Downstream Blade Row Research Articles

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.

Efficient analysis of unsteady flows within multi-stage turbomachines using the coupled time and passage spectral method

A coupled time and passage spectral method has been proposed very recently for tracking blade wakes penetrating the immediate downstream blade row and reaching far downstream blade rows. To achieve an efficient numerical analysis, the number of time and space modes to be retained has to be limited, as the computational cost of such an analysis is at least proportional to the number of modes. In this study, time and space modes related to downstream propagation of blade wakes reaching beyond their immediate downstream blade row are ranked according to their amplitudes of flow quantities through a time domain harmonic balance analysis using a domain consisting of multiple blade passages for the third row to capture the wakes of the first row of a two-stage fan. Modes with significant amplitudes are identified and they are really sparse. This sparsity of significant modes provides the premise for an efficient analysis using the coupled time and passage spectral method. A guideline for <italic>a priori</italic> selection of time and space modes has been developed by analyzing the frequencies and nodal diameters of those significant modes. The guideline is subsequently verified through four different coupled time and passage spectral analyses with different levels of accuracy by including different time and space modes.

Interacting Effects in a Multistage Axial Compressor Using Shrouded and Cantilevered Stators

This paper describes the analysis of the flowfield in a four-stage low-speed axial compressor at the design point investigated both numerically and experimentally, with particular emphasis on the impact of stator hub flow leakage in a section of the machine featuring a change of stator configuration. The goal of the work is to give insight into the effect of the stator hub configuration on the aerodynamic performance in terms of loss and efficiency of downstream blade rows as well as the resulting stage matching effects. The investigation is focused on aerodynamic behavior and stage interaction, which the stator hub configuration of a given stage induces in a downstream stage with respect to its performance and flowfield. Two shrouded and one cantilevered stator hub configurations, such as those commonly employed in industry, are investigated and discussed with respect to local flow phenomena and stagewise aerodynamic effects. The results show that the stator hub configuration of a given stage has a drastically high impact on the downstream rotor aerodynamic performance, whereas the downstream stator as well as the upstream rotor are only slightly influenced by the stator hub configuration.

Steady and Unsteady RANS Modeling of Wake Effects and Grid Resolution Requirements in a Low-Pressure Turbine Cascade

Due to relative motion between rotors and stators in aircraft engines, periodic wakes are present in downstream blade rows, which exert significant influence at flow loss and engine efficiency. To quantify and reproduce this influence, a low-pressure turbine cascade is computed using steady and unsteady RANS methods in the flow solver TRACE by DLR and MTU Aero Engines. A thorough grid study is carried out and various aspects of grid resolution requirements are investigated for the setups and considered performance metrics respectively. A steady state transition model extension that has been developed and published by the authors is applied to the cascade flow at a number of operating points and validated with experimental data while being compared to the unsteady results as well as the steady state results without the wake effect extension. Further, a variation of wake-related parameters is carried out while discussing the effects modeled in the unsteady setup as well as the ability of the two steady state setups (with and without wake extension) to capture the trends identified by the unsteady results. A sufficiently accurate reproduction by the wake model extension enables steady simulations of the inherently unsteady effects in the aerodynamic design of the turbine, which results in an enormous saving of computational time and effort.

A Comparison of the Wake Effects Generated by the Biased Triangle Bar and Traditional Cylinder Bar to the Boundary Layer on Suction Surface of LPT Blade

Abstract Considering the asymmetry of the low pressure turbine blade (LPT) wake at a low Reynolds number, the influence of asymmetric wakes which are similar to LPT wakes on the boundary layer of downstream blade rows in the near field is studied in the present paper, in order to increase wake flow prediction accuracy of the downstream blade without increasing the difficulty of the experiment or calculation load. Packb high-lift LPT airfoil was studied with CFX software. Following the analysis of the similarities between the wake generated by the cylinder bar and the triangle bar and the LPT blade wake in the near-field, the boundary layer flow characteristics on the suction surface under the different wakes were compared. In this research, it was found that the wakes of biased triangle bar shared more similarities with the LPT blade wake in the near field than the cylinder bar. Furthermore, the biased triangle bar wake was asymmetrical in terms of its centerline, and the separation bubble was suppressed while the calming effect was reduced after the wake-induced transition due to the asymmetry. And the time-averaged momentum thickness decreased by 7 % compared to the cylinder wake.

