Pressure-side Corner Research Articles

Effects of rotor squealer tip with non-uniform heights on heat transfer characteristic and flow structure of turbine stage

Squealer tip has a significant influence on both aerodynamic and heat transfer characteristics of the high-pressure turbine. Among the geometric parameters of the squealer, squealer height is one of the essential parameters in the tip design. However, due to the complexity of parameterization and meshing of the squealer, the related research is usually carried out on the squealer with a constant height. In this paper, a parameterization strategy generates squealer of assigned heights at four key positions of the blade, the leading edge-pressure side, the leading edge-suction side, the trailing edge-pressure side, and the trailing edge-suction side. An in-house mesh generation platform (NuFlux) is adopted to automatically generate the structured meshes. The aerothermal performance of a transonic turbine stage is assessed using steady Reynolds-averaged Navier–Stokes simulations with the k−ω shear stress transport model for the turbulence closure. The main purpose is to obtain the squealer tip configuration with the lowest heat transfer coefficient. The results show that non-uniform squealer further reduces the cavity floor heat transfer on the basis of uniform squealer by changing the interaction process between the asymmetric vortex pair (the pressure-side corner vortex and the casing-driven scraping vortex), which provides a valuable reference for the design of the squealer tip of advanced high-pressure turbines.

The loss mechanisms and vortex dynamics of rotor platform step geometries in a low-speed research compressor stage

The construction and assembly of an axial compressor blade row introduces real geometric features, which leads to the deviation of flow characteristics from the intended design. According to the realistic assembly errors observed in the four-stage low-speed research compressor developed by Shanghai Jiao Tong University, two representative step geometries at the rotor platform were abstracted. Unsteady simulations were carried out in the stage environment to study the influence of hub discontinuities on compressor performance and loss mechanisms. The comparisons of smooth hub, elevated hub (EH), and a wedged hub (WH) demonstrated that all the step geometries resulted in increased loss inside both rotor and stator rows. The loss induced by WH was more sensitive to step heights than EH. Under the influence of forward facing step, rotor incidence was reduced and horseshoe vortex (HV) was exacerbated, leading to the accumulation and separation of low momentum fluids at the pressure side corner. This led to a redistribution of losses across the span. In the lower half span, the relative total pressure loss increased almost linearly with step heights, while in the tip region, it reduced slightly. The backward facing step generated the step corner vortices (SCV) that shed and propagated downstream. The SCV interacted with the HV and cavity leakage vortex in the stator passage. Coupled with the augmented stator inflow angle, the corner separation at the suction side was dramatically intensified and led to a significant loss increase.

Effects of the Slotted-Cavity Tip on the Vortices and Aerothermal Performance in the High-Pressure Turbine

To enhance aerodynamic performance in high-pressure turbines, cavity tips are commonly employed. However, these cavity tips tend to generate more vortices than flat tips, leading to an increased thermal load on the blade tips. This study introduces a novel configuration aimed at suppressing vortex intensity by incorporating a slot at the cavity bottom. Through numerical simulations, the impact of slot geometry parameters on vortices and aerothermal performance is investigated. The findings reveal that a slot height of 0–0.1 mm effectively suppresses the pressure side corner vortex within the cavity, while a slot height of 0.2–0.3 mm eliminates it. A slot height of 0.3 mm reduced average heat transfer and the high heat transfer region at the cavity bottom by 22.6% and 92.6%, respectively; however, it was observed that the leakage flow increased by 15.9%. Therefore, this study further combines the slot structure with a winglet to control the leakage flow. The combined structures exhibit superior overall performance compared to the single-slot cavity tip. Compared to the conventional cavity tips, the combined structure achieves a maximum reduction of 27.3% in average heat transfer and 7.9% in leakage flow, along with a decline of 1.9% in total pressure loss.

