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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS2015

Time-Frequency Analysis of Boundary-Layer Instabilites Generated by Freestream Laser Perturbations

A controlled disturbance is generated in the freestream of the Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT) by focusing a high-powered Nd:YAG laser to create a laser-induced breakdown plasma. The plasma then cools, creating a freestream thermal disturbance that can be used to study receptivity. The freestream disturbance convects down-stream in the Mach-6 wind tunnel to interact with a flared cone model. The adverse pressure gradient created by the flare of the model is capable of generating second-mode instability waves that grow large and become nonlinear before experiencing natural transition in quiet flow. The freestream laser perturbation generates a wave packet in the boundary layer at the same frequency as the natural second mode, complicating time-independent analyses of the effect of the laser perturbation. The data show that the laser perturbation creates an instability wave packet that is larger than the natural waves on the sharp flared cone. The wave packet is still difficult to distinguish from the natural instabilities on the blunt flared cone.

Chou, Amanda↗
NASA NTRS2017

Wall-Resolved Large-Eddy Simulation of Flow Separation Over NASA Wall-Mounted Hump

This paper reports the findings from a study that applies wall-resolved large-eddy simulation to investigate flow separation over the NASA wall-mounted hump geometry. Despite its conceptually simple flow configuration, this benchmark problem has proven to be a challenging test case for various turbulence simulation methods that have attempted to predict flow separation arising from the adverse pressure gradient on the aft region of the hump. The momentum-thickness Reynolds number of the incoming boundary layer has a value that is near the upper limit achieved by recent direct numerical simulation and large-eddy simulation of incompressible turbulent boundary layers. The high Reynolds number of the problem necessitates a significant number of grid points for wall-resolved calculations. The present simulations show a significant improvement in the separation-bubble length prediction compared to Reynolds-Averaged Navier-Stokes calculations. The current simulations also provide good overall prediction of the skin-friction distribution, including the relaminarization observed over the front portion of the hump due to the strong favorable pressure gradient. We discuss a number of problems that were encountered during the course of this work and present possible solutions. A systematic study regarding the effect of domain span, subgrid-scale model, tunnel back pressure, upstream boundary layer conditions and grid refinement is performed. The predicted separation-bubble length is found to be sensitive to the span of the domain. Despite the large number of grid points used in the simulations, some differences between the predictions and experimental observations still exist (particularly for Reynolds stresses) in the case of the wide-span simulation, suggesting that additional grid resolution may be required.

Uzun, Ali↗
NASA NTRS2018

The Effect of Forward-Facing Steps on Stationary Crossflow Instability Growth and Breakdown

The effect of a forward-facing step on stationary crossflow transition was studied using standard stereo particle image velocimetry (PIV) and time-resolved PIV. Step heights ranging from 53 to 71% of the boundary-layer thickness were studied in detail. The steps above a critical step height of approximately 60% of the boundary-layer thickness had a significant impact on the stationary crossflow growth downstream of the step. For the critical cases, the stationary crossflow amplitude grew suddenly downstream of the step, decayed for a short region, then grew again. The adverse pressure gradient upstream of the step resulted in a region of crossflow reversal. A secondary set of vortices, rotating in the opposite direction to the primary vortices, developed underneath the uplifted primary vortices. The wall-normal velocity disturbance (V' ) created by these secondary vortices impacted the step, and is believed to feed into the strong vortex that developed downstream of the step. A large but very short negative crossflow region formed for a short region downstream of the step due to a sharp inboard curvature of the streamlines near the wall. For the larger step height cases, a crossflow-reversal region formed just downstream of the strong negative crossflow region. This crossflow reversal region is believed to play an important role in the growth of the stationary crossflow vortices downstream of the step, and may be a good indication of the critical forward-facing step height.