The Impact of Combustor Turbulence on Turbine Loss Mechanisms

Abstract A blade row that is located downstream of a combustor has an extremely high turbulence intensity at the inlet, typically above 10%. The peak turbulent length scale is also high, at around 20% of the chord of the downstream blade row. In a combustor, the turbulence is created by impinging jets in crossflow. This may result in the turbulence being anisotropic in nature. The aim of this paper is to investigate the effect of combustor turbulence on the loss mechanisms which occur in a turbine blade row. The paper has a number of important findings. The combustor turbulence is characterized and is shown to be isotropic in nature. It shows that, when no pressure gradient is present, combustor turbulence increases the loss of a turbulent boundary layer by 22%. The mechanism responsible for this change is shown to be a deep penetration of the turbulence into the boundary layer. It shows that the presence of combustor turbulence increases the profile loss and endwall loss in the turbine cascade studied by 37% and 47%, respectively. The presence of combustor turbulence also introduces a freestream loss resulting in the total loss of the turbine cascade rising by 47%. When these loss mechanisms were applied to the vane alone, of an engine representative high-pressure turbine, it was found to result in a 1.3% reduction in stage efficiency.

Generation of unsteady incoming wakes for wake/blade interaction

In this study, an effective method is proposed for the generation of unsteady incoming wakes to facilitate numerical simulation of wake/blade interaction. The wakes are specified through the inflow boundary conditions of a blade cascade, which includes the mean velocity defect as well as the velocity fluctuation induced by the vortex street in wakes. Compared to the large-eddy simulation (LES) of turbine cascades, the proposed method can be used to obtain unsteady wakes with different scales of fluctuation. Further, the analysis reveals that the velocity defect can be well fitted by the Gaussian model, and the velocity perturbation induced by wake vortex can be described by the modified Lamb-Oseen vortex or the Taylor Vortex. The vortices in wakes exhibit self-similarity behavior, which also confirms the feasibility of the proposed method. The wake/blade interaction is also simulated within T106 cascade with incoming wakes generated by the proposed method, and the results provide detailed information on the wake transportation in downstream bladerow.

Numerical Research on Inlet Total Pressure Distortion in a Transonic Compressor with Non-Axisymmetric Stator Clearance

Inlet total pressure distortion has great adverse effects on the aero-engine performance. The distorted flow passes through the compressor and becomes non-uniform in the downstream blade rows. Different from previous studies based on the assumption of circumferential uniformity, this study aims to improve circumferential non-uniform flow with the non-axisymmetric structure. Non-axisymmetric stator clearance was adopted to resolve the effects of non-uniform flow caused by inlet total pressure distortion in this paper. The 9 stators with tip clearance were installed in the distorted region and the flow field structure and performance under different operating conditions was studied. The study finds that the non-axisymmetric compressor with 9 tip clearance stators can ensure compressor efficiency while improving compressor stability margin. What’s more, the separation range and strength in the distorted region can be reduced significantly and the anti-distortion capability of compressor can be enhanced.

