Inlet Endwall Boundary Layer Research Articles

Prediction of the Influence of the Inlet End-Wall Boundary Layer on the Secondary Flow of a Low Pressure Turbine Airfoil Using RANS and Large-Eddy Simulations

Abstract This paper analyzes the effect of the inlet end-wall boundary layer on the secondary flow of a low pressure turbine airfoil cascade at Reynolds number 2 × 105 using RANS and implicit large-eddy simulations (LES). The results are compared against experimental data obtained at two low-speed linear cascade facilities, one located at the Whittle Laboratory of the University of Cambridge and the other at the Polytechnic University of Madrid. The RANS turbulence model is the k−ω−γ−Reθt and no sub-grid scale model has been used in the LES. An unstructured mesh of hexahedra and prisms is used, with high order elements used in the boundary layer region to better describe the airfoil shape in the LES. Two inlet end-wall boundary layers that produce different secondary flow patterns are analyzed: a laminar thin velocity profile and a turbulent thick velocity profile with several inlet turbulent intensities. The agreement between LES numerical predictions and experimental measurements of the position and intensity of the secondary vortices is very good for both cases. RANS simulations are much cheaper in terms of computational cost and reasonably predict most of the flow features, except when the inlet turbulence is low and turbulent transition prediction becomes critical. The effect of the inlet velocity profiles and inlet turbulence on the secondary flow structure is quite pronounced. The velocity profile thickness determines the spanwise penetration of the passage vortex, and different inlet turbulence intensities modify its mixing. Higher inlet turbulence intensities lead to a decrease of the secondary losses due to the passage vortex and an increase of end-wall losses.

The Effects of Inlet Boundary Layer Condition and Periodically Incoming Wakes on Secondary Flow in a Low Pressure Turbine Cascade

AbstractA particular turbine cascade design is presented with the goal of providing a basis for high quality investigations of endwall flow under high-speed conditions with unsteady inflow. The key feature of the design is an integrated two-part flat plate serving as a cascade endwall at part-span, which enables a variation of the inlet endwall boundary layer conditions. The new design is applied to the T106A low pressure turbine cascade for endwall flow investigations in the High-Speed Cascade Wind Tunnel of the Institute of Jet Propulsion at the Bundeswehr University Munich. Measurements are conducted under realistic flow conditions (M2th = 0.59, Re2th = 2 · 105) in three cases of varying endwall boundary layer conditions with and without periodically incoming wakes. The endwall boundary layer is characterized by 1D-CTA measurements upstream of the blade passage. Secondary flow is evaluated by five-hole-probe measurements in the turbine exit flow. A strong similarity is found between the time-averaged effects of unsteady inflow conditions and the effects of changing inlet endwall boundary layer conditions regarding the attenuation of secondary flow. Furthermore, the experimental investigations show that all design goals for the improved T106A cascade are met.

Validation Studies of Linear Oscillating Compressor Cascade and Use of Influence Coefficient Method

