Search

The performance of turbomachinery is largely affected by the nonuniform boundary conditions caused by the coupling between neighboring parts, such as the inlet distortion and hot streak. Existing works study this problem by comparing the flow fields with uniform and nonuniform boundary conditions, which is cost extensive. In this work a new and efficient method is developed by computing the sensitivities of arbitrary performance metric relative to the boundary conditions and this method quantitatively describes how the disturbance on the boundary surface affects the performance. The adjoint method popular in the design optimization is extended to current field for the efficient computation of the sensitivity. The extra cost is independent of the number of parameters describing the nonuniform distribution. Accuracy of the adjoint boundary sensitivity is validated against analytical results for a converging–diverging nozzle case. Its applications in a representative case are detailed and discussed in this paper. Current method provides a new tool, also direct guides to minimize the negative effect and improve the performance utilizing the adjoint boundary sensitivity.

The tip leakage flow passed over the tip clearance makes the flow very complicated near the tip gap, and the interaction of the tip leakage vortex and endwall vortex enhances the instability of the flow. Accurately capturing detailed flow structures and investigating the relationship between the flow structures and loss are beneficial for understanding the flow physics and providing guidance on reducing the loss. Due to the conventional Reynolds Averaged Navier-Stokes (RANS) methods is limited to predict the complex turbulence structures of the tip clearance flow, high fidelity simulation approaches are needed. In this work, the hybrid RANS/Large Eddy Simulation (LES) is adopted to simulate the tip leakage flow in linear cascade and demonstrates its ability to capture the small-scale flow structures. With the POD method, the time-averaged flow field and the dominating modes are obtained. Based on the analysis of the POD modes, it is found that the induced vortex generated by the interaction between the leakage vortex and the endwall vortex has strong turbulence characteristics. Based on the entropy generation rates, viscous loss mechanism is further analyzed. It is found that the shear strain rates dominate the viscous dissipation losses, and the fluctuation dissipation has a strong local enhancement effect.

In modern high-load high-pressure turbine, the secondary flow in the blade channel is very strong and occupies a large spanwise region. Although high-quality experimental data at the stage interfaces have been obtained in previous research, the influence of the clocking position on the secondary flow patterns is not fully understood. This paper investigates the clocking effect in a 1.5-stage high-pressure turbine and focuses on the variations of secondary flow patterns and their effect on the turbine performance. The detailed flow fields of various clocking positions were obtained by carrying out unsteady flow simulations using an in-house code. Among the four clocking positions in this work, the highest entropy generation was observed when the wakes from stator 1 hit the leading edges of stator 2, which is opposite to the well-known conclusion for the turbine with high-aspect-ratio blades. Detailed flow analysis showed that the wakes and the near tip secondary vortices from stator 1 showed different traces when entering the stator 2 channel and the secondary vortices clearly have a more important influence in determining the performance. The different behaviors of the secondary vortices explained the performance variations due to the clocking effect.

With the relative movement of neighboring blade rows, flows in multi-stage turbomachinery are unsteady and periodic in time at the design condition. As an alternative to the widely used time domain time marching method, the harmonic balance (HB) method has been successfully applied to simulate the essentially unsteady flow of multi-stage turbomachinery. By modelling various number of harmonics, the accuracy of this method could be adjusted at different level of computational cost. In this article, accuracy of the harmonic balance method is not only validated against the time domain time marching method, as in most previous works on this topic, but also against the data from an experiment campaign of a two-stage high-pressure turbine where strong tip leakage flow exists. Efficiency of this method is also assessed in detail by adjusting the number of harmonics and comparing with time domain time marching solution results. Results show that the harmonic balance method is a flexible tool with adjustable accuracy for fast-turnaround unsteady flow simulation of multi-stage turbomachinery. Results from this work can provide a guidance in applying the harmonic balance method with balance between accuracy and computational cost.

