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At least 289 records · Page 16

All sources NASA NTRS DOE OSTI

NASA NTRS1991

The role of tip clearance in high-speed fan stall

A numerical experiment has been carried out to define the near-stall casing endwall flowfield of a high-speed fan rotor. The experiment used a simulation code incorporating a simple clearance model, whose calibration is presented. The results of the simulation show that the interaction of the tip leakage vortex and the in-pasage shock plays a major role in determining the fan flow range. More specifically, the computations imply that it is the area increase of this vortex as it passes through the in-passage shock, which is the source of the blockage associated with stall. In addition, for fans of this type, it is the clearance over the forward portion of the fan blade which controls the flow processes leading to stall.

Adamczyk, J. J.↗

NASA NTRS1991

Deterministic blade row interactions in a centrifugal compressor stage

The three-dimensional viscous flow in a low speed centrifugal compressor stage is simulated using an average passage Navier-Stokes analysis. The impeller discharge flow is of the jet/wake type with low momentum fluid in the shroud-pressure side corner coincident with the tip leakage vortex. This nonuniformity introduces periodic unsteadiness in the vane frame of reference. The effect of such deterministic unsteadiness on the time-mean is included in the analysis through the average passage stress, which allows the analysis of blade row interactions. The magnitude of the divergence of the deterministic unsteady stress is of the order of the divergence of the Reynolds stress over most of the span, from the impeller trailing edge to the vane throat. Although the potential effects on the blade trailing edge from the diffuser vane are small, strong secondary flows generated by the impeller degrade the performance of the diffuser vanes.

Kirtley, K. R.↗

NASA NTRS1991

Three-dimensional analysis of the Pratt and Whitney alternate design SSME fuel turbine

The three dimensional viscous time-mean flow in the Pratt and Whitney alternate design space shuttle main engine fuel turbine is simulated using the average passage Navier-Stokes equations. The migration of secondary flows generated by upstream blade rows and their effect on the performance of downstream blade rows is studied. The present simulation confirms that the flow in this two stage turbine is highly three dimensional and dominated by the tip leakage flow. The tip leakage vortex generated by the first blade persists through the second blade and adversely affects its performance. The greatest mixing of the inlet total temperature distortion occurs in the second vane and is due to the large leakage vortex generated by the upstream rotor. It is assumed that the predominant spanwise mixing mechanism in this low aspect ratio turbine is the radial transport due to the deterministically unsteady vortical flow generated by upstream blade rows. A by-product of the analysis is accurate pressure and heat loads for all blade rows under the influence of neighboring blade rows. These aero loads are useful for advanced structural analysis of the vanes and blades.

Kirtley, K. R.↗

NASA NTRS1992

Aeroloads and secondary flows in a transonic mixed flow turbine stage

A numerical simulation of a transonic mixed flow turbine stage has been carried out using an average passage Navier-Stokes analysis. The mixed flow turbine stage considered here consists of a transonic nozzle vane and a highly loaded rotor. The simulation was run at the design pressure ratio and is assessed by comparing results with those of an established throughflow design system. The three-dimensional aerodynamic loads are studied as well as the development and migration of secondary flows and their contribution to the total pressure loss. The numerical results indicate that strong passage vortices develop in the nozzle vane, mix out quickly, and have little impact on the rotor flow. The rotor is highly loaded near the leading edge. Within the rotor passage, strong spanwise flows and other secondary flows exist along with the tip leakage vortex. The rotor exit loss distribution is similar in character to that found in radial inflow turbines. The secondary flows and non-uniform work extraction also tend to significantly redistribute a non-uniform inlet total temperature profile by the exit of the stage.

Kirtley, K. R.↗

NASA NTRS1994

Generating A Strobed Laser Light Sheet

An optoelectronic system generating synchronous, strobed sheet of laser light developed for use in making visible flow of air about model helicopter rotor. Used in wind-tunnel tests to determine actual locations of vortices for comparison with locations predicted by mathematical models to validate models. Each blade tip produces vortex. By establishing successive vortex locations, researcher determines trajectory of vortex pattern. Light-sheet strobe circuits provide selection of blade positions, strobe-pulse durations, and multiple pulses per revolution for rotors having two to nine blades. To make flow visible, vaporizing propylene glycol injected upstream of model. System also provides calibrated trigger delay of strobe pulses, adjustable strobe-pulse durations, selectable number of blades, and slip-sync mode to make flow visible as though in slow motion.