Abstract Advanced predictions of blade flutter have been continually pursued. It is noted however that validation cases of unsteady CFD methods against experimental cases with detailed 3D unsteady pressures are still rather lacking. The main objectives of the present work are two-folds. First, validate and understand the characteristics of blade tip clearance, as well as a bubble-type flow separation for an unsteady CFD solver against a 3D oscillating cascade experiment. And second, examine the applicability of the influence coefficient method (ICM) as widely used in an oscillating linear cascade setup. In the first part, the capability of a widely used commercial solver (CFX) for unsteady flows induced by a 3D oscillating compressor cascade is examined. The present computations have shown consistently a destabilizing effect of increasing blade tip clearance, in agreement with the experiment. More remarkably, the computational analyses reveal a distinctive interplay between the inlet endwall boundary layer and the tip clearance in relation to the aerodynamic damping. Different inlet endwall boundary layer thicknesses are shown to lead to qualitatively different aeroelastic stability characteristics in relation to tip clearance. The aero-damping variation with the tip clearance under the influence of the inlet endwall boundary layer seems to correlate closely to a balancing act between the passage vortex and the tip leakage vortex. The tip clearance aeroelastic behavior seems also in line with a simple quasi-steady analysis. On the other hand, the mid-chord laminar bubble separation on suction surface, though with a clear signature in the local aero-damping, has negligible effects on the overall stability. The second part aims to examine computationally the applicability of the influence coefficient method in a linear cascade setup. The comparison between the cascade-based ICM data and a baseline “tuned cascade” shows that the differences in the sensitivity to the far-field treatment can be significant, depending on inter-blade phase angles. On the other hand, non-linearity effects closely relevant to the basic linear assumption of the ICM are shown to only have a small influence. The present results suggest that extra caution should be exercised when comparing a CFD-based tuned cascade model with a finite cascade-based ICM model, at conditions close to acoustic resonance. The resultant discrepancies may well arise from the inherently different far-field sensitivities between the two models, rather than those typical numerical and physical modeling aspects of interest (e.g., meshing, spatial and temporal discretization errors as well as turbulence modeling).

Numerical aero-thermal study of high-pressure turbine nozzle guide vane: Effects of inflow conditions

Accurate predictability of high-pressure turbine nozzle guide vane aero-thermal performance is highly desired in the development campaign due to the exposure of the component to a frequent and high heat load. In this paper, the representative vane profile in modern aero-engines is numerically studied. Aerodynamics and aero-thermal validations of the blade profile have been performed in comparison with the available experimental data. It has been shown that a satisfactory agreement could be achieved with the use of the transitional turbulence model shear stress transport γ–θ due to its superiority in capturing the laminar–turbulent transition. Sensitivity studies on the increase in the inlet turbulence intensity, inlet endwall boundary layer thickness, and inlet total temperature profile have been performed to understand the impact of inflow conditions’ uncertainty on the aero-thermal predictability. Increasing the inlet turbulence intensity increases the pressure surface heat transfer coefficient and induces an earlier transition onset on the suction surface. Due to the rapid decay of turbulence intensity in the numerical model, the use of an artificially high inlet turbulence intensity has been shown to be effective in the prediction improvement. On the other hand, the change in the inlet boundary layer thickness influences the formation and strength of the secondary flow, namely, horseshoe vortex and passage vortex. These secondary flow phenomena affect the local blade surface heat transfer coefficient in the near-endwall region although the most significant rise in heat transfer is found on the endwall. The temperature distortion amplitude of a hot streak and its relative clocking position with the vane significantly affect the heat flux distribution. In contrast, the heat transfer coefficient is less sensitive to the change in hot streak conditions. However, it has been shown that increasing the temperature distortion amplitude could induce a larger difference among different clocking configurations. In addition, decreasing the difference between the fluid and wall temperature would delay the transition onset and stabilize the boundary layer. Further analysis of the unsteady effects has been carried out by comparing the steady and time-averaged flow solutions. It has been observed that the discrepancy between these solutions is attributed to the flow field nonlinearity. Thus, a significant discrepancy can be found in the laminar–turbulent transition as well as in the trailing edge region. However, since the contribution of these regions on the total area-averaged heat transfer is small, their influence on the total vane heat transfer is limited.

