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At least 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Calculation of unsteady linearized Euler flows in cascades using harmonically deforming grids

A method for calculating unsteady, inviscid, compressible flows in cascades is presented. Using the linearized Euler technique, the flow is decomposed into a steady or mean flow plus a harmonically varying small disturbance flow. The equations that describe the small disturbance flow are linear variable coefficient equations, and are solved using a pseudo-time time marching Lax-Wendroff technique. Unlike previous linearized methods, however, the solution is computed on a harmonically deforming computational grid that conforms to the motion of the vibrating airfoils. The mean flow and perturbation flow solutions are defined in the deforming coordinate system rather than in a coordinate system fixed in space. Hence, no extrapolation terms are required to implement the upwash boundary conditions at the airfoil surfaces, significantly improving the accuracy of the method. For transonic flow calculations, unsteady shock motions are modelled using shock capturing. The unsteady loads due to the shock motion are then seen as pressure impulses. Representative computational results are presented for transonic channel flows and subsonic and transonic cascade flows.

Hall, Kenneth C.↗
NASA NTRS1993

Supersonic vortex breakdown over a delta wing in transonic flow

The effects of freestream Mach number and angle of attack on the leading-edge vortex breakdown due to the terminating shock on a 65-degree, sharp-edged, cropped delta wing are investigated computationally, using the time-accurate solution of the laminar unsteady compressible full Navier-Stokes equations with the implicit upwind flux-difference splitting, finite-volume scheme. A fine O-H grid consisting of 125 x 85 x 84 points in the wrap-around, normal, and axial directions, respectively, is used for all the flow cases. Keeping the Reynolds number fixed at 3.23 x 10 exp 6, the Mach number is varied from 0.85 to 0.9 and the angle of attack is varied from 20 to 24 deg. The results show that, at 20-deg angle of attack, the increase of the Mach number from 0.85 to 0.9 results in moving the location of the terminating shock downstream. The results also show that, at 0.85 Mach number, the increase of the angle of attack from 20 to 24 deg results in moving the location of the terminating shock upstream. The results are in good agreement with the experimental data.

Kandil, Hamdy A.↗
NASA NTRS2011

Performance Tests of Snow-Related Variables Over the Tibetan Plateau and Himalayas Using a New Version of NASA GEOS-5 Land Surface Model that Includes the Snow Darkening Effect

The snow darkening effect, i.e. the reduction of snow albedo, is caused by absorption of solar radiation by absorbing aerosols (dust, black carbon, and organic carbon) deposited on the snow surface. This process is probably important over Himalayan and Tibetan glaciers due to the transport of highly polluted Atmospheric Brown Cloud (ABC) from the Indo-Gangetic Plain (IGP). This effect has been incorporated into the NASA Goddard Earth Observing System model, version 5 (GEOS-5) atmospheric transport model. The Catchment land surface model (LSM) used in GEOS-5 considers 3 snow layers. Code was developed to track the mass concentration of aerosols in the three layers, taking into account such processes as the flushing of the compounds as liquid water percolates through the snowpack. In GEOS-5, aerosol emissions, transports, and depositions are well simulated in the Goddard Chemistry Aerosol Radiation and Transport (GO CART) module; we recently made the connection between GOCART and the GEOS-5 system fitted with the revised LSM. Preliminary simulations were performed with this new system in "replay" mode (i.e., with atmospheric dynamics guided by reanalysis) at 2x2.5 degree horizontal resolution, covering the period 1 November 2005 - 31 December 2009; we consider the final three years of simulation here. The three simulations used the following variants of the LSM: (1) the original Catchment LSM with a fixed fresh snowfall density of 150 kg m-3 ; (2) the LSM fitted with the new snow albedo code, used here without aerosol deposition but with changes in density formulation and melting water effect on snow specific surface area, (3) the LSM fitted with the new snow albedo code as same as (2) but with fixed aerosol deposition rates (computed from GOCART values averaged over the Tibetan Plateau domain [Ion.: 60-120E; lat.: 20-50N] during March-May 2008) applied to all grid points at every time step. For (2) and (3), the same setting on the fresh snowfall density as in (1) was used.