Flutter analysis of a one-and-a-half-stage fan at low speed using nonlinear harmonic method

Multirow effects on flutter stability of a 1.5-stage fan at low speed are investigated using nonlinear harmonic method in a one-way coupled fashion. In the first part, the mesh-independence verification and validation of nonlinear harmonic method to simulate multirow effects are performed. In the second part, multirow effects are separated into two parts including acoustic reflection and rotor–stator interaction induced by relative motion between rotor and stators with each part investigated individually. Effect of acoustic reflection from upstream and downstream blade rows is investigated separately using a harmonic truncation method to avoid the change of time-mean flow. The results show that acoustic reflection can have a large effect on flutter stability of rotor blade. The simulation of the rotor–stator interaction effect indicates that the rotor–stator interaction does not significantly affect the flutter stability of rotor blade in this case. Lastly, the variation of aerodynamic modal damping ratio with the size of gap between inlet guide vane and rotor is investigated. Aerodynamic modal damping ratio at a nodal diameter whose fundamental mode is cut-on varies periodically with gap size. Wave splitting method is employed to further investigate the relation between the phase difference between incoming and outgoing wave and aero damping, which can be used to improve the flutter stability at the design stage.

Separation of Up and Downstream Forced Response Excitations of an Embedded Compressor Rotor

The aerodynamic interaction of upstream and downstream blade rows can have a significant impact on the forced response of the compressor. Previously, the authors carried out the forced response analysis of a three-row stator-rotor-stator (S1-R2-S2) configuration from a 3.5-stage compressor. However, since the stator vane counts in both the stators (S1 and S2) were the same, it was not possible to separate the excitations from both the rows as they excited the rotor at the same frequency. Hence, a new configuration was developed and tested in which the stator 1 blade count was changed to 38 and stator 2 blade count was maintained at 44 in order to study the individual influences of the stator on the embedded rotor. By using this method, the excitations from both rows can be determined, and the excitations can be quantified to determine the row having the maximum influence on the overall forcing. To achieve this, two sets of simulations were carried out. The three-row stator-rotor (S1-R2-S2) simulation was carried out at both the 38EO (engine order) and 44EO crossings at the peak efficiency (PE) operating condition. The two-row stator-rotor analysis (S1-R2) was carried out at the 38EO crossing, and the other two-Row (R2-S2) analyses were carried out at the 44EO crossing. The steady aerodynamics was preserved in both the cases. A study was done to determine the contribution of wave reflections from the stator inlet and exit planes to the forcing function. Two conclusions drawn from this study are as follows: (1) the modal force value decreased after the upstream stator was removed, which proved that wave reflections from this stator were significant and (2) the increase in modal force was in-line with experimental observations.

Improvement of Steam Turbine Stage Efficiency by Controlling Rotor Shroud Leakage Flows—Part I: Design Concept and Typical Performance of a Swirl Breaker

In high and intermediate pressure (HIP) steam turbines with shrouded blades, it is well known that shroud leakage losses contribute significantly to overall losses. Shroud leakage flow with a large tangential velocity creates a significant aerodynamic loss due to mixing with the mainstream flow. In order to reduce this mixing loss, two distinct ideas for rotor shroud exit cavity geometries were investigated using computational fluid dynamics (CFD) analyses and experimental tests. One idea was an axial fin placed from the shroud downstream casing to reduce the axial cavity gap, and the other was a swirl breaker placed in the rotor shroud exit cavity to reduce the tangential velocity of the leakage flow. In addition to the conventional cavity geometry, three types of shroud exit cavity geometries were designed, manufactured, and tested using a 1.5-stage air model turbine with medium aspect ratio blading. Test results showed that the axial fin and the swirl breaker raised turbine stage efficiency by 0.2% and 0.7%, respectively. The proposed swirl breaker was judged to be an effective way to achieve highly efficient steam turbines because it not only reduces the mixing losses but also improves the incidence angle distribution onto the downstream blade row. This study is presented in two papers. The basic design concept and typical performance of the proposed swirl breaker are presented in this part I, and the effect of axial distance between a swirl breaker and rotor shroud on efficiency improvement is discussed in part II [8].