Secondary Flow Control in Low Aspect Ratio Vanes Using Splitters

Low aspect ratio vanes, often the result of overall engine architecture constraints, create strong secondary flows and high end-wall loss. In this paper, a splitter concept is demonstrated that reduces secondary flow strength and improves stage performance. An analytic conceptual study, corroborated by inviscid computations, shows that the total secondary kinetic energy (SKE) of the secondary flow vortices is reduced when the number of passages is increased and, for a given number of vanes, when the inlet end-wall boundary layer is evenly distributed between the passages. Viscous computations show that, for this to be achieved in a splitter configuration, the pressure-side leg of the low aspect ratio vane horseshoe vortex, must enter the adjacent passage (and not “jump” in front of the splitter leading edge). For a target turbine application, four vane designs were produced using a multi-objective optimization approach. These designs represent current practice for a low aspect ratio vane, a design exempt from thickness constraints, and two designs incorporating splitter vanes. Each geometry is tested experimentally, as a sector, within a low-speed turbine stage. The vane designs with splitter geometries were found to reduce the measured secondary kinetic energy, by up to 85%, to a value similar to the design exempt from thickness constraints. The resulting flow field was also more uniform in both the circumferential and radial directions. One splitter design was selected for a full annulus test where a mixed-out loss reduction, compared to the current practice design, of 15.3% was measured and the stage efficiency increased by 0.88%.

Effects of Inlet Turbulence and End-Wall Boundary Layer on Aerothermal Performance of a Transonic Turbine Blade Tip

Most of the previous researches of inlet turbulence effects on blade tip have been carried out for low speed situations. Recent work has indicated that for a transonic turbine tip, turbulent diffusion tends to have a distinctively different impact on tip heat transfer than for its subsonic counterpart. It is hence of interest to examine how inlet turbulence flow conditioning would affect heat transfer characteristics for a transonic tip. The present work is aimed to identify and understand the effects of both inlet freestream turbulence and end wall boundary layer on a transonic turbine blade tip aerothermal performance. Spatially-resolved heat transfer data are obtained at aerodynamic conditions representative of a high-pressure turbine, using the transient infrared thermography technique with the Oxford High-Speed Linear Cascade research facility. With and without turbulence grids, the turbulence levels achieved are 7%–9% and 1%, respectively. On the blade tip surface, no apparent change in heat transfer was observed with high and low inlet turbulence intensity levels investigated. On the blade suction surface, however, substantially different local heat transfer distributions for the suction side near tip surface have been observed, indicating a strong local dependence of the local vortical flow structure on the freestream turbulence. These experimentally observed trends have also been confirmed by CFD examinations using the Rolls-Royce HYDRA. A further CFD analysis suggests that the level of inflow turbulence alters the balance between the passage vortex associated secondary flow and the over tip leakage (OTL) flow. Consequently, an enhanced inertia of near wall fluid at a higher inflow turbulence weakens the cross-passage flow. As such, the weaker passage vortex leads the tip leakage vortex to move further into the mid passage, with the less spanwise coverage on the suction surface, as consistently indicated by the heat transfer signature. Different inlet end wall boundary layer profiles are employed in the computational study with HYDRA. All CFD results indicate the inlet boundary layer thickness has little impact on the heat transfer over the tip surface as well as the pressure side near-tip surface. However, noticeable changes in heat transfer are observed for the suction side near-tip surface. Similar to the inlet turbulence effect, such changes can be attributed to the interaction between the passage vortex and the OTL flow.

Measurements of Losses and Reynolds Stresses in the Secondary Flow Downstream of a Low-Speed Linear Turbine Cascade

Experimental measurements of the mean and turbulent flow field were preformed downstream of a low-speed linear turbine cascade. The influence of turbulence on the production of secondary losses is examined. Steady pressure measurements were collected using a seven-hole pressure probe and the turbulent flow quantities were measured using a rotatable x-type hotwire probe. Each probe was traversed downstream of the cascade along planes positioned at three axial locations: 100%, 120%, and 140% of the axial chord (Cx) downstream of the leading edge. The seven-hole pressure probe was used to determine the local total and static pressure as well as the three mean velocity components. The rotatable x-type hotwire probe, in addition to the mean velocity components, provided the local Reynolds stresses and the turbulent kinetic energy. The axial development of the secondary losses is examined in relation to the rate at which mean kinetic energy is transferred to turbulent kinetic energy. In general, losses are generated as a result of the mean flow dissipating kinetic energy through the action of viscosity. The production of turbulence can be considered a preliminary step in this process. The measured total pressure contours from the three axial locations (1.00, 1.20, and 1.40Cx) demonstrate the development of the secondary losses. The peak loss core in each plane consists mainly of low momentum fluid that originates from the inlet endwall boundary layer. There are, however, additional losses generated as the flow mixes with downstream distance. These losses have been found to relate to the turbulent Reynolds stresses. An examination of the turbulent deformation work term demonstrates a mechanism of loss generation in the secondary flow region. The importance of the Reynolds shear stresses to this process is explored in detail.

