Endwall Region Research Articles

UNSTEADY LOSS MECHANISMS IN LOW-PRESSURE TURBINES WITH DIVERGING END-WALLS STUDIED VIA HIGH-FIDELITY SIMULATION

Abstract High-fidelity simulations are used to conduct controlled numerical experiments to investigate the effect of periodically incoming wakes on profile and three-dimensional loss mechanisms. The present work considers the MTU-T161 cascade with spanwise-diverging end-walls, representative of a high-lift, low-pressure turbine blade. All simulations are carried out at engine-relevant conditions, with exit Reynolds number of 90,000 and exit Mach number of 0.6. Upstream moving bars are used to generate incoming wakes which impinge on the blade and potentially alter its aerodynamic performance. Unlike previous studies, the incoming wakes are subjected to an additional axial pressure gradient when convecting through the passage, due to the divergence of the spanwise end-walls. Following validation against available experimental data, a systematic variation of flow coefficient and reduced frequency extends the parametric space studied to encompass engine-realistic operating conditions. The high-fidelity simulations reveal the impact of incoming wakes on blade boundary layer losses and wake-induced losses both at the mid-span and within the end-wall regions. Secondary losses incurred in the end-wall region show little sensitivity towards unsteadiness associated with incoming wakes and are rather prone to the turbulence levels in the passage. On the other hand, profile losses show high dependency on bar wakes in the absence of wake fogging. Whilst profile losses can be minimized by certain combinations of flow coefficients and reduced frequencies, they remain the dominant source of unsteady loss generation.

Numerical study on flow control mechanism of endwall fence in a high-lift turbine rotor with open separation

The enhanced cross pressure gradient and the interference effect between secondary flow and separation bubble on the suction side make the refinement of flow organization in the endwall region a key role in improving the performance of low Reynolds number turbine stages. This influence is further amplified in high-lift turbines with open separation. The flow control method of applying endwall fence on a high-lift rotor of a low-speed turbine stage has been numerically studied. By analyzing the transformation of separation bubble on the suction surface and the development of vortices in the endwall region, the flow control mechanism of the endwall fence at low Reynolds numbers is presented. The influence of different fence design parameters on aerodynamic performance has been summarized. The research results indicate that the induced generation of fence vortex can effectively suppress passage vortex. After installing the fence, the intensification of blockage in the endwall region near the leading edge effectively delays the occurrence of laminar separation. Due to the introduction of additional endwall losses, the flow control effect of the fence mainly comes from the suppression of separation. The flow field in the endwall region and flow control effect are significantly affected by the pitchiwse position and height of the fence, while the width of the fence has a slight impact. The flow control effect can be more effectively achieved by designing the height variation of the fence reasonably. In addition, numerical results indicate that the optimized fence all exhibit good control effectiveness under different operating conditions.

Unsteady vane-rotor secondary flow interaction of the endwall region in a 1.5-stage variable-geometry turbine

Variable-geometry turbines present significant improvements in both off-design cycle efficiency and enhanced engine responsiveness for future adaptive cycle engines. The adjustable vanes regulate the through-flow capacity by varying the installation angle and thus unavoidably require endwall clearance, resulting in profound implications for vane and downstream blade rows. Unsteady numerical simulations are carried out to investigate the vane-rotor flow interaction under the zero and partial clearance type of the adjustable vane in a 1.5-stage variable-geometry turbine. First, specific vortical structures and loss characteristics of upstream adjustable vane leakage flow at turning angles of −5°, 0°, and +5° are described. Then, the basic mechanisms of vane-rotor secondary flow interaction between acquired upstream leakage flow and downstream blade rows secondary flow is explored by the time-resolved method and quantitative detailed analysis method. The results show that adjustable vane vortex core area is the primary cause of internal loss within the vortex, and the mixing and interaction between the leakage vortex and other vortex systems contribute to secondary loss in the vane. Due to the downstream blade shear action caused by different flow velocity on the suction and pressure surfaces, the vane leakage flow fragments upon entering the R1 blade passage and develops downstream of the endwall region. Compared to the vane with zero clearance case at same turning angle, the intensity of the R1 tip leakage vortex at the design turning angle and open turning angle is reduced by 47.4 % and 23.66 % with the upstream partial clearance, respectively. The unsteady vane-rotor secondary flow interaction under different turning angles has large influence on the flow structure and loss characteristics of the downstream blade rows, and results in significant differences in the variation pattern and trend direction of the turbine performance.

