High Pressure Turbine Cascade Research Articles

Numerical Study on the Influence of Inlet Turbulence Intensity on Turbine Cascades

The turbulence intensity of the high-pressure turbine inlet plays an important role in the development of secondary flow and boundary layer evolution in the turbine passage. Unfortunately, current research often overlooks the coupling effect between boundary layer separation and endwall secondary flow, and lacks comprehensive exploration of loss variation. To complement existing research, this study utilizes numerical simulation techniques to investigate the evolution of the boundary layer and secondary flow in high-pressure turbine cascades under varying turbulence intensities, with experimental research results as the basis. Furthermore, the relationship between profile loss and endwall secondary flow loss is analyzed. The research results indicate that higher turbulence intensity can enhance the anti-separation ability of the boundary layer, thereby reducing the blade profile loss caused by separation in the boundary layer. However, higher turbulence intensities (Tu) enhance the development of secondary flow within the cascade, leading to a significant increase in secondary flow losses. Q-criterion methods are employed to display the vortex structures within the passage, while loss decomposition is leveraged to uncover the variations of different losses under different turbulence intensities (Tu of 1% and 6%).With the exception of a minor decrease in total pressure loss (Yp) when Tu increases from 1% to 2%, Yp demonstrates a substantial increase in all other cases with increasing Tu. Additionally, this study explains why the loss of high-pressure turbine cascades shows a trend of first increasing and then decreasing with the increase in inlet turbulence intensity.

Enhancing Aerodynamic Performance of a Non-Axisymmetric Endwall Contoured Cascade Through Section Profiling Method

Abstract Endwall contouring and 3D blade configurations are effective techniques for enhancing turbomachinery performance. The integration of these technologies is an important area of investigation. To continue the numerical and experimental validations of the non-axisymmetric endwall contoured cascade (Du et al., 2024, “Endwall Contouring for Improving Aerodynamic Performance in a High-Pressure Turbine Cascade,” ASME J. Turbomach., 146(10), p. 101001), the present study redesigns the cascade via section profiling using optimization methods to maximize performance gains. To improve the control effect of endwall contouring, blade section parameters are strategically adjusted during optimization. The experimental flow field traverses at cascade exits demonstrate the control of the cascade flow field by the optimized design. The redesigned cascade exhibits decreased overturning near the shroud and reduced underturning near the hub, as well as reduced profile loss in the mainstream, compared to the endwall contoured cascade. The modification of section parameters enhances the control of losses and the secondary flow intensity. Computational fluid dynamics is used to provide a detailed analysis of the flow field, providing insights into how flow structures vary when the blade profile is modified. The numerical results further elucidate the effect of blade section profiling on both the performance characteristics of the cascade and the configuration of the secondary flows.

Endwall Contouring for Improving Aerodynamic Performance in a High-Pressure Turbine Cascade

Abstract The endwall contouring has proven to be an effective technique in controlling the impacts of secondary flow within turbomachinery. A baseline cascade with the original axisymmetric endwall (BASE) has been redesigned with a non-axisymmetric contoured shroud endwall (EW-B1), and both configurations are investigated through experimental and numerical methods. The endwall is profiled through the B-spline surface method with the purpose of reducing the overall total pressure loss at the exit. The experimental studies involve flow field traverses at the exits of both cascades. Numerical calculations are conducted to gain a deeper understanding of the effects of endwall contouring. The numerical results exhibit good agreement with the experimental results, both in loss and flow angle. The results demonstrate a significant reduction in losses and secondary kinetic energy in EW-B1. The causes of the high-loss region and the overturning near the shroud are analyzed using computational fluid dynamics (CFD), as well as the changes in pressure fields and vortex structures within the cascade passage. The weakening of the horseshoe vortex and the radial movement of the passage vortex are confirmed to be the primary reasons for losses reduction.