Yasunari, Tppei J.↗
NASA NTRS1983

Wind turbine generator interaction with diesel generators on an isolated power system

The results of a dynamic interaction investigation to characterize any disturbances caused by interfacing the Mod 0A wind turbine (150 kW configuration) with the Block Island utility diesel generator grid are reported. The tests were run when only two diesel generators were on line, and attention was given to power, frequency, and voltage time profiles. The interconnected system was examined in the start-up and synchronization phase, normal shutdown and cut-out of the wind turbine, during fixed pitch generation, and during variable pitch operation. Governors were installed on the diesel generators to accommodate the presence of wind-derived electricity. The blade pitch control was set to maintain power at 150 kW or below. Power and voltage transients were insignificant during start-up and shutdown, and frequency aberrations were within the range caused by load fluctuations. It is concluded that wind turbine generation can be successfully implemented by an isolated utility, even with a significant penetration to the total grid output.

Scott, G. W.↗
NASA NTRS1982

Flight evaluation of Loran-C for general aviation area navigation

This paper reports on a flight evaluation of Loran-C which was part of a long range study of area navigation systems for general aviation. Tests involved two different Loran-C receivers, 6 different aircraft, and a variety of antennas. Uncorrected position fixes were typically accurate to one quarter mile. With measured corrections, repeatability was good to within 200 ft. Signal reliability was 99.7%. The receiver was not sensitive to atmospheric noise. The time difference grid demonstrated a long term stability of 0.3 microsecond. Vertical whip and ADF E-field antennas were found suitable for airborne use. Loran-C was found satisfactory for instrument approaches to runways at general aviation airports where published latitude-longitude coordinates were available. Accuracy was further improved by using locally measured Loran-C time difference coordinates.

Hollister, W. M.↗
NASA NTRS1998

Computational Investigation and Validation of Twin-Tail Buffet Response Including Dynamics and Control

Multidisciplinary tools for prediction of single rectangular-tail buffet are extended to single swept-back-tail buffet in transonic-speed flow, and multidisciplinary tools for prediction and control of twin-tail buffet are developed and presented. The configuration model consists of a sharp-edged delta wing with single or twin tails that are oriented normal to the wing surface. The tails are treated as cantilevered beams fixed at the root and allowed to oscillate in both bending and torsion. This complex multidisciplinary problem is solved sequentially using three sets of equations on a dynamic single or multi-block grid structure. The first set is the unsteady, compressible, Reynolds-averaged Navier-Stokes equations which are used for obtaining the flow field vector and the aerodynamic loads on the tails. The Navier-Stokes equations are solved accurately in time using the implicit, upwind, flux-difference splitting, finite volume scheme. The second set is the coupled bending and torsion aeroelastic equations of cantilevered beams which are used for obtaining the bending and torsion deflections of the tails. The aeroelastic equations'are solved accurately in time using, a fifth-order-accurate Runge-Kutta scheme. The third set is the grid-displacement equations and the rigid-body dynamics equations, which are used for updating the grid coordinates due to the tail deflections and rigid-body motions. The tail-buffet phenomenon is predicted for highly-swept, single vertical tail placed at the plane of geometric symmetry, and for highly-swept, vertical twin tails placed at three different spanwise separation distances. The investigation demonstrates the effects of structural inertial coupling and uncoupling of the bending and torsion modes of vibration, spanwise positions of the twin-tail, angle of attack, and pitching and rolling dynamic motions of the configuration model on the tail buffet loading and response. The fundamental issue of twin-tail buffet alleviation is addressed using two active flow-control methods. These methods are the tangential leading-edge blowing and the flow suction from the leading-edge vortex cores along their paths. Qualitative and quantitative comparisons with the available experimental data are presented. The comparisons indicate that the present multidisciplinary aeroelastic analysis tools are robust, accurate and efficient.

