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At least 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS1995

Fractional Step and Pseudocompressibility Methods for the Solution of Incompressible Navier-Stokes Equations

The fractional step and the pseudocompressibility methods for the solution of the incompressible Navier-Stokes equations are outlined. The fractional step method is based on finite-volume formulation and uses the pressure and the volume fluxes across the faces of each cell as dependent variables. The momentum equations are solved implicitly and the Poisson equation for the pressure is solved by using the multigrid method. The pseudocompressibility approach uses an implicit-higher-order-upwind differencing scheme for the convective terms together with the Gauss-Seidel line relaxation method. The dependent variables in the pseudocompressibility approach are the pressure and the cartesian velocity components in unstaggered mesh orientation. The 90-degree square duct flow, the wing-tip vortex wake flow and unsteady turbulent flows over an oscillating NACA 0015 airfoil are computed using both the fractional step and the pseudocompressibility methods. The results obtained from two different schemes are compared against experimental measurements.

Kiris, Cetin↗
NASA NTRS1995

Computational Flow Analysis of a Left Ventricular Assist Device

Computational fluid dynamics has been developed to a level where it has become an Indispensable part of aerospace research and design. Technology developed foe aerospace applications am also be utilized for the benefit of human health. For example, a flange-to-flange rocket engine fuel-pump simulation includes the rotating and non-rotating components: the flow straighteners, the impeller, and diffusers A Ventricular Assist Device developed by NASA Johnson Space Center and Baylor College of Medicine has a design similar to a rocket engine fuel pump in that it also consists of a flow straightener, an impeller, and a diffuser. Accurate and detailed knowledge of the flowfield obtained by incompressible flow calculations can be greatly beneficial to designers in their effort to reduce the cost and improve the reliability of these devices. In addition to the geometric complexities, a variety of flow phenomena are encountered in biofluids Then include turbulent boundary layer separation, wakes, transition, tip vortex resolution, three-dimensional effects, and Reynolds number effects. In order to increase the role of Computational Fluid Dynamics (CFD) in the design process the CFD analysis tools must be evaluated and validated so that designers gain Confidence in their use. The incompressible flow solver, INS3D, has been applied to flow inside of a liquid rocket engine turbopump components and extensively validated. This paper details how the computational flow simulation capability developed for liquid rocket engine pump component analysis has bean applied to the Left Ventricular Assist Device being developed jointly by NASA JSC and Baylor College of Medicine.

Kiris, Cetin↗
NASA NTRS2001

Full-Span Tiltrotor Aeroacoustic Model (TRAM) Overview and 40- by 80-Foot Wind Tunnel Test

Most helicopter data trends cannot be extrapolated to tiltrotors because blade geometry and aerodynamic behavior, as well as rotor and fuselage interactions, are significantly different for tiltrotors. A tiltrotor model has been developed to investigate the aeromechanics of tiltrotors, to develop a comprehensive database for validating tiltrotor analyses, and to provide a research platform for supporting future tiltrotor designs. The Full-Span Tiltrotor Aeroacoustic Model (FS TRAM) is a dual-rotor, powered aircraft model with extensive instrumentation for measurement of structural and aerodynamic loads. This paper will present the Full-Span TRAM test capabilities and the first set of data obtained during a 40- by 80-Foot Wind Tunnel test conducted in late 2000 at NASA Ames Research Center. The Full-Span TRAM is a quarter-scale representation of the V-22 Osprey aircraft, and a heavily instrumented NASA and U.S. Army wind tunnel test stand. Rotor structural loads are monitored and recorded for safety-of-flight and for information on blade loads and dynamics. Left and right rotor balance and fuselage balance loads are monitored for safety-of-flight and for measurement of vehicle and rotor aerodynamic performance. Static pressure taps on the left wing are used to determine rotor/wing interactional effects and rotor blade dynamic pressures measure blade airloads. All of these measurement capabilities make the FS TRAM test stand a unique and valuable asset for validation of computational codes and to aid in future tiltrotor designs. The Full-Span TRAM was tested in the NASA Ames Research Center 40- by 80-Foot Wind Tunnel from October through December 2000. Rotor and vehicle performance measurements were acquired in addition to wing pressures, rotor acoustics, and Laser Light Sheet (LLS) flow visualization data. Hover, forward flight, and airframe (rotors off) aerodynamic runs were performed. Helicopter-mode data were acquired during angle of attack and thrust sweeps for a variety of tunnel speeds. Wake geometry images were acquired using LLS photographs and suggest dual tip vortex formation at low thrust conditions. The full paper will include comparisons to isolated-rotor TRAM data acquired at the Duits-Nederlandse Windtunnel (DNW) in 1998. The FS TRAM has been established as a valuable national asset for tiltrotor research. Data reduction and analysis of the 40- by 80-Foot Wind Tunnel test results are underway. Follow-on testing of the FS TRAM is currently being planned for the NASA Ames 80- by 120-Foot Wind Tunnel in late 2001.