This paper presents a novel theory regarding the blade loading and the passage flow field within general turbomachineries. The basic philosophy is to establish an analytical relation between the loading, the flow angle, and the blade geometry based on the conservation of energy. Detailed validations and analyses will be carried out to provide a general scope regarding the theory itself as well as its advantages and limitations in common applications. The paper includes the theoretical derivation of the target relation. The starting point is the standard RANS equations. From that, with the aid of the passage-average operator, the relation between the loading and the passage flow field is derived under the energy balance. Theoretical analyses regarding the validity of the relation are performed based on the simulation results and test data on different cascades. Discussions are conducted regarding the assumption and potential applications of the theory. Conclusions are drawn on the applicability of the theory to introduce its potential applications in general turbomachineries.

Accurate predictions of Shock Waves and Boundary Layer Interaction (SWBLI) and strong Shock Waves and Wake Vortices Interaction (SWWVI) in a highly-loaded turbine propose challenges to the currently widely used Reynolds-Averaged Navier-Stokes (RANS) model. In this work, the SWBLI and the SWWVI in a highly-loaded Nozzle Guide Vane (NGV) are studied using a hybrid RANS/LES strategy. The Turbulence Kinetic Energy (TKE) budget and the Proper Orthogonal Decomposition (POD) method are used to analyze flow mechanisms. Results show that this hybrid RANS/LES method can obtain detailed flow structures for flow mechanisms analysis. Strong shock waves induce boundary layer separation, while the presence of a separation bubble can in turn lead to a Mach reflection phenomenon. The shock waves cause trailing-edge vortices to break clearly, and the wakes, in turn, can change the shocks intensity and direction. Furthermore, the Entropy Generation Rate (EGR) is used to analyze the irreversible loss. It turns out that the SWWVI can reduce the flow field loss. There are several weak shock waves in the NGV flow field, which can increase the irreversible loss. This work offers flow mechanisms analysis and presents the EGR distribution in SWBLI and SWWVI areas in a transonic turbine blade. Keywords: Boundary layer, Hybrid RANS/LES, NGV, Shock waves, Wakes

Efficient method to handle the geometric constraints in the optimization of turbomachinery blade profile is required. Without constraints on the blade thickness, optimal designs typically yield thinner blade to reduce the friction loss, however, at the risk of degraded strength and stiffness. This issue is seldom discussed and existing literature always treat the blade thickness constraint in an indirect manner. In this work, two different geometric constraints on the blade thickness are proposed and applied in the adjoint optimization: one is on the maximum blade thickness and the other is on the blade area. Methods to compute sensitivities of both constraints are proposed and they are integrated into an optimization system based on a finite volume code and a solver for the discrete adjoint equation. Adjoint optimization is conducted to minimize the entropy production in a transonic compressor stage. Results for the adjoint optimization without geometry constraint and two comparative cases are detailed. It is indicated that three cases yield similar performance improvement; however, if geometry constraints are properly handled, the optimal designs have almost the same maximum thickness as the original design, compared to a thinner blade profile with 14% reduction of maximum thickness in the case without geometry constraint. The cases considering geometry constraints also consume slightly reduced Central Processing Unit (CPU) cost. Result of this work verifies the effectiveness of the proposed method to treat geometric constraints in adjoint optimization.

An optimization framework is developed for the film cooling hole shape design to achieve highest film cooling effectiveness with affordable cost. A key to this design method is the development and applications of a multi-fidelity model (MFM) to improve the efficiency of surrogate model construction. To construct multi-fidelity model, empirical correlations from existing literature are utilized as low-fidelity data and Reynolds-Averaged Navier-Stokes (RANS) simulations provide high-fidelity data. Compared with single-fidelity surrogate models, multi-fidelity model developed in this work combines vast existing knowledge about shaped film cooling hole and three dimensional numerical simulation. As a result, the computation cost has been substantially decreased by 64.5% while accuracy is maintained with similar level, which provides a firm basis for the optimization. Considering three geometric parameters as design variables, i.e. laidback angel, lateral angle, and hole length-diameter ratio, the shaped hole is optimized with genetic algorithm (GA) combined with sequential quadratic programming (SQP) algorithm. After the optimization, an experiment campaign is further carried out to validate the optimization result using pressure sensitive paint (PSP) technology. With current multi-fidelity model, the spatially averaged film cooling effectiveness has been improved by 39 % compared to the baseline, which is an already well-designed shaped hole and the improvement is also verified by experiment. This work verifies the effectiveness of current multi-fidelity model for the design and optimization of shaped film cooling hole for better performance.