The Impact of Inlet Distortion and Reduced Frequency on the Performance of Centrifugal Compressors

This paper discusses the impact of inlet flow distortions on centrifugal compressors based upon a large experimental data base in which the performance of several impellers in a range of corrected flows and corrected speeds have been measured after been coupled with different inlet plenums technologies. The analysis extends to centrifugal compressor inlets including a side stream, typical of liquefied natural gas applications. The detailed measurements allow a thorough characterization of the flow field and associated performance. The results suggest that distortions can alter the head by as much as 3% and efficiency of around 1%. A theoretical analysis allowed to identify the design features that are responsible for this deviation. In particular, an extension of the so-called “reduced-frequency,” a coefficient routinely used in axial compressors and turbine aerodynamics to weigh the unsteadiness generated by upstream to downstream blade rows, allowed to determine a plenum-to-impeller reduced frequency that correlates very well with the measured performance. The theory behind the new coefficient is discussed together with the measurement details and validates the correlation that can be used in the design phase to determine the best compromise between the inlet plenum complexity and impact on the first stage.

Analysis of the Effect of Multirow and Multipassage Aerodynamic Interaction on the Forced Response Variation in a Compressor Configuration—Part I: Aerodynamic Excitation

The aeroelastic prediction of blade forcing is still a very important topic in turbomachinery design. Usually, the wake from an upstream airfoil and the potential field from a downstream airfoil are considered as the main disturbances. In recent years, it became evident that in addition to those two mechanisms, Tyler–Sofrin modes, also called scattered or spinning modes, may have a significant impact on blade forcing. It was recently shown in literature that in multirow configurations, not only the next but also the next but one blade row is very important as it may create a large circumferential forcing variation, which is fixed in the rotating frame of reference. In the present paper, a study of these effects is performed on the basis of a quasi three-dimensional (3D) multirow and multipassage compressor configuration. For the analysis, a harmonic balancing code, which was developed by DLR Cologne, is used for various setups and the results are compared to full-annulus unsteady calculations. It is shown that the effect of the circumferentially different blade excitation is mainly contributed by the Tyler–Sofrin modes and not to blade-to-blade variation in the steady flow field. The influence of various clocking positions, coupling schemes and number of harmonics onto the forcing is investigated. It is also shown that along a speed-line in the compressor map, the blade-to-blade forcing variation may change significantly. In addition, multirow flutter calculations are performed, showing the influence of the upstream and downstream blade row onto aerodynamic damping. The effect of these forcing variations onto random mistuning effects is investigated in the second part of the paper.

Analysis of the Effect of Multirow and Multipassage Aerodynamic Interaction on the Forced Response Variation in a Compressor Configuration—Part II: Effects of Additional Structural Mistuning

The prediction of aerodynamic blade forcing is a very important topic in turbomachinery design. Usually, the wake from the upstream blade row and the potential field from the downstream blade row are considered as the main causes for excitation, which in conjunction with relative rotation of neighboring blade rows, give rise to dynamic forcing of the blades. In addition to those two mechanisms, the so-called Tyler–Sofrin (or scattered or spinning) modes, which refer to the acoustic interaction with blade rows further up- or downstream, may have a significant impact on blade forcing. In particular, they lead to considerable blade-to-blade variations of the aerodynamic loading. In Part I of the paper, a study of these effects is performed on the basis of a quasi-three-dimensional multirow and multipassage compressor configuration. Part II of the paper proposes a method to analyze the interaction of the aerodynamic forcing asymmetries with the already well-studied effects of random mistuning stemming from blade-to-blade variations of structural properties. Based on a finite element (FE) model of a sector, the equations governing the dynamic behavior of the entire bladed disk can be efficiently derived using substructuring techniques. The disk substructure is assumed as cyclically symmetric, while the blades exhibit structural mistuning and linear aeroelastic coupling. In order to avoid the costly multistage analysis, the variation of the aerodynamic loading is treated as an epistemic uncertainty, leading to a stochastic description of the annular force pattern. The effects of structural mistuning and stochastic aerodynamic forcing are first studied separately and then in a combined manner for a blisk of a research compressor without and with aeroelastic coupling.