Uncertainty quantification of turbine endwall aero-thermal performance under combustor-turbine interface louver coolant and cavity

The uncertainty characteristics have a significant influence on turbine cooling performance and its reliability, especially for the combustor-turbine interface region which faces more complicated flow interactions. Previous researches under the certainty framework are not able to reflect the realistic cooling characteristics of the interface region under real operation circumstances. Considering the random fluctuations of the major structure and flow parameters, uncertainty quantification of turbine endwall cooling performance under realistic combustor-turbine interface conditions is conducted in this paper using the polynomial chaos expansion method. The mix characteristic of combustor louver coolant and interface cavity leakage in the turbine endwall region is investigated, and the influences of parameter uncertainties on endwall cooling, vane phantom cooling, and cascade aerodynamic characteristics are analyzed. The results indicate that the uncertainties at the combustor-turbine interface largely influence the endwall cooling by modifying the strength and position of the secondary flows. Under the random fluctuations of interface geometry parameters, the local mean value of endwall cooling effectiveness can reach up to 0.25, and the standard deviation of endwall cooling effectiveness can reach 0.1. The geometrical parameters of interface width and misalignment height are the decisive factors for endwall cooling uncertainty. The random fluctuations of interface width and misalignment height separately cause great cooling effectiveness uncertainties on the pressure side endwall and suction side hot ring. The uncertainties in mainstream and structure parameters exhibit a reduced propensity to propagate onto vane phantom cooling. The decisive factors for vane phantom cooling uncertainty are interface width and misalignment height, and the major influenced area is the vane middle part. Four mainstream and structure parameters all make great contributions to the aerodynamic uncertainty at the strong secondary flow region and vane wake flow region. This paper can provide a theoretical basis for the design of combustor-turbine interface cooling structures to accommodate the actual turbine operation uncertainty conditions.

Sensitivity of the endwall flow in a linear vane cascade to blade fillet geometry

Based on the blade chord and inlet velocity, the current computational study uses a linear vane cascade with a large filleted blade-endwall junction with a 2.01 x 105 Reynolds number. Three fillets with related profiles are explored. To evaluate the upshots of geometric differences in a fillet attached to the endwall flow-field, the height and endwall-width of the fillets are changed. The RANS k-ω turbulent model is used in the computations, and the results are compared to experimental results from a similar cascade without the fillet. The computed results of the secondary flow-field in the endwall region along the cascade are compared for baseline (no fillet) and filleted passages. As a result of diminished leading-edge and passage vortices, the fillets lower pitchwise pressure gradients, flow separation, axial vorticity, and overall pressure losses when compared to the baseline. The pros of fillets on endwall secondary flows are however unaffected by fillet's geometric changes.

Unsteady Interaction Between Purge Flow and Secondary Flow in High-Lift Low Pressure Turbine

Abstract The coupling effect between the complex vortex system and the rim purge flow in the endwall region of a high-lift low pressure turbine (LPT) will significantly influence the evolution of secondary flow and the corresponding losses. This paper focused on the unsteady interaction mechanism between the rim purge flow and the secondary flow inside the high-lift LPT under the periodic wake passing. The large eddy simulation (LES) method was used to reveal the influence mechanism of rim purge flow, rotor-stator cavity interaction and unsteady wakes on the secondary flow. Detailed experimental measurement was carried out for the flow field of high-lift LPT under the influences of static purge flow. The results showed that the rim purge flow significantly increased the overturning and underturning downstream of the endwall region, resulting in aggravating secondary loss. The rotating-rim increased the difference of the circumferential velocity between the mainstream and the purge flow, aggravating the Kelvin-Helmholtz (K-H) instability, inducing the K-H vortex structure with stronger energy amplitude, and further increasing the endwall flow loss. The incoming wakes weakened the energy amplitude of the K-H vortex at the rim purge outlet, thinned the thickness of the low-energy fluid at the blade leading edge. Additionally, the interaction between the wakes and the secondary vortices further suppressed the development of the secondary flow. Nevertheless, the incoming wakes caused additional mixing losses and made a negative impact on the overall aerodynamic performance of the high-lift LPT.

