Endwall Flow Research Articles

Trajectory and effectiveness of film cooling in endwall crossflow

The crossflow, one of the most important flow structures in the turbine endwall, has significant effects on the endwall film cooling. However, there is a lack of quantitative analysis on it due to the complexity of the endwall flow field. In the current study, a turning channel is utilized to generate the crossflow free from vortex structures. Experiments are conducted and numerical results are utilized to supplement flow-field details. Effects of the momentum ratio (I), crossflow intensity, and main-flow Reynolds number (ReD) are analyzed. Theories of three-dimensional boundary layer flow are introduced to help reveal underlying mechanisms. It is observed that coolant trajectories obey linearity in the cylindrical coordinate. Then, the deflection of coolant can be quantified using the straight-line slope, and I is demonstrated to well scale the coolant migration under the crossflow, which guides predictions on coolant trajectories. Reversely, the disturbance of cooling injections on the crossflow is non-negligible, mainly suppressing the crossflow development. Besides, the crossflow enhances the coolant spreading, and improves the cooling performance. ReD has non-negligible effects on the cooling performance by altering the crossflow profile in the boundary layer. Lower ReD leads to a distorted coolant spatial distribution, resulting in a severer main gas/coolant mixture.

Experimental study of cooling air effect on overall cooling of laminated configuration at a turbine vane end-wall

Turbine vane end-wall cooling management is complicated by the sophisticated cascade flow, generating the non-uniform distribution of metal temperature at end-wall surface, a negative factor of component durability. The double-wall laminated cooling has a huge potential to achieve the two-side thermal protection for end-wall, as well as the minimum cooling air amount. However, the design guidance of laminated cooling end-wall is still scarce, since that the reported flat-plate studies ignored the influence of end-wall flow. In present work, the cooling air effect on the metal overall effectiveness was extracted from the infrared thermography measurements under different cooling air flowrates, indicating that there is an utmost of cooling air consumption for the reduction of metal temperature. The two-dimensional contour of effectiveness induced by cooling air was selected as the new design principle. A case study indicated that the modified laminated cooling layout can acquire the uniform distribution of metal temperature over the entire end-wall surface. A comparison of metal overall effectiveness of original and modified end-walls revealed that in leading edge of pressure side, the novel film cooling design is more proper; and the dense layout of impingement-hole with large diameters is effective in the corner of pressure side.

Loss reduction in a high-speed compressor cascade using an L-shaped endwall groove to generate a streamwise vortex

Reducing the endwall loss is essential for the development of high-efficiency aero-engine compressors. An advanced vortex generation technique, using an L-shaped groove in the endwall, has been applied to a high-speed linear compressor cascade to manipulate the complex endwall flow and hence reduce associated aerodynamic loss by means of an induced streamwise vortex. A geometry generation method for the L-shaped endwall groove has been developed for the purpose of facilitating the design of the L-shaped groove. Hereafter, the L-shaped endwall groove has been optimized through a surrogate-based optimization framework. The numerical results demonstrated that the optimal L-shaped groove could control the endwall flow effectively and achieve a significant loss reduction at all operating points that were considered. At the design point, the calculated total pressure loss coefficient was decreased by 17.9%, with the static pressure rise increasing by 5.53%. The flow underturning and loss away from the endwall were reduced, while the flow overturning and loss near the endwall increased. The analysis of the flow-control mechanism indicated that the groove separation vortex, i.e. the streamwise vortex generated by the L-shaped endwall groove, presented a barrier to the endwall cross flow, mitigating the interaction of the endwall fluid with the suction side of the blade.

Influence of the inlet boundary layer on non-axisymmetric endwall contouring effects in a linear compressor cascade

The modern aero-engine compressors are commonly designed to be highly loaded which can introduce strong three-dimensional flow separation, and effective flow control is of vital importance in improving compressor performance. Non-axisymmetric endwall contouring nowadays is conducted on axial compressors to control the flow separation. In this paper, a contoured endwall is designed based on a linear CDA compressor cascade. The genetic algorithm and ANN surrogate model are used in the optimization process of the hub contouring, based on RANS simulations. Three inlet boundary layers which follow the fully turbulent boundary layer assumption are generated with different thickness, and the influence mechanism is numerically investigated. The optimal endwall with streamwise concave near the suction side provides great performance improvement at all inlet boundary layers at the aerodynamic design point. The endwall flow near the suction side is accelerated by the contoured endwall with cross-flow pressure gradient altered, suppressing the corner separation. The main pressure loss is closely related to one streamwise vortex that blocks the endwall flow. A thicker inlet boundary layer can induce early forming and increased strength on the streamwise vortex, enhancing the cross-flow movement and its interaction with the blade surface flow. In addition, the pressure loss near endwall in the inlet boundary layer is found more dominant on the endwall contouring effect than the boundary layer thickness.

