Endwall Secondary Flow Research Articles

Effect of Local Concave Non-axisymmetric Endwall Profiling on Endwall Secondary Flows of a Highly-Loaded Turbine Cascade

Non-axisymmetric endwall profiling (NAEP) has been widely utilized in reducing secondary flow loss of turbines. However, most of NAEP are designed for the endwall of the entire blade passage, which presents challenges to the design of cooling structures of turbine endwall. This study aims to explore the local concave non-axisymmetric endwall profiling (LCNP) method with the same effect as whole passage NAEP, and reveal the influence mechanism of LCNP on endwall secondary flow structures of a highly-loaded turbine cascade. Under the condition that the maximum depth of LCNP is unchanged, the axial length effect and pitchwise location effect of LCNP are studied, the influence mechanism of LCNP on the secondary flow loss is analyzed, and the genetic algorithm is utilized to optimize LCNP at the optimal position. Results show that as the axial length of the LCNP increases and the pitchwise location gets closer to the suction surface, the intensity and range of the passage vortex are decreased, and the total pressure loss coefficient (loss coefficient) of the turbine cascade is decreased. When LCNP is 100% axial chord in length and at the position of 2/9 pitch, the loss coefficient is reduced by 5.49%. LCNP was optimized at the optimal position, and the optimal LCNP reduced the loss coefficient of the turbine cascade by 6.73%. After the local concave endwall profiling, the loading in the middle of the endwall of the turbine cascade is reduced, and the intensity of the passage vortex is effectively inhibited, which is the mechanism that the loss coefficient of the turbine cascade is reduced. However, after the local concave endwall profiling, the loading in the trailing of the endwall of the turbine cascade is increased, the transverse migration of the new boundary layer in the endwall is accelerated, and the loss coefficient of the near endwall is increased.

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

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

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.

Numerical investigation on film cooling and aerodynamic performance for gas turbine endwalls with upstream vane-type and cascade-type slots

To further improve the thermal efficiency of gas turbine, the turbine inlet temperature (TIT) increases up to 2000 K, leading to an extremely high heat load on the first-stage vane. Taking into consideration that there are complex endwall secondary flow and the passage vortices which have significant effect on the flow field and thermal behavior near the endwall, it is essential to study the aero-thermal characteristics of endwall and develop efficient endwall cooling approach. In this work, two novel upstream slot structures, referred to as the vane-type slot and cascade-type slot, are proposed, and a comprehensive numerical analysis is conducted to reveal their mechanism and superiority in improving the film cooling and aerodynamic performance of first-stage vane compared with the conventional slot structure. The results indicate that compared with the conventional rectangular straight slot, the coolant jets from two novel slots have a more significant inhibition effect on the development of the endwall secondary flow due to their lower outflow angle and thus higher momentum in streamwise direction, which contributes to reducing the strength of the endwall secondary flow. This is beneficial to provide a lower cascade aerodynamic loss and a higher film cooling effectiveness of the endwall near the suction side leading edge and the higher film cooling effectiveness along the trajectory of the horseshoe vortex on the more downstream region. Moreover, the vane-type slot provides the largest coolant coverage and cooling effectiveness of endwall and the lowest total pressure loss within slot and cascade aerodynamic loss, followed by the cascade-type slot and conventional slot.

Turbine Blade Endwall Heat Transfer and Film Cooling Performance With Multigap Jets and Film Holes