Leighty, Bradley D.↗

NASA NTRS1994

An Experimental Study of the Flowfield on a Semispan Rectangular Wing with a Simulated Glaze Ice Accretion

Wind tunnel experiments were conducted in order to study the effect of a simulated glaze ice accretion on the flowfield of a semispan, reflection-plane, rectangular wing at Re = 1.5 million and M = 0.12. A laser Doppler velocimeter was used to map the flowfield on the upper surface of the model in both the clean and iced configurations at alpha = 0, 4, and 8 degrees angle of attack. At low angles of attack, the massive separation bubble aft of the leading edge ice horn was found to behave in a manner similar to laminar separation bubbles. At alpha = 0 and 4 degrees, the locations of transition and reattachment, as deduced from momentum thickness distributions, were found to be in good agreement with transition and reattachment locations in laminar separation bubbles. These values at y/b = 0.470, the centerline measurement location, matched well with data obtained on a similar but two dimensional model. The measured velocity profiles on the iced wing compared reasonably with the predicted profiles from Navier-Stokes computations. The iced-induced separation bubble was also found to have features similar to the recirculating region aft of rearward-facing steps. At alpha = 0 degrees and 4 degrees, reverse flow magnitudes and turbulence intensity levels were typical of those found in the recirculating region aft of rearward-facing steps. The calculated separation streamline aft of the ice horn at alpha = 4 degrees, y/b = 0.470 coincided with the locus of the maximum Reynolds normal stress. The maximum Reynolds normal stress peaked at two locations along the separation streamline. The location of the first peak-value coincided with the transition location, as deduced from the momentum thickness distributions. The location of the second peak was just upstream of reattachment, in good agreement with measurements of flows over similar obstacles. The intermittency factor in the vicinity of reattachment at alpha = 4 degrees, y/b = 0.470, revealed the time-dependent nature of the reattachment process. The size and extent of the separation bubble were found to be a function of angle of attack and the spanwise location. Three dimensional effects were found to be strongest at alpha = 8 degrees. The calculated separation and stagnation streamlines were found to vary little with spanwise location at alpha = 0 degrees. The calculated separation streamlines at alpha = 4 degrees revealed that the bubble was largest near the centerline measurement plane, whereas the tip-induced vortex flow and the model root-tunnel wall boundary-layer interaction reduced the size of the bubble. These effects were found to be most dramatic at alpha = 8 degrees.

Khodadoust, Abdollah↗

NASA NTRS2000

Turbulence Measurements and Computations for the Predication of Broadband Noise in High Bypass Ratio Fans

Work was performed under this grant with a view to providing the experimental and computational results needed to improve the prediction of broadband stator noise in large bypass ratio aircraft engines. The central hypothesis of our study was that a large fraction of this noise was generated by the fan tip leakage vortices. More specifically, that these vortices are a significant component of the fan wake turbulence and they contain turbulent eddies of a type that can produce significant broadband noise. To test this hypothesis we originally proposed experimental work and computations with the following objectives: (1) to build a large scale two-dimensional cascade with a tip gap and a stationary endwall that, as far as possible, simulates the fan tip geometry, (2) to build a moving endwall for use with the large scale cascade, (3) to measure, in detail, the turbulence structure and spectrum generated by the blade wake and tip leakage vortex, for both endwall configurations, (4) to use the CFD to compute the flow and turbulence distributions for both the experimental configurations and the ADP fan, (5) to provide the experimental and CFD results for the cascades and the physical understanding gained from their study as a basis for improving the broadband noise prediction method. In large part these objectives have been achieved. The most important achievements and findings of our experimental and computational efforts are summarized below. The bibliography at the end of this report includes a list of all publications produced to date under this project. Note that this list is necessarily incomplete the task of publication (particularly in journal papers) continues.

