Suction Side Boundary Layer Research Articles

Experimental Study of Surface Roughness Effects on a Turbine Airfoil in a Linear Cascade— Part II: Aerodynamic Losses

Abstract The present experimental study is part of a comprehensive analysis accounting for heat transfer and aerodynamic losses on a highly loaded low pressure turbine blade with varying surface roughness. Whereas Part I focuses on heat transfer measurements at airfoil midspan with different deterministic surface roughnesses, Part II investigates surface roughness effects on aerodynamic losses of the same airfoil. A set of different arrays of deterministic roughness (the same as used in Part I) is investigated in these experiments. The height and eccentricity of the roughness elements are varied, showing the combined influence of roughness height and anisotropy on the losses produced in the boundary layers. It is shown that the boundary layer loss is dominated by the suction side. Therefore, the investigations focus on measurements of the suction side boundary layer thickness at midspan directly upstream of the trailing edge. The experiments are conducted at several freestream turbulence levels (Tu1=1.4–10.1%) and different Reynolds numbers. The measurements reveal that suction side boundary layer thickness is increased by up to 190% if surface roughness shifts the transition onset upstream. However, in some cases, at low Reynolds numbers and freestream turbulence, surface roughness suppresses boundary layer separation and decreases the trailing edge boundary layer thickness by up to 30%.

Large Eddy Simulation of a Controlled Diffusion Compressor Cascade

In this research a Controlled Diffusion (CD) compressor cascade stator blade is simulated at a Reynolds number of ∼700,000, based on inflow velocity and chord length, using Large Eddy Simulation (LES). A wide range of flow inlet angles are computed, including conditions near the design angle, and at high negative and positive incidence. At all inlet angles the surface pressure distributions are well-predicted by the LES. Near the design angle the computed suction side boundary layer thickness agrees well with experimental data, whilst the pressure side boundary layer is poorly predicted due to the inability of LES to capture natural boundary layer transition on the present grid. A good estimation of the loss is computed near the design angle, whilst at both high positive and negative incidences the loss is less well predicted owing to discrepancies between the computed and experimental boundary layer thickness. At incidences above the design angle a laminar separation bubble forms near the leading edge of the suction surface, which undergoes a transition to turbulence. Similar behaviour is noted on the pressure surface at negative incidence. At high negative incidence contra-rotating vortex pairs are found to form around the leading edge in response to an unsteady stagnation line across the span of the blade. Such structures are not apparent in time-averaged statistical data due to their highly-transient nature.

Advanced Nonaxisymmetric Endwall Contouring for Axial Compressors by Generating an Aerodynamic Separator—Part II: Experimental and Numerical Cascade Investigation

Modern methods for axial compressor design are capable of shaping the blade surfaces in a three-dimensional way. Linking these methods with automated optimization techniques provides a major benefit to the design process. The application of nonaxisymmetric contoured endwalls is considered to be very successful in turbine rotors and vanes. Concerning axial compressors, nonaxisymmetric endwalls are still a field of research. This two-part paper presents the recent development of a novel endwall design. A vortex created by a nonaxisymmetric endwall groove acts as an aerodynamic separator, preventing the passage vortex from interacting with the suction side boundary layer. This major impact on the secondary flow results in a significant loss reduction by means of load redistribution, reduction in recirculation areas, and suppressed corner separation. Part I of this paper deals with the endwall design and its compressor application. The resulting flow phenomena and physics are described and analyzed in detail. The second paper presents the detailed experimental and numerical investigation of the developed endwall groove. The measurements carried out at the transonic cascade wind tunnel of DLR in Cologne, demonstrated a considerable influence on the cascade performance. A loss reduction and redistribution of the cascade loading were achieved at the aerodynamic design point, as well as near the stall condition of the cascade. This behavior is well predicted by the numerical simulation. The combined analysis of experimental and numerical flow patterns allows a detailed interpretation and description of the resulting flow phenomena. In this context, high fidelity 3D-Reynolds-averaged Navier–Stokes flow simulations are required to analyze the complex blade and endwall boundary layer interaction.