An Active Phantom Cooling Concept for Turbine Endwall Cooling From Pressure-Surface Film Coolant Injection

Abstract Cooling of the endwall of a turbine vane should receive special attention due to its uniqueness of near-wall complex secondary flows and concomitant challenge of offering film-coverage for cooling the endwall pressure-side corner regions. The use of internal enhanced cooling at the endwall backside could be an option, but it increases manufacturing cost, adds weight to the component, causing excessive pressure losses in the secondary air system. Novel film cooling concepts are, therefore, required to provide effective cooling for these difficult-to-cool regions. This study proposes an active cooling concept effected by placing a row of film cooling holes on the vane pressure surface near the endwall with the intention of utilizing second-order cooling (or phantom cooling) from pressure-surface film-coolant injection to provide increased cooling effectiveness and enlarge coverage on the endwall. The effects of hole diameter, injection angle, and compound angle, as well as coolant injection rate are investigated. Detailed phantom cooling effectiveness over the endwall is documented using Pressure-Sensitive Paint (PSP). To provide a description of the flow physics driving the cooling process, computational modeling is carried out to document mixing of coolant with the freestream. Experiments show that significant cooling occurs in the endwall pressure-side corner and extends beyond the passage throat. Higher coolant injection rates and an optimized pressure-surface injection geometry maximize endwall phantom cooling. An effectiveness correlation for the active cooling is developed to provide a straightforward tool for designers to apply in turbine design.

Improving turbine endwall overall cooling effectiveness using curtain cooling and redistributed film-hole layouts: an experimental and computational study

Abstract To enhance the cooling deficiency that occurs in a baseline endwall using axially-arranged cooling holes, this paper proposes a new locally-enhanced hole layout using curtain cooling and fan-shaped film holes being arranged on iso-Mach lines. The objective of cooling hole re-design is to minimize secondary flows and thus to provide better film coverage. In experiments, Infrared thermography techniques are applied to validate overall cooling effectiveness of the newly-designed endwall, and aero-thermal fields at the cascade exit are detected by five-hole and thermocouple probes. Additionally, computational fluid dynamic simulations are performed to provide complementary flow insights. A comparison with the baseline hole layout reveals that for a given total coolant flow rate, the newly-designed endwall significantly improves the cooling performance by up to 27% without a noticeable aerodynamic penalty, resulting in a lower and more uniform temperature field. Curtain coolant effectively suppresses the development of horseshoe vortex and provides adequate thermal protection for leading-edge junctures and pressure-side corner regions. The redistribution of fan-shaped film holes reinforces the cooling performance in the passage throat and trailing edge regions. At low and high total mass flow rates, the coolant split between various cooling sources has a substantial impact on cooling performance.

Vortical characteristics of the corner separation flow in a high-speed compressor cascade

In aero-engine compressors, corner separations cause blockage, loss generation, and flow instabilities. The aim of the current study is to gain better knowledge of the vortical structures in corner separation flows and their resultant blockage and loss. Numerical simulations at different incidences have been performed for a high-speed linear compressor cascade. First, vortical structures have been visualized with the help of the Q-criterion and eigenvector method, and those associated with the corner separation have been focused on. It was found that the corner separation occurring on either the pressure surface or suction surface corner region was characterized by a large-scale streamwise vortex, named as the corner separation vortex in this paper. The pressure-side corner separation vortex is observed at a great negative incidence initiated from a focus point on the pressure surface, while the suction-side corner separation vortex observed at all concerned incidences originated from a focus on the endwall near the suction surface. Then, the mechanism by which corner separation vortices resulted in blockage and loss has been uncovered. The results showed that low-momentum fluids inside the core region of corner separation vortices led to the flow blockage and the viscous friction between them, and mainstream was responsible for the entropy generation. One major contributor of low-momentum fluids in the suction-side corner region was an off-wall reversed-flow region which was attributed to the breakdown of the suction-side corner separation vortex. Finally, multi-vortex structures in the leading region of the suction-side corner separation at two high positive incidences have been analyzed. It was found that the horseshoe vortex influenced the formation of the suction-side corner separation vortex, which induced a multi-vortex structure, and prevented the onset location of suction-side corner separation vortex from moving upstream with the increased incidence.