Eppink, Jenna L.↗
NASA NTRS2019

Wind Tunnel Testing of Active Flow Control on High-Lift Common Research Model

A 10%-scale high-lift version of the Common Research Model (CRM-HL) and an Active Flow Control (AFC) version of the model equipped with a simple-hinged flap (CRM-SHLAFC) were successfully tested. The tests were performed in the 14- by 22-Foot Subsonic Tunnel (14x22) at the NASA Langley Research Center (LaRC). The CRM-HL has a set of 37° inboard and outboard single-element Fowler flaps. The CRM-SHL-AFC has a set of 50° inboard and 55° outboard simple-hinged flaps equipped with integrated modular AFC cartridges on the flap shoulder. Both high-lift configurations share the same 30° slats and engine nacelle. Three new types of AFC devices were examined: the Double-Row Sweeping Jets (DRSWJ), the Alternating Pulsed Jets (APJ), and the High Efficiency Low Power (HELP) actuators. The DRSWJ and the APJ actuators used two rows of unsteady jets, whereas the HELP actuators used a combination of unsteady and steady jets, to overcome strong adverse pressure gradients while minimizing the mass flow usage. Nozzle pressure ratio, mass flow consumption and the power coefficient, which takes account of both supply air pressure and mass flow usage for the actuators, were used for judging the performance efficiency of the AFC devices. A prestall lift performance degradation for the CRM-HL configuration was resolved with a properly placed nacelle chine. The configuration with nacelle chine was chosen as the representative reference conventional high-lift case for comparison with the CRMSHL- AFC. The AFC-induced lift coefficient increment (DCL) was maintained for the entire lift curve over the CRM-SHL-AFC case with no AFC for almost all flow-control cases examined. The lift curve of the reference CRM-HL have a slightly steeper slope compared to those of the CRM-SHL-AFC configurations. The HELP actuation concept was extremely effective in controlling flow separation in the “linear region” of the curves comparing lift coefficient to mass flow rate. The HELP actuation achieved a targeted DCL of 0.50 using a moderate amount of mass flow and supply air pressure. The CRM-SHL-AFC configuration equipped with HELP actuation was able to match or exceed the lift performance of the reference conventional high-lift configuration (i.e., CRM-HL equipped with a nacelle chine), thus meeting the NASA Advanced Air Transport Technology (AATT) project goal.

Lin, John C.↗
NASA NTRS2020

Wind Tunnel Testing of High Efficiency Low Power (HELP) Actuation for Active Flow Control

A High Efficiency Low Power (HELP) methodology has been successfully developed and tested on an active flow control (AFC) version of the 10%-scale high-lift Common Research Model (CRM-HL) at the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel (14x22). The AFC variant of the CRM-HL, designated as CRM-SHL-AFC, was integrated with modular HELP actuator cartridges on the shoulder of its highly deflected (≥50°) simplehinged flaps. A reference conventional CRM-HL configuration equipped with Fowler flaps and a nacelle chine was chosen to provide the targeted lift enhancement goals and for comparison with the results of the CRM-SHL-AFC. The current data are presented with the Transonic Wall Interference Correction System (TWICS) method applied. The HELP actuators, which use a combination of unsteady sweeping jets and steady discrete jets in tandem, were designed to overcome strong adverse pressure gradients, while minimizing the pneumatic power usage. The power coefficient (C(sub π)), which takes account of both supply air pressure and mass flow usage for the AFC actuators, is a useful parameter for judging the actuators’ performance efficiency and perhaps for scaling-up of the AFC system. Full HELP actuation coverage with a constant spanwise nozzle pressure ratio (NPR) was able to achieve the lift enhancement goals for the entire lift curve and was found to be the most effective AFC case. For the most effective case, mass flow rates greater than ~0.91 lbm/s and NPR values greater than ~1.8 (corresponding to C(sub π) = 0.18) are needed to achieve the lift enhancement goal at α = 9°, whereas mass flow rates greater than ~1.23 lbm/s and NPR values greater than ~2.3 (corresponding to C(sub π) = 0.3) are needed to achieve the lift enhancement goal at α = 17°. Surface pressure distributions indicate that the HELP actuation on the flap shoulder increased the suction pressures (and flow circulation) globally in both the streamwise and the spanwise directions, and thereby enhanced the lift over the entire high-lift system. The CRM-SHLAFC configuration equipped with HELP actuation was able to match or exceed the lift performance of the reference conventional CRM-HL, thus meeting the objective of the research.