Efficiency enhancement in transonic compressor rotor blades using synthetic jets: A numerical investigation

Several passive and active techniques were studied and developed by compressor designers with the aim of improving the aerodynamic behavior of compressor blades by reducing, or even eliminating, flow separation. Fluidic-based methods, in particular, have been investigated for a long time, including both steady and unsteady suction, blowing and oscillating jets. Recently, synthetic jets (zero mass flux) have been proposed as a promising solution to reduce low-momentum fluid regions inside turbomachines. Synthetic jets, with the characteristics of zero net mass flux and non-zero momentum flux, do not require a complex system of pumps and pipes. They could be very efficient because at the suction part of the cycle the low-momentum fluid is sucked into the device, whereas in the blowing part a high-momentum jet accelerates it. To the authors’ knowledge, the use of synthetic jets has never been experimented in transonic compressor rotors, where this technique could be helpful (i) to reduce the thickness and instability of blade suction side boundary layer after the interaction with the shock, and (ii) to delay the arising of the low-momentum region which can take place from the shock-tip clearance vortex interaction at low flow operating conditions, a flow feature which is considered harmful to rotor stability. Therefore, synthetic jets could be helpful to improve both efficiency and stall margin in transonic compressor rotors. In this paper, an accurate and validated CFD model is used to simulate the aerodynamic behavior of a transonic compressor rotor with and without synthetic jets. Four technical solutions were evaluated, different for jet position and velocity, and one was investigated in detail.

Large eddy simulation of free-stream turbulence effects on heat transfer to a high-pressure turbine cascade

Large eddy simulation of an uncooled, transonic, linear high-pressure turbine vane cascade has been pursued using a high-order numerical method with about 83 million grid points. Heat transfer and boundary layer development on the surface of the blade are investigated in the presence of inflow turbulence. Quantitative comparisons with the experiments of Arts et al. (Tech. Note 174, von Karman Institute for Fluid Dynamics, Belgium, 1990) show excellent agreement. The distortion of vortices from the inflow turbulence in the cascade passage leads to the formation of long streamwise streaky structures, leading to transition of suction-side boundary layer. Significant heat transfer augmentation is seen on both the pressure and suction sides because of the presence of these structures.

Radial fluid migration and endwall blockage in axial flow rotors

The article presents an analytical model for systematic investigation of the effects stimulating or retarding radial outward migration of fluid in the suction side (SS) boundary layer of the blades of axial flow rotors, contributing to increased near-tip loss and promoting tip stalling. With application of the model to an experimental case study, coupled with evaluation of detailed flow measurement data, it has been pointed out that the outward migration and near-tip accumulation of high-loss fluid are intensified by the controlled vortex design (CVD) style. This phenomenon is related to the SS outward flow associated with the vortices shed from the blade of spanwise changing circulation. It has been found that the outward migration, and consequently, the endwall blockage at the casing, becomes more pronounced as the intensity of shed vortices due to CVD increases. Based on experimental data, it has been pointed out that the increase in axial displacement thickness is approximately proportional to the radial gradient of blade circulation. Based on the purposeful use of the analytical model, forward blade sweep was found to be especially beneficial for the rotors of CVD, as a means for moderating outward migration along the suction surface.

Separated Flow Measurements on a Highly Loaded Low-Pressure Turbine Airfoil

Boundary layer separation, transition, and reattachment have been studied on a new, very high lift, low-pressure turbine airfoil. Experiments were done under low freestream turbulence conditions on a linear cascade in a low speed wind tunnel. Pressure surveys on the airfoil surface and downstream total pressure loss surveys were documented. Velocity profiles were acquired in the suction side boundary layer at several streamwise locations using hot-wire anemometry. Cases were considered at Reynolds numbers (based on the suction surface length and the nominal exit velocity from the cascade) ranging from 25,000 to 330,000. In all cases, the boundary layer separated, but at high Reynolds number the separation bubble remained very thin and quickly reattached after transition to turbulence. In the low Reynolds number cases, the boundary layer separated and did not reattach, even when transition occurred. This behavior contrasts with previous research on other airfoils, in which transition, if it occurred, always induced reattachment, regardless of Reynolds number.