Abstract Due to manufacturing, assembly, and operational conditions, gaps are commonly presented in the turbine components, such as platform blade-to-blade gap (slashface gap) and rim cavity platform gap (slot gap). The coolant ejected through these gaps not only limits hot mainstream gas ingestion into the disk cavity but also has the potential to provide endwall film cooling coverage. Nevertheless, the interaction of these gap flows, predesigned film hole jets, and endwall secondary flows, significantly impacts endwall heat transfer and film cooling performance and leads to a more complicated flow field near endwall. To further understand the endwall flow physics behavior in this complicated flow field, the combined effects of film hole jets and multigap leakages (slashface jet and slot jet) were experimentally studied in a transient wind tunnel with a six-blade linear cascade (nonrotating). The endwall heat transfer was measured and recorded by the IR technique at inlet average turbulence intensity (Tu) of 7.5% and an exit Mach number (Maex) of 0.4. In addition, detailed numerical predictions were also performed to discuss the flow physics near endwall and the multigap leakages behavior. Results indicated that the slashface jet is an essential contributor to endwall film cooling performance in this endwall configuration. The slashface jet delays the development of the film hole jets, leading to an enhancement of film cooling performance downstream of the film holes. Increasing the mass flow ratio of the slashface jet (MFRslashface) can prevent the high-temperature mainstream ingestion, reduce the thermal failure risks, and increase the peak of film cooling effectiveness downstream of the endwall (167% at MFRslashface = 0.5%). Due to the limitation and hindrance impacts of the disk vortex (DV), the cavity vortex (CV), and the suction side leg of the horse-shoe vortex (HV,s), the slot flow is confined to a small region downstream of the slot gap.

Effects of proactive phantom cooling on turbine endwall aerodynamic and film cooling performances considering aggressive swirling inflow

The turbine vane endwall downstream of the lean premixed combustor operates in an extremely complicated environment resulting from strong swirling inflow. Consequently, advanced cooling methods should be developed considering aggressive swirling inflow conditions. In the current work, the numerical results indicated that the swirling inflow seriously deteriorated the film cooling level by more than 30%, especially near the pressure side endwall, and it increased the total pressure loss of the vane passage by up to 80%. According to the endwall aerothermal performance under the swirling inflow conditions, the proactive phantom cooling method was presented. With the help of the endwall secondary flow, the proactive phantom coolant could obviously increase the film cooling effectiveness on the endwall and the vane pressure side surface at the same time. Furthermore, the effects of the outlet height and the mass flow rate on the performance of proactive phantom cooling were investigated. The design with h/H = 4.0% and MFR = 0.30% was found to be the best choice in this study with the largest film cooling coverage area and a 44% increase of the overall film cooling effectiveness. This paper provides a novel cooling method to increase the endwall film cooling performance under swirling inflow conditions.

Effect investigation of single-slotted and double-slotted configurations on the corner separation and aerodynamic performance in a high-load compressor cascade

Controlling the corner separation through blade-end slotting has been widely proven advantageous in compressors. This study proposed a new double-slotted configuration and deeply revealed its benefits and flow control mechanisms by compared with the single-slotted configuration. Based on a high-load compressor cascade, the Slotted_A scheme with a single blade-end slot1 and the Slotted_B scheme with a smaller slot2 behind the slot1 were established and detailedly evaluated. The results showed that, since the slot jet with high momentum has efficacious re-energizing effects for the local low-energy fluid, both the two slotted schemes can prevent the original cascade from entering the corner stall condition. But the slot2 jet can further control the end-wall secondary flow developing and separating behind the slot1 in the Slotted_A cascade. Therefore, the Slotted_B scheme behaves better in controlling the corner separation than the Slotted_A scheme, resulting in larger turning angles, smaller vortex structures, and weaker flow blockage. As a whole, the Slotted_A scheme and Slotted_B scheme can respectively increase the turning angle by an average of 0.93° and 1.53°, improve the static pressure coefficient by 18.2% and 21.7% on average, and achieve an average reduction of 38.3% and 43.7% in the total pressure loss. In contrast, the Slotted_B scheme is more advantageous in improving the pressure diffusion capacity and stability margin for the high-load compressor blade.