John C Lin↗
NASA NTRS2021

Numerical Simulation of Pressure-induced Separation of Turbulent Flat-plate Boundary Layers: Definition and Overview of New Cases with Suction-only Transpiration and a Step in Reynolds Number

Results from two new direct numerical simulations (DNSs) are added to the database described in Coleman, Rumsey & Spalart (2018) (henceforth CRS18), and available at the NASA TMR website (https://turbmodels.larc.nasa.gov). The new flows, Cases D and E, are similar to the earlier ones (Cases A–C) in that a transpiration profile above a flat plate creates a prolonged adverse pressure gradient (APG) that drives a canonical turbulent zero-pressure-gradient (ZPG), flat-plate boundary layer towards separation; they differ in that for the new cases (1) the APG is not followed by a favorable gradient (FPG), since only suction transpiration is employed, and (2) the mean shear stress does not cross zero, so that the width of the mean bubble, strictly defined, is essentially zero. However, the probability of reversed skin friction exceeds 49%. The motivation was to produce a flow that is more akin to technological flows, including shock-boundary-layer interaction, with a gradual turbulence-controlled recovery. One of the new simulations is also at a significantly higher Reynolds number than the earlier flows.

Turbulence↗
NASA NTRS2022

Semispan Test Results of an Active Flow Control Enabled High-Lift Common Research Model in Landing Configuration

A 10%-scale semispan, Active Flow Control (AFC) enabled, simplified high-lift version of the Common Research Model (CRM-SHL-AFC) was tested in the 14- by 22-Foot Subsonic Tunnel at the NASA Langley Research Center. The main objective of the test was to develop an AFC system that can provide the necessary lift recovery on a simple-hinged flap high-lift system while minimizing its pneumatic power requirement. Three new types of AFC approaches were examined: Double-Row Sweeping Jets (DRSWJ), Alternating Pulsed Jets (APJ), and High Efficiency Low Power (HELP) actuators. The DRSWJ and the APJ actuators used two rows of unsteady jets, whereas the HELP actuators used an upstream row of sweeping jets combined with a downstream row of steady jets to overcome strong adverse pressure gradients. The test was conducted mostly at a freestream Mach number of 0.20. For exploration purposes, a limited number of runs were made at lower Mach numbers or using vortex generators (VGs). Minimal sensitivity to Mach number or VGs, for the cases evaluated, were observed. The AFC-induced lift coefficient increment was maintained over the AFC-off case for most flow-control cases examined. The CRM-SHL-AFC configuration equipped with HELP actuation was the only actuator configuration able to match or exceed the targeted lift performance of a reference conventional high-lift configuration. The presented aerodynamic data include lift, drag, and pitching moment coefficients as a function of angle of attack, with and without the Transonic Wall Interference Correction System (TWICS) method applied. Lift increments as a function of AFC pneumatic power usage (i.e., nozzle pressure ratio, mass flow, momentum coefficient, and power coefficient) are also presented at a lower angle of attack (α = 8.9°) and at maximum lift (α= 17.1°). At the two angles of attack, the surface pressure distributions and autospectral densities from the unsteady pressure transducers for the AFC-off case and the best HELP actuation case are compared.

high-lift↗
NASA NTRSDate not supplied

Liquid Jet Response to Detonation Waves in a Linear Detonation Combustor

The impact of periodic detonation wave impact on a liquid fuel jet is investigated in a linear detonation combustor. The linear detonation combustor operated with gaseous natural gas and oxygen generates sustained, self-excited detonation waves that propagate along its length at approximately 8 kHz, representing a wave propagation frequency in typical rocket rotating detonation engines. The effect of the detonation wave on the dynamic injection and break-up of a single diesel jet injected into the combustion chamber at varying injection pressures is evaluated with chemiluminescence, fuel planar laser induced fluorescence and Mie scattering measurements at 100 kHz. The detonation wave significantly impacts the liquid jet trajectory with its deflection in both windward and leeward direction as the adverse pressure gradient across it changes between wave passages. The maximum recovery height of the liquid jet is observed to be consistent across all operating conditions, and dependent on the detonation wave strength in the chamber.