Implementation of Rectangular Vortex Generator Pairs to Improve Film Cooling Effectiveness on Transonic Rotor Blade Endwall

Abstract The rectangular vortex generator pairs (RVGPs) are arranged upstream of the film cooling holes to achieve better coolant coverage on the endwall near the pressure-side corner area. The endwall film cooling effectiveness distributions under transonic flow conditions are numerically calculated for the single RVGP and double rows of RVGPs cases. At first, the effects of three geometrical parameters (i.e., the distance between RVGP and cooling hole, height of RVGP and attack angle of RVGP) on endwall film cooling effectiveness are studied with a single hole and RVGP at different mainstream inlet Reynolds numbers and blowing ratios. Then, the double rows of RVGPs are applied to further enhance the overall film cooling effectiveness on the blade endwall. The results show that the implementation of RVGPs significantly enhances the film cooling effect on transonic blade endwall at a pressure-side corner area. With the increase of RVGP height, the lateral coolant coverage on the endwall corner area is improved. However, by decreasing the distance between the vortex generator pair and cooling hole, the film cooling effectiveness downstream of the cooling holes is increased. The attack angle of RVGP mainly affects the shape of coolant spreading on endwall surface. The RVGP with optimum dimensions and arrangement is able to suppress the coolant from lifting off the endwall and increase the coolant diffusion near the endwall. Compared with no vortex generator case, the area-averaged film cooling effectiveness on endwall with double rows of RVGPs is improved by 13.16%.

Large-Eddy Simulation of a Single Airfoil Tip-Leakage Flow

A large-eddy simulation is performed on a single nonrotating airfoil with tip gap in order to highlight the noise sources in the tip-leakage flow. The geometry is composed of a NACA 5510 airfoil (5% camber, 10% thickness) mounted between two flat plates, with a 10 mm tip gap between the airfoil and one of the plates. The Reynolds number based on the chord is 960,000 and the Mach number is 0.2. The aerodynamic results from the simulation show the formation of the tip-leakage vortex and the interaction of the turbulent structures that exit the gap with the external flow near the trailing edge. Then, a Ffowcs Williams and Hawkings analogy is used to compute the far-field noise levels and an acoustic decomposition is performed to highlight the main surface contributions to the tip-leakage noise. The numerical simulation shows very good agreement compared with experiments and allows identification of the main noise source in the tip gap. The main noise source results from the interaction of the turbulent structures created in the tip gap close to the pressure side corner with the suction side tip edge. The scraping of turbulent structures along the tip edge is shown to be an efficient noise mechanism.

Effect of Non-Axisymmetric Endwall Profiling on Heat Transfer and Film Cooling Effectiveness of a Transonic Rotor Blade

Abstract Effects of non-axisymmetric endwall profiling on total pressure loss, heat transfer, and film cooling effectiveness of a transonic rotor blade were numerically investigated. The numerical methods, including the turbulence model and grid sensitivity, were validated with the existing experimental data. To reduce the thermal load on endwall, non-axisymmetric endwall profiling near leading edge and at pressure-side corner area was performed with a range of contour amplitudes. Heat transfer and flow fields near the profiled endwalls were analyzed and also compared with the plain endwall configuration. On the profiled endwall, three kinds of cooling holes, i.e., cylindrical holes, rounded-rectangular holes, and elliptical holes, were arranged, and film cooling effect was investigated at three blowing ratios. Results indicate that, with endwall profiling, the area-averaged Stanton number on endwall is reduced by 7.71% and total pressure loss in cascade is reduced by 11.07%. Among three kinds of cooling holes, the arrangement of the elliptical hole performs the best film cooling effect on the profiled endwall. Compared with the plain endwall, non-axisymmetric endwall with elliptical cooling holes improves film cooling coverage by 10.87%, reduces the Stanton number by 8.88%, and increases the net heat flux reduction performance by 4% at M = 0.7.

Effects of the squealer winglet structures on the heat transfer characteristics and aerodynamic performance of turbine blade tip

The heat transfer characteristics and aerodynamic performance of three different squealer winglet structures based on GE-E3 turbine rotor squealer tip are numerically investigated, which is compared with the conventional squealer tip. The Reynolds-Averaged Navier-Stokes (RANS) solutions and standard k-ω turbulence model are solved to analyze the aerothermal performance of rotor tip with different squealer winglet structures. The numerical method is verified with experimental data. The numerical results show that the area-averaged heat transfer coefficient of squealer winglet structure at the pressure side, suction side and both the pressure and suction side are respectively reduced by 12.2%, 17.1% and 19.8% compared to the conventional squealer tip. The squealer winglet structure on the blade tip mostly influence the flow structures inside the cavity with the decreased strength of the pressure side corner vortex and the scraping vortex, which reduce the heat transfer coefficient and the leakage flow rate of the blade tip. The squealer winglet structure at the pressure side is found to increase the total pressure loss coefficient of the rotor blade by 8.5% in contrast to the conventional squealer tip. However, the squealer winglet structure at the suction side and both sides effectively reduce the total pressure loss coefficient by 8.5% and 2.5% respectively. The results reveal that squealer winglet structure at the suction side has the tradeoff of the heat transfer characteristics and aerodynamic performance among four different turbine rotor squealer tip profiles.

