Development of a two-stream mixing layer from tripped and untripped boundary layers
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Bell, James H.↗
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Direct Numerical Simulations of High-Speed Turbulent Boundary Layers over Riblets
Direct numerical simulations (DNS) of spatially developing turbulent boundary layers over riblets with a broad range of riblet spacings are conducted to investigate the effects of riblets on skin friction at high speeds. Zero-pressure gradient boundary layers under two flow conditions (Mach 2:5 with T(sub w)/T(sub r) = 1 and Mach 7:2 with T(sub w)/T(sub r) = 0:5) are considered. The DNS results show that the drag-reduction curve (delta C(sub f)/C(sub f) vs l(sup +)(sub g )) at both supersonic speeds follows the trend of low-speed data and consists of a `viscous' regime for small riblet size, a `breakdown' regime with optimal drag reduction, and a `drag-increasing' regime for larger riblet sizes. At l l(sup +)(sub g) approx. 10 (corresponding to s+ approx 20 for the current triangular riblets), drag reduction of approximately 7% is achieved at both Mach numbers, and con rms the observations of the few existing experiments under supersonic conditions. The Mach- number dependence of the drag-reduction curve occurs for riblet sizes that are larger than the optimal size, with smaller slopes of (delta C(sub f)/C(sub f) for larger freestream Mach numbers. The Reynolds analogy holds with 2(C(sub h)=C(sub f) approximately equal to that of at plates for both drag-reducing and drag-increasing configurations.
Duan, Lian↗
Transducer Resolution Effect on Pressure Fluctuations Beneath Hypersonic Turbulent Boundary Layers
The size of a pressure transducer is known to affect the accuracy of measurements of wall-pressure fluctuations beneath a turbulent boundary layer because of spatial averaging over the sensing area of the transducer. In this paper, the effect of finite transducer size is investigated by applying spatial averaging or wavenumber filters to a database of hypersonic wall pressure generated from a direct numerical simulation (DNS) that simulates the turbulent portion of the boundary layer over a sharp 7° half-angle cone at nominally Mach 8. Here, a good comparison between the DNS and the experiment in the Sandia Hypersonic Wind Tunnel at Mach 8 is achieved after spatial averaging is applied to the DNS data over an area similar to the sensing area of the transducer. The study shows that a finite sensor size similar to that of the PCB132 transducer can cause significant attenuation in the root-mean-square and power spectral density (PSD) of wall-pressure fluctuations, and the attenuation effect is identical between cone and flat plate configurations at the same friction Reynolds number. The Corcos theory is found to successfully compensate for the attenuated high-frequency components of the wall-pressure PSD.
42 ENGINEERING↗
Measurements in Separated and Transitional Boundary Layers Under Low-Pressure Turbine Airfoil Conditions
Detailed velocity measurements were made along a flat plate subject to the same dimensionless pressure gradient as the suction side of a modern low-pressure turbine airfoil. Reynolds numbers based on wetted plate length and nominal exit velocity were varied from 50,000 to 300,000, covering cruise to takeoff conditions. Low and high inlet free-stream turbulence intensities (0.2% and 7%) were set using passive grids. The location of boundary-layer separation does not depend strongly on the free-stream turbulence level or Reynolds number, as long as the boundary layer remains non-turbulent prior to separation. Strong acceleration prevents transition on the upstream part of the plate in all cases. Both free-stream turbulence and Reynolds number have strong effects on transition in the adverse pressure gradient region. Under low free-stream turbulence conditions transition is induced by instability waves in the shear layer of the separation bubble. Reattachment generally occurs at the transition start. At Re = 50,000 the separation bubble does not close before the trailing edge of the modeled airfoil. At higher Re, transition moves upstream, and the boundary layer reattaches. With high free-stream turbulence levels, transition appears to occur in a bypass mode, similar to that in attached boundary layers. Transition moves upstream, resulting in shorter separation regions. At Re above 200,000, transition begins before separation. Mean velocity, turbulence and intermittency profiles are presented.