In modern aero-engine and gas turbine, the hot streak and residual swirl caused by the combustor have strong impact on both aero and thermal performances of the high-temperature turbine. Due to different design philosophies of combustors and turbine, one combustor usually matches two or more vanes and the resultant mismatching effect is numerically investigated with unsteady simulation where one nonuniform core faces two Nozzle Guide Vane (NGV) blades. Results indicate that the nonuniformity clocking position dominates the averaged-temperature, while the swirl orientation dominates the radial distribution. The mismatching effect strongly impacts both time-averaged and instantaneous performances, in that neighboring blades of NGV have averaged-temperature difference of 10.5 K, and it also halves the base frequency and amplifies unsteady variation in the rotor passage. By analyzing both aero and thermal results from a matrix of four cases, the LE-pos configuration has the best performance. The complex transportation of the nonuniformity in the LE-pos case is detailed and the unsteady interactions between the inlet nonuniformity and the complex flow topology are analyzed. The residual swirl carries the hot streak fluctuating radially in rotor passages and redistributes the temperature on rotor surface. Results from this work highlight the necessity of considering the mismatching effect in gas turbine design optimization.

The performance of film cooling is influenced by many parameters, and the nonuniform flow caused by the internal cooling system is found to largely affect the film cooling, which further complicates the in-hole flow and draws new difficulties in predicting the cooling performance. In this study, we find a very interesting phenomenon that there always exists an in-hole interface, on which distributions of many parameters, including the velocity and kinetic energy, are seldom affected by the mainstream. The existence of this specific interface can be observed for both cylindrical and shaped film cooling holes under most operating conditions. The theoretical analysis of this interface is conducted in this study based on the characteristic decomposition of the Navier–Stokes equation, and this interface is named as the characteristic interface. Theoretical analysis and numerical observations suggest the film cooling system can be simplified to two weakly coupled regions separated by this interface. It also explains why existing source term models for film cooling may fail. Based on these findings, a new prediction model is developed, which uses the convolutional neural networks (CNN) model to predict the boundary conditions on the characteristic interface. The new model outperforms existing source term models and yields similar accuracy as full-mesh computational fluid dynamics (CFD), while reducing the computational cost by one order of magnitude. This model is further evaluated in large eddy simulation (LES), showing moderate success. To sum up, the current work reports the characteristic interface phenomenon in the film cooling hole, based on which a new and efficient prediction model is developed and verified.

One-dimensional laterally averaged adiabatic film cooling effectiveness η¯lat-based correlations have been widely employed in the cooling design of the modern gas turbine and aero-engine; however, the flow field of the discrete film cooling is fully three dimensional, and thus, the cooling effectiveness distribution on the solid surface is two dimensional. Accurate prediction of the cooling effectiveness distribution in the lateral direction would help to optimize the film cooling design, but few paid attention to this issue in the literature. In this study, a simple yet accurate scalar diffusion equation based model is proposed to extend the one-dimensional correlation into two dimensional. The model is proved to be accurate and efficient. According to the accuracy analysis, the R2 value is larger than 0.95 for the two-dimensional prediction and over 0.93 along the centerline. With given input parameters, the calculation cost for solving a certain case is in the magnitude of 1 × 10−3s in time using the space-marching method. There is only the effective diffusion coefficient left to be modeled in the control equation. It represents the balance between the diffusion and the passive transportation by the main flow. Analyses conducted within the typical experimental range show that κ~eff is only dependent on the velocity ratio and the main-flow turbulence.