Experimental study on flow phenomena induced by trailing edge shapes for a transonic high pressure turbine cascade

Experimental study on flow phenomena induced by trailing edge shapes for a transonic high pressure turbine cascade

A Numerical Test Rig for Turbomachinery Flows Based on Large Eddy Simulations With a High-Order Discontinuous Galerkin Scheme—Part II: Shock Capturing and Transonic Flows

Abstract In the second paper of this three-part series, we focus on the simulation of transonic test cases for turbomachinery applications using a high-order discontinuous Galerkin spectral element method (DGSEM). High-fidelity simulations of transonic compressors and turbines are particularly challenging, as they typically occur at high Reynolds numbers and require additional treatment to reliably capture the shock waves characterizing such flows. A recently developed finite-volume subcell shock capturing scheme tailored for the DGSEM is applied and evaluated with regard to the shock sensor. To this end, we conduct implicit large eddy simulations of a high-pressure turbine cascade from the public literature and a transonic compressor cascade measured at the German Aerospace Center, both at a high Reynolds number above 106. Based on the results, we examine modal-energy and flow-feature based shock indicator functions, compare the simulation data to experimental and numerical studies, and present an analysis of the unsteady features of the flows.

Aerodynamic Performance and Leakage Flow in Turbine Cascades With Sweeping Jet Actuators

Abstract In recent years, unsteady flow-control technology has been considered as more promising for reducing secondary flow loss in turbines. Sweeping jet actuators (SJAs), in particular, have steadily been used in turbomachinery to reduce loss due to their wide sweeping range. However, most of the preliminary studies were focused mainly on computational fluid dynamics calculations, with very few on wind tunnel experiments to verify the flow-control effectiveness. To fill this gap, an innovative attempt to investigate experimentally the influences of SJAs on the aerodynamic performance and leakage flow is presented in a high-pressure turbine cascade. Comparing SJAs and hole-type steady jet actuators (HSJAs), the influence on the tip leakage flow and the loss characteristics at various incidences and injection frequencies are discussed in detail. The results indicate that the SJAs effectively reduce the magnitude and extension of the leakage vortex and the passage vortex, and weaken the interactions between them. The loss at the cascade outlet drops by up to 10.3% when the jet flow of an SJA equals 0.2% of the inlet flow, which is significantly better than that with an HSJA (6.3%).

Effects of the multicavity tip coolant injection on the blade tip and the over tip casing aerothermal performance in a high-pressure turbine cascade

In high performance gas turbine engines, the over tip leakage flow driven by the lateral pressure gradient is inevitably induced inside the blade tip gap in the high-pressure stage turbine due to the freestanding airfoil design methodology. To obtain an increasing level of thermal efficiency, the turbine inlet temperature is gradually increased in terms of the Brayton cycle. Hence, the blade tip and the over tip casing are subjected to high thermal load. In real turbine blade, cavity tips are widely used to decrease the over tip leakage flow and the thermal load on the blade tip and over tip casing. In the present study, the numerical simulations were conducted to investigate the effects of the multicavity coolant injection on the blade tip and the over tip casing aerothermal performance. Three-dimensional (3-D) Reynolds-averaged Navier–Stokes (RANS) equations and standard turbulence model were solved together in the simulations. The results indicate the ribs inside the tip cavity alter the distribution of film cooling efficiency by changing flow structure within the tip gap. Most of the coolant is limited in each little cavity owing to the blockage of ribs. Here, the swirling action of each cavity vortex results in the coolant providing a wider film cooling coverage. Therefore, the increase in film cooling effectiveness on the blade tip surface is more efficient than that of over-tip casing. The blade with four tip cavities film cooling (4CFC) obtains the largest area-averaged film cooling effectiveness, which is augmented by 14.3% in comparison with the case with a single tip cavity film cooling (1CFC).

Effects of suction surface near tip coolant injection on the aerodynamics of tip leakage flow in a high-pressure turbine cascade

The current study explores the possibility of using the coolant injected from the near tip suction side of a turbine blade to reduce the tip leakage loss. Each case studied has a cooling hole injecting a coolant stream. To model the mixing of the coolant at different axial locations, each of four cases is investigated with a cooling hole located at 10% blade axial chord [Formula: see text], 20% [Formula: see text], 30% [Formula: see text] or 40% [Formula: see text] from the leading edge. In computational fluid dynamics simulation, the exit area of the cooling hole is defined with mass flow inlet boundary condition to simulate a coolant injection. It is found that the tip leakage loss of the case with a cooling hole located at 10% [Formula: see text] is the lowest and is 3.2% lower than that of the uncooled baseline case. The tip leakage loss increases as the cooling hole located further downstream. For the case with cooling hole at 40% [Formula: see text], the loss is 1.5% higher than the uncooled baseline case. The effect of the blowing ratio on the tip leakage loss is also investigated. For the case with the cooling hole located at 10% [Formula: see text], the tip leakage loss first decreases and then increases as the blowing ratio increases. When the tip leakage loss of one case is less than that of baseline case, it is found that mixing of the coolant and the low-energy fluid in the tip leakage vortex core reduces the stagnation pressure difference between mainstream and the fluid within the vortex core, and thus reducing the mixing loss within the blade passage.