Kandil, Osama A.↗
NASA NTRS1978

Mounting structure

A mounting platform for heat producing instruments operated in a narrow equilibrium temperature range comprises a grid-like structure with relatively large openings therein. The instruments are secured to and thermally coupled with the grid surface facing the instruments. Excess heat from the instruments is selectively radiated to the ambient through openings in the grid, the grid surfaces at these openings exhibiting low thermal emissivity and adsorptivity. The remainder of the grid is maintained at the equilibrium temperature and is covered with a thermal insulating blanket. Thus, the entire system including the platform and instruments is maintained substantially isothermal, whereby the instruments remain in fixed physical relationship to one another.

Ganssle, Eugene Robert↗
NASA NTRS2014

Aeroelastic Analyses of the SemiSpan SuperSonic Transport (S4T) Wind Tunnel Model at Mach 0.95

Detailed aeroelastic analyses of the SemiSpan SuperSonic Transport (S4T) wind tunnel model at Mach 0.95 with a 1.75deg fixed angle of attack are presented. First, a numerical procedure using the Computational Fluids Laboratory 3-Dimensional (CFL3D) Version 6.4 flow solver is investigated. The mesh update method for structured multi-block grids was successfully applied to the Navier-Stokes simulations. Second, the steady aerodynamic analyses with a rigid structure of the S4T wind tunnel model are reviewed in transonic flow. Third, the static analyses were performed for both the Euler and Navier-Stokes equations. Both the Euler and Navier-Stokes equations predicted a significant increase of lift forces, compared to the results from the rigid structure of the S4T wind-tunnel model, over various dynamic pressures. Finally, dynamic aeroelastic analyses were performed to investigate the flutter condition of the S4T wind tunnel model at the transonic Mach number. The condition of flutter was observed at a dynamic pressure of approximately 75.0-psf for the Navier-Stokes simulations. However, it was observed that the flutter condition occurred a dynamic pressure of approximately 47.27-psf for the Euler simulations. Also, the computational efficiency of the aeroelastic analyses for the S4T wind tunnel model has been assessed.

Hur, Jiyoung↗
NASA NTRS1991

Design optimization of transonic airfoils

Numerical optimization procedures were considered for the design of airfoils in transonic flow based on the transonic small disturbance (TSD) and Euler equations. A sequential approximation optimization technique was implemented with an accurate approximation of the wave drag based on the Nixon's coordinate straining approach. A modification of the Euler surface boundary conditions was implemented in order to efficiently compute design sensitivities without remeshing the grid. Two effective design procedures producing converged designs in approximately 10 global iterations were developed: interchanging the role of the objective function and constraint and the direct lift maximization with move limits which were fixed absolute values of the design variables.

Joh, C.-Y.↗
NASA NTRS1985

Computation of unsteady transonic aerodynamics with steady state fixed by truncation error injection

A novel technique is introduced for efficient computations of unsteady transonic aerodynamics. The steady flow corresponding to body shape is maintained by truncation error injection while the perturbed unsteady flows corresponding to unsteady body motions are being computed. This allows the use of different grids comparable to the characteristic length scales of the steady and unsteady flows and, hence, allows efficient computation of the unsteady perturbations. An example of typical unsteady computation of flow over a supercritical airfoil shows that substantial savings in computation time and storage without loss of solution accuracy can easily be achieved. This technique is easy to apply and requires very few changes to existing codes.

Fung, K.-Y.↗
NASA NTRS1991

Reynolds number effects on supersonic asymmetrical flows over a cone at high angle of attack

The supersonic viscous flow over a 5-degree half-angle cone at an angle of attack of four times the cone half-angle is studied computationally using both the conical and the three-dimensional Navier-Stokes equations. The numerical solutions were obtained with an implicit, upwind-biased algorithm. Asymmetrical flowfields of the absolute-instability type are found using the conical-flow equations which agree with published results. However, the absolute instabilities of the originally symmetric flow found with the conical equations do not occur in the three-dimensional simulations, although spurious asymmetric three-dimensional flows for symmetric bodies arise if the grid resolution is insufficient in the nose region. The asymmetric flows computed with the three-dimensional equations are convective instabilities and are possible if the local Reynolds number exceeds a critical value and a fixed geometric asymmetry is imposed. A continuous range of asymmetries can be developed, depending on the size of the disturbance and the Reynolds number. As the Reynolds number is increased, the asymmetries demonstrate a bistable behavior at levels of side force consistent with those predicted using the conical equations. Below a certain critical Reynolds number, any flow asymmetries arising from a geometrical asymmetry are damped with increasing distance downstream from the geometrical asymmetry.