The source term model is an effective way to reduce the calculation consumption in predicting film cooling performance. The aim of this work is to develop a new source term model to consider the non-uniform in-hole flows. The internal cross flow that occurs in air-cooled blades and vanes has been shown to distinctly affect the cooling performance. In this work, the in-hole flow mechanism with the presence of the internal cross flow is studied. A topological model that divides in-hole flows into three parts and several regions is put forward. It is shown that, compared with the parallel ones, the perpendicular internal cross flows dominate the cooling performance. What is more, flows within the cooling hole are not affected by the main flow, especially in the region under a certain plane (named the R plane). This feature makes it possible for a source term model to work under various main flow conditions. Based on these understandings, a non-uniform source term model is established. It models and solves the three-dimensional in-hole flows by a two-dimensional transient solver and describes the effect of the internal cross flow as initial conditions using a vortex model. The result of the transient solver is then transformed into the axial velocity distribution at the R plane and is used as the source term. The prediction of the source term model is compared with full CFD under various VR, VRi and β, showing the ability of this model to predict the effect of the internal cross flow with discrepancy increasing with VR. Accurate modeling of the three-dimensional in-hole flows is the key point in this source term model and needs further development.

Multi-row film cooling is widely applied in modern gas turbine cooling designs. However, there are few effective methods to predict local cooling effectiveness and the Sellers model based on the existing 1-D correlation is popular in the community but its prediction accuracy is questionable in some cases. The primary work of the current study is to analyze the prediction error of the Sellers model and to reveal the underlying flow mechanism. Considering the lateral uniformity assumption of the 1-D Sellers model, the effect of pitchwise row-to-row misalignment can hardly be included and the 2-D Sellers model is, therefore, proposed. Numerical simulations under density ratio near 1.0 are conducted on the double-row film cooling. Configurations of inline and staggered arrangements and blowing ratios from 0.5 to 2.0 are included. Cooling effectiveness can be theoretically decomposed into two parts contributed by separate rows, and this method has been validated. The footprint of coolant can, therefore, be tracked. Based on vortical interaction analysis, conclusions are obtained that pitchwise row-to-row spacing determines the overestimation or underestimation of the model-predicted result and the blowing ratio affects the absolute value of the prediction error. By comparing prediction accuracy of the two rows, the conclusion can be drawn that the prediction error is, to a greater extent, contributed by the row #2 because it is more likely to be affected by the row-to-row interaction.

Film cooling is a typical three-dimensional fluid phenomenon, where the coolant with lower temperature is ejected from discrete holes to protect metal walls from being burnt by the hot mainstream. It is a great challenge for Reynolds-averaged Navier–Stokes (RANS) methods to accurately predict the coolant coverage on the wall because the turbulent thermal diffusion tends to be under-predicted due to inherent assumptions behind RANS models. In this paper, a framework of integrated field inversion and machine learning is built to enhance RANS prediction of turbulent thermal diffusion. A neural network (NN) is trained in this framework to predict the spatially varying turbulent Prandtl number ([Formula: see text]) and to improve the prediction of RANS models. The temperature distribution obtained from the large eddy simulation is used as the learning target, and the discrete adjoint method is used as the inverse model that helps calculate derivatives of mean square error of the temperature distribution to NN parameters. The training process of NN shows good convergence properties. The results show that the obtained NN effectively increases the insufficient turbulent thermal diffusion by predicting much lower [Formula: see text] than the commonly used value of 0.9. The NN-enhanced RANS provides significant improvements on predicting experimental temperature distributions compared with general RANS models not only on the training data but also on the unseen testing data. In addition, the obtained NN can be implemented into general-purpose software with minimal effort and no numerical stability problem.