McCluer, Megan S.↗
NASA NTRS1997

Incompressible Navier-Stokes Computations in Aerospace and Beyond

Challenges in applying CFD to real world problems will be discussed using incompressible flow examples. Despite the remarkable progress made in CFD technology during the past two decades, new challenges are ahead of us. Some of these will be discussed in conjunction with the incompressible Navier-Stokes solver development and their applications. Computed examples include wing tip vortex formation and propagation, flow simulation of an advanced rocket pump, and an extension of the similar technology to biofluid analysis and design. Numerical issues dealing with these applications will be discussed both from algorithm development and from application point of view. The material will be printed in a bound volume after the symposium.

Kwak, Dochan↗
NASA NTRS1998

Numerical Experiment with Time and Spatial Accuracy of Navier-Stokes Computation For Helicopter Problems

Helicopter flowfields are highly unsteady, nonlinear and three-dimensional. In forward flight and in hover, the rotor blades interact with the tip vortex and wake sheet developed by either itself or the other blades. This interaction, known as blade-vortex interactions (BVI), results in unsteady loading of the blades and can cause a distinctive acoustic signature. Accurate and cost-effective computational fluid dynamic solutions that capture blade-vortex interactions can help rotor designers and engineers to predict rotor performance and to develop designs for low acoustic signature. Such a predictive method must preserve a blade's shed vortex for several blade revolutions before being dissipated. A number of researchers have explored the requirements for this task. This paper will outline some new capabilities that have been added to the NASA Ames' OVERFLOW code to improve its overall accuracy for both vortex capturing and unsteady flows. To highlight these improvements, a number of case studies will be presented. These case studies consist of free convection of a 2-dimensional vortex, dynamically pitching 2-D airfoil including light-stall, and a full 3-D unsteady viscous solution of a helicopter rotor in forward flight In this study both central and upwind difference schemes are modified to be more accurate. Central difference scheme is chosen for this simulation because the flowfield is not dominated by strong shocks. The feature of shock-vortex interaction in such a flow is less important than the dominant blade-vortex interaction. The scheme is second-order accurate in time and solves the thin-layer Navier-Stokes equations in fully-implicit manner at each time-step. The spatial accuracy is either second and fourth-order central difference or third-order upwind difference using Roe-flux and MUSCLE scheme. This paper will highlight and demonstrate the methods for several sample cases and for a helicopter rotor. Preliminary computations on a rotor were performed by using this method and are in the process of documentation.

Ahmad, Jasim↗
NASA NTRS2002

Centrifugal Compressor Aeroelastic Analysis Code

Centrifugal compressors are very widely used in the turbomachine industry where low mass flow rates are required. Gas turbine engines for tanks, rotorcraft and small jets rely extensively on centrifugal compressors for rugged and compact design. These compressors experience problems related with unsteadiness of flowfields, such as stall flutter, separation at the trailing edge over diffuser guide vanes, tip vortex unsteadiness, etc., leading to rotating stall and surge. Considerable interest exists in small gas turbine engine manufacturers to understand and eventually eliminate the problems related to centrifugal compressors. The geometric complexity of centrifugal compressor blades and the twisting of the blade passages makes the linear methods inapplicable. Advanced computational fluid dynamics (CFD) methods are needed for accurate unsteady aerodynamic and aeroelastic analysis of centrifugal compressors. Most of the current day industrial turbomachines and small aircraft engines are designed with a centrifugal compressor. With such a large customer base and NASA Glenn Research Center being, the lead center for turbomachines, it is important that adequate emphasis be placed on this area as well. Currently, this activity is not supported under any project at NASA Glenn.