Mechanisms of oscillating suction in controlling the tip leakage flow in a high-load compressor cascade

To control tip leakage flow and reduce related loss in axial-flow compressors, the oscillating suction scheme was introduced in a high-load compressor cascade with a tip clearance, and its flow control mechanisms on the tip leakage flow were studied numerically. It is found that the flow control effect of oscillating suction is the best when the unsteady excitation frequency is equal or harmonic to the feature frequency of the flow field related to the flow separation. The loss of the cascade controlled by the oscillating suction with a reduced frequency of 0.39 and excitation amplitude of 10 kPa was reduced by about 3.4% compared with the steady suction scheme, and about 11.14% less than that of the non-suction scheme. Oscillating suction discretized the concentrated tip leakage vortex into a series of scale-reduced vortex structures by the periodic unsteady excitation, which can be identified by the analysis using proper orthogonal decomposition (POD). The feature frequencies became consistent with the oscillating excitation frequency or its multiple, and the fluctuation generated by the tip leakage flow was enhanced. The energy proportion of high-order POD modes was increased by oscillating excitation, which indicated the vortex structures with weak fluctuation intensity were more active. The mixing and energy exchange among the mainstream, separated flow, endwall secondary flow, and tip leakage flow were enhanced, so that a better flow control effect could be achieved. Furthermore, flow capacity of tip passage was modified, the overall aerodynamic performance of the compressor cascade was improved thereby.

Numerical Investigations on Effect of Inflow Parameters on Development of Secondary Flow Field for Linear Low-Pressure Turbine Cascade

Abstract Endwall flows contribute the most crucial role in loss generation for axial flow turbine and compressor blades. These losses lead to modifying the blade loading and overall performance in terms of stable operating range. The present study aimed to determine the endwall flow streams in a low-speed low-pressure linear turbine cascade vane using a numerical approach. The study includes two sections. The first section includes an attempt to understand different secondary flow streams available at the endwall. The location of the horseshoe vortex and subsequent vortex patterns are identified in the section. The selection of a suitable turbulence model among shear stress transport (SST) k–ω and SST γ–θ to identify endwall flow streams is studied prior to the section. The steady-state numerical study is performed using Reynolds Averaged Navier–Stokes equations closed by the SST γ–θ turbulence model. The computational results are validated with experimental results available in the literature and are found to be in good agreement. The study is extended for different inflow conditions in a later section. The second section includes the effect of flow incidence and turbulence intensity on the endwall secondary flow field. Four different inflow incidences are considered for the study. The inlet turbulence intensities are varied by 1% and 10% for each case. The results revealed different secondary flow patterns at an endwall and found the change in behavior with inflow conditions. SST γ–θ turbulence model with lower turbulence intensity is more suitable to identify such flow behavior.

Research on the Influence of Inflow Conditions on the Aerodynamic Performance of a Tandem Cascade

We perform a thorough numerical analysis of the impact of inflow conditions on the aerodynamic performance of a tandem cascade. In particular, we investigate the effects of the incidence angle and the inlet boundary layer (IBL) thickness on the three-dimensional flow field structure and aerodynamic performance. Our results show that the gap flow strength of the tandem cascade decreases with the increase of incidence angle, and it can effectively reduce the mixing of the wakes of the forward blade (FB) and rear blade (RB). In turn, this prevents the passage vortex (PV) in the RB passage from developing along the circumferential direction. The occurrence of IBL does not modify the effects of the incidence angle on the tandem cascade, however, it reduces the load of the RB and the gap flow strength near the endwall. Under all incidence angles, IBL increases the total pressure loss of the tandem cascade, and decreases the static pressure rise (except for an incidence angle equal to -6°). The maximum loss increment is at 2° incidence angle, and the maximum static pressure rise decrement is at 6° incidence angle (Thick-IBL condition) or 7° incidence angle (Thin-IBL condition). Furthermore, we found that the presence of IBL changes the minimum loss condition from 0° (design condition) to -2° incidence angle. Our results thus indicate that in the practical engineering application of the tandem cascade, the reality that IBL degrades the tandem cascade performance in the full incidence angle range should be considered. And the strong endwall secondary flow effect caused by IBL should be considered in the tandem cascade three-dimensional design, so that the tandem cascade two-dimensional performance advantage can be better played.