Large-Scale Detached-Eddy Simulation Analysis of Stall Inception Process in a Multistage Axial Flow Compressor

This paper describes the flow mechanisms of rotating stall inception in a multistage axial flow compressor of an actual gas turbine. Large-scale numerical simulations of the unsteady have been conducted. The compressor investigated is a test rig compressor that was used in the development of the Kawasaki L30A industrial gas turbine. While the compressor consists of a total of 14 stages, only the front stages of the compressor were analyzed in the present study. The test data show that the fifth or sixth stages of the machine are most likely the ones leading to stall. To model the precise flow physics leading to stall inception, the flow was modeled using a very dense computational mesh, with several million cells in each passage. A total of 2 × 109 cells were used for the first seven stages (3 × 108 cells in each stage). Since the mesh was still not fine enough for large-eddy simulation (LES), a detached-eddy simulation (DES) was used. Using DES, a flow field is calculated using LES except in the near-wall where the turbulent eddies are modeled by Reynolds-averaged Navier–Stokes. The computational resources required for such large-scale simulations were still quite large, so the computations were conducted on the K computer (RIKEN AICS in Japan). Unsteady flow phenomena at the stall inception were analyzed using data mining techniques such as vortex identification and limiting streamline drawing with line integral convolution (LIC) techniques. In the compressor studied, stall started from a separation on the hub side rather than the commonly observed leading-edge separation near the tip. The flow phenomenon first observed in the stalling process is the hub corner separation, which appears in a passage of the sixth stator when approaching the stall point. This hub corner separation grows with time, and eventually leads to a leading-edge separation on the hub side of the stator. Once the leading-edge separation occurs, it rapidly develops into a rotating stall, causing another leading-edge separation of the neighboring blade. Finally, the rotating stall spreads to the upstream and downstream blade rows due to its large blockage effect.

Comparison of Two-Dimensional and Three-Dimensional Turbine Airfoils in Combination With Nonaxisymmetric Endwall Contouring

Tangential endwall contouring (TEWC) is intended to improve the turbomachinery blading efficiency. This paper summarizes the experimental and numerical investigation of a test turbine with endwall contoured vanes and blades. Constant section (2D) airfoils as well as optimized compound lean (3D) high pressure steam turbine blading in baseline and endwall contoured configurations have been examined. Brush seals (BSs) are implemented within the casing sided cavities to minimize the leakage flow near the tip endwalls, where the contouring is also applied. The pressure and temperature data that are recorded in three axial measuring planes are plotted to visualize the change in flow structure. This shows that the efficiency is increased for 2D airfoils by means of endwall contouring. However, the efficiency of the first stage suffers, and the endwall contouring is still beneficial for the overall performance of the engine. Both phenomena (an efficiency loss in stage one and an improvement of the performance in stage two) have also been measured for the optimized 3D configurations; thus, it can be expected that the endwall contouring has also a beneficial impact on the performance of multirow turbines. The numerical investigations demonstrate in detail, how the secondary flow phenomena are influenced by end-wall contouring and a description of the changes in vortex formations as well as blade loading are given for the various configurations. It has been found that for steady computational fluid dynamics (CFD) simulations the use of stage interfaces suppresses the positive effects of the endwall contour onto the downstream blade row.