Enhanced Prediction of Three-Dimensional Finite Iced Wing Separated Flow Near Stall

Icing on three-dimensional wings causes severe flow separation near stall. Standard improved delayed detached eddy simulation (IDDES) is unable to correctly predict the separating reattaching flow due to its inability to accurately resolve the Kelvin-Helmholtz instability. In this study, a shear layer adapted subgrid length scale is applied to enhance the IDDES prediction of the flow around a finite NACA (National Advisory Committee for Aeronautics) 0012 wing with leading edge horn ice. It is found that applying the new length scale contributes to a more accurate prediction of the separated shear layer (SSL). The reattachment occurs earlier as one moves towards either end of the wing due to the downwash effect of the wing tip vortex or the influence of end-wall flow. Consequently, the computed surface pressure distributions agree well with the experimental measurements. In contrast, standard IDDES severely elongates surface pressure plateaus. For instantaneous flow, the new length scale helps correctly resolve the rollup and subsequent pairing of vortical structures due to its small values in the initial SSL. The computed Strouhal numbers of vortical motions are approximately 0.2 in the initial SSL based on the vorticity thickness and 0.1 around the reattachment based on the separation bubble length. Both frequencies increase when moving towards the wing tip due to the downwash effect of the tip vortex. In comparison, the excessive eddy viscosity levels from the standard IDDES severely delay the rollup of spanwise structures and give rise to "overcoherent" structures.

Effective strategy of non-axisymmetric endwall contouring in a linear compressor cascade

The corner separation and the related secondary flow have great impact on the compressor performance, and non-axisymmetric endwall contouring is proved effective in improving compressor efficiency. The aim of the study is to improve the compressor performance by two local endwall contouring strategies at the design and off-design conditions. The endwall is parameterized and the Bezier curve is used to loft the endwall surface. The design of the contoured endwall is based on a multi-point optimization method to minimize the aerodynamic pressure loss. In order to identify the influence of the contoured endwall, a detailed flow analysis is conducted on four effective contoured endwall designs. The selected endwall geometries exhibit great control ability on the corner separation and significantly reduce the pressure loss at the two operating conditions. The directional concave near the leading edge can induce strong streamwise pressure gradient and accelerate the endwall flow, greatly reducing the cross-passage pressure gradient. The convex structures near the concave edge and at the outlet can block the cross-flow and prevent the interaction between the cross-flow and the suction corner flow. The benefit of the contoured endwall is mainly due to the re-distributed endwall static pressure and blocking of the cross-flow movement. In terms of the control effect, the shape of the concave also matters and better control effect is observed on the deep and wide concave. The flow will be guided by the concave, and the best suppression on corner separation is observed on the concave which follows the suction side. The results also indicate that the relief of the hub corner separation slightly increases the shroud pressure loss.

Comparison of Endwall Loss Reduction Techniques in a High-Lift Turbine Passage

Abstract The development of techniques that reduce the losses in the endwall region is an important area of research as it supports an increase in the turbine design space through the use of higher lift blade designs while maintaining high efficiency. Several active and passive shape contouring methods that reduce losses generated by the secondary flow have been developed and investigated on a high-lift front-loaded low-pressure turbine research profile in a low-speed linear cascade configuration. This paper summarizes and compares alterations to the three-dimensional secondary flow field by the application of three different techniques: blade profile contouring, optimized endwall shape contouring, and localized low mass coefficient jets. Each method was applied to identical research blade profiles and compared in the same linear cascade wind tunnel resulting in a unique perspective into the loss reduction mechanisms associated with each technique. The design strategy will be discussed along with a detailed description of changes to the secondary flow field using in- and out-of-passage total pressure loss measurements and high-speed stereoscopic particle image velocimetry. The key findings include the loss reduction mechanisms associated with each approach and the manipulation of key endwall flow structures such as the passage vortex and a strong suction surface corner separation.

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.