Numerical investigation and loss estimation of high-pressure turbine cascade flow with contoured endwall and incoming wakes

With the aim of suppressing the irreversible losses, the secondary flow at endwall and incoming vortices interactions are assumed to be the most concerned physics in gas turbine for modern aero-engine and military vehicles. However, the combining effect of contoured endwall and incoming periodic wakes were seldom discussed previously. In the present work, the research subject is the high-pressure turbine cascade with non-axisymmetric endwall and incoming wakes induced by moving bar fulfilled with periodic boundary conditions in pitchwise direction. Firstly, the aerodynamics and thermodynamics in the turbine cascade with secondary flow control technique, which denotes to non-axisymmetric endwall, were experimentally and numerically investigated. For what concerns the profile aerodynamic loadings, total pressure loss coefficients and vortices coefficient were surveyed to quantify the influence of interaction between incoming wakes and secondary flow at contoured endwall. The cumulative difference of entropy production rate associated with dissipation is then employed to evaluate the loss under different inflow condition. Combined with dynamic mode decomposition (DMD) method, the energy loss features associated with incoming wakes at different passing frequencies was captured. The results indicated that the flow separation at the suction wall with the axial position of 70%∼80% Cax upstream trailing edge was suppressed in the cases with incoming wakes for high pressure turbine cascade. And stronger incoming wakes tends show better suppression effect on the flow separation. However, when combined with contoured endwall, the negative influence of non-axisymmetric endwall at the middle passage was enlarged. And the general aerodynamics of the cascade flow may not be necessarily improved in term of energy loss characteristics with the increase of passing frequency of incoming wakes. The present work preserves guiding significance for the design and flow control of high pressure turbine cascade.

Numerical Investigation of the Wake Vortex-Related Flow Mechanisms in Transonic Turbines

As the core equipment of the power generation system, a gas turbine is an indispensable energy-converting device in the national industry. The flow inside a high-pressure turbine (HPT) is highly unsteady, which has a great influence on the aerothermal performance and structural strength. To better clarify the flow mechanism and guide the advanced design, the basic flow characteristics of transonic turbines are investigated in the paper by a modified scale-adaptive simulation (SAS) model based on the shear stress transport (SST) turbulence model. The numerical results reveal the formation and development of the secondary flow structures such as wake vortex, pressure wave, shock wave, and the interactions among them. The length and frequency characteristics of wake are in good agreement with the large eddy simulation (LES) and the experimental data. Based on the detailed flow information, the local loss analysis is performed using the entropy generation rate. In summary, the wake vortex-related flow is the main origin of unsteadiness and entropy loss in high-pressure turbine cascade.

Effect of non-axisymmetric endwall and periodic upstream wakes on the aero-thermal dynamics in turbine cascade

The thermal management and heat transfer behavior considering both the contoured endwall and upstream wakes were seldom investigated. In the present work, the experimental measurements of aero-thermal characteristics regarding to a high-pressure turbine cascade with flat and non-axisymmetric endwall, which was optimized with the aim of minimum aerodynamic losses, were carried out. The numerical analysis of the cascade flow with and without upstream wakes and contoured endwall was performed. Then the field synergy theory and entropy production rate concerning heat transfer were introduced. Furthermore, the combining influence of incoming wakes and secondary flow at endwall on the heat transfer characteristics was investigated. Results show that the non-axisymmetric endwall suppresses the passage vortex by weakening the cross flow and surface vortex with concave profiling. And the underturning effect of yaw angle at the cascade exit is reduced due to lower spanwise positon of passage vortex and weaker secondary flow. The positive influence of contoured endwall on losses induced by temperature gradients is mainly localized at the passage downstream, and the leading edge area for contoured endwall, however, has the potential of overheating due to significantly strengthened horse shoe vortex at suction side. The periodic wakes scatter and suppress the negative influence of the contoured endwall on the heat transfer at the leading edge by mixing with the horse shoe vortex. And when wakes and passage vortex meet, the risk of overheating at the rear part of the cascade is reduced. The present work may guide shed now lights into the application of cooling technique on the high-pressure turbine cascade.