Film cooling effect of upstream jump coolant on turbine endwall with combustor-turbine interface misalignment

The investigation of jump cooling performance on turbine endwall independently causes its actual cooling effectiveness to deviate from the design values. In this paper, the endwall jump cooling structure with combustor-turbine interface cavity and combustor liners is modeled. Taking the realistic combustor-turbine interface features into account, the effect of jump coolant on turbine endwall film cooling and heat transfer characteristics is numerically studied. Under different endwall misalignment modes, the aerodynamic interaction mechanisms of realistic combustor outflow profile, jump coolant jet and cascade secondary flows at turbine endwall region are revealed. The modification effects of combustor-turbine interface features on jump cooling of the endwall are investigated. The results show that the combustor outflow severely ingests into the combustor-turbine interface cavity after flowing across several steps. Two branches of cavity vortex are subsequently generated in the middle-pitch region of cascade to affect the jump coolant jet. At low coolant blowing ratio, the attachment and separation sides of cavity vortex individually result in high and low cooling regions, while the horseshoe vortex leads to a large wedge-shaped cooling region. At z/Cax < 0, the backward step leads to a higher cooling effectiveness than forward step by 0.2. The jump coolant only leads to phantom cooling effect on downstream part of the vane suction side surface. The net heat flux reduction (NHFR) between two cavity vortexes and at pressure side junction are individually 0<NHFR<0.2 and −0.2<NHFR < −0.1. At high blowing ratio, the jump coolant can cover the pressure side junction with the highest cooling effectiveness of 0.5. The jump coolant can flow across the horseshoe vortex and directly upwash the vane surface. The weak protection regions between two cavity vortexes and near the attachment line of the pressure side horseshoe vortex enlarges. Compared with the forward step, the backward step achieves a lower NHFR by up to 0.1 and a higher Nusselt number. This paper provides an in-depth analysis of the aero-thermal physics of endwall jump cooling concerning realistic combustor-turbine interface conditions.

Integrated optimization of a turbine stage at a low Reynolds number via NURBS surface and machine learning

As the turbine load increases, both the blade profile losses and secondary flow losses in the endwall region cannot be ignored for turbines operating at low Reynolds numbers. To fully explore the flow control effects of the blade and endwall shapes, a viable integrated parameterization method and optimization method are formulated for turbine stages. NURBS surfaces are utilized for parameterizing the endwall and blade geometries. Combined with machine learning methods and improved chaos particle swarm optimization (ICPSO), an efficient surrogate model optimization strategy for high-dimensional small-sample problems is discussed. The above methods are applied to the integrated optimization of the endwalls and suction surfaces of a low-speed turbine stage, which is operated at a low Reynolds number. Under the constraint of nearly unchanged output power, the isentropic efficiency of the turbine stage is improved by 1.33%. The flow control effects of blade reshaping and endwall contouring are separately compared. The results show that the flow structures near the suction surfaces of the blades can be adjusted by blade reshaping. Moreover, the flow near the endwalls can be reorganized by using contoured endwalls and optimized blades, and the secondary flow in the endwall region can thus be significantly weakened.

Investigation of Transition Flow on Suction Surface of Highly Loaded Compressor Cascade Controlled by Plasma Actuator

A practical method for improving the performance of highly loaded compressor cascades is to delay the onset of transition. Plasma actuators are used on the suction surface to manage the transition from laminar to turbulent by injecting momentum into the boundary layer. An experimental study was conducted in a low-speed wind tunnel to validate the actuator's control capability and the numerical simulation model. Both experimental and numerical results show that actuators located at various positions can improve the performance of the cascade and reduce flow loss to varying degrees. The most effective method is to place the actuator upstream of the separation bubble to delay the onset of the transition from laminar to turbulent flow, which results in a 6.32% decrease in total pressure loss and a 2.5% increase in static pressure rise. The control scheme at the transition position restores laminar flow locally, causing a secondary transition. However, it has the lowest control capability due to the dissipation of the actuator momentum. The downstream control scheme only suppresses the degree of turbulent separation without delaying laminar separation, resulting in weaker control effectiveness compared to the upstream laminar separation scheme. The study also examined the interaction between the plasma actuator on the suction surface and the flow in the endwall region, which showed that the accumulation of low-energy fluid in the corner region increases, resulting in a slight elevation of shear losses between the midspan boundary layer and the corner region.