Keith, Theo G., Jr.↗
NASA NTRS2009

Development of a Large Field-of-View PIV System for Rotorcraft Testing in the 14- x 22-Foot Subsonic Tunnel

A Large Field-of-View Particle Image Velocimetry (LFPIV) system has been developed for rotor wake diagnostics in the 14-by 22-Foot Subsonic Tunnel. The system has been used to measure three components of velocity in a plane as large as 1.524 meters by 0.914 meters in both forward flight and hover tests. Overall, the system performance has exceeded design expectations in terms of accuracy and efficiency. Measurements synchronized with the rotor position during forward flight and hover tests have shown that the system is able to capture the complex interaction of the body and rotor wakes as well as basic details of the blade tip vortex at several wake ages. Measurements obtained with traditional techniques such as multi-hole pressure probes, Laser Doppler Velocimetry (LDV), and 2D Particle Image Velocimetry (PIV) show good agreement with LFPIV measurements.

Jenkins, Luther N.↗
NASA NTRS2011

FUN3D and CFL3D Computations for the First High Lift Prediction Workshop

Two Reynolds-averaged Navier-Stokes codes were used to compute flow over the NASA Trapezoidal Wing at high lift conditions for the 1st AIAA CFD High Lift Prediction Workshop, held in Chicago in June 2010. The unstructured-grid code FUN3D and the structured-grid code CFL3D were applied to several different grid systems. The effects of code, grid system, turbulence model, viscous term treatment, and brackets were studied. The SST model on this configuration predicted lower lift than the Spalart-Allmaras model at high angles of attack; the Spalart-Allmaras model agreed better with experiment. Neglecting viscous cross-derivative terms caused poorer prediction in the wing tip vortex region. Output-based grid adaptation was applied to the unstructured-grid solutions. The adapted grids better resolved wake structures and reduced flap flow separation, which was also observed in uniform grid refinement studies. Limitations of the adaptation method as well as areas for future improvement were identified.

Park, Michael A.↗
NASA NTRS2011

Characterization of Axial Inducer Cavitation Instabilities via High Speed Video Recordings

Sub-scale water tests were undertaken to assess the viability of utilizing high resolution, high frame-rate digital video recordings of a liquid rocket engine turbopump axial inducer to characterize cavitation instabilities. These high speed video (HSV) images of various cavitation phenomena, including higher order cavitation, rotating cavitation, alternating blade cavitation, and asymmetric cavitation, as well as non-cavitating flows for comparison, were recorded from various orientations through an acrylic tunnel using one and two cameras at digital recording rates ranging from 6,000 to 15,700 frames per second. The physical characteristics of these cavitation forms, including the mechanisms that define the cavitation frequency, were identified. Additionally, these images showed how the cavitation forms changed and transitioned from one type (tip vortex) to another (sheet cavitation) as the inducer boundary conditions (inlet pressures) were changed. Image processing techniques were developed which tracked the formation and collapse of cavitating fluid in a specified target area, both in the temporal and frequency domains, in order to characterize the cavitation instability frequency. The accuracy of the analysis techniques was found to be very dependent on target size for higher order cavitation, but much less so for the other phenomena. Tunnel-mounted piezoelectric, dynamic pressure transducers were present throughout these tests and were used as references in correlating the results obtained by image processing. Results showed good agreement between image processing and dynamic pressure spectral data. The test set-up, test program, and test results including H-Q and suction performance, dynamic environment and cavitation characterization, and image processing techniques and results will be discussed.