High-Resolution Measurements of Heat Transfer, Near-Wall Intermittency, and Reynolds-Stresses Along a Flat Plate Boundary Layer Undergoing Bypass Transition

Abstract A new experimental dataset focusing on the influence of high freestream turbulence and large pressure gradients on boundary layer transition is presented. The experiments are conducted in a new wind tunnel equipped with a flat plate test section and a new kind of turbulence generator, which allows for a continuous variation of turbulence intensity. The flat plate is mounted midway between contoured top and bottom walls. Two different wall contours can be implemented to create pressure distributions on the flat plate that are typical for the pressure and suction side of high pressure turbine cascades. A large variation of Reynolds number from 3.0 × 105 to 7.5 × 105 and inlet turbulence intensity between 1.1% and 8% is realized, resulting in more than 100 test cases. Measurements comprise highly resolved heat transfer, near-wall intermittency and freestream Reynolds stress distributions. Near-wall intermittency is measured using a traversable hotfilm sensor while freestream Reynolds stresses are measured simultaneously at the same position with a revolvable X-wire probe. Additionally, turbulent length scales are analyzed using the X-wire signal along the flat plate. Results show that heat transfer and near-wall intermittency distributions are in good agreement and that heat transfer at high turbulence levels increases prior to the formation of first turbulence spots. Transition onset is found to be influenced by the turbulence Reynolds number, i.e., turbulent length scales. At constant inlet turbulence intensity, transition onset moves upstream, when the turbulent Reynolds number is decreased.

Implicit Discontinuous Galerkin Solution on Unstructured Mesh for Turbine Blade Secondary Flow

Abstract To enhance solution accuracy with high-order methods, an implicit solver using the discontinuous Galerkin (DG) discretization on unstructured mesh has been developed. The DG scheme is favored chiefly due to its distinctive feature of achieving a higher-order accuracy by simple internal sub-divisions of a given mesh cell. It can thus relieve the burden of mesh generation for local refinement. The developed method has been assessed in several test cases. First, an examination on the mesh-dependency with different DG orders for discretization has demonstrated the expected mesh convergence order of accuracy correlation. The flow around a cylinder solved with up to the fifth-order accuracy demonstrates the need for a high-order geometrical representation corresponding to the high-order flow field discretization. A calculated turbulent flat plate boundary layer demonstrates the capability of the high-order DG in capturing the laminar sub-layer for a given coarse mesh (y + >20) without a wall function, in a clear contrast to a conventional second-order scheme. The focal test case of the present work is a high-pressure (HP) turbine cascade with the flow losses being strongly influenced by both the laminar separation induced transition on the blade suction surface and the secondary flow development in the endwall region. Fully turbulent RANS solutions consistently overpredict the losses, which can be significantly reduced by an empirically specified transition at a mid-chord point. Of particular interest is a contrasting behavior of pure laminar flow solutions. A steady laminar flow solution is difficult to converge, and even when converging, tends to give an unrealistically large separation. On the other hand when the laminar solution is run in a time-accurate unsteady mode, a qualitatively different flow pattern emerges. The separated shear layer is then shown to lead to coherently shed unsteady vortices with a net entrainment from the main stream to the near-wall region. The resultant time-averaged near-wall flow then effectively becomes a reattached boundary layer. Both overall and distributed losses computed from unsteady laminar solutions are shown to be consistently much better than the fully turbulent RANS solutions. It is remarkable and quite unexpected that for such a seemingly complex 3D flow problem, the straightforward unsteady laminar solutions with no empirical input seem to be comparable with (or slightly better than) the tripped RANS with empirical input.