Vortex dynamics characteristics in the tip region based on Wray–Agarwal model

In order to solve the blockage effect and energy dissipation phenomenon caused by cavitation in the low-pressure vortex core region, this paper analyzes the spatial evolution of vorticity intensity and turbulent kinetic energy intensity under different cavitation conditions based on the Wray–Agarwal (WA) model. First, the tip leakage flow characteristics are studied, the evolution of vorticity and vorticity intensity is analyzed, then the distribution of turbulent kinetic energy distribution in the blade tip region is studied, and finally, the vorticity transport characteristics of the tip region are analyzed. It is found that the tip leakage rate is less affected by the vortex cavitation of the tip leakage, and there is a strong interaction between the leakage flow at the tip leading edge and the trailing edge, and the separation vortices and low-speed regions formed in the end-wall region cause blockage of the flow passage. Low pressure causes cavitation to cover most regions of the suction surface, inhibiting the formation and development of the tip leakage vortices. The distribution range of high turbulent kinetic energy region is almost the same as that of high-vorticity region, and there is a positive correlation between the two intensities. Severe cavitation causes the high turbulent kinetic energy region at the outlet of the flow passage to develop in the radial and axial directions of the impeller, which increases the turbulent dissipation and energy loss. The change of vorticity transport intensity caused by cavitation is mainly reflected in the expansion contraction term, and the Coriolis force term plays a dominant role in the vorticity transport process. This paper provides a reference for further improving the performance of mixed-flow pumps.

A Numerical Test Rig for Turbomachinery Flows Based on Large Eddy Simulations With a High-Order Discontinuous Galerkin Scheme—Part III: Secondary Flow Effects

Abstract In this final paper of a three-part series, we apply the numerical test rig based on a high-order discontinuous Galerkin scheme to the MTU T161 low-pressure turbine with diverging end walls at off-design Reynolds number of 90,000, Mach number of 0.6, and inflow angle of 41 deg. The inflow end wall boundary layers are prescribed in accordance with the experiment. Validation of the setup is shown against recent numerical references and the corresponding experimental data. Additionally, we propose and conduct a purely numerical experiment with upstream bar wake generators at a Strouhal number of 1.25, which is well above what was possible in the experiment. We discuss the flow physics at midspan and in the end wall region and highlight the influence of the wakes from the upstream row on the complex secondary flow system using instantaneous flow visualization, phase averages, and modal decomposition techniques.

Effects of Gas Thermophysical Properties on the Full-Range Endwall Film Cooling of a Turbine Vane

To protect turbine endwall from heat damage of hot exhaust gas, film cooling is the most significant method. The complex vortex structures on the endwall, such as the development of horseshoe vortices and transverse flow, affects cooling coverage on the endwall. In this study, the effects of gas thermophysical properties on full-range endwall film cooling of a turbine vane are investigated. Three kinds of gas thermophysical properties models are considered, i.e., the constant property gas model, ideal gas model, and real gas model, with six full-range endwall film cooling holes patterns based on different distribution principles. From the results, when gas thermophysical properties are considered, the coolant coverage in the pressure side (PS)-vane junction region is improved in Pattern B, Pattern D, Pattern E, and Pattern F, which are respectively designed based on the passage middle gap, limiting streamlines, heat transfer coefficients (HTCs), and four-holes pattern. Endwall η distribution is mainly determined by relative ratio of ejecting velocity and density of the hot gas and the coolant. For the cooling holes on the endwall with an injection angle of 30°, the density ratio is more dominant in determining the coolant coverage. At the injection angle of 45°, i.e., the slot region, the ejecting velocity is more dominant in determining the coolant coverage. When the ejecting velocity Is large enough from the slot, the coolant coverage on the downstream endwall region is also improved.

Experimental decoupled-analysis of overall cooling effectiveness for a turbine endwall with internal and external cooling configurations

High-efficient composite cooling structures that are comprised of external and internal cooling techniques, are one of the key technologies to guarantee the aerothermal performance and reliability of a turbine vane, especially the endwall region due to unique complex flow structures. It is thereby critical to understand the flow mechanism and conjugate heat transfer characteristics in this region for further optimal design. In this paper, a representative endwall model which matches non-dimensional parameters to real turbine operating conditions is investigated. Thermal protection for the endwall is provided by internal jet impingement and external purge flow and discrete hole injection cooling. Comprehensive measurements of overall cooling effectiveness over the endwall, total pressure loss coefficient, and non-dimensional temperature at the cascade exit are performed by using an infrared (IR) thermography technique, a five-hole probe, and a temperature probe. Through decoupled analysis, heat transfer and flow characteristics of film cooling only, both jet impingement and film cooling, upstream purge cooling only, and combined cooling are analyzed, respectively, and their optimal ranges of cooling air are obtained. The measurements reveal that with film cooling only, only the regions around the film holes are covered. Using both film and jet impingement cooling, cooling coverage is expanded, and cooling effectiveness is much higher with a more uniform distribution pattern. Purge flow achieves thermal protection for upstream regions of endwalls, and the combined cooling scheme that involves all cooling sources provides the best cooling performance. Globally, those cooling schemes have no obvious effects on overall cascade aerodynamic performance but affect the development and distribution of secondary flows.