Devenport, William J.↗

NASA NTRS2002

Contrail Formation in Aircraft Wakes Using Large-Eddy Simulations

In this work we analyze the issue of the formation of condensation trails ("contrails") in the near-field of an aircraft wake. The basic configuration consists in an exhaust engine jet interacting with a wing-tip training vortex. The procedure adopted relies on a mixed Eulerian/Lagrangian two-phase flow approach; a simple micro-physics model for ice growth has been used to couple ice and vapor phases. Large eddy simulations have carried out at a realistic flight Reynolds number to evaluate the effects of turbulent mixing and wake vortex dynamics on ice-growth characteristics and vapor thermodynamic properties.

Paoli, R.↗

NASA NTRS2003

F/A-18 Performance Benefits Measured During the Autonomous Formation Flight Project

The Autonomous Formation Flight (AFF) project at the NASA Dryden Flight Research Center (Edwards, California) investigated performance benefits resulting from formation flight, such as reduced aerodynamic drag and fuel consumption. To obtain data on performance benefits, a trailing F/A-18 airplane flew within the wing tip-shed vortex of a leading F/A-18 airplane. The pilot of the trail airplane used advanced station-keeping technology to aid in positioning the trail airplane at precise locations behind the lead airplane. The specially instrumented trail airplane was able to obtain accurate fuel flow measurements and to calculate engine thrust and vehicle drag. A maneuver technique developed for this test provided a direct comparison of performance values while flying in and out of the vortex. Based on performance within the vortex as a function of changes in vertical, lateral, and longitudinal positioning, these tests explored design-drivers for autonomous stationkeeping control systems. Observations showed significant performance improvements over a large range of trail positions tested. Calculations revealed maximum drag reductions of over 20 percent, and demonstrated maximum reductions in fuel flow of just over 18 percent.

Vachon, M. Jake↗

NASA NTRS1979

Airport Safety Aid

An airplane generates a wake like that of a ship, except that the airplane's wake is invisible. The plane's movement through the air creates-at each wing tip-a "vortex," a turbulent wind moving in a circular pattern like an air whirlpool. Large aircraft, such as commercial jetliners, produce powerful vortices which can be hazardous to small planes following closely be hind. For that reason, the Federal Aviation Administration requires a five mile spacing between large and small aircraft approaching a runway, to allow time for the vortices to dissipate. At busy terminals, this spacing requirement restricts landing operations and contributes to airport congestion. With an eye toward reducing the spacing requirement while assuring lightplane safety, the Department of Transportation is conducting research on the characteristics of aircraft wake. The spinoff system pictured, atop the checkerboard van in the lower photo and in close-up at upper right, is producing valuable information which will enable airport controllers to determine when it is safe for a lightplane to land. Developed by Lockheed Missiles and Space Company, Sunnyvale, California, it is called a Laser Doppler Velocimeter, or LDV.

Source record↗

NASA NTRS2008

Pre-Stall Behavior of a Transonic Axial Compressor Stage via Time-Accurate Numerical Simulation

CFD calculations using high-performance parallel computing were conducted to simulate the pre-stall flow of a transonic compressor stage, NASA compressor Stage 35. The simulations were run with a full-annulus grid that models the 3D, viscous, unsteady blade row interaction without the need for an artificial inlet distortion to induce stall. The simulation demonstrates the development of the rotating stall from the growth of instabilities. Pressure-rise performance and pressure traces are compared with published experimental data before the study of flow evolution prior to the rotating stall. Spatial FFT analysis of the flow indicates a rotating long-length disturbance of one rotor circumference, which is followed by a spike-type breakdown. The analysis also links the long-length wave disturbance with the initiation of the spike inception. The spike instabilities occur when the trajectory of the tip clearance flow becomes perpendicular to the axial direction. When approaching stall, the passage shock changes from a single oblique shock to a dual-shock, which distorts the perpendicular trajectory of the tip clearance vortex but shows no evidence of flow separation that may contribute to stall.