Reduction of endwall secondary flow losses with leading-edge fillet in a highly loaded low-pressure turbine

A detailed experimental and numerical investigation of the effect of leading-edge fillet on the endwall secondary flow was performed in a highly loaded low-pressure turbine cascade. The objective is to gain further understanding on the reduction mechanism of endwall secondary flow losses by leading-edge fillet. The results suggest that the numerical results offer a reliable prediction for the endwall secondary flow structure together with suction surface separation bubble. A pressure-side corner separation vortex is found in the front part of pressure surface/endwall junction, where an inverse flow goes upward and over the leading edge to suction side of the same blade. This inverse flow also takes part in the interaction between endwall secondary flow and separation bubble. However, the leading-edge fillet can fill the region of corner separation vortex, and the inverse flow is eliminated. Combined with the reduction of cross-pressure gradient achieved by this fillet, the loss coefficient and flow angle deviation at exit plane are decreased. Besides, the leading-edge fillet will reduce the intensity of leading-edge horseshoe vortex and thus decrease the loss of passage vortex and shed vortex. It is these three aspects introduced by leading-edge fillet that achieves an effective reduction of the endwall secondary loss in turbine.

Turbine Blade Tip Heat Transfer in Low Speed and High Speed Flows

In this paper, high and low speed tip flows are investigated for a high-pressure turbine blade. Previous experimental data are used to validate a computational fluid dynamics (CFD) code, which is then used to study the tip heat transfer in high and low speed cascades. The results show that at engine representative Mach numbers, the tip flow is predominantly transonic. Thus, compared with the low speed tip flow, the heat transfer is affected by reductions in both the heat-transfer coefficient and the recovery temperature. The high Mach numbers in the tip region (M>1.5) lead to large local variations in recovery temperature. Significant changes in the heat-transfer coefficient are also observed. These are due to changes in the structure of the tip flow at high speed. At high speeds, the pressure side corner separation bubble reattachment occurs through supersonic acceleration, which halves the length of the bubble when the tip-gap exit Mach number is increased from 0.1 to 1.0. In addition, shock/boundary-layer interactions within the tip gap lead to large changes in the tip boundary-layer thickness. These effects give rise to significant differences in the heat-transfer coefficient within the tip region compared with the low speed tip flow. Compared with the low speed tip flow, the high speed tip flow is much less dominated by turbulent dissipation and is thus less sensitive to the choice of turbulence model. These results clearly demonstrate that blade tip heat transfer is a strong function of Mach number, an important implication when considering the use of low speed experimental testing and associated CFD validation in engine blade tip design.

Numerical investigation of active tip-clearance control through tip cooling injection in an axial turbine cascade

This paper presents a numerical investigation of an active tip-clearance control method based on cooling injection from the blade tip surface. It aims to study the influences of air injection on controlling tip clearance flow, with emphasis on the effects of the injection location on secondary flow and the potential thermal benefits from the cooling jet. The results show that injection location plays an important role in the redistribution of secondary flow within the cascade passage. Injection located much closer to the pressure-side corner performs better in reducing tip clearance massflow and its associated losses. However, it also intensifies tip passage vortex, due to less restraint deriving from the reduced tip clearance vortex. Lower plenum total pressure is required to inject equivalent amount of cooling air, but the heat transfer condition on the blade tip surface is a bit worse than that with injection from the reattachment region. Thus the optimum location of air injection should be at the tip separation vortex region.