Research on the Influence of Steam Turbine Seal Leakage

CFD was employed to simulate the steam flow in 1.5-stage cascades with three different seal clearances. When the seal clearance is 0, there is no steam seal leakage, no obvious secondary flow in the cascade, and the stage efficiency reaches 88.27%. When the clearance of diaphragm seal and rotor tip seal is 1mm, the leakage of diaphragm seal strongly interferes with the flow in the cascade, which promotes the formation and development of the end-wall secondary flow near rotor hub, and the rotor tip leakage flow has little effect on next stage, and the stage efficiency drops by about 2 percentage points. When the seal clearance is 4mm, the end-wall secondary flow near rotor hub and next stator casing is strengthened significantly, and the attack angle loss increases, so the stage efficiency decreases by about 13 percentage points.

원형 실린더 전단에 분절 핀 설치에 따른 바닥면의 유동 및 열전달 특성 연구

원형 실린더 전단에 분절 핀 설치에 따른 바닥면의 유동 및 열전달 특성 연구

Impact of Boundary Layer Suction on Clearance Leakage Flow in a Cantilever Stator of Transonic Compressor

To control secondary flow effects and enhance the aerodynamic performance of the compressor, the flow control effects of the flow suction at the endwall with different circumferential positions and at the blade tip were numerically investigated in the cantilever stator of an axial single-stage transonic compressor. The main purpose was to gain a better understanding of the application of boundary layer suction and the associated control mechanisms in the cantilever stator. The studies show that the optimal position of the endwall suction slot should be located up the stator blade, in terms of the leakage flow structures and the blade tip unloading effect. In addition, the flow control effects of the suction at the blade tip on leakage flow upstream is better than that of the endwall flow suction with the same structure. Further, the studies of compressor aerodynamic performance curves illustrate that the efficiency and pressure ratio is increased by 0.34% and 1.09% at the peak efficiency point, and are increased by 0.39% and 0.14% at the near stall point, respectively.

Analysis of upstream, double-row, cylindrical holes on primary and secondary effects of endwall flow and film cooling

Flow features and film cooling performance of five configurations of double-row, cylindrical holes, upstream of an E3 vane, in a linear cascade are numerically investigated. This simulation is completed using a verified turbulence model at four blowing ratios (M = 0.5, 1.0, 1.5, 2.0). The first three configurations have two rows of cylindrical holes, each row with the same compound angle (β=-45°, 0° or 45°), while the other two have two rows with opposite compound angles (β=-45°, 45° and β=45°, -45°), which are also referred to as double-jet film cooling (DJFC) holes. The primary effects on the downstream endwall and the secondary effects on the nearby airfoil of the cooled passage are analyzed and discussed in detail. Results show that at low blowing ratios the movement of the coolant is denominated by the interaction between the jets and vortices resulting in similar film coverage on both the endwall and airfoil. The effect of vortices is reduced at high blowing ratios. It is also shown that the movement of the coolant is determined by the initial velocity direction, as well as the film cooling configuration.

A Novel Vortex Identification Technique Applied to the 3D Flow Field of a High-Pressure Turbine

Abstract The efficiency of modern axial turbomachinery is strongly driven by the secondary flows within the vane or blade passages. The secondary flows are characterized by a complex pattern of vortical structures that origin, interact, and dissipate along the turbine gas path. The endwall flows are responsible for the generation of a significant part of the overall turbine loss because of the dissipation of secondary kinetic energy and mixing out of nonuniform momentum flows. The understanding and analysis of secondary flows requires a reliable vortex identification technique to predict and analyze the impact of specific turbine designs on the turbine performance. However, the literature shows a remarkable lack of general methods to detect vortices and to determine the location of their cores and to quantify their strength. This paper presents a novel technique for the identification of vortical structures in a general 3D flow field. The method operates on the local flow field, and it is based on a triple decomposition of motion proposed by Kolář. In contrast to a decomposition of velocity gradient into the strain and vorticity tensors, this method considers a third, pure shear component. The subtraction of the pure shear tensor from the velocity gradient remedies the inherent flaw of vorticity-based techniques, which cannot distinguish between rigid rotation and shear. The triple decomposition of motion serves to obtain a 3D field of residual vorticity whose magnitude is used to define vortex regions. The present method allows to locate automatically the core of each vortex, to quantify its strength, and to determine the vortex bounding surface. The output may be used to visualize the turbine vortical structures for the purpose of interpreting the complex three-dimensional viscous flow field and to highlight any case-to-case variations by quantifying the vortex strength and location. The vortex identification method is applied to a high-pressure turbine with three optimized blade tip geometries. The 3D flow field is obtained by computational fluid dynamics (CFD) computations performed with Numeca FINE/Open. The computational model uses steady-state Reynolds-averaged Navier–Stokes (RANS) equations closed by the Spalart-Allmaras turbulence model. Although developed for turbomachinery applications, the vortex identification method proposed in this work is of general applicability to any three-dimensional flow field.