Chen, Jen-Ping↗

NASA NTRS2009

Large Eddy Simulation of Transonic Flow Field in NASA Rotor 37

The current paper reports on numerical investigations on the flow characteristics in a transonic axial compressor, NASA Rotor 37. The flow field was used previously as a CFD blind test case conducted by American Society of Mechanical Engineers in 1994. Since the CFD blind-test exercise, many numerical studies on the flow field in the NASA Rotor 37 have been reported. Although steady improvements have been reported in both numerical procedure and turbulence closure, it is believed that all the important aspects of the flow field have not been fully explained with numerical studies based on the Reynolds Averaged Navier-Stokes (RANS) solution. Experimental data show large dip in total pressure distribution near the hub at downstream of the rotor at 100% rotor speed. Most original numerical solutions from the blind test exercise did not predict this total pressure deficit correctly. This total pressure deficit at the rotor exit was attributed to a hub corner flow separation by the author. Several subsequent numerical studies with different turbulence closure model also calculated this dip in total pressure rise. Also, several studies attributed this total pressure deficit to a small leakage flow coming from the hub in the test article. As the experimental study cannot be repeated, either explanation cannot be validated. The primary purpose of the current investigation is to investigate the transonic flow field with both RANS and a Large Eddy Simulation (LES). The RANS approach gives similar results presented at the original blind test exercise. Although the RANS calculates higher overall total pressure rise, the total pressure deficit near the hub is calculated correctly. The numerical solution shows that the total pressure deficit is due to a hub corner flow separation. The calculated pressure rise from the LES agrees better with the measured total pressure rise especially near the casing area where the passage shock interacts with the tip clearance vortex and flow becomes unsteady due to this interaction. The LES simulation also calculates the total pressure rise deficit near the hub and it agrees well with the measured data.

Hah, Chunill↗

NASA NTRS2011

Steady and Unsteady Flow Effects of Circumferential Grooves Casing Treatment in a Transonic Compressor Rotor

The current paper reports on an investigation of steady and unsteady flow effects of circumferential grooves casing treatment in a transonic compressor rotor. Circumferential grooves casing treatment is used mainly to increase stall margin in axial compressors with a relatively small decrease in aerodynamic efficiency. It is widely believed that flow mechanisms of circumferential grooves casing treatment near stall conditions are not yet well understood even though this treatment has been used widely in real engines. Numerical analysis based on steady Reynolds-averaged Navier-Stokes (RANS) has been the primary tool used to understand flow mechanism for circumferential grooves casing treatment. Although steady RANS explains some flow effects of circumferential grooves casing treatment, it does not calculate all the measured changes in the compressor characteristics. Therefore, design optimization of circumferential grooves with steady RANS has not been very successful. As a compressor operates toward the stall condition, the flow field becomes transient. Major sources of self-generated flow unsteadiness are shock oscillation and interaction between the passage shock and the tip leakage vortex. In the present paper, an unsteady Reynolds-averaged Navier-Stokes (URANS) approach is applied to study the effects of circumferential grooves in a transonic compressor. The results from URANS are compared with the results from RANS and measured data. The current investigation shows that there are significant unsteady flow effects on the performance of the circumferential grooves casing treatment. For the currently investigated rotor, the unsteady effects are of the same magnitude as the steady effects in terms of extending the compressor stall margin.

Hah, Chunill↗

NASA NTRS1985

Wide-field shadowgraph flow visualization of tip vortices generated by a helicopter rotor

The vortex trajectory and vortex wake generated by helicopter rotors were visualized using a wide-field shadowgraph technique. Use of a retro-reflective Scotchlite screen made it possible to investigate the flow field generated by full-scale tail rotors. Tip vortex trajectories were visible in shadowgraphs for a range for tip Mach number of 0.38 - 0.60. The effect of the angle of attack was substantial. At an angle of attack greater than 8 degrees, the visibility of the vortex core was significant even at relatively low tip Mach numbers. The theoretical analysis of the visibility was carried out for a rotating blade. This analysis demonstrated that the visibility decreases with increasing dimensionless core radius (ro/c) and increases with increasing tip Mach number. Based on this investigation, it is concluded that the wide-field shadowgraph flow visualization technique should be feasible to study the flow field generated by a large main rotor in a wind tunnel and in an outdoor full-scale test stand. Of note is that the shadowgraph technique could easily be used with other on-going tests such as aerodynamic performance study, noise measurements, velocity measurements with LDV or hot-wire anemometer, local pressure measurement on the rotor surface, etc.