Purpose Adopting large eddy simulation (LES) to simulate the complex flow in turbomachinery is appropriate to overcome the limitation of current Reynolds-Averaged Navier–Stokes modelling and it provides a deeper understanding of the complicated transitional and turbulent flow mechanism; however, the large computational cost limits its application in high Reynolds number flow. This study aims to develop a three-dimensional GPU-enabled parallel-unstructured solver to speed up the high-fidelity LES simulation. Design/methodology/approach Compared to the central processing units (CPUs), graphics processing units (GPUs) can provide higher computational speed. This work aims to develop a three-dimensional GPU-enabled parallel-unstructured solver to speed up the high-fidelity LES simulation. A set of low-dissipation schemes designed for unstructured mesh is implemented with compute unified device architecture programming model. Several key parameters affecting the performance of the GPU code are discussed and further speed-up can be obtained by analysing the underlying finite volume-based numerical scheme. Findings The results show that an acceleration ratio of approximately 84 (on a single GPU) for double precision algorithm can be achieved with this unstructured GPU code. The transitional flow inside a compressor is simulated and the computational efficiency has been improved greatly. The transition process is discussed and the role of K-H instability playing in the transition mechanism is verified. Practical/implications The speed-up gained from GPU-enabled solver reaches 84 compared to original code running on CPU and the vast speed-up enables the fast-turnaround high-fidelity LES simulation. Originality/value The GPU-enabled flow solver is implemented and optimized according to the feature of finite volume scheme. The solving time is reduced remarkably and the detail structures including vortices are captured.

In modern aero-engine and gas turbine, the highly loaded turbine leads to strong secondary flow by which the main flow and the film cooling jets are driven from the pressure side to the suction side. As a result, the cross flow has considerable influences on the performance of film cooling jets in the endwall region. In this work, the effects of cross flow on the shaped-hole film cooling performance are investigated both experimentally and numerically. Experimental studies are conducted in a curved channel which simulates the real turbine blade environment with the blowing ratio from 0.5 to 2.5. Numerical studies are also conducted to obtain detailed information about vortical structures. Results show that the cross flow changes the relative position of the two pairs of kidney vortices generated at the outlet of the shaped hole. The negative kidney vortices are stuck to the endwall due to the lift force and they decay faster downstream. The coolant is driven by the cross flow to migrate along the radial pressure gradient direction. And the coolant near the endwall migrates further. The peak value of the cooling effectiveness is suppressed; however, the laterally averaged cooling effectiveness is increased, especially at high blowing ratio. An analytical model describing the flow topology about coolant jets with cross flow is proposed and results from this work may help to understand the film cooling behaviour affected by real complicated flow structures in the engine environment.

The swirl generator is widely used in lean-burn combustor to guarantee flame stability and reduce NOx emissions. Thus, the non-uniformities induced by the swirler would affect and damage film-cooling effectiveness on the turbine components, even blow-off coolant coverage, to some certain extent. The arrangement of film-cooling holes was normally designed to be perpendicular to the axial direction and in standard straight row. In this work we experimentally studied the effects of inlet swirl and mass flow ratio on traditional film cooling holes arrangement and a new arrangement pattern whose holes are located along the isobars. Results indicated that the swirl perturbation would damage film coverage. The film-cooling effect of endwall on which the holes are along the isobars, is invariably more promising than that of endwall on which hole arrangement is perpendicular to axial with the same mass flow of coolant, whether the inlet conditions is uniform or not. Keywords: Films-cooling, Holes locating, Isobaric, Swirling flow, Turbine design

This work develops a new and effective method to predict the 2-D cooling effectiveness distribution. The assumption in the existing 1-D method is abandoned and replaced by a scalar diffusion equation to model the lateral spreading of the coolant. The proposed model can be used for predicting 2-D multi-row cooling performance and is validated over four arrangements of double-row cooling holes against in-house experimental data. The paper further presents the advantages of the current method over the classic 1-D method. 2-D distributions for different multi-row configurations are successfully obtained and the relative error is reduced to 19.77% as opposed to that of 45.65% with the 1-D model. The time consumption by the proposed method is on the order of 1 × 1 0 − 3 s which is exceedingly efficient for the cooling design process. To sum up, the proposed method is accurate, capable of 2-D multi-row modeling and of simple form and high efficiency. Hence it is potentially an effective tool for film cooling design.