propulsion↗
NASA NTRSDate not supplied

Characterization of Large Drop Velocity in the NASA Icing Research Tunnel

This paper presents experimental work conducted in the Icing Research Tunnel at NASA Glenn Research Center to characterize the velocity of large drops in the test section. Some icing spray clouds with large drops were generated with Mod1 nozzles at low nozzle air pressure of 2 to 4 psig for various tunnel air speeds. Drop diameters and drop velocities were measured via high-resolution imaging with a Particle Imaging Particle Tracking Velocimetry probe developed by Artium Technologies. The probe was mounted at four different locations aligned with the centerline of the test section from near the end of the contraction to the constant height test section part of the tunnel. CFD analyses were performed. It showed that the probe head geometry affects the local air flow in the measurement area of the probe between the prongs and in front of the probe. Initial analysis of the air velocity data during the test also indicated that the probe mounting stand has blockage effect on the local tunnel air velocity measurement by a pitot-static probe affixed on the mounting plate next to the stand. Those findings were later verified in the Icing Research Tunnel using a new pitot-static probe design with a linear motion system to measure the local tunnel air velocity with and without the probe. As a result, additional drop trajectory simulations were performed with the airflow moving towards the probe head with a Langmuir-D 7-bin distribution cloud containing large drops. The simulation results helped identify a critical drop-size threshold of 300 μm above which the velocities of larger drops are negligibly affected by the adverse pressure gradient generated by the probe head due to their large drop inertia. From the dimensional analysis of the drop velocity measurement data obtained, it showed that at the tunnel test section reference location a generalized empirical correlation was developed for the non-dimensional drop velocity as a function of the non-dimensional corrected drop diameter independent of the actual spraybar pressure settings and the tunnel air speeds. The generalized curve-fit correlation showed that the drop velocity was universally asymptotic to about 86 percent of the corresponding tunnel air speed at the test section reference location for the largest drop diameter captured by the probe. Further evaluation of this correlation is recommended to assess its applicability for Supercooled Large Drop icing scaling applications in the Icing Research Tunnel.

SLD↗
NASA NTRSDate not supplied

Characterization of Large Drop Velocity in the NASA Icing Research Tunnel

This presentation presents experimental work conducted in the Icing Research Tunnel at NASA Glenn Research Center to characterize the velocity of large drops in the tunnel test section. Some icing spray clouds with large-sized drops were generated with Mod1 nozzles at low nozzle air pressure of 2 to 4 psig for various tunnel air speeds. Drop diameters and drop velocities were measured via high-resolution imaging with a Particle Imaging Particle Tracking Velocimetry probe developed by Artium Technologies. The probe was mounted at four different locations aligned with the centerline of the test section from near the end of the contraction to the constant height test section part of the tunnel. CFD analyses were performed. It showed that the probe head geometry affects the local air flow between the prongs and in front of the probe. Initial analysis of the air velocity data during the test also indicated that the probe mounting stand has blockage effect on the local tunnel air velocity measurement by a Pitot static probe affixed on the mounting plate. Those findings were later verified in the Icing Research Tunnel using a new Pitot probe design with a linear motion system to measure the local tunnel air velocity with and without the probe. As a result, additional drop trajectory simulations as airflow moving towards the probe head were run for a Langmuir-D 7-bin drop size distribution of a spray cloud with a nominally large value of medium volumetric diameter. The simulation results helped identify a critical drop-size threshold of 300 µm above which the velocities of larger drops are not affected by the adverse pressure gradient generated by the probe head due to their large drop inertia. From the dimensional analysis of the drop velocity measurement data obtained, it showed that at the tunnel test section reference location a generalized empirical correlation was developed for the non-dimensional drop velocity as a function of the non-dimensional corrected drop diameter independent of the actual spraybar pressure settings and the tunnel air speeds. The generalized curve-fit correlation showed that the drop velocity was universally asymptotic to about 86 percent of the corresponding tunnel air speed at the test section reference location for the largest drop diameter captured by the probe. Further evaluation of this correlation is recommended to assess its applicability for Supercooled Large Drop icing scaling applications in the Icing Research Tunnel.

SLD↗
DOE OSTI2026

Validation of OpenPronghorn for Periodic Hill Flow Separation

OpenPronghorn is an open-source, MOOSE-based thermal-hydraulics simulation tool used for advanced reactor analysis. As an open-source code, it offers a transparent framework for validating governing equations, assumptions, and numerical methods against established benchmarks. This study evaluates OpenPronghorn's RANS turbulence model against the ERCOFTAC Case 81 periodic hill benchmark, a standard test case for separated turbulent flow featuring curved-wall separation, recirculation, shear-layer development, and reattachment. Streamwise velocity profiles predicted by OpenPronghorn were compared to reference LES data at multiple x/h locations. Results show that OpenPronghorn captures the overall trend of the velocity profiles, but the largest discrepancies occur in the separated-flow region, where turbulence is highly anisotropic and strongly affected by adverse pressure gradients and wall curvature—conditions that are inherently difficult for standard RANS models to resolve. Future work will test alternative k-e model variants and correction terms to improve prediction accuracy in this region.