Parthasarathy, S. P.↗

NASA NTRS1975

Design and preliminary tests of a blade tip air mass injection system for vortex modification and possible noise reduction on a full-scale helicopter rotor

Full-scale tests were conducted on the Langley helicopter rotor test facility as part of a study to evaluate the effectiveness of a turbulent blade tip air mass injection system in alleviating the impulsive noise (blade slap) caused by blade-vortex interaction. Although blade-slap conditions could not be induced during these tests, qualitative results from flow visualization studies using smoke showed that the differential velocity between the jet velocity and the rotor tip speed was a primary parameter controlling the vortex modification.

Pegg, R. J.↗

NASA NTRS1976

Subsonic loads on wings having sharp leading edges and tips

A vortex-lattice method for predicting the aerodynamics of wings having separation at the sharp edges in incompressible flows is extended to compressible subsonic flows using a modified Prandtl-Glauert transformation. Numerical results showing the effect of freestream Mach number on the aerodynamic coefficients are compared with available experimental data for several planforms. It is shown that the proposed method is suitable for predicting the aerodynamic loads on low-aspect wings at moderate angles of attack for high subsonic freestream Mach number. The method is limited to angles of attack up to 12 deg for high subsonic freestream Mach number and to angles of attack up to 20 deg for Mach number not exceeding 0.5.

Kandil, O. A.↗

NASA NTRS1985

Evolution of a vortical structure associated with the bursting event in a channel flow

The temporal evolution of a horseshoe vortex structure in a channel is investigated by a numerical simulation. A spectral numerical method is employed to integrate the time-dependent, three-dimensional Navier-Stokes equations. The initial vortical structure is obtained by applying a conditional sampling technique to a data base generated from a direct simulation of a turbulent channel flow. The evolution of this vortical structure under the influence of the self-induced motion and the mean shear is presented. It is shown that the initial sheet-like vortical structure rolls up into a vortex tube as it is convected downstream. Turbulence structures associated with the vortex stretching are investigated. Production of vorticity stretching is high inside the vortex legs, although it is also substantial in the tip region and above the legs. High Reynolds shear stress is produced near the tip of the vortex.

Kim, J.↗

NASA NTRS1994

Numerical study of the trailing vortex of a wing with wing-tip blowing

Trailing vortices generated by lifting surfaces such as helicopter rotor blades, ship propellers, fixed wings, and canard control surfaces are known to be the source of noise, vibration, cavitation, degradation of performance, and other hazardous problems. Controlling these vortices is, therefore, of practical interest. The formation and behavior of the trailing vortices are studied in the present research. In addition, wing-tip blowing concepts employing axial blowing and spanwise blowing are studied to determine their effectiveness in controlling these vortices and their effects on the performance of the wing. The 3D, unsteady, thin-layer compressible Navier-Stokes equations are solved using a time-accurate, implicit, finite difference scheme that employs LU-ADI factorization. The wing-tip blowing is simulated using the actuator plane concept, thereby, not requiring resolution of the jet slot geometry. Furthermore, the solution blanking feature of the chimera scheme is used to simplify the parametric study procedure for the wing-tip blowing. Computed results are shown to compare favorably with experimental measurements. It is found that axial wing-tip blowing, although delaying the rolling-up of the trailing vortices and the near-field behavior of the flowfield, does not dissipate the circulation strength of the trailing vortex farther downstream. Spanwise wing-tip blowing has the effect of displacing the trailing vortices outboard and upward. The increased 'wing-span' due to the spanwise wing-tip blowing has the effect of lift augmentation on the wing and the strengthening of the trailing vortices. Secondary trailing vortices are created at high spanwise wing-tip blowing intensities.

Lim, Hock-Bin↗