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At least 343 records · Page 19

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NASA NTRS1986

Computation of turbulent rotating channel flow with an algebraic Reynolds stress model

An Algebraic Reynolds Stress Model has been implemented to modify the Kolmogorov-Prandtl eddy viscosity relation to produce an anisotropic turbulence model. The eddy viscosity relation becomes a function of the local turbulent production to dissipation ratio and local turbulence/rotation parameters. The model is used to predict fully-developed rotating channel flow over a diverse range of rotation numbers. In addition, predictions are obtained for a developing channel flow with high rotation. The predictions are compared with the experimental data available. Good predictions are achieved for mean velocity and wall shear stress over most of the rotation speeds tested. There is some prediction breakdown at high rotation (rotation number greater than .10) where the effects of the rotation on turbulence become quite complex. At high rotation and low Reynolds number, the laminarization on the trailing side represents a complex effect of rotation which is difficult to predict with the described models.

Warfield, M. J.↗

NASA NTRS1990

A perspective on coherent structures and conceptual models for turbulent boundary layer physics

Direct numerical simulations of turbulent boundary layers have been analyzed to develop a unified conceptual model for the kinematics of coherent motions in low Reynolds number canonical turbulent boundary layers. All classes of coherent motions are considered in the model, including low-speed streaks, ejections and sweeps, vortical structures, near-wall and outer-region shear layers, sublayer pockets, and large-scale outer-region eddies. The model reflects the conclusions from the study of the simulated boundary layer that vortical structures are directly associated with the production of turbulent shear stresses, entrainment, dissipation of turbulence kinetic energy, and the fluctuating pressure field. These results, when viewed from the perspective of the large body of published work on the subject of coherent motions, confirm that vortical structures may be considered the central dynamic element in the maintenance of turbulence in the canonical boundary layer. Vortical structures serve as a framework on which to construct a unified picture of boundary layer structure, providing a means to relate the many known structural elements in a consistent way.

Robinson, Stephen K.↗

NASA NTRS1991

A unified viscous theory of lift and drag of 2-D thin airfoils and 3-D thin wings

A unified viscous theory of 2-D thin airfoils and 3-D thin wings is developed with numerical examples. The viscous theory of the load distribution is unique and tends to the classical inviscid result with Kutta condition in the high Reynolds number limit. A new theory of 2-D section induced drag is introduced with specific applications to three cases of interest: (1) constant angle of attack; (2) parabolic camber; and (3) a flapped airfoil. The first case is also extended to a profiled leading edge foil. The well-known drag due to absence of leading edge suction is derived from the viscous theory. It is independent of Reynolds number for zero thickness and varies inversely with the square root of the Reynolds number based on the leading edge radius for profiled sections. The role of turbulence in the section induced drag problem is discussed. A theory of minimum section induced drag is derived and applied. For low Reynolds number the minimum drag load tends to the constant angle of attack solution and for high Reynolds number to an approximation of the parabolic camber solution. The parabolic camber section induced drag is about 4 percent greater than the ideal minimum at high Reynolds number. Two new concepts, the viscous induced drag angle and the viscous induced separation potential are introduced. The separation potential is calculated for three 2-D cases and for a 3-D rectangular wing. The potential is calculated with input from a standard doublet lattice wing code without recourse to any boundary layer calculations. Separation is indicated in regions where it is observed experimentally. The classical induced drag is recovered in the 3-D high Reynolds number limit with an additional contribution that is Reynold number dependent. The 3-D viscous theory of minimum induced drag yields an equation for the optimal spanwise and chordwise load distribution. The design of optimal wing tip planforms and camber distributions is possible with the viscous 3-D wing theory.

Yates, John E.↗

NASA NTRS1996

Application of an Upwind High Resolution Finite-Differencing Scheme and Multigrid Method in Steady-State Incompressible Flow Simulations

The analysis and design of a submarine propulsor requires the ability to predict the characteristics of both laminar and turbulent flows to a higher degree of accuracy. This report presents results of certain benchmark computations based on an upwind, high-resolution, finite-differencing Navier-Stokes solver. The purpose of the computations is to evaluate the ability, the accuracy and the performance of the solver in the simulation of detailed features of viscous flows. Features of interest include flow separation and reattachment, surface pressure and skin friction distributions. Those features are particularly relevant to the propulsor analysis. Test cases with a wide range of Reynolds numbers are selected; therefore, the effects of the convective and the diffusive terms of the solver can be evaluated separately. Test cases include flows over bluff bodies, such as circular cylinders and spheres, at various low Reynolds numbers, flows over a flat plate with and without turbulence effects, and turbulent flows over axisymmetric bodies with and without propulsor effects. Finally, to enhance the iterative solution procedure, a full approximation scheme V-cycle multigrid method is implemented. Preliminary results indicate that the method significantly reduces the computational effort.