Entropy Generation of Secondary Flow in a Turning Passage with Different Boundary Layer Characteristics

The development of secondary flow along a curved channel is a fundamental flow phenomenon occurring in a wide range of engineering applications, including turbomachinery, aerospace, heating, ventilation, air conditioning, etc. The underlying flow physics about end-wall secondary flows has been well-documented in the open literature, while the interaction between a secondary flow and a side-wall boundary layer, which is critical to the aerothermal performance of a side-wall surface, has not been comprehensively studied. In this study, the entropy generation of secondary flow and the interaction between an end-wall passage vortex and a side-wall boundary layer were numerically investigated by Reynolds-averaged Navier–Stokes (RANS) CFD for a 90° curved channel. The transportation effect of secondary flow and the generation mechanism of an induced vortex pair on the side wall is reported. It was also found that the growth of the secondary flow can be suppressed due to the displacement effect of the side-wall boundary layer. Furthermore, it was found that the interaction between a secondary flow and a side-wall boundary layer provides a suppression effect on side-wall boundary layer separation.

Uncertainty Quantification on the Cooling Performance of a Transonic Turbine Vane With Upstream Endwall Misalignment

Abstract With increasing aerodynamic and thermal loads, film cooling has been a popular technology integrated into the design of the modern gas turbine vane endwall, especially for the first-stage vane endwall. A staggering amount of research has been completed to quantify the effect of operating conditions and cooling hole geometrical properties. However, most of these investigations did not address the influence of the manufacturing tolerances, assembly errors, and operation degradations on the endwall misalignment. In this paper, uncertainty quantification (UQ) analysis was performed to quantify the impacts of upstream endwall misalignment uncertainties on the endwall film cooling performance and vane surface phantom cooling performance. The upstream endwall misalignment, step geometry with various step heights, commonly exists between the combustor exit and the first-stage vane endwall. Based on the non-intrusive polynomial chaos expansion (NIPC) and the uniform probability distribution assumption, the deviation (step height) uncertainties of the upstream endwall misalignment were quantified. To predict the endwall secondary flow and film cooling effectiveness in the transonic linear vane passage, the commercial computational fluid dynamic solver ANSYS FLUENT was used to numerically solve the three-dimensional steady-state Reynolds-Averaged Navier–Stokes (RANS) equations. The robustness analysis of endwall film cooling performance and phantom cooling performance to the upstream endwall misalignment was conducted for three design upstream step heights (ΔH): a baseline configuration (ΔH = 0 mm), two misaligned configurations with forward step (ΔH = −5 mm) and backward step (ΔH = 5 mm), respectively. Results show that the actual cooling performance has a high probability of deviating from the nominal value for the baseline configuration. The critical regions that are most sensitive to the upstream step misalignment are also identified by variances. The UQ results also show that the design geometry with a forward step has a more robust film cooling performance on endwall and phantom cooling performance on the vane pressure side surface, which means a smaller variance and a better expectation than the no-step configuration. In contrast, the design geometry with a backward step induces the reductions in the expectation of the film cooling effectiveness and coolant coverage and the amplification of performance fluctuations. This work provides a certain guiding direction for the optimization design for the upstream step geometry.

A spanwise loss model for axial compressor stator based on machine learning

In the early stage of aircraft engine design, the through-flow method is an important tool for designers. The accuracy of the through-flow method depends heavily on the accuracy of the loss model. However, most existing models cannot (or cannot well) provide the spanwise loss distribution. To construct an effective spanwise loss model, both turbomachinery knowledge and machine learning skills were used in this paper. A large number of numerical simulations were carried out to build a database containing more than 1000 compressor cascade numerical samples. Secondary flow intensity was introduced as the independent variable to carry out feature engineering. A model containing a selector based on support vector machine regression and estimators based on K-nearest neighbor regression was constructed. Numerical test set and design data of two former high-pressure core compressors were used for validation. Results suggest that the spanwise loss model show good consistency with both numerical test set and data of two former compressors. It can reflect the influence of secondary flow and can also predict both value and trend of total pressure loss coefficient well, with mean absolute error general around or less than 1% and R2 (coefficient of determination) more than 0.8 on the test set. Especially when dealing with loss coefficient at mid-span position, the model shows even better performance, with R2 over 0.97 on the test set. And the selector of the model can well classify the samples, predict the intensity of secondary flow and help estimators to capture the phenomenon that end-wall secondary flow extends to the mid-span.