Volino, Ralph J.↗
Analysis of in-flight boundary-layer state measurements on a subsonic transport wing in high-lift configuration
Flight experiments on NASA Langley's B737-100 (TSRV) airplane have been conducted to document flow characteristics in order to further the understanding of high-lift flow physics, and to correlate and validate computational predictions and wind-tunnel measurements. The project is a cooperative effort involving NASA, industry, and universities. In addition to focusing on in-flight measurements, the project includes extensive application of various computational techniques, and correlation of flight data with computational results and wind-tunnel measurements. Results obtained in the most recent phase of flight experiments are analyzed and presented in this paper. In-flight measurements include surface pressure distributions, measured using flush pressure taps and pressure belts on the slats, main element, and flap elements; surface shear stresses, measured using Preston tubes; off-surface velocity distributions, measured using shear-layer rakes; aeroelastic deformations of the flap elements, measured using an optical positioning system; and boundary-layer transition phenomena, measured using hot-film anemometers and an infrared imaging system. The analysis in this paper primarily focuses on changes in the boundary-layer state that occurred on the slats, main element, and fore flap as a result of changes in flap setting and/or flight condition. Following a detailed description of the experiment, the boundary-layer state phenomenon will be discussed based on data measured during these recent flight experiments.
vanDam, C. P.↗
Coupling Between Alfvén Wave and Kelvin–Helmholtz Waves in the Low Latitude Boundary Layer
The Kelvin–Helmholtz (KH) instability of magnetohydrodynamic surface waves at the low latitude boundary layer is examined using both an eigenfrequency analysis and a time-dependent wave simulation. The analysis includes the effects of sheared flow and Alfvén velocity gradient. When the magnetosheath flows are perpendicular to the ambient magnetic field direction, unstable KH waves that propagate obliquely to the sheared flow direction occur at the sheared flow surface when the Alfvén Mach number is higher than an instability threshold. Including a shear transition layer between the magnetosphere and magnetosheath leads to secondary KH waves (driven by the sheared flow) that are coupled to the resonant surface Alfvén wave. There are remarkable differences between the primary and the secondary KH waves, including wave frequency, the growth rate, and the ratio between the transverse and compressional components. The secondary KH wave energy is concentrated near the shear Alfvén wave frequency at the magnetosheath with a lower frequency than the primary KH waves. Although the growth rate of the secondary KH waves is lower than the primary KH waves, the threshold condition is lower, so it is expected that these types of waves will dominate at a lower Mach number. Because the transverse component of the secondary KH waves is stronger than that of the primary KH waves, more efficient wave energy transfer from the boundary layer to the inner magnetosphere is also predicted.
79 ASTRONOMY AND ASTROPHYSICS↗
An Experimental Study of Roughness-Induced Instabilities in a Supersonic Boundary Layer
Progress on an experimental study of laminar-to-turbulent transition induced by an isolated roughness element in a supersonic laminar boundary layer is reported in this paper. Here, the primary focus is on the effects of roughness planform shape on the instability and transition characteristics. Four different roughness planform shapes were considered (a diamond, a circle, a right triangle, and a 45 degree fence) and the height and width of each one was held fixed so that a consistent frontal area was presented to the oncoming boundary layer. The nominal roughness Reynolds number was 462 and the ratio of the roughness height to the boundary layer thickness was 0.48. Detailed flow- field surveys in the wake of each geometry were performed via hot-wire anemometry. High- and low-speed streaks were observed in the wake of each roughness geometry, and the modified mean flow associated with these streak structures was found to support a single dominant convective instability mode. For the symmetric planform shapes - the diamond and circular planforms - the instability characteristics (mode shapes, growth rates, and frequencies) were found to be similar. For the asymmetric planform shapes - the right-triangle and 45 degree fence planforms - the mode shapes were asymmetrically distributed about the roughness-wake centerline. The instability growth rates for the asymmetric planforms were lower than those for the symmetric planforms and therefore, transition onset was delayed relative to the symmetric planforms.