NAVIER STOKES EQUATIONS↗

DOE OSTI2023

On the convergence of statistics in simulations of stationary incompressible turbulent flows

When reporting statistics from simulations of statistically stationary chaotic phenomenon, it is important to verify that the simulations are time-converged. This condition is connected with the statistical error or number of digits with which statistics can be reliably reported. In this work we consider numerical experiments of low Reynolds number incompressible homogeneous and isotropic turbulence as a model problem to investigate statistical convergence over finite simulation times. Specifically, we investigate the time integration requirements that allow meaningful reporting of the statistical error associated with finiteness of the temporal domain. We address two key questions: (1) How long should a simulation be performed in terms of large eddy time, and (2) How should the simulation time be divided among temporal windows over which a quantity of interest is estimated so that its statistical error could be reliably reported? We find that reliable reporting of statistical errors requires simulations on the order of 10 4 large eddy times, which is orders of magnitude longer than typically performed. Additionally, data post-processing should employ windows of at least ten times the large eddy time scale, with the most robust computation of statistical error of the mean requiring window sizes of an additional factor of ten. For practical simulations, we demonstrate that it is possible to estimate the statistical error within a factor of two under a less stringent condition in which a minimum of four windows with size at least ten large eddy times are used. In conclusion, our observations for homogeneous isotropic turbulence are also shown to hold in turbulent channel flow.

42 ENGINEERING↗

NASA NTRS1966

Binary fluid amplifier solves stability and load problems

Digital fluid amplifier has load intensity, high stability, and operates at low reynolds numbers. It contains specially designed nozzles to provide uniform exit-velocity profiles and to ensure jets of low turbulence.

Larkin, B. D.↗

NASA NTRS1990

An experimental investigation of nozzle-exit boundary layers of highly heated free jets

An experimental investigation of the effects of nozzle operating conditions on the development of nozzle-exit boundary layers of highly heated air free jets is reported in this paper. The total pressure measurements in the nozzle-exit boundary layer were obtained at a range of jet Mach numbers from 0.1 to 0.97 and jet total temperatures up to 900 K. The analysis of results shows that the nozzle-exit laminar boundary-layer development depends only on the nozzle-exit Reynolds number. For the nozzle-exit turbulent boundary layer, however, it appears that the effects of the jet total temperature on the boundary-layer integral characteristics are independent from the effects of the nozzle-exit Reynolds number. This surprizing finding has not yet been reported. Further, laminar boundary-layer profiles were compared with the Pohlhausen solution for a flat-wall converging channel and an acceptable agreement was found only for low Reynolds numbers. For turbulent boundary layers, the dependence of the shape factor on relative Mach numbers at a distance of one momentum thickness from the nozzle wall resembles Spence's prediction. Finally, the calculated total pressure loss coefficient was found to depend on the nozzle-exit Reynolds number for the laminar nozzle-exit boundary layer, while for the turbulent exit boundary layer this coefficient appears to be constant.