A new approach of endwall recirculation in axial compressors

This study reports on a novel design of endwall recirculation in a transonic axial flow compressor rotor. The bled air from downstream of the rotor is injected over the tip clearance of the blades in the circumferential direction opposite to the blade rotation. Results reveal that the treated casing effectively increases the operating range of the rotor at the expense of some efficiency loss. It is shown that the endwall recirculation applied in this work does not reduce the incidence angle upstream of the blade but reduces the pressure difference between the pressure and suction surface and pushes the leakage vortex and the passage shock downstream. The endwall flow was found to be responsible for stall initiation in the treated casing (similar to the smooth casing).

Near-Wall Flow in the Blade Cascades Representing Last Rotor Root Sections of Large Output Steam Turbines

This paper investigates the flow past two variants of root section profile cascades for a last stage rotor considering three-dimensional flow structures in the near-wall region. Analyses were drawn based on RANS numerical simulations of both variants and on the experimental data obtained by the 3D traversing in the exit flow field of one of the variants. Extent of 3D structures at two different regimes and its influence on aerodynamic characteristics of the blade cascades was assessed. The distributions of Mach number along the profiles were compared with 2D optical measurements and its distortion due to the presence of the sidewall was explored. The interaction between main vortical structures was described and its influence on the loading of the blades, mechanical energy losses and exit flow angle was discussed. The results showed that for a front loaded blade the vortical structures appeared earlier and at a larger extent than for an aft loaded variant. However, due to different Mach number distribution, contribution of end wall flow to the energy losses was lower in the case of the aft loaded variant. The influence of the near wall flow on the loading was found to be rather weak while the deviation of the exit flow angle appeared to be comparable for both of the variants.

A new curvature-controlled stacking-line method for optimization design of compressor cascade considering surface smoothness

The paper presents an advanced parametric method of blade stacking lines in terms of sweep and lean based on controlled curvature. To the knowledge of the authors, there is no related approach reported in open literature that uses Bezier spline as the radial curvature distribution to improve the smoothness of the blade surface; most previous studies ignored the discontinuous slopes of curvature of the parametric curves. The parametric method called curvature-controlled stacking-line method (CCSLM) is performed by changing the magnitude of the sweep or lean. A fourth Bezier spline is adopted to define the curvature of spanwise stacking line directly ensuring surface smoothness. Then, the redesign cascades are created by sectional profiles stacked along the radial stacking lines which are obtained by twice integrating the Bezier spline. Then, the advanced method is conducted to optimize a high-subsonic controlled diffusion airfoil at design point, where the blade shape is generated in terms of lean. A single-objective optimization is performed using Kriging model and genetic algorithm to optimize total pressure loss, and the optimized geometry is obtained. The optimization results show that the blade design CCSLM has significant effects on the endwall flow vortex as well as radial loading distribution. The reduction of total pressure loss and secondary flow is also observed, and the aerodynamic performance is well improved compared with the original cascade.

Large-Eddy Simulation and RANS Analysis of the End-Wall Flow in a Linear Low-Pressure Turbine Cascade, Part I: Flow and Secondary Vorticity Fields Under Varying Inlet Condition

Abstract In low-pressure turbines (LPTs), around 60–70% of losses are generated away from end-walls, while the remaining 30–40% is controlled by the interaction of the blade profile with the end-wall boundary layer. Experimental and numerical studies have shown how the strength and penetration of the secondary flow depends on the characteristics of the incoming end-wall boundary layer. Experimental techniques did shed light on the mechanism that controls the growth of the secondary vortices, and scale-resolving computational fluid dynamics (CFD) allowed to dive deep into the details of the vorticity generation. Along these lines, this paper discusses the end-wall flow characteristics of the T106 LPT profile at Re = 120 K and M = 0.59 by benchmarking with experiments and investigating the impact of the incoming boundary layer state. The simulations are carried out with proven Reynolds-averaged Navier–Stokes (RANS) and large-eddy simulation (LES) solvers to determine if Reynolds-averaged models can capture the relevant flow details with enough accuracy to drive the design of this flow region. Part I of the paper focuses on the critical grid needs to ensure accurate LES and on the analysis of the overall time-averaged flow field and comparison between RANS, LES, and measurements when available. In particular, the growth of secondary flow features, the trace and strength of the secondary vortex system, and its impact on the blade load variation along the span and end-wall flow visualizations are analyzed. The ability of LES and RANS to accurately predict the secondary flows is discussed together with the implications this has on design.