Strong secondary flow results in substantial aerodynamic loss for highly-loaded turbine. Accurate prediction of the complicated flow structures proposes challenges to the widely used Reynolds Average Navier-Stokes (RANS) approach. This work employs the hybrid RANS/Large Eddy Simulation (LES) method to study the unsteady flows for a highly-loaded turbine blade, with both flat endwall and optimized contoured endwall. Evolution of the unsteady flows in the endwall region is analyzed, with emphasis on the loss generation mechanism. Results show that for the flat endwall, the horseshoe vortex system is highly unsteady and is a significant source of unsteadiness in both the passage and the wake region. It contributes to the unsteady passage vortex, also the earlier breakdown of the trailing edge shedding vortex. The Probability Density Function (PDF) histogram of velocity in the wake region is bimodal, implying the perturbations from two mechanisms. For the contoured endwall, the unsteady evolution of the horseshoe vortex is blocked, which results in significantly reduced unsteadiness, also the merge between the horseshoe vortex with the passage vortex is prevented. Effect of the flow unsteadiness on the loss generation is assessed based on the entropy generation rates contributed by the time-averaged flow and the fluctuations. The effect of unsteadiness is two-fold: unsteady perturbations trigger the vortex breakdown into small-scale structures and thus weakened wake velocity deficit and loss generation; however, in both the passage and the wake region the entropy generation by the fluctuations are remarkable. For the contoured endwall, due to the reduced flow unsteadiness, loss generation contributed by the fluctuations are much smaller compared to the flat endwall. The results highlight the importance of including the loss generation by the fluctuations, also a possible mechanism to reduce the secondary loss by attenuating the flow unsteadiness with the contoured endwall.

With the development of computational power, large eddy simulation (LES) method is increasingly used in simulating complex flow. However, there still exist many factors affecting the LES quality and appropriate mesh resolution is among one of them. This work aims to develop an automatic procedure to refine the LES mesh by combining adaptive mesh refinement (AMR) and LES quality criteria. An LES refinement criterion is developed by estimating the proper grid length scale which meets the accuracy requirement of LES method. With this criterion, the baseline mesh is automatically refined with the AMR method. In this work, an efficient one-shot refinement strategy is also proposed to reduce the overall simulation time. Current AMR-based LES method is verified with the typical LES test case about the flow past circular cylinder at Re D = 3900. Results show that the automatically refined mesh provides systematically better agreement with experimental results and with current method the balance between ac...

Summary Adaptive Mesh Refinement (AMR) shows attractive properties in automatically refining the flow region of interest, and with AMR better prediction can be obtained with much less labour work and cost compared to manually re-meshing or the global mesh refinement. Cartesian AMR is well established; however, AMR on hybrid unstructured mesh which is heavily employed in the high-Reynolds number flow simulation, is less matured and existing methods may result in degraded mesh quality, which mostly happens in the boundary layer or near the sharp geometric features. User intervention or additional constraints, such as freezing all boundary layer elements or refining the whole boundary layer, are required to assist the refinement process. In this work, a novel AMR strategy is developed to handle existing difficulties. In the new method, high-order unstructured elements are first generated based on the baseline mesh; then the refinement is conducted in the parametric space; at last the mesh suitable for the solver is output. Generating refined elements in the parametric space with high-order elements is the key of this method and this helps to guarantee both the accuracy and robustness. With the current method, three-dimensional hybrid unstructured mesh of huge size and complex geometry can be automatically refined, without user intervention nor additional constraints. With test cases including the two-dimensional airfoil and three-dimensional full aircraft, the current AMR method proves to be accurate, simple and robust. This article is protected by copyright. All rights reserved.

This work presents the implementation and study of the quadratic constitutive relation nonlinear eddy-viscosity model with representative compressor application, for which the corner separation has been poorly predicted with the widely used linear Boussinesq eddy-viscosity model. With the introduction of the Reynolds stress anisotropy, the secondary flow of the second kind and its effect on the corner flow can be well captured and this results in greatly improved prediction of pressure coefficient, total pressure loss coefficient and the corner separation size. Without the quadratic constitutive relation model, the separation size and loss are generally over-estimated. The mechanism of the improvement is studied using both the vortex dynamics and the momentum equation. It is proved that quadratic constitutive relation model consumes low CPU time and provides much improved compressor corner separation prediction without worsening the convergence property.