Arellano, Patrick↗
NASA NTRS2015

A Comparative Study of Simulated and Measured Gear-Flap Flow Interaction

The ability of two CFD solvers to accurately characterize the transient, complex, interacting flowfield asso-ciated with a realistic gear-flap configuration is assessed via comparison of simulated flow with experimental measurements. The simulated results, obtained with NASA's FUN3D and Exa's PowerFLOW® for a high-fidelity, 18% scale semi-span model of a Gulfstream aircraft in landing configuration (39 deg flap deflection, main landing gear on and off) are compared to two-dimensional and stereo particle image velocimetry measurements taken within the gear-flap flow interaction region during wind tunnel tests of the model. As part of the bench-marking process, direct comparisons of the mean and fluctuating velocity fields are presented in the form of planar contour plots and extracted line profiles at measurement planes in various orientations stationed in the main gear wake. The measurement planes in the vicinity of the flap side edge and downstream of the flap trailing edge are used to highlight the effects of gear presence on tip vortex development and the ability of the computational tools to accurately capture such effects. The present study indicates that both computed datasets contain enough detail to construct a relatively accurate depiction of gear-flap flow interaction. Such a finding increases confidence in using the simulated volumetric flow solutions to examine the behavior of pertinent aer-odynamic mechanisms within the gear-flap interaction zone.

Khorrami, Mehdi R.↗
NASA NTRS2016

Effect of AFT Rotor on the Inter-Rotor Flow of an Open Rotor Propulsion System

The effects of the aft rotor on the inter-rotor flow field of an open rotor propulsion rig were examined. A Particle Image Velocimetry (PIV) dataset that was acquired phase locked to the front rotor position has been phase averaged based on the relative phase angle between the forward and aft rotors. The aft rotor phase was determined by feature tracking in raw PIV images through an image processing algorithm. The effect of the aft rotor potential field on the inter-rotor flow were analyzed and shown to be in good agreement with Computational Fluid Dynamics (CFD) simulations. It was shown that the aft rotor had no substantial effect on the position of the forward rotor tip vortex but did have a small effect on the circulation strength of the vortex when the rotors were highly loaded.

Counter Rotating↗
NASA NTRS2016

Effect of Aft Rotor on the Inter-Rotor Flow of an Open Rotor Propulsion System

The effects of the aft rotor on the inter-rotor flow field of an open rotor propulsion rig were examined. A Particle Image Velocimetry (PIV) dataset that was acquired phase locked to the front rotor position has been phase averaged based on the relative phase angle between the forward and aft rotors. The aft rotor phase was determined by feature tracking in raw PIV images through an image processing algorithm. The effects of the aft rotor potential field on the inter-rotor flow were analyzed and shown to be in reasonably good agreement with Computational Fluid Dynamics (CFD) simulations. The aft rotor position was shown to have a significant upstream effect, with implications for front rotor interaction noise. It was found that the aft rotor had no substantial effect on the position of the forward rotor tip vortex but did have a small effect on the circulation strength of the vortex when the rotors were highly loaded.

Counter Rotating↗
NASA NTRS2021

Computational Fluid Dynamics Simulations of the Transonic Dynamics Tunnel Airstream Oscillator System

This paper presents computational fluid dynamics simulations of the flow in the NASA Langley Research Center Transonic Dynamics Tunnel (TDT) with the airstream oscillator system (AOS) in operation. Test section flow angle can be modulated using sinusoidally oscillating vanes which are located in pairs within the contraction section of the tunnel circuit. A compilation of previously published experimental data for validation has been reviewed that contains hot-wire anemometer and fast-response probe data acquired in an empty test section with the vanes oscillating in-phase and out-of-phase. These experimental data have not previously been compared with computational results. The present results provide an opportunity for validation of the simulations of the sinusoidal AOS flow field in the TDT test section. Computed tunnel centerline, lateral and cross-section contour data of flow angle and vorticity show that resolving the vane tip vortex is critical to accurately predicting the test section center flow angle. The computed results show, as do the experimental data, what appears to be a resonance at higher frequencies. It is currently not known what is the cause of that resonance. Further refinement of the mesh is expected to improve the comparisons with experimental data