Thomas, J. L.↗
NASA NTRS1992

Analysis of an advanced ducted propeller subsonic inlet

A time marching Navier-Stokes code called PARC (PARC2D for 2-D/axisymmetric and PARC3D for 3-D flow simulations) was validated for an advanced ducted propeller (ADP) subsonic inlet. The code validation for an advanced ducted propeller (ADP) subsonic inlet. The code validation was implemented for a non-separated flow condition associated with the inlet operating at angles-of-attack of 0 and 25 degrees. The inlet test data were obtained in the 9 x 15 ft Low Speed Wind Tunnel at NASA Lewis Research Center as part of a cooperative study with Pratt and Whitney. The experimental study focused on the ADP inlet performance for take-off and approach conditions. The inlet was tested at a free stream Mach number of 0.2, at angles-of-attack between O and 35 degrees, and at a maximum propeller speed of 12,000 RPM which induced a corrected air flow rate of about 46 lb/sec based on standard day conditions. The computational grid and flow boundary conditions (BC) were based on the actual inlet geometry and the funnel flow conditions. At the propeller face, two types of BC's were applied: a mass flow BC and a fixed flow properties BC. The fixed flow properties BC was based on a combination of data obtained from the experiment and calculations using a potential flow code. Comparison of the computational results with the test data indicates that the PARC code with the propeller face fixed flow properties BC provided a better prediction of the inlet surface static pressures than the predictions when the mass flow BC was used. For an angle-of-attack of 0 degrees, the PARC2D code with the propeller face mass flow BC provided a good prediction of inlet static pressures except in the region of high pressure gradient. With the propeller face fixed flow properties BC, the PARC2D code provided a good prediction of the inlet static pressures. For an angle-of-attack of 25 degrees with the mass flow BC, the PARC3D code predicted statis pressures which deviated significantly from the test data; however, with the fixed flow properties BC, a good comparison with the test data was obtained.

Iek, Chanthy↗
NASA NTRS1998

Effect of Configuration Pitching Motion on Twin Tail Buffet Response

The effect of dynamic pitch-up motion of delta wing on twin-tail buffet response is investigated. The computational model consists of a delta wing-twin tail configuration. The computations are carried out on a dynamic multi-block grid structure. This multidisciplinary problem is solved using three sets of equations which consists of the unsteady Navier-Stokes equations, the aeroelastic equations, and the grid displacement equations. The configuration is pitched-up from zero up to 60 deg. angle of attack, and the freestream Mach number and Reynolds number are 0.3 and 1.25 million, respectively. With the twin tail fixed as rigid surfaces and with no-forced pitch-up motion, the problem is solved for the initial flow conditions. Next, the problem is solved for the twin-tail response for uncoupled bending and torsional vibrations due to the unsteady loads on the twin tail and due to the forced pitch-up motion. The dynamic pitch-up problem is also solved for the flow response with the twin tail kept rigid. The configuration is investigated for inboard position of the twin tail which corresponds to a separation distance between the twin tail of 33% wing chord. The computed results are compared with the available experimental data.