Influences of large fillets on endwall flows in a vane cascade with upstream slot film-cooling

Investigations in cascade employ filleted blades to influence the near endwall secondary flows and total-pressure losses. The secondary flows aggravate the aerodynamic losses and endwall thermal stresses in the gas turbine passages. Investigations of different configurations of the slot-bleed flow from the endwall of cascade upstream show significant influences on the near endwall flows. In the present paper, the near endwall flow-field in a linear cascade employing filleted vanes and bleed flow from the upstream endwall slots is measured experimentally. Two fillet profiles are tested, one is larger than the other. The objectives are to quantify the combined effects of endwall fillet, fillet profiles, and film-cooling flow on the endwall region flow-field. The fillet covers the junction of vane and endwall upstream of the cascade throat region. The discontinuous bleed-slots near the cascade entrance provide the film-cooling flow on endwall and simulate the gaps between combustor/nozzle-vane discs or stator/rotor discs in the gas turbine. The inlet Reynolds number of 2.0E+05 is based on the chord length of vane profile. The density and temperature ratios of the coolant flow to mainstream are both about 1.0. The inlet blowing ratio of the film-cooling flow varies between 1.0 and 2.8. The flow-field is measured through the distributions of flow temperature, yaw angle, axial vorticity, and total-pressure losses along the vane passage. The effects on the flow-field are then presented by comparing the cases of filleted vanes with the cases of un-filleted vanes. The results of flow yaw angle and axial vorticity in the filleted passage are smaller than those in the passage without the fillet. The yaw angles responsible for the endwall secondary flows are the smallest for the smaller fillet (Fillet-2). The temperature field indicates the pitchwise distributions of the coolant on endwall specially near the pressure side are better when the smaller fillet is employed. Also, the weakened passage vortex of the endwall secondary flows in the filleted passage reduces the total-pressure losses. Although the total-pressure losses decrease at very high coolant mass flux with and without the fillet, the losses are always smaller for the filleted passage than for the un-filleted passage. The present investigation is beneficial for improving and optimizing the endwall film-cooling in the gas turbine passage.

Passive Separation Control with Blade-End Slots in a Highly Loaded Compressor Cascade

The proposed passive control method, blade-end slots, is implemented to reorganize the blade-end flow in a highly loaded compressor cascade under a wide incidence range at . Both experimental and numerical investigations are conducted to validate the control effect and reveal the control mechanism. Experimental results show that the proposed method can effectively suppress the corner separation and broaden the effective operating range. The total pressure loss in the measurement plane is significantly reduced, ranging as high as 18.4, 20.6, 24.3, and 39.4% under the incidence angle of , 0°, 2° and 4°, respectively. Flow details indicate that the blade-end slots can generate self-adaptive high-momentum jet flows through the pressure difference between blade pressure and suction surface. The self-adaptive jet flow reenergizes the local low-momentum fluid to reduce separation, and restrain its migration toward midspan. The low-momentum flows near the endwall on the pressure side are aspirated, thus weakening the cross-passage migration of low-momentum endwall flow into suction surface corner. Proportion of unseparated regions in the blade midspan is enlarged; consequently, flow capacities are improved. The attachment at the blade-end regions is enhanced, and the insufficient circulation originally found at the blade extremity is improved.

Impact of hub gap leakage on stator endwall flow in an axial compressor stage with casing treatment

Abstract In an axial compressor stage equipped with slot-type casing treatment, simulation without hub gap between rotors and stators achieves much lower stall margin enhancement than experiment, which results from the over-predicted corner stall in stators. Aiming at clarifying the combined effect of hub leakage flow and casing treatment on operating range and the flow mechanism of leakage flow interacting with stator flow, a detailed numerical research is performed in this compressor stage. The effects of hub leakage mass flow are elaborated in terms of compressor performance and flow fields. Unsteady simulations with appropriate leakage mass flow condition are able to predict the stall margin improvement and total pressure characteristics preferably. Casing treatment slots tend to re-distribute the loading of downstream stator by decreasing loading above 70% span and increasing that below 70% span, which result in the exaggerated corner stall in stators. Hub leakage flow from cavity resists the tendency brought by casing treatment and succeeds in relieving the heavy loading near the hub. The flow mechanism for the effect of cavity leakage flow on stall margin is elucidated. Cavity hub leakage gives rise to a thickened boundary layer upstream stator inlet by producing a separation bubble and aggravating the boundary layer skewing. The thickened boundary layer produces a horseshoe vortex in the stator passage, which induces main flow to rolling up as the pressure-side branch and avoids the emergence of corner stall.