Kegerise, Michael A.↗
Measurements in the near-wall region of a relaxing three-dimensional low speed turbulent air boundary layer
An experimental investigation was conducted at selected locations of the near-wall region of a three dimensional turbulent air boundary layer relaxing in a nominally zero external pressure gradient behind a transverse hump (in the form of a 30 deg swept, 5-foot chord wing-type model) faired into the side wall of a low speed wind tunnel. Wall shear stresses measured with a flush-mounted hot-film gage and a sublayer fence were in very good agreement with experimental data obtained with two Preston probes. With the upstream unit Reynolds number held constant at 325,000/ft. approximately one-fourth of the boundary layer thickness adjacent to the wall was surveyed with a single rotated hot-wire probe mounted on a specially designed minimum interference traverse mechanism. The boundary layer (approximately 3.5 in thick near the first survey station where the length Reynolds number was 5.5 million) had a maximum crossflow velocity ratio of 0.145 and a maximum crossflow angle of 21.875 deg close to the wall.
Hebbar, K. S.↗
Intermittency distributions in turbulent wall boundary layers with and without pressure gradients
Intermittency distributions in turbulent wall boundary layers with and without pressure gradients
BOUNDARY LAYER FLOW↗
The influence of ablation on stagnation region convective heating for dissociated and partially ionized boundary-layer flows.
Convective heat transfer through high temperature boundary layers with and without ablation products
BOUNDARY LAYER↗
A non-isothermal wall boundary layer computer program with corner effects
Nonisothermal wall boundary layer computer program with corner effects
WALL↗
Turbulent measurements in supersonic boundary layer flows using laser velocimetry
Laser velocimetry has been applied successfully in the measurement of mean velocities and Reynolds stresses in a zero pressure gradient, Mach 2.9 boundary layer, and upstream and downstream of a two-dimensional, shock-wave boundary layer interaction for the same Mach number. The Reynolds stresses were obtained using a one-component, dual-scatter laser velocimeter operated in a manner analogous to the 'slanted hot-wire' technique. Signal processing was of the single-particle counting type which permitted measurements to be made using only the naturally occurring particles in the tunnel air stream for light scattering. The results for the shock-wave boundary layer interaction are presented along with hot-wire anemometer and pitot-static pressure measurements obtained in the same flow. Also, a data acquisition scheme for use with a two-component laser velocimeter is presented which should provide even better accuracies in future studies.
Johnson, D. A.↗
Variable Sweep Transition Flight Experiment (VSTFE)-Parametric Pressure Distribution Boundary Layer Stability Study and Wing Glove Design Task
The Variable Sweep Transition Flight Experiment (VSTFE) was initiated to establish a boundary-layer transition data base for laminar flow wing design. For this experiment, full-span upper-surface gloves will be fitted to a variable sweep F-14 aircraft. The results of two initial tasks are documented: a parametric pressure distribution/boundary-layer stability study and the design of an upper-surface glove for Mach 0.8. The first task was conducted to provide a data base from which wing-glove pressure distributions could be selected for glove designs. Boundary-layer stability analyses were conducted on a set of pressure distributions for various wing sweep angles, Mach numbers, and Reynolds number in the range of those anticipated for the flight-test program. The design procedure for the Mach 0.8 glove is described, and boundary-layer stability calculations and pressure distributions are presented both at design and off-design conditions. Also included is the analysis of the clean-up glove (smoothed basic wing) that will be flight-tested initially and the analysis of a Mach 0.7 glove designed at the NASA Langley Research Center.
Rozendaal, Rodger A.↗
Three-dimensional boundary-layer instability and separation induced by small-amplitude streamwise vorticity in the upstream flow
We consider the effects of a small-amplitude, steady, streamwise vorticity field on the flow over an infinitely thin flat plate in an otherwise uniform stream. We show how the initially linear perturbation, ultimately leads to a small-amplitude but nonlinear cross flow far downstream from the leading edge. This motion is imposed on the boundary-layer flow and eventually causes the boundary layer to separate. The streamwise velocity profiles within the boundary layer become inflexional in localized spanwise regions just upstream of the separation point. The flow in these regions is therefore susceptible to rapidly growing inviscid instabilities.