Nonuniform height endwall fence optimization of a low-pressure turbine cascade

As the lift of turbine increases, the secondary flow in the endwall region seriously restricts improvement of the aerodynamic performance of low-pressure turbine. The particle swarm optimization is improved to pursue stronger optimization ability. An optimization platform is built that combines Bayesian optimization (BO), the eXtreme Gradient Boosting (XGBoost) algorithm and improved chaos particle swarm optimization. The test results show that the surrogate model optimization method proposed in this paper has more prominent application potential in the field of aerodynamic optimization. The above optimization method is used for the optimization design of the nonuniform height endwall fence (NHEF) of a high-lift low-pressure turbine cascade to achieve an efficient secondary flow control effect. The NURBS curve is used to realize the parameterization of the NHEF. The secondary flow control mechanism of traditional fences and NHEF has been compared and analysed in detail. The numerical results show that the NHEF can effectively prevent the cross migration of the boundary layer, and the fence vortex induced by the fence can suppress the development of the passage vortex. The total pressure loss at the outlet of the cascade installed NHEF is reduced by 14.82%. The secondary flow control effect of the NHEF is better than that of a traditional fence. The feature importance analysis results of input variables show that the pitchwise position of the fence has the most obvious influence on the loss of the cascade. In addition, based on the results of feature importance, a 53% length reduction fence is designed to reduce the extra weight brought by the fence. The design achieves the effect of comprehensively considering the weight and aerodynamic performance.

Vortex Structure Topology Analysis of the Transonic Rotor 37 Based on Large Eddy Simulation

Highly three–dimensional and complex flow structures are closely related to the aerodynamic losses occurring in the transonic axial–flow compressor. The large eddy simulation (LES) approach was adopted to study the aerodynamic performance of the NASA rotor 37 for the cases at the design, the near stall (NS), and the near choke (NC) flow rate. The internal flow vortex topology was analyzed by the Q–criterion method, the omega (Ω) vortex identification method, and the Liutex identification method. It was observed that the Q–criterion method was vulnerable to being influenced by the flow with high–shear deformation rate, especially near the end–wall regions. The Ω method was adopted to recognize the three–dimensional vortex structure with a higher precision than that of the Q–criterion method. Meanwhile, the Liutex vortex identification method showed a good performance in vortex identification, and the corresponding contribution of Liutex components in the vortex topology was analyzed. The results show that the high–vortex fields around the separation line and reattachment line had high vortex components in the x–axis, the tip clearance vortices presented a high–vortex component in the y–axis, and the suction side corner vortex possessed high–vortex components in the y– and z–axes.

Film-cooling hole optimization and experimental validation considering the lateral pressure gradient

The flow in the turbine endwall region consists of the complicated secondary flow structures driven by the lateral pressure gradient, which heavily affects the performance of film cooling. In this work, the film-cooling hole design optimization is performed considering the existence of the lateral pressure gradient in the real flow environment. Results have shown that the optimal film-cooling hole design is heavily influenced by the lateral pressure gradient in the endwall region, especially the compound angle design is clearly different from the flat plate flow environment. The optimization results are further validated with experiments using the pressure-sensitive paint (PSP) technique, and the film cooling performance is shown to be improved by 42.9%. This work demonstrates the importance of considering the real flow environment in the film-cooling hole design and also can provide guidance to the film-cooling hole design in the endwall region.

Control mechanism of secondary flow in a turbine cascade with non-axisymmetric endwall profiling under Co-rotating incoming vortex

Abstract Upstream vortex has a significant effect on the secondary flow structure of the downstream turbine in the stage environment. This study investigates the secondary flow structure with non-axisymmetric endwall profiling (NAEW) under the interaction of co-rotating incoming vortex (Vic). A half-delta wing vortex generator is utilized to model Vic. The turbine cascade case which exhibited maximum reduction of the cascade loss with NAEW under no incoming vortex is studied. The mechanism of loss reduction with NAEW under the interaction of Vic is analysed. Vic could decrease the secondary flow near the endwall region by affecting the horseshoe vortex transport in the cascade. However, its loss reduction was lower than the loss increments of Vic itself. The arrival of Vic at the leading edge of the cascade increased the strength of the horseshoe vortex, resulting in a significant increase in loss. Under the interaction of Vic, NAEW decreased the blade loading near endwall region, which resulted in the reduction of cascade loss.