Robert E Bartels↗
NASA NTRSDate not supplied

Quantitative Approach for the Accurate CFD Simulation of Hover in Turbulent Flow

Time-dependent Navier-Stokes simulations have been carried out for a V22 rotor in hover using an improved, low-diffusion, HLLE++ upwind algorithm in the OVERFLOW code. Emphasis is placed on lessons learned over the past decade regarding the effects of high-order spatial accuracy, grid resolution, and the use of detached eddy simulation in predicting the rotor performance, i.e., figure-of-merit (FM). A quick-start procedure is described together with a statistical measure of FM convergence that reduces hover computations by 5 times, similar in computational work for forward flight. Furthermore, Cartesian adaptive mesh refinement in the rotor wake revealed a complex turbulent flow with worm-like structures of various scales. These turbulent worms, found computationally more than a decade ago, have recently been observed in a separate German Aerospace Center (DLR) experiment. Moreover, adaptive mesh refinement has been used to resolve the tip-vortex to its correct physical size.

Neal M. Chaderjian↗
NASA NTRS1971

Qualitative responses of a vortex core to tip blowing and intersecting airfoils

The qualitative responses of a trailing vortex core to change in its convective velocity produced by injection of air into the vortex core and by the flow field of a lifting surface in the path of the core are discussed. Flow pictures of the vortex core and vorticity measurements in the core show that an axial momentum injection of approximately 35% of the total wing drag alters the vortex structure and behavior quite drastically by effectively inducing a more rapid dispersion of the vorticity in the core. The data indicate that the phenomenon is governed by the rate of injection of the momentum rather than the mass flow rate. There also seems to be an optimum rate of injection beyond which increased injection does not bring about equally significant changes in the vortex core. Flow patterns of the region of interaction between the flow field of a lifting surface and the vortex core show two modes of vortex response; (1) the vortex core bends, following the streamline shape until it intercepts the wake of the wing where it is abruptly dispersed, or (2) the vortex core may be sliced into two similar vortices when the vortex hits the leading edge of the wing.

Kantha, H. L.↗
NASA NTRS1975

Wing-tip vanes as vortex attenuation and induced drag reduction devices

Analytical studies have been conducted to examine the feasibility of utilizing wing tip turbines to remove swirl from the wing trailing vortex, and hence reduce the potential for upset of following aircraft. Energy recovery from the turbines is also analyzed. A computer routine has been developed to permit rapid parametric studies of various tip turbine designs. It is shown that the optimum turbine is a non-rotating set of vanes which reduce swirl and recover energy in the form of reduced overall configuration induced drag. A specific case study indicates a 23% reduction in induced drag for a rectangular wing of aspect ratio 5.33, operated at a lift coefficient at 1.0.

Wentz, W. H., Jr.↗
NASA NTRS2015

Investigation of Unsteady Tip Clearance Flow in a Low-Speed One and Half Stage Axial Compressor with LES And PIV