Unsteady Simulation of an Axial Compressor Stage With Casing and Blade Passive Treatments

This paper deals with the numerical simulation of technologies to increase the compressor performances. The objective is to extend the stable operating range of an axial compressor stage using passive control devices located in the tip region. First, the behavior of the tip leakage flow is investigated in the compressor without control. The simulation shows an increase in the interaction between the tip leakage flow and the main flow when the mass flow is reduced, a phenomenon responsible for the development of a large flow blockage region at the rotor leading edge. A separation of the rotor suction side boundary layer is also observed at near stall conditions. Then, two approaches are tested in order to control these flows in the tip region. The first one is a casing treatment with nonaxisymmetric slots. The method showed a good ability to control the tip leakage flow but failed to reduce the boundary layer separation on the suction side. However, an increase in the operability was observed but with a penalty for the efficiency. The second approach is a blade treatment that consists of a longitudinal groove built in the tip of each rotor blade. The simulation pointed out that the device is able to control partially all the critical flows with no penalty for the efficiency. Finally, some recommendations for the design of passive treatments are presented.

Experimental Investigation of the Clocking Effect in a 1.5-Stage Axial Turbine—Part II: Unsteady Results and Boundary Layer Behavior

Abstract Comprehensive experimental investigations were conducted to get deeper insight into the physics of stator clocking in turbomachines. Different measurement techniques were used to investigate the influence of varying clocking positions on the highly unsteady flow field in a 1.5-stage axial low-pressure (LP) turbine. A Reynolds number typical for LP turbines as well as a two-dimensional blade design were chosen. Stator 2 was developed as a high-lift profile with a separation bubble on the suction side. This paper presents the results that were obtained by means of unsteady x-wire measurements upstream and downstream of Stator 2 and surface mounted hot-film measurements on the Stator 2 suction side. It was found that for the case when the Stator 1 wakes impinge close to the leading edge of Stator 2 the interaction between the Stator 1 and the rotor vortical structures takes place in proximity of the Stator 2 boundary layer, which leads to a shift of the transition point in the upstream direction. The major loss parameter concerning the Stator 2 aerodynamic performance could be attributed to the strength of the periodic fluctuations within the Stator 2 suction side boundary layer. A phase shift in the quasiwall shear stress signal in the front region of the Stator 2 vane was observed for different clocking positions.

An experimental investigation of the dissipation mechanisms in the suction side boundary layer of a turbine blade

The present work is part of an extensive experimental activity carried out by the authors in recent years aimed at investigating the boundary layer transition phenomenon in turbine blades. The large scale of the cascade and the use of advanced LDV instrumentation and precision probe traversing mechanism resulted in high degree of spatial resolution and high accuracy of measurements. The main dissipation mechanism determining the profile losses in turbomachinery blades is the work of deformation of the mean motion within the boundary layer operated by both viscous and turbulent shear stresses. In the present paper, the local viscous and turbulent deformation works have been directly evaluated from the detailed measurements of boundary layer mean velocity and Reynolds shear stress. The results show the distributions and the relative importance of the viscous and turbulent contributions to the loss production, in relation with the boundary layer states occurring along the turbine profile.