Effects of incidence angle on endwall convective transport within a high-turning turbine rotor passage

Effects of incidence angle on the endwall convective transport within a high-turning turbine rotor passage have been investigated. Surface flow visualizations and heat/mass transfer measurements at off-design conditions are carried out at a fixed inlet Reynolds number of 2.78 × 10 5 for the incidence angles of −10°, −5°, 0, 5°, and 10°. The result shows that the incidence angle has considerable influences on the endwall local transport phenomena and on the behaviors of various endwall vortices. In the negative incidence case, convective transport is less influenced by the leading edge horseshoe vortex and by the suction-side corner vortex along their loci but is increased along the pressure-side corner vortex. In the case of positive incidence, however, convective transport is augmented remarkably along the leading edge horseshoe vortex, and is much influenced by the suction-side corner vortex. Moreover, heat/mass transfer is enhanced significantly along the pressure-side leading edge corner vortex. Local endwall convective transport in the area other than the endwall vortex sites is influenced significantly by the cascade inlet-to-exit velocity ratio which depends strongly on the incidence angle.

고선회 터빈 동익 팁 표면에서의 열전달 특성

Abstract The heat/mass transfer characteristics on the plane surface of a high-turning first-stage turbine rotor blade has been investigated by employing the naphthalene sublimation technique. At the Reynolds number of 2.09×10 5 , heat/mass transfer coefficients are measured for the gap height-to-chord ratio, h/c , of 2.0% at turbulence levels of Tu = 0.3 and 14.7%. A tip-surface flow visualization is also performed for h/c = 2.0% at Tu = 0.3%. The results show that there exists a strong flow separation/re-attachment process, which results in severe local thermal load along the pressure-side corner, and a pair of named tip gap vortices in this study is identified along the pressure- and suction-side corners near the leading edge. The loci and subsequent development of the pressure- and suction-side gap are discussed in detail. The combustor-level high inlet turbulence, which increases the tip-surface heat/mass transfer, provides more uniform thermal-load distribution.기호설명

Aerodynamic character of partial squealer tip arrangements in an axial flow turbine. Part I: Detailed aerodynamic field modifications via three dimensional viscous flow simulations around baseline tip

This paper deals with the viscous flow simulations of the complex tip leakage flow in Axial Flow Turbine Research Facility (AFTRF). Special attention is paid to the 3D structure of the tip leakage flow mechanisms in a baseline tip configuration with no desensitisation. Although past experimental studies provide much insight into the physical understanding of the tip region aerodynamics, there are still many areas of the flow-field in which experiments are extremely difficult to perform. Fine details of the entrance flow near the pressure side where the tip leakage jet starts to form, the leakage jet formation between the tip surface and outer casing, the re-circulatory flow zone very near the pressure side corner, the interaction area of the tip vortex with the conventional passage flow, the influence of the relative motion of the outer casing and leakage flow reversal can be visualised with excellent resolution. After the presentation of the measured inlet boundary conditions and a grid independency study, the baseline tip flow simulations are discussed in detail.

Heat Transfer and Effectiveness on Film Cooled Turbine Blade Tip Models

In unshrouded axial turbine stages, a small but generally unavoidable clearance between the blade tips and the stationary outer seal allows a clearance gap leakage flow to be driven across the blade tip by the pressure-to-suction side pressure difference. In modern high-temperature machines, the turbine blade tips are often a region prone to early failure because of the presence of hot gases in the gap and the resultant added convection heating that must be counteracted by active blade cooling. The blade tip region, particularly near the trailing edge, is often very difficult to cool adequately with blade internal coolant flow, and film cooling injection directly onto the blade tip region can be used in an attempt to reduce the heat transfer rates directly from the hot clearance flow to the blade tip. An experimental program has been designed and conducted to model and measure the effects of film coolant injection on convection heat transfer to turbine blade tips. The modeling approach follows earlier work that found the leakage flow to be mainly a pressure-driven flow related strongly to the airfoil pressure loading distribution and only weakly, if at all, to the relative motion between blade tip and shroud. In the present work the clearance gap and blade tip region is thus modeled in stationary form with primary flow supplied to a narrow channel simulating the clearance gap above a plane blade tip. Secondary film flow is supplied to the tip surface through a line array of discrete normal injection holes near the upstream or pressure side. Both heat transfer and effectiveness are determined locally over the test surface downstream of injection through the use of thin liquid crystal coatings and a computer vision system over an extensive test matrix of clearance heights, clearance flow Reynolds numbers, and film flow rates. The results of the study indicate that film injection near the pressure-side corner on plane turbine blade tips can provide significant protection from convection heat transfer to the tip from the hot clearance gap leakage flow.