Lepicovsky, J.↗

NASA NTRS1997

[Low-Frequency Flow Oscillation

The results of the research conducted under this grant are presented in detail in three Master theses, by Heinrich, Balow, and Broeren. Additional analysis of the experimental data can be found in two AIAA Journal articles and two conference papers. Citations for all of the studies' publications can be found in the bibliography which is attached. The objective of Heinrich's study was to document the low-frequency flow oscillation on the LRN-1007 airfoil, which had been previously observed at low Reynolds number, to determine its origin, and explore the phenomenon at higher Reynolds number. Heinrich performed detailed flow visualization on the airfoil using surface fluorescent oil and laser-sheet off-body visualization. A large leading-edge separation bubble and trailing-edge separation was identified on the airfoil just prior to the onset of the unsteady stall flow oscillation. From the laser-sheet data, the unsteady flow appeared as a massive boundary-layer separation followed by flow reattachment. Hot-wire data were taken in the wake to identify the presence of the flow oscillation and the dominant frequency. The oscillation was found in the flow from a Reynolds number of 0.3 to 1.3 x 10 exp 6. The Strouhal number based on airfoil projected height was nominally 0.02 and increased slightly with increasing Reynolds number and significantly with increasing airfoil angle of attack. Balow focused his research on the leading-edge separation bubble which was hypothesized to be the origin of the low-frequency oscillation. Initially, experimental measurements in the bubble at the onset of the low-frequency oscillation were attempted to study the characteristics of the bubble and explain possible relationships to the shear-layer-flapping phenomena. Unfortunately, the bubble proved to be extremely sensitive to the probe interference and it drastically reduced the size of the bubble. These detailed measurements were then abandoned by Balow. However, this led to a series of tests where the leading-edge bubble and trailing-edge separation were altered and the affect on the flow-oscillation studied. Balow found that by tripping the airfoil boundary-layer with "zigzag" tape ahead of bubble separation, the bubble was effectively eliminated mid the oscillation suppressed. Wake survey drag measurements showed a drastic reduction in airfoil drag when the bubble and oscillation were eliminated. Using the "zigzag" tape, the trailing-edge separation was moved downstream approximately 5 percent chord. This was found to reduce the amplitude of the oscillation, particularly in the onset stage at low angle of attack (around 14 degrees). Through detailed analysis of the wake behind the airfoil during the unsteady flow oscillation, Balow provided a better understanding of the wake flowfield. Broeren studied the oscillating flowfield in detail at Reynolds number equal 3 x 10 exp 5 and an angle of attack of 15 degrees using laser Doppler velocimetry (LDV). Two-dimensional LDV data were acquired at 687 grid points above the model upper surface while hot-wire data were taken simultaneously in the wake. Using the hot-wire signal, the LDV data were phase averaged into 24 bins to represent a single ensemble average of one oscillation cycle. The velocity data showed a flowfield oscillation that could be divided into three flow regimes. In the first regime, the flow over the airfoil was completely separated initially, the flowfield reattached from the leading edge and the reattachment point moved downstream with increasing time or phase. Broeren referred to this as the reattachment regime. The bubble development regime followed, where a leading-edge separation bubble formed at the leading edge and grew with increasing time. During the initial part of this regime the trailing-edge separation continued to move downstream. However, during the last 30 degrees of phase the trailing-edge separation moved rapidly forward and appeared to merge with the leading-edge bubble. During the third regime, the separation regime, the flow was segmented from the airfoil leading edge and did not reattach to the airfoil surface. The reverse flow was seen to grow in vertical extent up from the model surface as the phase increased. Next reattachment began again at the leading edge signaling the start of the reattachment regime, and so the cycle continued. From Broeren's work, the details of the unsteady flowfield over the airfoil were seen for the first time. From this research a great deal has been learned about the low-frequency flow oscillation which naturally occurs on the LRN-1007 airfoil near stall. The oscillation was seen to persist at higher Reynolds number, the dependence of the Strouhal number on angle of attack and Reynolds number were discovered, the critical role played by the laminar bubble was shown and the entire upper surface flowfield during a flow oscillation cycle was measured and analyzed. What still eludes understanding is the scaling of the flow oscillation and why certain airfoils, such as the LRN, have a very strong low-frequency mode and other airfoils exhibit no organized low-frequency oscillation at all.