42 - ENGINEERING↗
NASA NTRS2014

Flight Tests of a Supersonic Natural Laminar Flow Airfoil

IR thermography was used to characterize the transition front on a S-NLF test article at chord Reynolds numbers in excess of 30 million Changes in transition due to Mach number, Reynolds number, and surface roughness were investigated - Regions of laminar flow in excess of 80% chord at chord Reynolds numbers greater than 14 million IR thermography clearly showed the transition front and other flow features such as shock waves impinging upon the surface A series of parallel oblique shocks, of yet unknown origin, were found to cause premature transition at higher Reynolds numbers. NASA has a current goal to eliminate barriers to the development of practical supersonic transport aircraft Drag reduction through the use of supersonic natural laminar flow (S-NLF) is currently being explored as a means of increasing aerodynamic efficiency - Tradeoffs work best for business jet class at M<2 Conventional high-speed designs minimize inviscid drag at the expense of viscous drag - Existence of strong spanwise pressure gradient leads to crossflow (CF) while adverse chordwise pressure gradients amplifies and Tollmien-Schlichting (TS) instabilities Aerion Corporation has patented a S-NLF wing design (US Patent No. 5322242) - Low sweep to control CF - dp/dx < 0 on both wing surfaces to stabilize TS - Thin wing with sharp leading edge to minimize wave drag increase due to reduction in sweep NASA and Aerion have partnered to study S-NLF since 1999 Series of S-NLF experiments flown on the NASA F-15B research test bed airplane Infrared (IR) thermography used to characterize transition - Non-intrusive, global, good spatial resolution - Captures significant flow features well

laminar flow↗
NASA NTRS1972

Analytical and experimental investigation of circulation control by means of a turbulent Coanda jet

An analytical and experimental investigation of circulation control on a circular cylinder by means of tangential blowing (Coanda effect) is presented. The analytical method developed has also been used to estimate the blowing coefficients required for achieving potential flow on airfoils with flaps. The analysis is presented for conditions for which the flow in the boundary layer ahead of the jet exit is turbulent. The turbulent boundary layer and the jet layer on the upper surface, and the turbulent boundary layer on the lower surface are computed by a multi-strip integral method. The region of integration is between the correponding transition and separation points on each surface. Longitudinal curvature effects, which give rise to a radial pressure gradient across the jet layer and to an additional adverse tangential pressure gradient just upstream of the separation point, are included in the jet layer analysis in an approximate manner. The longitudinal curvature effect is found to have a pronounced influence on the separation of the jet layer.

Levinsky, E. S.↗
DOE OSTI2023

A universal velocity transformation for boundary layers with pressure gradients

The logarithmic law of the wall does not capture the mean flow when a boundary layer is subjected to a strong pressure gradient. In such a boundary layer, the mean flow is affected by the spatio-temporal history of the imposed pressure gradient; and accounting for history effects remains a challenge. This work aims to develop a universal mean flow scaling for boundary layers subjected to arbitrary adverse or/and favourable pressure gradients. Here, we derive from the Navier–Stokes equation a velocity transformation that accounts for the history effects and maps the mean flow to the canonical law of the wall. The transformation is tested against channel flows with a suddenly imposed adverse or favourable pressure gradient, boundary layer flows subjected to an adverse pressure gradient, and Couette–Poiseuille flows with a streamwise pressure gradient. It is found that the transformed velocity profiles follow closely the equilibrium law of the wall.

97 MATHEMATICS AND COMPUTING↗
NASA NTRS1994

Large-eddy simulation of flow through a plane, asymmetric diffuser

A challenge for traditional turbulence modeling, based on the Reynolds averaged Navier-Stokes equations, remains the accurate prediction of 'mild', adverse pressure-gradient driven separation from a smooth surface. With this study we want to explore the capability of large-eddy simulation to predict the separation which occurs on the deflected wall of an asymmetric, plane diffuser with opening angle of 10 deg. The flow through the plane diffuser exhibits some additional interesting physical phenomena which make it a challenging test case. In addition to 'mild' separation about halfway down the deflected ramp, the flow is characterized by a small backflow zone with stalled fluid in the rear part of the expanding section. The turbulent flow entering the diffuser is subject to combined adverse and radial pressure gradients stemming from the convex curvature. Finally the flow recovers into a developed, turbulent channel flow in the outlet section. Obi et al. provide measurements of mean flow, Reynolds stresses, and pressure recovery, which were obtained by means of LDV in a wind tunnel. The objective of this study is to investigate whether LES with the standard dynamic model is able to accurately predict the flow in the one-sided diffuser and to explore the resolution requirements and associated costs.