Inner Workings of Shrouded and Unshrouded Transonic Fan Blades

This paper reports on the numerical assessment of the differences in aerodynamic performance between part span shrouded and unshrouded fan blades generally found in the first stage of multistage fans in low bypass ratio aircraft engines. Rotor flow fields for both blade designs were investigated at two operating conditions using a three-dimensional viscous flow analysis. Although designed to the same radius ratio, aspect ratio, and solidity, the unshrouded fan rotor had a slightly increased tip speed (+3%) and somewhat lower pressure ratio (−3.2%) due to engine cycle requirements. Even when allowing for these small differences, the analysis reveals interesting differences in the level and in the radial distribution of efficiency between these two rotors. The reason for the improved performance of the shrouded rotor in part can be attributed to the shroud blocking off the radial migration of boundary layer fluid centrifuged from the hub on the suction side. As a result, the shock boundary layer interaction seems to be improved on the shrouded blade. At the cruise condition, the efficiency is the same for both rotors. The slightly better efficiency of the shrouded blade in the outer panel is nullified by the large efficiency penalty in the vicinity of the shroud. As there is no significant radial migration of fluid in the suction side boundary layer, as indicated by the analysis at this condition relative to the design speed case, the benefit due to the shroud is greatly reduced. At this speed and at lower speeds, the shroud becomes a net additional loss for the blade. Also of interest from the numerical results is the indication that significant blade ruggedization penalties to performance can be reduced in the case of the unshrouded blade through custom tailoring of its mean camber line.

Numerical investigation of the tone noise mechanism over laminar airfoils

This paper presents the first numerical investigation via direct numerical simulation of the tone noise phenomenon occurring in the flow past laminar airfoils. This phenomenon corresponds to the radiation of discrete acoustic tones in some specific flow conditions, and has received much attention since the 1970s, and several experimental studies have been carried out to identify and understand the underlying physical mechanisms. However, several points remain to be clarified in order to provide a complete explanation of its origin. The flow around a two-dimensional NACA0012 airfoil is considered in order to have a deeper understanding of the tone noise phenomenon. Consistently with previous experimental studies, it is shown that depending on the Reynolds number and angle of attack, two different types of acoustic spectrum are observed: one which exhibits a broadband contribution with a dominant frequency together with a sequence of regularly spaced discrete frequencies, while the other one is only characterized by a simple broadband contribution. The first configuration is typical of the tone noise phenomenon. The present work shows that in this case, the mean flow on the pressure side of the airfoil exhibits a separation bubble near the trailing edge and the main tone frequency is close to the most amplified frequency of the boundary layer. The mechanism proposed in previous works for the main tone generation – which implies the existence of a separation bubble at the pressure side – is therefore validated by numerical simulation. On the other hand, the analysis of the suction side boundary layer reveals that there is no separation and that the most amplified frequency is different from the main tonal one. However, the suction side boundary layer is highly receptive to the tone frequency. Finally, an original explanation for the existence of the secondary discrete frequencies observed in the radiated pressure spectrum is given. They are associated to a bifurcation of the airfoil wake from a symmetric to a non-symmetric vortex pattern. A possible explanation for the existence of this bifurcation is the interaction between the disturbances which are the most amplified by the suction side boundary layer and those originating in the forcing of the suction side flow by the main tone noise mechanism.

Modeling of Wake-Induced Transition in Linear Low-Pressure Turbine Cascades

An unsteady Reynolds-averaged Navier-Stokes (URANS) strategy is applied to the problem of wake-induced transition at high freestream turbulence on the suction side of two blades representative of those used in low-pressure turbines. Experimentally, the blades are arranged in high-aspect-ratio linear cascades, with upstream circular bars generating passing wakes, and two-dimensional flow conditions are, thus, assumed. The strategy combines an explicit algebraic Reynolds-stress turbulence model with transition-specific modifications targeted at capturing the effects of high freestream turbulence and of pretransitional laminar fluctuations. Close attention is paid to numerical accuracy, and grids of up to 140,000 cells are used in combination with 800 time steps per pitchwise traverse to resolve small-scale features in the blade boundary layers that are associated with the unsteady interaction. The computational results demonstrate that the combined model returns a good representation of the response of the suction-side boundary layer to the passing wakes in both blades. Specifically, in the boundary layer of one of the two blades, the wakes are observed to cause a periodic upstream shift in the transition onset and, thus, correspondingly periodic attachment and calming. In the other, no separation occurs, and the wakes are shown to produce a significant periodic reduction in shape factor and increase in skin friction in the blade boundary layer, again as a consequence of the upstream shift in the transition location.