Research on the flow field and film cooling effectiveness of the endwall with swirling film cooling

In order to obtain the effect of endwall secondary flow on the swirling film cooling, a geometric model of cascade is established to research the endwall swirling film cooling and swirling flow induced by prismatic jet impingement configurations. Numerical simulation is applied to investigate three jet flow configurations on the endwall film cooling performance at the compound angles of film hole γ = 0° and 30° and blowing ratios M = 0.5–2.0. The influence of complex vortex structures near endwall for jet flow is analyzed in detail; the strong transverse cross flow near the endwall is the main reason affecting the film cooling effectiveness. The variation laws of endwall film cooling effectiveness with the compound angle of film hole, jet flow configuration, and the blowing ratio are obtained. As the blowing ratio increases, the spanwise average film cooling effectiveness increases first and then decreases. While the blowing ratio is M =1.0, the endwall film cooling effectiveness is the best. Increasing the compound angle of the film hole leads to a decrease in the endwall cooling effectiveness. The spanwise average cooling effectiveness of γ = 30° decreases by 35% compared to the γ = 0°.

Numerical Investigation into a New Method of Non-Axisymmetric End Wall Contouring for Axial Compressors

To deal with corner separation in high-load axial compressors, this paper proposes a new end wall contouring method aimed at controlling the end wall secondary flow in more than one local area, generating a geometry with fewer control variables that is applicable for multiple working conditions. The new method defines more than one surface unit function, with different effects on end wall secondary flow. Then, the geometry of these surface unit functions will be superposed to generate the end wall contouring, to combine their flow control effects. After applying the new method to a bi-objective optimization design process, with 15 design variables aimed at minimizing the loss of cascade at 0° and 4° incidence, the optimal design reduces the total pressure loss of the high-load cascade by 5% under the former incidence and by 3% under the latter. The most effective design rule is constructing an end wall surface with the rising suction side and sinking pressure side in the blade channel, while locally raising the SS corner with a gentle upstream slope. According to the analysis, the design variables of the new method show an intuitive influence on the variation of end wall geometry and the movement of secondary flow. The corner separation has been effectively suppressed, with fewer control variables than before. It, thus, indicates the advantage of the newly developed end wall contouring method compared with previous studies.

The Effects of Axisymmetric Convergent Contouring and Blowing Ratio on Endwall Film Cooling and Vane Pressure Side Surface Phantom Cooling Performance

Abstract To reduce the intensity of endwall secondary flow, the axisymmetric convergent contoured endwall is commonly designed in the first nozzle guide vane (NGV) passage of the real gas turbine engines. This endwall contouring can obviously alter flow field near endwall and affect the coolant flow through the upstream double-row discrete film holes. This leads to a significant influence on the endwall film cooling performance, vane surface phantom cooling, and vane passage aerodynamic performance. In this paper, a detailed numerical investigation on the endwall film cooling and vane pressure side surface phantom cooling was performed, at the simulated realistic gas turbine operating conditions (high inlet freestream turbulence level of 16%, exit Mach number of 0.85, and exit Reynolds number of 1.7 × 106). Based on a double coolant temperature model, a novel numerical method for the predictions of adiabatic wall film cooling effectiveness was proposed. This numerical method was validated by comparing the predicted results with experimental data of endwall Nusselt number, endwall film cooling effectiveness, and flow visualization near endwall. The results indicate that the present numerical method can accurately predict endwall thermal load distributions, endwall film cooling distributions, and vane surface phantom cooling distributions. The endwall heat transfer coefficient, endwall film cooling effectiveness, phantom cooling effectiveness of the vane pressure side surface, and total pressure loss coefficients (TPLC) were predicted and compared for two endwall contouring shapes (flat endwall and axisymmetric convergent contoured endwall) at three different blowing ratios (low blowing ratio of BR = 1.0, design blowing ratio of BR = 2.5, and high blowing ratio of BR = 3.5) with a constant density ratio of DR = 1.2, based on the present novel numerical method. Results show that the axisymmetric convergent endwall contouring leads to a slight enhancement (maximum enhancement level less than 20%) of endwall heat transfer in the entire vane passage (0 < x < 0.65Cx). The axisymmetric convergent endwall contouring has a significantly desired effect on endwall film cooling performance (maximum enhancement level of 67%), phantom cooling performance of the vane pressure side surface (maximum increase level approximately 100%), and aerodynamic loss (maximum reduction level of 1.45%) for all blowing ratio cases, but the benefit enhancement level is obviously affected by the blowing ratio values. This suggests that the optimum of endwall contouring shapes is an effective technical way to improve endwall film cooling performance and decrease the depletion of coolant; the coupled effects of the appropriate axisymmetric convergent endwall contouring and the optimum blowing ratio should be considered in the design process of advance endwall cooling schemes.