Kaltenbach, Hans-Jakob↗
NASA NTRS1992

Navier-Stokes analysis and experimental data comparison of compressible flow within ducts

Many aircraft employ ducts with centerline curvature or changing cross-sectional shape to join the engine with inlet and exhaust components. S-ducts convey air to the engine compressor from the intake and often decelerate the flow to achieve an acceptable Mach number at the engine compressor by increasing the cross-sectional area downstream. Circular-to-rectangular transition ducts are used on aircraft with rectangular exhaust nozzles to connect the engine and nozzle. To achieve maximum engine performance, the ducts should minimize flow total pressure loss and total pressure distortion at the duct exit. Changes in the curvature of the duct centerline or the duct cross-sectional shape give rise to streamline curvature which causes cross stream pressure gradients. Secondary flows can be caused by deflection of the transverse vorticity component of the boundary layer. This vortex tilting results in counter-rotating vortices. Additionally, the adverse streamwise pressure gradient caused by increasing cross-sectional area can lead to flow separation. Vortex pairs have been observed in the exit planes of both duct types. These vortices are due to secondary flows induced by pressure gradients resulting from streamline curvature. Regions of low total pressure are produced when the vortices convect boundary layer fluid into the main flow. The purpose of the present study is to predict the measured flow field in a diffusing S-duct and a circular-to-rectangular transition duct with a full Navier-Stokes computer program, PARC3D, and to compare the numerical predictions with new detailed experimental measurements. The work was undertaken to extend previous studies and to provide additional CFD validation data needed to help model flows with strong secondary flow and boundary layer separation. The S-duct computation extends the study of Smith et al, and Harloff et al, which concluded that the computation might be improved by using a finer grid and more advanced turbulence models. The present study compares results for both the Baldwin-Lomas and k-epsilon turbulence models and is conducted with a refined grid. For the transition duct, two inlet conditions were considered, the first with straight flow and the second with swirling flow. The first case permits examination of the effects of the geometric transition on the flow field, while the second case includes the rotational flow effect characteristic of a gas turbine engine.

Harloff, G. J.↗
NASA NTRS1987

An efficient iteration strategy for upwind/relaxation solutions to the thin-layer Navier-Stokes equations

A previously developed upwind/relaxation algorithm for solving the unsteady, compressible, thin-layer Navier-Stokes equations is presently modified so that the downstream influence of the subsonic part of the boundary layer in an otherwise supersonic flow is suppressed by restricting the streamwise pressure gradient. A 'parabolized' solution is then efficiently obtained by marching downstream and iterating locally in each crossflow plane until achieving convergence. This parabolized solution is an excellent final one for problems without large adverse streamwise pressure gradients.

Newsome, Richard W.↗
NASA NTRS1984

An experimental study of a supercritical trailing-edge flow

An experimental study has been conducted of a transonic, turbulent, high-Reynolds-number blunt trailing-edge flow. The model shape and the surface pressure distribution are characteristics of a modern supercritical airfoil under shock-free conditions. Reynolds number and pressure gradient scaling of the boundary layer are relevant to airfoil applications. The data set is exceptionally accurate and consistent, with the momentum balance accounting for the flux of momentum to within 1 percent, except in the immediate vicinity of the blunt trailing edge. The experimental flow exhibits strong viscous-inviscid interaction and higher-order boundary-layer effects including strong adverse streamwise pressure gradient, significant normal pressure gradients associated with surface and streamline curvature, and significant wake curvature. Navier-Stokes calculations with a two-equation K-epsilon turbulence model predict the correct pressure distribution which demonstrates the utility of these engineering tools. The experiment approaches separation at the strailing edge. However, in comparison to the experiment, the calculations predict too high skin friction and insufficient displacement thickness growth. An analysis of the turbulent and mean flow fields reveals the turbulence model defects are likely in modeling the dissipation source and sink terms, and in the eddy viscosity relation.

Brown, J. L.↗