The primary focus of this paper is to investigate how a rotor's unsteady tip clearance flow structure changes in a low speed one and half stage axial compressor when the rotor tip gap size is increased from 0.5 mm (0.49% of rotor tip blade chord, 2% of blade span) to 2.4 mm (2.34% chord, 4% span) at the design condition are investigated. The changes in unsteady tip clearance flow with the 0.62 % tip gap as the flow rate is reduced to near stall condition are also investigated. A Large Eddy Simulation (LES) is applied to calculate the unsteady flow field at these three flow conditions. Detailed Stereoscopic PIV (SPIV) measurements of the current flow fields were also performed at the Johns Hopkins University in a refractive index-matched test facility which renders the compressor blades and casing optically transparent. With this setup, the unsteady velocity field in the entire flow domain, including the flow inside the tip gap, can be measured. Unsteady tip clearance flow fields from LES are compared with the PIV measurements and both LES and PIV results are used to study changes in tip clearance flow structures. The current study shows that the tip clearance vortex is not a single structure as traditionally perceived. The tip clearance vortex is formed by multiple interlaced vorticities. Therefore, the tip clearance vortex is inherently unsteady. The multiple interlaced vortices never roll up to form a single structure. When phased-averaged, the tip clearance vortex appears as a single structure. When flow rate is reduced with the same tip gap, the tip clearance vortex rolls further upstream and the tip clearance vortex moves further radially inward and away from the suction side of the blade. When the tip gap size is increased at the design flow condition, the overall tip clearance vortex becomes stronger and it stays closer to the blade suction side and the vortex core extends all the way to the exit of the blade passage. Measured and calculated unsteady flow fields inside the tip gap agree fairly well. Instantaneous velocity vectors inside the tip gap from both the PIV and LES do show flow separation and reattachment at the entrance of tip gap as some earlier studies suggested. This area at the entrance of tip gap flow (the pressure side of the blade) is confined very close to the rotor tip section. With a small tip gap (0.5mm), the gap flow looks like a simple two-dimensional channel flow with larger velocity near the casing for both flow rates. A small area with a sharp velocity gradient is observed just above the rotor tip. This strong shear layer is turned radially inward when it collides with the incoming flow and forms the core structure of the tip clearance vortex. When tip gap size is increased to 2.4 mm at the design operation, the radial profile of the tip gap flow changes drastically. With the large tip gap, the gap flow looks like a two-dimensional channel flow only near the casing. Near the rotor top section, a bigger region with very large shear and reversed flow is observed.

compressor↗
NASA NTRS2015

Investigation of Unsteady Tip Clearance Flow in a Low-Speed One and Half Stage Axial Compressor with LES And PIV

The primary focus of this paper is to investigate how a rotor's unsteady tip clearance flow structure changes in a low speed one and half stage axial compressor when the rotor tip gap size is increased from 0.5 mm (0.49% of rotor tip blade chord, 2% of blade span) to 2.4 mm (2.34% chord, 4% span) at the design condition are investigated. The changes in unsteady tip clearance flow with the 0.62 % tip gap as the flow rate is reduced to near stall condition are also investigated. A Large Eddy Simulation (LES) is applied to calculate the unsteady flow field at these three flow conditions. Detailed Stereoscopic PIV (SPIV) measurements of the current flow fields were also performed at the Johns Hopkins University in a refractive index-matched test facility which renders the compressor blades and casing optically transparent. With this setup, the unsteady velocity field in the entire flow domain, including the flow inside the tip gap, can be measured. Unsteady tip clearance flow fields from LES are compared with the PIV measurements and both LES and PIV results are used to study changes in tip clearance flow structures. The current study shows that the tip clearance vortex is not a single structure as traditionally perceived. The tip clearance vortex is formed by multiple interlaced vorticities. Therefore, the tip clearance vortex is inherently unsteady. The multiple interlaced vortices never roll up to form a single structure. When phased-averaged, the tip clearance vortex appears as a single structure. When flow rate is reduced with the same tip gap, the tip clearance vortex rolls further upstream and the tip clearance vortex moves further radially inward and away from the suction side of the blade. When the tip gap size is increased at the design flow condition, the overall tip clearance vortex becomes stronger and it stays closer to the blade suction side and the vortex core extends all the way to the exit of the blade passage. Measured and calculated unsteady flow fields inside the tip gap agree fairly well. Instantaneous velocity vectors inside the tip gap from both the PIV and LES do show flow separation and reattachment at the entrance of tip gap as some earlier studies suggested. This area at the entrance of tip gap flow (the pressure side of the blade) is confined very close to the rotor tip section. With a small tip gap (0.5mm), the gap flow looks like a simple two-dimensional channel flow with larger velocity near the casing for both flow rates. A small area with a sharp velocity gradient is observed just above the rotor tip. This strong shear layer is turned radially inward when it collides with the incoming flow and forms the core structure of the tip clearance vortex. When tip gap size is increased to 2.4 mm at the design operation, the radial profile of the tip gap flow changes drastically. With the large tip gap, the gap flow looks like a two-dimensional channel flow only near the casing. Near the rotor top section, a bigger region with very large shear and reversed flow is observed.

unsteady↗