Endwall Heat Transfer and Cooling Performance of a Transonic Turbine Vane With Upstream Injection Flow

Abstract Flow fields near the turbine vane endwall region are very complicated due to the presence of high three-dimensional passage vortices and endwall secondary flows. This makes it challenging for the endwall to be effectively cooled by using traditional endwall cooling methods, such as impingement cooling combined with local film cooling inside the vane passage. One effective endwall cooling scheme: coolant injection flow through discrete holes upstream of the vane leading edge on the endwall, has been considered by many gas turbine companies. The present paper focuses on endwall film cooling effectiveness evaluation with upstream coolant injection through discrete holes. Detailed experimental and numerical studies on endwall heat transfer and cooling performance with coolant injection flow through upstream discrete holes are presented in this paper. High-resolution heat transfer coefficient (HTC) and adiabatic film cooling effectiveness values were measured using a transient infrared thermography technique on an axisymmetric contoured endwall. The endwall tested was a scaled up inner endwall of an industrial transonic turbine vane with double-row discrete cylindrical film cooling holes located 0.39Cx upstream of the vane leading edge. The tests were performed in a transonic linear cascade blowdown wind tunnel facility. Conditions were representative of a land-based power generation turbine with exit Mach number of 0.85 corresponding to exit Reynolds number of 1.5 × 106, based on exit condition and axial chord length. A high turbulence level of 16% with an integral length scale of 3.6%P was generated using inlet turbulence grid to reproduce the typical turbulence conditions in real turbine. Low temperature air was used to simulate the typical coolant-to-mainstream condition by controlling two parameters of the upstream coolant injection flow: mass flowrate to determine the coolant-to-mainstream blowing ratio (BR = 2.5, 3.5), and gas temperature to determine the density ratio (DR = 1.2). To highlight the interactions between the upstream coolant flow and the passage secondary flow combined with the influence on the endwall heat transfer and cooling performance, a comparison of CFD predictions with experimental results was performed by solving steady-state Reynolds-averaged Navier–Stokes (RANS) using the commercial CFD solver ansys fluent V.15. A detailed numerical method validation was performed for four different Reynolds-averaged turbulence models. The realizable κ−ε model was validated to be suitable to obtain reliable numerical solution. The influences of a wide range of coolant-to-mainstream blowing ratios (BR = 1.0, 1.5, 1.9, 2.5, 3.0, 3.5) were numerically studied. Complex interactions between coolant injections and secondary flows in vane passage were presented and discussed. An optimal value of the blowing ratio for the present upstream discrete film hole is also suggested based on the current study. Results indicate that for lower values of BR, the endwall coolant coverage from the upstream double-row discrete holes is strongly controlled by the passage secondary flow, thus the cooling effectiveness is very poor. As the BR increases, the strong secondary flow in vane passage can be suppressed by the coolant injections and begin to be almost eliminated when BR increases to a critical value (BR = 2.5–3.0). Beyond the critical BR, most of the injected coolant begins to lift off from the endwall and penetrate significantly into the mainstream flow, yielding inefficient endwall cooling performance.

