion ramp experiment, simulating the nozzle on a hypersonic air-breather from the NASA Ames 3.5 Foot Hypersonic wind tunnel are summarized. Similar work on expansion nozzles conducted in the NASA Langley hypersonic wind tunnel complex is cited. Free-jet air-frame propulsion integration and configuration stability experiments conducted at Langley in the hypersonic wind tunnel complex on a small generic model are also summarized.

Arnold, James O.↗
NASA NTRS1998

Experimental and Theoretical Study of Flow Fields Around Ducted-Nacelle Models

The flow field near four small-scale ducted-nacelle bodies of revolution has been analytically and experimentally studied to determine exterior and interior mass-flow characteristics, and to measure flow-field overpressures generated by the nacelle's forebody shape. Four nacelle models with the same profile, but of different sizes, were used in the study. Shadowgraph pictures showed inlet shocks attached to the cowl lip (indicating unchoked flow) on all four models, at all the test Mach numbers, through an angle of attack range of 0.0 to 6.0 degrees. Pressure signatures measured in the flow field of the largest of the four nacelle models were compared with those predicted by corrected and uncorrected Whitham theory. At separation distances greater than 3.0 to 4.0 inlet diameters, good agreement was found. Poorer agreement was found at extreme near-field separation distances, but this was attributed to pressure-gage limitations and probe-flow field interactions. The overall favorable results supported a conclusion that corrected Whitham theory was sufficiently accurate to make the nacelle-wing interference-lift code useful for sonic-boom analysis and the preliminary design of supersonic-cruise conceptual aircraft.

Mack, Robert J.↗
NASA NTRS1961

The Effect of Air Bleed on the Heat Transfer and Pressure Distribution on 30 Conical Flares at a Mach Number of 6.8

The effect of a boundary-layer bleed at the start of a 30 deg half-angle flare upon the shape of the flow boundaries, the pressure distribution on the flare, and the heat transfer to the flare was studied at a Mach number of 6.8. The forebody was an ogive cylinder. Test Reynolds numbers, based on forebody length, ranged from 1 x 10(exp 6) to 7.4 x 10(exp 6). Schlieren photographs showed the effect produced upon the flow boundaries by varying the dimensions of the bleed in both the radial and axial directions and by blunting the lip at the leading edge of the flare. The heat transfer and pressure distribution on the flare were correlated with the shape and nature of the flow boundaries.

Crawford, Davis H.↗
NASA NTRS1998

Lineage Analysis of Axis Formation Under Novel Gravity

Recent intriguing work by Cooke ('86) and Neff, et al. ('93) suggests that there are subtle developmental changes in the Xenopus laevis embryos subjected to novel gravitational fields. These changes include the position of the third cleavage plane, the dorsal lip of the blastopore, and also the size of the head and eyes. However, compensation occurred later in development, so that by the tadpole stages there is no apparent difference between experimental and control embryos. How these early morphological changes are corrected is not clear. Through this project, we plan to determine whether the distribution of cytoplasmic morphogenetic determinants, and thus the developmental fate of blastomeres, is altered by novel gravitational fields by either tilting them or rotating them in a horizontal clinostat. We then plan to compare the control and experimental embryos with respect to blastomere fate (by lineage tracing with fluorescent dextrans), blastomere commitment and autonomous differentiation potential (by transplantation and culture), and distribution of cytoplasmic morphogens (by in situ hybridization). These three approaches, when applied in tandem, will provide a definitive test of the hypothesis that the distribution of cytoplasmic morphogenetic determinants and thus the developmental fate of blastomeres can be altered by novel gravitational fields.

Huang, Sen↗
NASA NTRS1999

Far Ultraviolet Imaging from the Image Spacecraft: System Design - 1

Direct imaging of the magnetosphere by the IMAGE spacecraft A,ill be supplemented by observation of the global aurora, the footprint of magnetospheric regions. To assure the simultaneity of these observations and tile measurement of the magnetospheric back-round neutral gas density, the IMAGE satellite instrument complement includes three Far Ultraviolet (FUV) instruments. In tile wavelength region 120-1 90 nm. a downward-viewing aurora imager is only minimally contaminated sunlight, scattered from clouds and ground, and radiance of the aurora observed in a nadir viewing geometry can be observed in the presence of the high-latitude day-low. Tile Wideband Imaging Camera (WIC) will provide broadband ultraviolet images of the aurora for maximum spatial and temporal resolution by imaging the LBH N2 bands of the aurora. The Spectrographic Imager (SI), a monochromatic imager, will image different types of aurora, filtered by wavelength. By measuring the Doppler-shifted Lyman-alpha, the proton-induced component of the aurora will be imaged separately. Finally, the GEO instrument will observe the distribution of the geocoronal emission, which is a measure of the neutral background density source of the charge exchange in the magnetosphere. The FUV instrument complement looks radially outward from the rotating IMAGE satellite and, therefore, it spends only a short time observing tile aurora and the Earth during, each spin. Detailed descriptions of the WIC, Si, GEO, and their individual performance validations are discussed in companion papers. This paper summarizes the system requirements and system design approach taken to satisfy the science requirements. One primary requirement is to maximize photon collection efficiency and use efficiently tile short time available foe exposures. The FUV auroral imagers WIC and SI both have wide fields of view and take data continuously as the auroral region proceeds through the field of view. To minimize data volume, multiple images are taken and electronically co-added by suitably shifting each image to compensate for the spacecraft rotation. In order to minimize resolution loss, the images have to be distortion-corrected in real time foe both WIC and Si prior to co-adding. The distortion correction is accomplished using high-speed look Lip tables that are pre-generated by least square fitting to polynomial functions by the on-orbit processor. The instruments were calibrated individually while on stationary platforms, mostly in vacuum chambers as described in the companion papers. Extensive -round-based testing was performed with visible and near UV simulators mounted oil a rotating platform to estimate their on-orbit performance. The predicted instrument system performance is summarized and some of the preliminary data formats are shown.

Mende, S. B.↗
NASA NTRS1999

Digging for Fossils in the Hertzsprung Gap

Objective was to conduct deep (approx. 250 ks) pointings on two EUV sources, the early-F giant beta Cas and the mid-G giant mu Velorum; to obtain spectra in the range 70-300 A and to record Deep Survey light curves over the extensive duration of each observation. We have analyzed the DS lightcurve and the SW spectrum, breaking the latter lip into time slices corresponding to key phases of the observation: pre-flare, flare rise, and two segments of the flare decay.

Ayres, Thomas R.↗