1단 터빈 내 앞전 변형의 영향 하에 공력 특성에 대한 비정상 수치해석적 연구

The important problems that arise in the design and performance of the axial flow turbine are the prediction and control of secondary flows. Some progresses have been made on understanding flow conditions that occur when the inlet endwall boundary layer separates at the point in the endwall and rolls up into the horseshoe vortex. And the flows though an axial turbine tend to be extremely complex due to its inherent unsteady and viscous phenomena. The passing wakes generated from the trailing edge of the stator make an interaction with the rotor. Unsteady flow should be considered rotor/stator interactions. The main purpose of this research is control of secondary flow and improvement efficiency in turbine by leading edge modification in unsteady state. When the wake from the stator ran into the modified leading edge of the rotor, the leading edge generated the weak pressure fluctuation by complex passage flows. In conclusion, leading edge modification(bulb2) results in the reduced total pressure loss in the flow field.

Endwall Boundary Layer Development in an Engine Representative Four-Stage Low Pressure Turbine Rig

Research by de la Rosa Blanco et al. (“Influence of the State of the Inlet Endwall Boundary Layer on the Interaction Between the Pressure Surface Separation and the Endwall Flows,” Proc. Inst. Mech. Eng., Part A, 217, pp. 433–441) in a linear cascade of low pressure turbine (LPT) blades has shown that the position and strength of the vortices forming the endwall flows depend on the state of the inlet endwall boundary layer, i.e., whether it is laminar or turbulent. This determines, amongst other effects, the location where the inlet boundary layer rolls up into a passage vortex, the amount of fluid that is entrained into the passage vortex, and the interaction of the vortex with the pressure side separation bubble. As a consequence, the mass-averaged stagnation pressure loss and therefore the design of a LPT depend on the state of the inlet endwall boundary layer. Unfortunately, the state of the boundary layer along the hub and casing under realistic engine conditions is not known. The results presented in this paper are taken from hot-film measurements performed on the casing of the fourth stage of the nozzle guide vanes of the cold flow affordable near term low emission (ANTLE) LPT rig. These results are compared with those from a low speed linear cascade of similar LPT blades. In the four-stage LPT rig, a transitional boundary layer has been found on the platforms upstream of the leading edge of the blades. The boundary layer is more turbulent near the leading edge of the blade and for higher Reynolds numbers. Within the passage, for both the cold flow four-stage rig and the low speed linear cascade, the new inlet boundary layer formed behind the pressure leg of the horseshoe vortex is a transitional boundary layer. The transition process progresses from the pressure to the suction surface of the passage in the direction of the secondary flow.

Effect of the Leakage Flows and the Upstream Platform Geometry on the Endwall Flows of a Turbine Cascade

This work describes the effect that the injection of leakage flow from a cavity into the mainstream has on the endwall flows and their interaction with a large pressure surface separation bubble in a low-pressure turbine. The effect of a step in hub diameter ahead of the blade row is also simulated. The blade profile under consideration is a typical design of modern low-pressure turbines. The tests are conducted in a low speed linear cascade. These are complemented by numerical simulations. Two different step geometries are investigated, i.e., a backward-facing step and a forward-facing step. The leakage tangential velocity and the leakage mass flow rate are also modified. It was found that the injection of leakage mass flow gives rise to a strengthening of the endwall flows independently of the leakage mass flow rate and the leakage tangential velocity. The experimental results have shown that below a critical value of the leakage tangential velocity, the net mixed-out endwall losses are not significantly altered by a change in the leakage tangential velocity. For these cases, the effect of the leakage mass flow is confined to the wall, as the inlet endwall boundary layer is pushed further away from the wall by the leakage flow. However, for values of the leakage tangential velocity around 100% of the wheel speed, there is a large increase in losses due to a stronger interaction between the endwall flows and the leakage mass flow. This gives rise to a change in the endwall flows’ structure. In all cases, the presence of a forward-facing step produces a strengthening of the endwall flows and an increase of the net mixed-out endwall losses when compared with a backward-facing step. This is because of a strong interaction with the pressure surface separation bubble.