BLADEROW INTERACTIONS, TRANSITION, AND HIGH-LIFT AEROFOILS IN LOW-PRESSURE TURBINES

▪ Abstract The flow in turbomachines is unsteady due to the relative motion of the rows of blades. In the low-pressure turbine, the wakes from the upstream bladerows provide the dominant source of unsteadiness. Because much of the blade-surface boundary-layer flow is laminar, one of the most important consequences of this unsteadiness is the interaction of the wakes with the suction-side boundary layer of a downstream blade. This is important because the blade suction–side boundary layers are responsible for most of the loss of efficiency and because the combined effects of random (wake turbulence) and periodic disturbances (wake velocity defect and pressure fields) cause the otherwise laminar boundary layer to undergo transition and eventually become turbulent. This article summarizes the process of wake-induced boundary-layer transition in low-pressure turbines and the loss generation processes that result. Particular emphasis is placed on how the effects of wakes may be exploited to control loss generation and how this has enabled successful development of ultra-high-lift low-pressure turbines.

Large-Eddy Simulation of Flow Around Low-Pressure Turbine Blade with Incoming Wakes

The flow around a low-pressure turbine rotor blade with periodically incoming wakes at a realistic Reynolds number is computed by means of large-eddy simulation (LES). The computed results are discussed in terms of phase-averaged and mean quantities. A comparison is made with an existing direct numerical simulation (DNS) for the same geometry and operating conditions. Particular attention is devoted to flow structures associated with the incoming wakes and their effect on the boundary layers. The analysis of the flowfield reveals patterns similar to those encountered in DNS and in LES of flow in the same geometry at a lower Reynolds number. Noticeable differences occur in the suction-side boundary layer, which exhibits a complete transition to turbulence for the present case

Experimental study of two AIRBUS/ONERA airfoils in near stall conditions. Part I: Boundary layers

The flow around two airfoils in the vicinity of stall has been experimentally investigated in the ONERA/CFM F2 wind tunnel. The aim of this study was to provide a detailed and well established database for CFD validation, in low speed conditions ( M ∞=0.15) for a fixed Reynolds number of 2.1×10 6 based on the chord. The 2D single element models were fitted with wall pressure taps allowing the measurement of pressure distributions and the determination of the lift coefficient. Measurements over a wide range of angles of attack provided a global view of the phenomena. The first airfoil typically exhibits an abrupt stall, while the second one is characterized by a very mild one. The main part of this study was a very careful survey of the suction side boundary layer and near wake, using the F2 LDA system in a 2D configuration. This first part deals with the boundary layer surveys. A companion paper presenting the wake results will be issued later.

Large Eddy Simulation of Flow Around an Airfoil Near Stall

A large eddy simulation (LES) of a turbulent flow past an airfoil near stall at a chord Reynolds number of 2.1 x 10 6 is performed and compared with wind-tunnel experiments. This configuration still constitutes a challenging test case for Reynolds-averaged Navier-Stokes (RANS) simulation and LES as a result of the complexity of the suction side boundary layer: an adverse pressure gradient creates successively a laminar separation bubble, a turbulent reattachment, and a turbulent separation near the trailing edge. To handle this high-Reynolds-number flow with LES on available supercomputers, a local mesh-refinement technique and a discretization of the convective fluxes are developed in a block-structured finite volume code to reduce the total number of grid points and the numerical dissipation acting on the small scales, respectively. Influence of subgrid scale modeling (SGS) is assessed through the comparisons of explicit selective mixed scale model (SMSM) and implicit monotone-integrated LES model results. Moreover, the solution sensitiveness to grid refinement and spanwise extent is investigated

Experimental Investigation of Turbulence Influence of Wake Passing on the Boundary Layer Development of Highly Loaded Turbine Cascade Blades