Bragg, Michael B.↗

NASA NTRS2018

Generation of Mars Helicopter Rotor Model for Comprehensive Analyses

The present research is aimed at providing a performance model for the Mars Helicopter (MH), to understand the complexity of the flow, and identify future regions of improvement. The low density of the Martian atmosphere and the relatively small MH rotor, result in very low chord-based Reynolds number flows. The low density and Reynolds numbers reduce the lifting force and lifting efficiency, respectively. The high drag coefficients in subcritical flow, especially for thicker sections, are attributed to laminar separation from the rear of the airfoil. In the absence of test data, efforts have been made to explore these effects using prior very low Reynolds number research efforts. The rotor chord-based Reynolds number range is observed to be subcritical, which makes boundary layer transition unlikely to occur. The state of the two-dimensional rotor boundary layer in hover is approximated by calculating the instability point, laminar separation point, and the transition location to provide understanding of the flow state in the high Mach-low Reynolds number regime. The results are used to investigate the need for turbulence modeling in Computational Fluid Dynamics (CFD) calculations afterwards. The goal is to generate a performance model for the MH rotor for a free wake analysis, because the computational budget for a complete Navier-Stokes solution for a rotating body-fitted rotor is substantial. In this study, a Reynolds-Averaged Navier-Stokes (RANS) based approach is used to generate the airfoil deck using C81Gen with stitched experimental data for very high angles of attack. A full Grid Resolution Study is performed and over 4,500 cases are completed to create the full airfoil deck. The laminar separation locations are predicted within the accuracy of the approximate method when compared with the CFD calculations. The model is presented through airfoil data tables (c81 files) that are used by comprehensive rotor analysis codes such as CAMRADII, or the mid-fidelity CFD solver RotCFD. Finally, the rotor performance is compared with experimental data from the 25ft Space Simulator at the NASA Jet Propulsion Laboratory (JPL) and shows good correlation for the rotor Figure of Merit over the available thrust range.

Model↗

NASA NTRS1992

An experimental study of a cold-wall hypersonic boundary layer

The boundary layer that developed on a one-meter long flat plate at a nominal Mach number of 8 and a Reynolds number of 10.2 million per meter was studied using a shock tunnel. The wall-to-stagnation temperature ratio was 0.35. The turbulent velocity profiles possessed very small wake components. The small wake component is characteristic of low Reynolds number flows and may indicate that a turbulent boundary layer was not fully developed. The magnitude of the surface pressure fluctuations in the turbulent part of the flow was found to be in agreement with semi-empirical predictions. These fluctuations were larger than those obtained at lower supersonic Mach numbers and indicated that pressure fluctuations may not be neglected at hypersonic Mach numbers. The fluctuations also possessed a convection velocity that increased with transducer spacing and followed the trend of previous investigations.

Chung, Kung-Ming↗

NASA NTRS1998

Aerodynamics of a Transitioning Turbine Stator Over a Range of Reynolds Numbers

Midspan aerodynamic measurements for a three vane-four passage linear turbine vane cascade are given. The vane axial chord was 4.45 cm. Surface pressures and loss coefficients were measured at exit Mach numbers of 0.3, 0.7, and 0.9. Reynolds number was varied by a factor of six at the two highest Mach numbers, and by a factor of ten at the lowest Mach number. Measurements were made with and without a turbulence grid. Inlet turbulence intensities were less than I% and greater than IO%. Length scales were also measured. Pressurized air fed the test section, and exited to a low pressure exhaust system. Maximum inlet pressure was two atmospheres. The minimum inlet pressure for an exit Mach number of 0.9 was one-third of an atmosphere, and at a Mach number of 0.3, the minimum pressure was half this value. The purpose of the test was to provide data for verification of turbine vane aerodynamic analyses, especially at low Reynolds numbers. Predictions obtained using a Navier-Stokes analysis with an algebraic turbulence model are also given.

Boyle, R. J.↗

NASA NTRS1980

Full-coverage film cooling. II - Heat transfer data and numerical simulation

Experimental research into heat transfer from full-coverage film-cooled surfaces with three injection geometries was described in Part I. This part has two objectives. The first is to present a simple numerical procedure for simulation of heat transfer with full-coverage film cooling. The second objective is to present some of the Stanton number data that was used in Part I of the paper. The data chosen for presentation are the low-Reynolds number, heated-starting-length data for the three injection geometries with five-diameter hole spacing. Sample data sets with high blowing ratio and with ten-diameter hole spacing are also presented. The numerical procedure has been successfully applied to the Stanton number data sets.