Influence of the state of the inlet endwall boundary layer on the interaction between pressure surface separation and endwall flows

This paper describes the effect of the state of the inlet boundary layer (laminar or turbulent) on the structure of the endwall flow on two different profiles of low-pressure (LP) turbine blades (solid thin and hollow thick). At present the state of the endwall boundary layer at the inlet of a real LP turbine is not known. The intention of this paper is to show that, for different designs of LP turbine, the state of the inlet boundary layer affects the performance of the blade in very different ways. The testing was completed at low speed in a linear cascade using area traversing, flow visualization and static pressure measurements. The paper shows that, for a laminar inlet boundary layer the two profiles have a similar loss distribution and structure of endwall flow. However, for a turbulent inlet boundary layer the two profiles are shown to differ significantly in both the total loss and endwall flow structure. The pressure side separation bubble on the solid thin profile is shown to interact with the passage vortex, causing a higher endwall loss than that measured on the hollow thick profile.

Mainstream Aerodynamic Effects Due to Wheelspace Coolant Injection in a High-Pressure Turbine Stage: Part I—Aerodynamic Measurements in the Stationary Frame

The relative aerodynamic and performance effects associated with rotor–NGV gap coolant injections were investigated in the Axial Flow Turbine Research Facility (AFTRF) of the Pennsylvania State University. This study quantifies the effects of the coolant injection on the aerodynamic performance of the turbine for radial cooling, impingement cooling in the wheelspace cavity and root injection. Overall, it was found that even a small quantity (1 percent) of cooling air can have significant effects on the performance character and exit conditions of the high pressure stage. Parameters such as the total-to-total efficiency, total pressure loss coefficient, and three-dimensional velocity field show local changes in excess of 5, 2, and 15 percent, respectively. It is clear that the cooling air disturbs the inlet end-wall boundary layer to the rotor and modifies secondary flow development, thereby resulting in large changes in turbine exit conditions.

Mainstream Aerodynamic Effects Due to Wheelspace Coolant Injection in a High-Pressure Turbine Stage: Part II—Aerodynamic Measurements in the Rotational Frame

The current paper deals with the aerodynamic measurements in the rotational frame of reference of the Axial Flow Turbine Research Facility (AFTRF) at the Pennsylvania State University. Stationary frame measurements of “Mainstream Aerodynamic Effects Due to Wheelspace Coolant Injection in a High Pressure Turbine Stage” were presented in Part I of this paper. The relative aerodynamic effects associated with rotor–nozzle guide vane (NGV) gap coolant injections were investigated in the rotating frame. Three-dimensional velocity vectors including exit flow angles were measured at the rotor exit. This study quantifies the secondary effects of the coolant injection on the aerodynamic and performance character of the stage main stream flow for root injection, radial cooling, and impingement cooling. Current measurements show that even a small quantity (1 percent) of cooling air can have significant effects on the performance and exit conditions of the high-pressure turbine stage. Parameters such as the total pressure coefficient, wake width, and three-dimensional velocity field show significant local changes. It is clear that the cooling air disturbs the inlet end-wall boundary layer to the rotor and modifies secondary flow development thereby resulting in large changes in turbine exit conditions. Effects are the strongest from the hub to midspan. Negligible effect of the cooling flow can be seen in the tip region.