Experimental investigations regarding the effects of wake passing on boundary layer development were performed on a highly loaded low pressure (LP) turbine cascade blade. Data was obtained for various flow conditions and is intended to be used for the validation of numerical models dealing with periodic unsteady transition. A phase shift between turbulence and velocity fluctuations in the wake path was observed in the near-wall boundary layer, which might have considerable impact on the use of integral-type transition criteria. The experimental data set is already publicly available for download on the web page of the institute. A moving bar type wake generator was employed to simulate the upstream blade row. Tests were performed at turbomachinery-like Mach and Reynolds number conditions in the High-Speed Cascade Wind Tunnel for different bar speeds, bar spacings and background turbulence levels. This paper focuses on the exemplary description of flow phenomena characteristic for unsteady flow for a single exit Reynolds number of Re2th = 200,000. Conventional measurement techniques like static pressure tappings on the suction and pressure side as well as five hole probe traverses were complemented with fast-response pressure transducers embedded into and hot film sensors glued onto the suction surface to study the time-resolved boundary layer development in detail. Additional studies using single hot wire traverses in the suction side boundary layer and triple hot wire probes upstream of the cascade inlet plane were deployed to complete the data sets. The results indicate that for the LP turbine the transition point moves periodically when subjected to wake passing, which greatly affects the loss generation in the suction side boundary layer. Laminar separation is suppressed or reduced whenever the wake is present.

Analysis of DNS and LES of Flow in a Low Pressure Turbine Cascade with Incoming Wakes and Comparison with Experiments

The flow around a low-pressure turbine rotor blade with incoming periodic wakes is computed by means of DNS and LES. The latter adopts a dynamic sub-grid-scale model. The computed results are compared with time-averaged and instantaneous measured quantities. The simulation sreveal the presence of elongated flow structures, stemming from the incoming wake vorticity, which interact with the pressure side boundary layer. As the wake approaches the upstream half of the suction side, its vortical structures are stretched and align with the main flow, resulting in an impingement at virtually zero angle of attack. Periodically, in the absence of impinging wakes, the laminar suction side boundary layer separates in the adverse pressure gradient region. Flow in the laminar separation bubble is found to undergo transition via a Kelvin–Helmholtz instability. Subsequent impingement of the wake inhibits separation and thus promotes boundary layer reattachment. LES provides a fair reproduction of the DNS results both in terms of instantaneous, phase-averaged, and time-averaged flow fields with a considerable reduction in computational effort.

The effect of turbulence intensity and length scale on low-pressure turbine blade aerodynamics

Abstract Unpredicted losses have been observed in low-pressure gas turbine stages during high altitude operation. These losses have been attributed to aerodynamic separation on the turbine blade suction surfaces. To gain insight into boundary layer transition and separation for these low Reynolds number conditions, the heat transfer distribution on a Langston turbine blade shape was measured in a linear cascade wind tunnel for turbulence levels of 0.8% and 10% and Reynolds numbers of 40–80 k . Turbulence levels of 10% were generated using three passive biplanar lattice grids with square-bar widths of 1.27, 2.54 and 6.03 cm to investigate the effect of turbulence length scale. The heat transfer was measured using a uniform heat flux (UHF) liquid crystal technique. As turbulence levels increased, stagnation heat transfer increased and the location of the suction-side boundary layer transition moved upstream toward the blade leading edge. For this turbine blade shape the transition location did not depend on turbulence length scale, the location is more dependent on pressure distribution, Reynolds number and turbulence intensity. For the 10% turbulence cases, the smaller length scales had a larger affect on heat transfer at the stagnation point. A laser tuft method was used to differentiate between boundary layer transition and separation on the suction surface of the blade. Separation was observed for all of the low turbulence (clean tunnel) cases while transition was observed for all of the 10% turbulence cases. Separation and transition locations corresponded to local minimums in heat transfer. Reattachment points did not correspond to local maximums in heat transfer, but instead, the heat transfer coefficient continued to rise downstream of the reattachment point. For the clean tunnel cases, streamwise streaks of varying heat transfer were recorded on the concave pressure side of the turbine blade. These streaks are characteristic of either Gortler vortices or a three-dimensional transition process. For the 10% turbulence cases, these streaks were not present. The results presented in this paper show that turbulence length scale, in addition to intensity have an important contribution to turbine blade aerodynamics and are important to CFD modelers who seek to predict boundary layer behavior in support of turbine blade design optimization efforts.