Passive Control with Blade-End Slots and Whole-Span Slot in a Large Camber Compressor Cascade

Suitable slot structure of the compressor blade can generate high-momentum jet flow through pressure difference between the pressure and suction surface, it has been proved that the slot jet flow can reenergize the local low-momentum fluid to effectively suppress the flow separation on the suction surface. In order to explore a slotted method for better comprehensive suppressing effects on the boundary layer separation near blade midspan and the three-dimensional corner separation, a diffusion stator cascade with large camber angle is selected as the research object. Firstly, the Slotted_1 and Slotted_2 whole-span slotted schemes are set up, then the Slotted_3 scheme with whole-span slot and blade-end slots is proposed, finally the performance of original cascade and slotted cascades is computed under a wide range of incidence angles at the Mach number of 0.7. The results show that: in the full range of incidence angles, compared with the whole-span slotted cascades, the development of the endwall secondary flow on the suction surface of Slotted_3 cascade is effectively suppressed, the degree of the mutual interference between the secondary flow and the main flow is reduced. Besides, on the suction surface of Slotted_3 cascade, the boundary layer separation near blade midspan and the corner separation are basically eliminated. As a result, compared with those of original cascade, the total pressure losses of Slotted_3 cascade are reduced in the full range of incidence angles, and its operating range of incidence angles is broadened. Moreover, compared with the whole-span slotted schemes, Slotted_3 scheme has a better adaptability to wide range of incidence angles.

Research on the Influence of Steam Turbine Seal Leakage

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

Aerodynamic performance of profiled endwalls with upstream slot purge flow in a linear turbine cascade having pressure side separation

In aeroengines, purge flow directly fed from the compressor (which bypasses the combustor) is introduced through the disk space between blade rows to prevent the hot ingress. Higher quantity of purge gas fed through the wheel space can provide additional thermal protection to the passage endwall and blade surfaces. However, the interaction of purge flow with the mainstream flow leads to higher secondary losses. Secondary losses inside a turbine blade passage can be reduced effectively by endwall contouring. This paper presents computational investigation on the influence of non-axisymmetric endwall contouring over endwall secondary flow modification in the presence of purge flow with the pressure side bubble (PSB). The experimental analysis was conducted for the base case without purge and base case with purge (BCP) configurations having flat endwalls. The total pressure loss coefficient and exit yaw angle deviation were measured with the help of a five-hole pressure probe. Static pressure distribution over the blade midspan was obtained by 16 channel Scanivalve. Aerodynamic performances of three different profiled endwalls are numerically analyzed and are compared against the BCP configuration. The effects of different contoured endwall geometries on endwall static pressure distribution and secondary kinetic energy were also discussed. Analysis shows that in the first contoured endwall configuration (EC1), the formation of stagnation zones at a contour valley close to the suction surface causes the exit total pressure loss coefficient to increase. The shifting of the contour valley near to the pressure surface (EC2 configuration) has resulted in local acceleration of the diverted pressure side leg of the horseshoe vortex over the hump toward the end of the passage. In the third configuration (EC3 configuration), reduced valley depth and optimum hump height have effectively redistributed the endwall pitchwise pressure gradient. The increased static pressure coefficient at the endwall near to the pressure surface has eliminated the PSB formation. In addition, computational results of unsteady Reynolds averaged Navier–Stokes simulations are obtained for analyzing transient behavior of PSB, with more emphasis on its migration on the pressure surface and transport across the blade passage. The additional work done by the mainstream fluid to transport the low momentum PSB fluid has caused higher aerodynamic penalty at the blade exit region. In this viewpoint, the implementation of contoured endwalls has shown beneficial effects by eliminating the PSB and related secondary vortices. At 27% of axial chord downstream of the blade trailing edge, a 4.1% reduction in the total pressure loss coefficient was achieved with endwall contouring.