Goldstein, M. E.↗
Numerical Simulation of Boundary Layer Ingesting (BLI) Inlet-Fan Interaction
Future civil transport designs may incorporate engine inlets integrated into the body of the aircraft to take advantage of efficiency increases due to weight and drag reduction. Additional increases in engine efficiency are predicted if the inlet ingests the lower momentum boundary layer flow. Previous studies have shown, however, that efficiency benefits of Boundary Layer Ingesting (BLI) ingestion are very sensitive to the magnitude of fan and duct losses, and blade structural response to the non-uniform flow field that results from a BLI inlet has not been studied in-depth. This paper presents an effort to extend the modeling capabilities of an existing rotating turbomachinery unsteady analysis code to include the ability to solve the external and internal flow fields of a BLI inlet. The TURBO code has been a successful tool in evaluating fan response to flow distortions for traditional engine/inlet integrations, such as the development of rotating stall and inlet distortion through compressor stages. This paper describes the first phase of an effort to extend the TURBO model to calculate the external and inlet flowfield upstream of fan so that accurate pressure distortions that result from BLI configurations can be computed and used to analyze fan aerodynamics and structural response. To validate the TURBO program modifications for the BLI flowfield, experimental test data obtained by NASA for a flushmounted S-duct with large amounts of boundary layer ingestion was modeled. Results for the flow upstream and in the inlet are presented and compared to experimental data for several high Reynolds number flows to validate the modifications to the solver. Quantitative data is presented that indicates good predictive capability of the model in the upstream flow. A representative fan is attached to the inlet and results are presented for the coupled inlet/fan model. The impact on the total pressure distortion at the AIP after the fan is attached is examined.
Inlet Fan Interactions↗
Eliminating Flow Separation and Reducing Viscous Drag Through Boundary Layer Analysis and Manipulation
As both computers and flow-analyzing equations have increased in sophistication, Computational Fluid Dynamics (CFD) has evolved into a fixture for advanced aircraft design. While CFD codes have improved in accuracy and efficiency, their ability to encompass viscous effects is lacking in certain areas. For example, current CFD codes cannot accurately predict or correct for the increased drag due to these viscous effects at some flow conditions. However, by analyzing an airfoil's turbulent boundary layer, one can predict not only flow separation via the shape factor parameter, but also viscous drag via the momentum thickness. Various codes have been written which can calculate turbulent boundary layer parameters. The goal of my research is to develop procedures for modifying an airfoil (via its local pressure distribution) to eliminate boundary layer separation and/or to reduce viscous drag. The modifications to the local pressure distribution necessary to achieve these objectives will be determined using a direct-iterative method installed into a turbulent boundary layer analyzer. Furthermore, the modifications should preserve the basic characteristics of the original airfoil.
Oser, Matthew D.↗
Shock wave-boundary layer interactions in rarefied gas flows
A numerical study is presented, using the direct simulation Monte Carlo (DSMC) method, of shock wave-boundary layer interactions in low density supersonic flows. Test cases include two-dimensional, axially-symmetric and three-dimensional flows. The effective displacement angle of the boundary layer is calculated for representative flat plate, wedge, and cone flows. The maximum pressure, shear stress, and heat transfer in the shock formation region is determined in each case. The two-dimensional reflection of an oblique shock wave from a flat plate is studied, as is the three-dimensional interaction of such a wave with a sidewall boundary layer.
Bird, G. A.↗
Unsteady laminar boundary-layer calculations on oscillating configurations including backflow. Part 2: Airfoil in high-amplitude pitching motion. Dynamic stall
A previously developed finite-difference procedure for calculating unsteady, incompressible, laminar boundary layers on an oscillating flat plate is applied to a wing section undergoing high-amplitude pitching oscillations about various mean incidences. To start the entire boundary-layer calculation, appropriate initial conditions and outer boundary conditions are specified, using a stagnation-point fixed frame of reference. The breakdown of the numerical calculation procedure in the x,t-domain is interpreted to coincide with unsteady separation. Details of the boundary-layer behavior in the vicinity of separation are investigated, and a close analogy between the present results and those for a three-dimensional steady separation is found.
Geissler, W.↗