Large eddy simulation of turbomachinery flows using a high-order implicit residual smoothing scheme

A recently developed fourth-order accurate implicit residual smoothing scheme (IRS4) is investigated for the large eddy simulation of turbomachinery flows, characterized by moderate to high Reynolds numbers and subject to severe constraints on the maximum allowable time step if an explicit scheme is used. For structured multi-block meshes, the proposed approach leads to the inversion of a scalar pentadiagonal system by mesh direction, which can be done very efficiently. On the other hand, applying IRS4 at each stage of an explicit Runge–Kutta time scheme allows to increase the time step by a factor 5 to 10, leading to substantial savings in terms of overall computational time. With respect to standard second-order fully implicit approaches, the IRS4 does not require approximate linearization and factorization procedures nor inner Newton-Raphson subiterations. As a consequence, it represents a better cost-accuracy compromise for the numerical simulations of turbulent flows where the maximum time step is controlled by the lifetime of the smallest resolved turbulent structures. Numerical results for the well-documented high-pressure VKI LS-89 planar turbine cascade illustrate the potential of IRS4 for significantly reducing the overall cost of turbomachinery large eddy simulations, while preserving an accuracy similar to the explicit solver even for sensitive quantities like the heat transfer coefficient and the turbulent kinetic energy field.

Influences of a multi-cavity tip on the blade tip and the over tip casing aerothermal performance in a high pressure turbine cascade

In modern gas turbine, the over tip leakage flow is inevitably generated in the tip gap of the first stage turbine blade due to the freestanding airfoil. In order to obtain a higher thermal efficiency, the turbine inlet temperature is gradually increased. Therefore, the over tip leakage flow induces significant aerodynamic losses and the blade tip and the over tip casing endure high level of thermal load. In the pursuit of a high-performance turbine engine, cavity tips are widely implemented in the turbine blade to reduce the over tip leakage flow and the thermal load on the blade tip and over tip casing. In the current study, the influences of the multi-cavity squealer tip on the blade tip and the over tip casing aerothermal performance were numerically investigated. Three-dimensional (3D) Reynolds-Averaged Navier-Stokes (RANS) equations and standard k-ω turbulence model were solved to conduct the simulations. The results indicate that the flat tip attains the largest thermal load on the blade tip, over-tip casing and induces the largest total pressure loss. However, the case with a single tip cavity (1CST) achieves the smallest total pressure loss and heat transfer coefficient on the over-tip casing, which are 7.6% and 19.6% lower than the flat tip, respectively. The case with a multi-cavity tip obtains a decreasing heat transfer coefficient on the blade tip. The case with five tip cavities (5CST) achieves the smallest heat transfer coefficient on the blade tip among all simulated cases, which is decreased by 17.5% relative to the flat tip.

Influence of Blade Loading Profile on Wake Dynamics in High-Pressure Turbine Cascades

The influences of blade loading profile on wake convection and wake/wake interaction were studied in two different blade designs for high-pressure (HP) turbines (front-loaded (FL) and aft-loaded (AL)), installed in linear cascades. A high-speed moving bar (HSMB) apparatus replicated wake shedding, and a closed loop wind tunnel produced engine-relevant Mach numbers (Ma = 0.7) and Reynolds numbers (Re = 3 × 105). The FL blades had approximately 10% greater total pressure loss when operated with unsteady wake passage. Phase conditioned particle image velocimetry (PIV) measurements were made in the aft portion of the blade channel and downstream of the blade trailing edge. The turbulence kinetic energy (TKE) in the wake was approximately 30% higher for the FL blades when the wake entered the measurement field-of-view. The pressure field in the upstream region of the FL blade design is believed to induce high magnitude strain rates—leading to increased TKE production—and more aggressively turn and dilate the unmixed wake—leading to increased mixing related losses. The higher TKE for the FL blades largely dissipated, being approximately equal to the AL wake by the time the wake reached the end of the blade passage. The interaction of the convected wake with the wake from the blade trailing edge caused periodic vortex shedding at the second harmonic of the convected wake frequency. This interaction also modulated the strength of the trailing edge wake. However, little difference was found in the modulation amplitudes between different cases due to similar strengths of the convected wakes in this region. The higher wake TKE in the upstream portion of the blade channel for the FL blades, therefore, is expected to be the cause of the higher total pressure loss.