Experimental study on the aerodynamic performance of variable geometry low-pressure turbine adjustable guide vanes

An annular cascade testing rig of a variable-geometry low-pressure turbine with a large expansion angle was designed and built, combining a five-hole probe, a surface static pressure test system, and a PIV(Particle Image Velocimetry) measurement system. The effects of exit Mach numbers (0.2, 0.3, 0.4), turning angles (−1.5°, 0°, 3°), and adjustable device diameters (30 mm, 40 mm) on turbine cascade aerodynamic performance were experimentally investigated. Numerical simulations were also carried out to gain more insight into the flow fundamentals. Results show that the endwall clearance and the adjustable device significantly influence the flow field, with leakage flows before, after, and around the disc. The leakage vortex and passage vortex formation resulted in two high-loss regions in the shroud and hub endwall regions. The aerodynamic loss structures at the exit section were similar at different Mach numbers and varied considerably at different turning angles. Besides, a larger area of the adjustable device (d = 40 mm) was found to reduce the leakage loss caused by upper and lower clearance leakage flows and inhibit the development of secondary flows, resulting in a more straightforward secondary flow structure and a smaller influence range. Quantitatively, the adjustable device with a diameter of 40 mm could decrease the average total pressure loss coefficients by 11.5%, 16.6%, and 6.3% at three turning angles (−1.5°, 0°, 3°), respectively, compare to the adjustable device with a diameter of 30 mm.

Integrated passage design based on extended free-form deformation and adjoint optimization

Inspired by the three-dimensional design of flow passages in turbomachinery, this study proposes the concept of integrated passage design. The capability of adjoint method for efficient optimization and the flexibility of the parameterization method based on extended free-form deformation have been considered to develop a feasible approach to design an integrated passage. This concept was applied to redesign a typical transonic fan, Rotor 67, and the results were analyzed by CFX. It is shown that the passage was adequately adjusted in all three dimensions and reduced the strength of shock wave and wake-induced flow. In particular, the secondary flow was appropriately reorganized and the corner separation was well controlled in the end wall region, leading to significant improvements in adiabatic efficiency and diffusion.

Turbine nozzle endwall aero-thermal characteristics under combustor louver coolant with interface cavity

Targeting on combustor-turbine interface cooling system of combustor louver coolant and interface cavity, the endwall film cooling and heat transfer characteristics as well as the vane phantom cooling performance of a nozzle guide vane are investigated by solving the Reynolds-Averaged Navier-Stokes equations, and the secondary flow structure at endwall region are analyzed. The results show that most of the louver coolant leads to the core cooling region on the endwall after recirculating in the cavity with the cavity vortex, while a minority of the louver coolant escapes from the cavity vortex and forms the outer endwall cooling region limited by the horseshoe vortex. Under a certain mainstream flow condition, the endwall cooling performance and net heat flux reduction (NHFR) only depend on the mass flow ratio (MFR) of louver coolant instead of blowing ratio (BR). With increasing louver BR, the coolant mainly leads to the increase of the width and film cooling effectiveness of the core region at high MFR, but it leads to increasing cooling effectiveness of the outer region at low MFR. When the louver MFR increases from 0.25 % to 1.91 %, the averaged endwall cooling effectiveness at z/Cax = 0.1 is increased from 0.04 to 0.21. The NHFR of the core and outer cooling region reaches 0.3 and 0.1, respectively. The pressure and suction side horseshoe vortexes cause two protection failure regions, and its NHFR further decreases with increasing BR. At high MFR, the heat transfer in the low Nusselt number regions resulted by cavity vortex separation is further weakened with the increase of BR, and the region is enlarged toward two sides of the cascade. When MFR = 1.91 %, the Nusselt number of the high heat transfer region at z/Cax = 0.7 decreases from 2200 to 2000, and the averaged NHFR of fore part endwall reaches up to 0.2. This paper provides guidance for endwall cooling configuration design by taking the compound effect of the combustor-turbine interface cooling scheme into account.