Crawford, M. E.↗

NASA NTRS1991

Project ARES 2: High-altitude battery-powered aircraft

A high-altitude, battery-powered, propeller-driven aircraft was designed and is being built by undergraduate students at California State University, Northridge. The aircraft will fly at an altitude of 104,000 ft at Mach 0.2 (190 ft/sec) and will be instrumented to record flight performance data, including low Reynolds number propeller and airfoil information. This project will demonstrate the feasibility of electric-powered flight in a low-density, low-temperature Earth environment that models the atmosphere of Mars. Data collected will be used to design a Mars aircraft to investigate the surface of Mars prior to manned missions. The instrumented payload and the mission profile for the high-altitude Earth flight were determined. Detailed aerodynamic and structural analyses were performed. Control, tracking, and data recording subsystems were developed. Materials were obtained and fabrication begun. The aircraft has a 32-ft wing span, a wing area of 105 sq ft, is 17.5 ft long, has a 12-in payload bay, and weighs 42 lb. It is composed primarily of lightweight materials, including Mylar, and composite materials, including graphite/epoxy and aramid core honeycomb sandwich. Low-altitude flight testing to check guidance and control systems and to calibrate data-gathering instruments will take place this summer, followed shortly by the 104,000-ft flight.

Source record↗

NASA NTRS1994

Crossflow instability control on a swept-wing: Preliminary studies

The pressure distribution on a swept wing causes the streamlines at the edge of the boundary layer to be curved. This pressure gradient normal to the external streamline creates a velocity component normal to the external streamline within the boundary layer which is referred to as the crossflow velocity. Because the crossflow velocity profile perpendicular to the wing surface has an inflection point, the profile is unstable. The stationary instability mode takes the form of crossflow vortices. Under these conditions, the boundary layer on the wing is extremely unstable and transition to turbulent flow takes place much closer to the leading edge of the wing than it would on an unswept wing. Higher skin friction drag is associated with turbulent flow, and so better aircraft performance could be obtained if the crossflow could be eliminated One method of controlling crossflow that is being investigated is boundary-layer suction. An extensive airfoil suction experiment in the 8 feet Transonic Pressure Tunnel (TPT) at NASA Langley Research Center will begin late in 1994. Because of the size, complexity, and expense associated with this test, a number of 'risk-reduction' tests are currently being conducted. The 20 x 28 in. Shear Flow Control Tunnel at NASA Langley is being used for some of these tests. Prior to the summer of 1994, a flat plate with a swept leading edge was installed in the 20 x 28 in. tunnel, with a displacement body mounted on the tunnel ceiling that created a pressure distribution on the plate similar to the pressure distribution on a swept wing. The flow over the plate was investigated during the summer of 1994 using a laser Doppler velocimeter (LDV) system. The LDV measurements indicated the possible presence of multiple disturbance modes, a rarely-seen phenomena since, in most tests, one disturbance mode dominates. The possible existence of multiple disturbance modes in the flat plate boundary layer, however, means that the flow in the 20 x 28 in. tunnel is of interest itself, and will be investigated more thoroughly in the future. With a view to these investigations, the boundary layer traverse mechanism in the 20 x 28 in. tunnel was modified to improve its performance, and strain gauges were mounted on the traverse in order to monitor its deflection during a test. Other preliminary work conducted in the 20 x 28 in. tunnel included the use of an infrared camera system. Previous work with this system showed that transition indeed could be detected, but the signal produced by the crossflow vortices was too weak to be detected. It was hoped that spraying the flat plate with naphthalene would augment the heat transfer associated with the crossflow vortices so that they would show up in a IR image; however, experiments showed that this would not work. Another set of tests was conducted in the 20 x 28 in. tunnel to determine the tripping requirements for a set of airfoil-shaped struts that will be used in the 8 feet TPT experiment. Since the Reynolds number associated with these struts is small, a laminar boundary layer would separate early, causing large fluctuations in the flow field. A turbulent boundary layer would remain attached further back, but tripping from laminar to turbulent flow at low Reynolds number is very difficult. However, trip strip configurations were found that should effectively trip the boundary layer at the required conditions. Currently underway is an investigation of the data acquisition requirements for the 8 feet TPT experiment, with the purpose of the finding the minimum amount of data needed to characterize sufficiently the swept-wing boundary layer. This study is being conducted using a numerically generated data set.