Scale-Resolving Simulations of Bypass Transition in a High-Pressure Turbine Cascade Using a Spectral Element Discontinuous Galerkin Method

The application of a new computational capability for accurate and efficient high-fidelity scale-resolving simulations of turbomachinery is presented. The focus is on the prediction of heat transfer and boundary layer characteristics with comparisons to the experiments of Arts et al. (1990, “Aero–Thermal Investigation of a Highly Loaded Transonic Linear Turbine Guide Vane Cascade,” von Karman Institute for Fluid Dynamics, Rhode St. Genese, Belgium, Technical Note No. 174.) for an uncooled, transonic, linear high-pressure turbine (HPT) inlet guide vane cascade that includes the effects of elevated inflow turbulence. The computational capability is based on an entropy-stable, discontinuous Galerkin (DG) spectral element approach that extends to arbitrarily high orders of spatial and temporal accuracy. The suction side of the vane undergoes natural transition for the clean inflow case, while bypass transition mechanisms are observed in the presence of elevated inflow turbulence. The airfoil suction-side boundary layer turbulence characteristics during the transition process thus differ significantly between the two cases. Traditional simulations based on the Reynolds-averaged Navier–Stokes (RANS) fail to predict these transition characteristics. The heat transfer characteristics for the simulations with clean inflow agree well with the experimental data, while the heat transfer characteristics for the bypass transition cases agree well with the experiment when higher inflow turbulence levels are prescribed. The differences between the clean and inflow turbulence cases are also highlighted through a detailed examination of the characteristics of the transitional and turbulent flow fields.

Second-Law Analysis: A Powerful Tool for Analyzing Computational Fluid Dynamics (CFD) Results

Second-law analysis (SLA) is an important concept in thermodynamics, which basically assesses energy by its value in terms of its convertibility from one form to another.[...]

Tip gap size effects on thermal performance of cavity-winglet tips in transonic turbine cascade with endwall motion

AbstractThe thermal performance of two cavity-winglet tips with endwall motion is investigated in a transonic high pressure turbine cascade, which operates at an engine representative exit Mach number of 1.2 and an exit Reynolds number of 1.7 × 106. The numerical method is first validated with experimental data and then used to investigate blade heat transfer at three different tip clearances of 1.1, 2.1 and 3.1% chord. The effects of relative endwall motion are considered. The present results show that as the size of the tip gap increases, the heat transfer coefficient and heat load on the tip increases. The winglet geometries on the blade tip mainly affect the tip flow structure close to them. At a larger tip clearance, the size of the separation bubble above the pressure side winglet increases. The heat transfer coefficient is high on the pressure side winglet due to the flow reattachment at all tip clearances. Within the tip gap, when the size of the tip clearance increases, the size of the cavity vortex increases and the cavity scraping vortex due to relative endwall motion becomes smaller. The impingement of the both two vortexes can lead to high heat transfer coefficient on the cavity floor surface. On the blade suction surface, when the size of the tip clearance increases, the heat transfer coefficient of the cavity tip increases, but those of the winglet tips decreases. The heat transfer coefficient is high on the side surface of the suction side winglet at all tip clearances because of the tip leakage flow impingement.

Effects of Tip Gap Size on the Aerodynamic Performance of a Cavity-Winglet Tip in a Turbine Cascade

The aerodynamic performance of a cavity-winglet tip is investigated in a high-pressure turbine cascade by experimental and numerical methods. The winglet tip has geometric features of a cavity and a suction side fore-part winglet. A cavity tip is studied as the baseline case. The aerodynamic performances of the two tips are investigated at three tip gaps of 0.8%, 1.7%, and 2.7% chord. At tip gaps of 1.7% and 2.7% chord, the loss near the blade tip is dominated by the tip leakage vortex (TLV) for both tips, and the winglet tip mainly reduces the loss generated by the tip leakage vortex. In the past, it was concerned that at a small tip gap, the winglet tip could introduce extra secondary loss and show little aerodynamic benefits. The winglet tip used in the current study is also found to be able to effectively reduce the loss at the smallest tip gap size of 0.8% chord. This is because at this small tip gap, the tip leakage vortex and the passage vortex (PV) appear simultaneously for the cavity tip. The winglet tip is able to reduce the pitchwise pressure gradient in the blade passage, which tends to suppress the formation of the passage vortex. The effects of the winglet tip on the flow physics and the loss mechanisms are explained in detail.