Sheta, Essam F.↗
NASA NTRS1987

Initial turbulence effect on jet evolution with and without tonal excitation

The effect of initial turbulence level on the development of a jet and on the susceptability of the jet to discrete tone excitation was experimentally investigated. Turbulence intensity was varied, over the range 0.15 to 5 percent, by using screens and grids placed upstream of an 8.8 cm diameter nozzle. Top-hat mean velocity profiles with approximately identical initial boundary layer states were ensured in all cases; the turbulence spectra were broadband. It was found, contrary to earlier reports, that the natural jet decay remained essentially unchanged for varying initial turbulence. For a fixed amplitude of the tonal excitation, increasing the initial turbulence damped out the growth of the instability wave; as a result, the excitability, assessed from the mean velocity decay on the axis, was found to diminish. However, the degree of damping in the amplification of the instability wave was only slight compared to the large increase in the initial turbulence. The jet with 5 percent turbulence could be measurably altered by excitation with a velocity perturbation amplitude as little as 0.25 percent of the jet velocity. The amplitude effect data indicate an upper bound of the extent to which a jet could be excited, and thus its plume shortened, by the plane wave, single frequency excitation. An additional data set with no grid or trip, yielding a nominally laminar boundary layer, re-emphasizes the profound effect of initial boundary layer state on jet evolution as well as on its excitability. This jet decayed the fastest naturally, and consequently, it was the least excitable in spite of its turbulence being the least.

Raman, G.↗
NASA NTRS1987

Initial turbulence effect on jet evolution with and without tonal excitation

The effect of initial turbulence level on the development of a jet and on the susceptibility of the jet to discrete tone excitation was experimentally investigated. Turbulence intensity was varied, over the range 0.15 to 5 percent, by using screens and grids placed upstream of an 8.8 cm diameter nozzle. Top-hat mean velocity profiles with approximately identical initial boundary layer states were ensured in all cases; the turbulence spectra were broadband. It was found, contrary to earlier reports, that the natural jet decay remained essentially unchanged for varying initial turbulence. For a fixed amplitude of the tonal excitation, increasing the initial turbulence damped out the growth of the instability wave; as a result, the excitability, assessed from the mean velocity decay on the axis, was found to diminish. However, the degree of damping in the amplification of the instability wave was only slight compared to the large increase in the initial turbulence. The jet with 5 percent turbulence could be measurably altered by excitation with a velocity perturbation amplitude as little as 0.25 percent of the jet velocity. The amplitude effect data indicate an upper bound of the extent to which a jet could be excited, and thus its plume shortened, by the plane wave, single frequency excitation. An additional data set with no grid or trip, yielding a nominally laminar boundary layer, re-emphasizes the profound effect of initial boundary layer state on jet evolution as well as on its excitability. This jet decayed the fastest naturally, and consequently, it was the least excitable in spite of its turbulence being the least.

Raman, G.↗
NASA NTRS2002

Aeroelastic Calculations of Quiet High- Speed Fan Performed

An advanced high-speed fan was recently designed under a cooperative effort between the NASA Glenn Research Center and Honeywell Engines & Systems. The principal design goals were to improve performance and to reduce fan noise at takeoff. Scale models of the Quiet High-Speed Fan were tested for operability, performance, and acoustics. During testing, the fan showed significantly improved noise characteristics, but a self-excited aeroelastic vibration known as flutter was encountered in the operating range. Flutter calculations were carried out for the Quiet High-Speed Fan using a three-dimensional, unsteady aerodynamic, Reynolds-averaged Navier-Stokes turbomachinery code named "TURBO." The TURBO code can accurately model the viscous flow effects that can play an important role in various aeroelastic problems such as flutter with flow separation, flutter at high loading conditions near the stall line (stall flutter), and flutter in the presence of shock and boundary-layer interaction. Initially, calculations were performed with no blade vibrations. These calculations were at a constant rotational speed and a varying mass flow rate. The mass flow rate was varied by changing the backpressure at the exit boundary of the computational domain. These initial steady calculations were followed by aeroelastic calculations in which the blades were prescribed to vibrate harmonically in a natural mode, at a natural frequency, and with a fixed interblade phase angle between adjacent blades. The AE-prep preprocessor was used to interpolate the in-vacuum mode shapes from the structural dynamics mesh onto the computational fluid dynamics mesh and to smoothly propagate the grid deformations from the blade surface to the interior points of the grid. The aeroelastic calculations provided the unsteady aerodynamic forces on the blade surface due to blade vibrations. These forces were vector multiplied with the structural dynamic mode shape to calculate the work done on the blade during one vibration period, then this result was converted to an aerodynamic damping. Flutter occurs when the aerodynamic damping becomes negative, if structural damping is ignored. The results of these aeroelastic calculations are summarized in this plot of aerodynamic damping versus mass flow rate at a constant rotational speed. As the backpressure is increased, the mass flow rate through the fan decreases and the fan operating point moves towards the stall line. The aeroelastic calculations showed that the aerodynamic damping decreases as the stall line is approached, as observed during testing. In addition, the aeroelastic calculations with the TURBO code correctly predicted the aeroelastic parameters: the most unstable vibration mode and interblade phase angle, as observed during testing. The Quiet High-Speed Fan demonstrated significant noise reductions during testing, but flutter imposed limits on its operating range. The accurate calculation of the aeroelastic characteristics using the TURBO code is a significant step toward eliminating flutter from the operating range and toward realizing the benefits of reduced fan noise. The aeroelastic calculations described here were performed under a grant by University of Toledo researchers in collaboration with Glenn's researchers.