In the modern highly-loaded gas turbine, due to the large pressure difference between the suction side and the pressure side of the turbine blade, strong cross flow is formed and it strongly affects the aerodynamic and cooling performances in the end-wall region. The film cooling behavior in the environment of strong cross flow is different from the straight channel environment widely studied in the literature. In this research, the effect of cross flow on film cooling is investigated by Large Eddy Simulation (LES) using subgrid-scale (SGS) model. Numerical simulation is carried out in a curved passage to simulate the turbine blade passage. Shaped cooling hole with blowing ratio 1 is studied. The time-averaged friction line results are compared with existing experimental ink trace results. The vortex structures, both time-averaged and instantaneous, are analyzed to study the effect of cross flow on film cooling. At the exit of the cooling hole, the hanging vortices with negative y-vorticity are more flat in shape and closer to the wall in position in contrast to hanging vortex with positive y-vorticity, which is caused by cross flow and results in the asymmetry of hairpin vortices downstream as well as the asymmetry of the distribution of coolant. It has been shown that the vortices from mainstream have a significant impact on the field near the exit of the cooling hole. Those vortices interact with the hairpin vortices from the cooling hole and directly lead to the asymmetry of the hairpin vortices. Proper Orthogonal Decomposition (POD) analysis is further conducted to extract the dominant flow structures and the physical mechanisms of primary POD modes are given to explain the distribution of film cooling effectiveness affected by cross flow. Based on the specific situation in this work, a fast incremental POD (iPOD) approach is adopted since the rank of the field matrix is far less than the rows, which is caused by the tall and thin character of the matrix, which makes the analysis less costly and more effective. This research helps to understand the cooling performance in the real turbine blade passage and to explain the coolant mixing process based on the instantaneous flow field obtained using high precision LES simulation and powerful iPOD.

It has been known that in previous optimization problems, surrogate models are widely used to improve the optimization efficiency. The optimization result depends heavily on the accuracy of surrogate model, while it also costs a lot to construct a high-accurate surrogate model. In order to obtain the optimal design of a shaped film cooling hole which is already well-designed, an optimization framework based on multi-fidelity model is proposed in this article. The characteristic of the multi-fidelity model in this work is that it combines the experimental design information and numerical simulation together. The optimization efficiency can be greatly improved by 64.5% with multi-fidelity model and furthermore the prediction accuracy is maintained with a high level. In this work, three geometry parameters are selected as design variables, which are lateral angle, laidback angle and hole length. A combined optimization method which includes GA (generic algorithm) and SQP (sequential quadratic programming) algorithm is used in the optimization process. A validation experiment is further conducted to validate the optimized hole design using PSP (pressure sensitive paint) and 3D printing technology. The optimization results have shown that the overall film cooling performance is 39% higher compared with baseline design. This work shows the merits of multi-fidelity model in high-efficiency optimization film cooling holes and it may provide further guide for other relative works.

The complex leakage flow structure in the tip region of unshrouded rotor is a main source of turbine aerodynamic loss. Due to the complex turbulence characteristics of the tip leakage flow, the widely used Reynolds Averaged Navier-Stokes (RANS) approach may fail to accurately predict the multi-scale turbulent flow and the related loss. In order to effectively improve the turbine efficiency, more insights into the turbulence characteristics and the loss mechanism in the tip leakage flow are required. In this work, a Delayed Detached Eddy Simulation (DDES) study is conducted to simulate the flow inside a high pressure turbine blade, with emphasis on the tip region. DDES results are in good agreement with the experiment and the comparison with RANS results verifies the advantages of DDES in resolving finer flow structures of leakage flow, also in capturing the complex turbulence characteristics. The snapshot Proper Orthogonal Decomposition (POD) method is used to extract the dominant flow features. The flow structures and the distribution of Reynolds stress help to reveal the process of leakage flow and its interaction with the secondary flow. Meanwhile, it is found that the separation vortex (SV) forms from leading edge to trailing edge, and the strong interactions between tip leakage vortex (TLV) and passage secondary vortex (PSV) significantly enhance the turbulence intensity. Based on the DDES results, loss analysis of tip leakage flow is conducted based on entropy generation rates. For the leakage flow related loss, the largest local entropy generation rate occurs at 50 % of axial chord, and the interaction between the leakage vortex and up passage vortex promotes the loss generation. To sum up, the current DDES study about the tip leakage flow provides helpful information about the loss generation mechanism and may guide the design of low-loss blade tip.