Design of Compressor Endwall Velocity Triangles

The operating range of a compressor is usually limited by the rapid growth of three-dimensional (3D) separations in the endwall flow region. In contrast, the freestream region is not usually close to its diffusion limit and has little effect on overall range. In light of these two distinct flow regions, this paper considers how velocity triangles in the endwall region should be designed to give a more balanced spanwise failure across the span of a blade row. In the first part of this paper, the sensitivity of 3D separations in a single blade row to variations in realistic multistage inlet conditions and endwall geometry is investigated. It is shown that a blade's 3D separation size is largely controlled by the dynamic pressure within the incoming endwall “repeating stage” boundary layer and not the detailed local geometry within the blade row. In the second part of this paper, the traditional design process is “flipped.” Instead of redesigning a blade's endwall geometry to cope with a particular inlet profile into the blade row, the endwall region is redesigned in the multistage environment to “tailor” the inlet profile into downstream blade rows, giving the designer a new extra degree-of-freedom. This extra degree-of-freedom is exploited to balance freestream and endwall operating range, resulting in a compressor having an increased operating range of ∼20%. If this increased operating range is traded with reduced blade count, it is shown that a design efficiency improvement of ∼0.5% can be unlocked.

Effects of Upgraded Cooling System and New Blade Materials on a Real Gas Turbine Performance

Abstract The aim of this paper is to study the effects on the performance of a real heavy-duty gas turbine of two different solutions for enhancing turbine blades thermal resistance. An upgrade of the first stator cooling system and the adoption of improved blade materials are simulated exploiting an in-house simulation tool (ESMS). The changes are studied separately in order to point out the positive effects as well as the related risks, such as the side effect of temperature increase on downstream blade rows, to be considered in service operations.

Numerical Study of Purge and Secondary Flows in a Low-Pressure Turbine

The secondary flow increases the loss and changes the flow incidence in the downstream blade row. To prevent hot gases from entering disk cavities, purge flows are injected into the mainstream in a real aero-engine. The interaction between purge flows and the mainstream usually induces aerodynamic losses. The endwall loss is also affected by shedding wakes and secondary flow from upstream rows. Using a series of eddy-resolving simulations, this paper aims to improve the understanding of the interaction between purge flows, incoming secondary flows along with shedding wakes, and mainstream flows on the endwall within a stator passage. It is found that for a blade with an aspect ratio of 2.2, a purge flow with a 1% leakage rate increases loss generation within the blade passage by around 10%. The incoming wakes and secondary flows increase the loss generation further by around 20%. The purge flow pushes the passage vortex further away from the endwall and increases the exit flow angle deviation. However, the maximum exit flow angle deviation is reduced after introducing incoming wakes and secondary flows. The loss generation rate is calculated using the mean flow kinetic energy equation. Two regions with high loss generation rate are identified within the blade passage: the corner region and the region where passage vortex interacts with the boundary layer on the suction surface. Loss generation rate increases dramatically after the separated boundary layer transitions. Since the endwall flow energizes the boundary layer and triggers earlier transition on the suction surface, the loss generation rate close to the endwall at the trailing edge (TE) is suppressed.

Effect of Purge Flow Swirl on Hot-Gas Ingestion into Turbine Rim Cavities

A simplified turbine rim cavity is considered, with a focus on understanding the effects of purge flow swirl on main gas ingestion. Recent work has shown that introducing swirl into purge flow upstream of a blade row is desirable insofar as reducing viscous losses associated with purge flow interaction with the hot-gas path. However, computational fluid dynamics results suggest that, in the presence of the rotating external pressure nonuniformity, due to the downstream blade row, swirled purge flow is less effective in preventing ingestion. This is reflected in higher purge air mass flow rates necessary to seal a given cavity, and that in turn diminishes the positive effect of preswirling purge flow in the first place. It is reasoned that swirled purge flow moves with the rotating pressure nonuniformity and responds to it more readily than nonswirled purge flow, which sees the averaged effect of multiple blade passing events. A flow model based on this physical principle is developed, showing good agreement with computational results. The model yields an ingestion criterion with a parametric dependence on purge flow parameters. The analysis is extended to an unsteady situation, whereby the effects of both stationary and rotating pressure nonuniformities, from vanes and blades, respectively, are taken into account. This unsteady flow model points to an optimal design space, in the context of minimizing purge flow losses while maintaining an appropriate margin with regard to hot-gas ingestion.