Bakhle, Milind A.↗
NASA NTRS2007

Comparison of Water Vapor Measurements by Airborne Sun Photometer and Near-Coincident in Situ and Satellite Sensors during INTEX/ITCT 2004

We have retrieved columnar water vapor (CWV) from measurements acquired by the 14-channel NASA Ames Airborne Tracking Sun photometer (AATS-14) during 19 Jetstream 31 (J31) flights over the Gulf of Maine in summer 2004 in support of the Intercontinental Chemical Transport Experiment (INTEX)/Intercontinental Transport and Chemical Transformation (ITCT) experiments. In this paper we compare AATS-14 water vapor retrievals during aircraft vertical profiles with measurements by an onboard Vaisala HMP243 humidity sensor and by ship radiosondes and with water vapor profiles retrieved from AIRS measurements during eight Aqua overpasses. We also compare AATS CWV and MODIS infrared CWV retrievals during five Aqua and five Terra overpasses. For 35 J31 vertical profiles, mean (bias) and RMS AATS-minus-Vaisala layer-integrated water vapor (LWV) differences are -7.1 percent and 8.8 percent, respectively. For 22 aircraft profiles within 1 hour and 130 km of radiosonde soundings, AATS-minus-sonde bias and RMS LWV differences are -5.4 percent and 10.7 percent, respectively, and corresponding J31 Vaisala-minus-sonde differences are 2.3 percent and 8.4 percent, respectively. AIRS LWV retrievals within 80 lan of J31 profiles yield lower bias and RMS differences compared to AATS or Vaisala retrievals than do AIRS retrievals within 150 km of the J31. In particular, for AIRS-minus-AATS LWV differences, the bias decreases from 8.8 percent to 5.8 percent, and the RMS difference decreases from 2 1.5 percent to 16.4 percent. Comparison of vertically resolved AIRS water vapor retrievals (LWVA) to AATS values in fixed pressure layers yields biases of -2 percent to +6 percent and RMS differences of -20 percent below 700 hPa. Variability and magnitude of these differences increase significantly above 700 hPa. MODIS IR retrievals of CWV in 205 grid cells (5 x 5 km at nadir) are biased wet by 10.4 percent compared to AATS over-ocean near-surface retrievals. The MODIS-Aqua subset (79 grid cells) exhibits a wet bias of 5.1 percent, and the MODIS-Terra subset (126 grid cells) yields a wet bias of 13.2 percent.

Livingston, J.↗
NASA NTRS1985

Mean Winds and Tides over Urbana, Illinois During November 16-22, 1981

The Urbana Meteor Radar uses a high power pulse transmitter to achieve high echo rates, and an interferometer to spatially locate the individual echoes. The irregularly spaced observations of range, radial velocity, and direction cosines are reduced to a rectangular grid, evenly spaced in time and height, of estimates of the zonal and meridional wind. This is accomplished using the method of GROVES whereby a constant horizontal wind is fitted, using least squares, to all echoes occurring within a fixed time/height interval. This interpolation region is then stepped in time and height. the size of the interpolation region is 2 hours by 5 km, the time step is approximately 11 minutes, and the height step is 2 km. The filtering of high temporal and spatial frequencies is somewhat reduced by weighting the echoes according to their distance from the center of the interpolation box.

Tetenbaum, D.↗