Shock wave and tip leakage are important flow features at small length scales. These flow phenomena and their interactions play important roles in the performance of modern transonic fans and compressors. In most numerical predictions of these features, mesh convergence studies are conducted using overall performance data as criteria. However, less effort is made in assessing the quality of the predicted small-scale features using a mesh that yields a fairly accurate overall performance. In this work, this problem is addressed using the adaptive mesh refinement (AMR) method, which automatically refines the local mesh and provides very high resolution for the small-scale flow feature, at much less cost compared with globally refining the mesh. An accurate and robust AMR system suitable for turbomachinery applications is developed in this work and the widely studied NASA Rotor-37 case is investigated using the current AMR method. The complex interactions between the shock wave and the boundary layer, as well as those between the shock wave and the tip vortex, are accurately captured by AMR with a very high local grid resolution, and the flow mechanisms are analyzed in detail. The baseline mesh, which is considered to be “acceptable” according to the commonly used mesh convergence study, is unable to capture the detailed interaction between the shock wave and the boundary layer. Moreover, it falsely predicts the tip leakage vortex breakdown, which is a consequence of inadequate resolution in the tip region. Current work highlights the importance of a careful check of the mesh convergence, if small-scale features are the primary concern. The AMR method developed in this work successfully captures the flow details in the transonic compressor in an automatic fashion, and has been verified to be efficient compared with the globally mesh refinement or manually mesh regeneration.

A bleeding system is necessary for gas turbines and may largely affect the compressor performance. In the existing literature the circumferentially uniform bleeding is widely studied; however, it is different from the industrial bleeding geometry and may not fully reflect the effects on the compressor. We conducted a numerical study to investigate the flow mechanism of the real industrial bleeding geometry in a single-stage high-speed axial compressor environment. Results show that bleeding effects increase the adiabatic efficiency and slightly reduce the total-to-static pressure ratio. Furthermore, it is discovered that a locally aperiodic flow field in the rotor passages is generated due to the real bleeding geometry. In the tip region the maximum incident angle difference can reach 1 degree. Locally aperiodic bleeding makes the tip leakage strength and blockage different in various rotor passages.

Film cooling is an important cooling method for the high-temperature components of modern aero-engines. There have been many previous studies on the influences of compound angle on film cooling performance, most of which are conducted in the ideal flat plate environment with zero pressure gradient, and the results indicate that at high blowing ratios, a large compound angle tends to improve the film cooling performance. However, in this work, it is found that a large compound angle tends to decrease the film cooling performance significantly with the influences of endwall lateral pressure gradient. A curved tunnel with constant width is employed to generate lateral pressure gradient similar to that in the endwall region of real turbine blade. It is shown that the crossflow near the endwall surface is the most obvious flow characteristic in this flow environment. Based on this discovery, we further conduct research on the influences of compound angles on film cooling performance in a curved tunnel both with numerical and experimental approaches. A well-designed laidback fan-shaped film cooling hole is used in this work. The experimental results show that for a film cooling hole with a large compound angle under blowing ratios ranging from 1.0 to 3.0, film cooling performance is reduced significantly under the environment of endwall lateral pressure gradient compared with zero pressure gradient environment. Detailed numerical results further indicate that due to the interaction between endwall crossflow and the strong single kidney vortex created by the large compound angle hole, large amounts of coolant are lifted away from endwall surface, which results in poor cooling performance. A topological model on the interaction between endwall crossflow and kidney vortices is proposed to help illustrate this phenomenon. The results of this work suggest that turbine cooling designers carefully consider the negative influence of large compound angles on endwall film cooling due to the lateral pressure gradient.