Bridges, David H.↗

NASA NTRS1960

Heat Transfer to 36.75 and 45 degree Swept Blunt Leading Edges in Free Flight at Mach Numbers from 1.70 to 2.99 and From 2.50 to 4.05

A flight investigation has been conducted to study the heat transfer to swept-wing leading edges. A rocket-powered model was used for the investigation and provided data for Mach number ranges of 1.78 to 2.99 and 2.50 to 4.05 with corresponding free-stream Reynolds number per foot ranges of 13.32 x 10(exp 6) to 19.90 x 10(exp 6) and 2.85 x 10(exp 6) to 4.55 x 10(exp 6). The leading edges employed were cylindrically blunted wedges ', three of which were swept 450 with leading-edge diameters of 1/4, 1/2, and 3/4 inch and one swept 36-750 with a leading-edge diameter of 1/2 inch. In the high Reynolds number range, measured values of heat transfer were found to be much higher than those predicted by laminar theory and at the larger values of leading-edge diameter were approaching the values predicted by turbulent theory. For the low Reynolds number range a comparison between measured and theoretical heat transfer showed that increasing the leading-edge diameter resulted in turbulent flow on the cylindrical portion of the leading edge.

ONeal, Robert L.↗

NASA NTRSDate not supplied

Introduction to the NATO AVT-298 Task Group and Development of the SWiFT Configuration

Reynolds number effects are a key development risk for blended wing body aircraft, particularly at high lift take-off and landing conditions. The full scale aerodynamics of these configurations are thought not to be well represented in industry standard low Reynolds number test facilities and the capability to predict Reynolds number effects using computational fluid dynamics or otherwise is not proven. The NATO AVT-298 research task group was established to investigate these issues and this paper sets out the background, motivations and the collaborative approach as well as the development of a new research configuration to serve as a focus for the research. The ‘SWiFT’ configuration, was developed to exhibit flow physics of interest to the AVT-298 research task group and to be suitable for an experimental campaign. The group collaboratively developed the research configuration to be relevant for related blended-wing-body aircraft programmes using computational fluid dynamics and wing design methods and considering the practicalities of testing. This research targets an improvement in our capability to design, optimise and mature future blended wing body aircraft.

SWiFT↗

NASA NTRS1990

High Reynolds Number Test of the Boeing TR77 Airfoil in the Langley 0.3-Meter Transonic Cryogenic Tunnel

A Boeing TR77 airfoil associated with the Advanced Technology Airfoil Test (ATAT) program was tested in the Langley 0.3 m Transonic Cryogenic Tunnel. Limited analysis of the data indicated that increasing Reynolds number for a fixed Mach number resulted in increased normal-force, nose-down pitching moment, and decreased drag coefficient. Increasing Mach number while keeping the Reynolds number constant yielded the expected increase in normal-force slopes, nose-down pitching moment coefficients, and decrease in angle of attack associated with maximum normal-force coefficient. Turbulent boundary layer flow was achieved over the airfoil at low Reynolds numbers for the test Mach number range using aluminum discs.

Chu, Julio↗

NASA NTRS1956

A Theory for Stability and Buzz Pulsation Amplitude in Ram Jets and an Experimental Investigation Including Scale Effects

From a theory developed on a quasi-one-dimensional-flow basis, it is found that the stability of the ram jet is dependent upon the instantaneous values of mass flow and total pressure recovery of the supersonic diffuser and immediate neighboring subsonic diffuser. Conditions for stable and unstable flow are presented. The theory developed in the report is in agreement with the experimental data of NACA-TN-3506 and NACA-RM-L50K30. A simple theory for predicting the approximate amplitude of small pressure pulsation in terms of mass-flow decrement from minimum-stable mass flow is developed and found to agree with experiments. Cold-flow tests at a Mach number of 1.94 of ram-jet models having scale factors of 3.15:1 and Reynolds number ratios of 4.75:1 with several supersonic diffuser configurations showed only small variations in performance between geometrically similar models. The predominant variation in steady-flow performance resulted from the larger boundary layer in the combustion chamber of the low Reynolds number models. The conditions at which buzz originated were nearly the same for the same supersonic diffuser (cowling-position angle) configurations in both large and small diameter models. There was no appreciable variation in stability limits of any of the models when the combustion-chamber length was increased by a factor of three. The unsteady-flow performance and wave patterns were also similar when considered on a reduced-frequency basis determined from the relative lengths of the model. The negligible effect of Reynolds number on stability of the off-design configurations was not anticipated in view of the importance of boundary layer to stability, and this result should not be construed to be generally applicable. (author)

Trimpi, Robert L↗