Endwall Of Cascade Research Articles

Experimental study of film cooling characteristics of annular cascade endwall with a partitioned film cooling layout

Based on the heat transfer characteristic of the endwall and the distribution law of the turbine cascade vortex system, a partitioned film cooling layout was proposed on an annular cascade endwall. By the passage vortex lift-off line and the passage throat line, the endwall is divided into four regions, which were respectively upstream passage wedge region, upstream pressure side region, downstream pressure side region, and downstream passage wedge region. The endwall film cooling performances of the leakage flow and the partitioned film holes were studied experimentally by the PSP technique. Results indicate that the upstream passage wedge region can be cooled highly when the MFR of leakage flow is more than 1%. Owing to the film superposition effect, the η of the upstream pressure side region is lower than that of two downstream partitions. As the increased coolant MFR, the η of the upstream partition was increased, however, there was an opposite trend for the downstream partition. In addition, it is found that the leakage flow enhanced dramatically the film cooling characteristics of the partitioned film holes on the region of X/Cax ≤ 0.5 in the study of endwall full-coverage film cooling.

Experimental Study of Film Cooling Characteristics of Annular Cascade Endwall with a Partitioned Film Cooling Layout

Experimental Study of Film Cooling Characteristics of Annular Cascade Endwall with a Partitioned Film Cooling Layout

Effect of leakage flow on the endwall cooling performance in high subsonic turbine cascade

Gas Turbine is widely used in many fields, for the purpose of achieving higher efficiency, the main flow temperature needs to be improved. In the combustion chamber, the flame profile is designed to be a flat pattern distribution to reduce NOx, as well as to decrease thermal load. Thermal protection for the hot-end components has become an important issue.Within the present study, experimental and numerical methods are used to investigate the effects of leakage flow on cascade endwall cooling performance for high subsonic flow condition. Test variable including blowing ratio and upstream slot inclination angle. Blowing ratio(M) ranged from 0.3 to 2, upstream slot inclination angles are 0°, 45°,60°.The results stated that when the blowing ratio increases, the magnitudes of adiabatic cooling effectiveness increases, but the growing slope decreases. It is showed that the optimal performance is observed at a test condition of blowing ratio around 1.7. Under low blowing ratio (0.3–0.6) condition, the adiabatic cooling effectiveness of suction side is significantly better than the pressure side. As the blowing ratio increases, the coverage of coolant on endwall becomes more uniform, and the endawall thermal loads decreases.When compared at a particular value of blowing ratio, a higher adiabatic cooling effectiveness zone approaches to the suction side with a 45° slot inclination angle compared to 30 and 60° inclination angle. Changing the upper stream slot inclination angle, the local distribution of adiabatic cooling effectiveness is varied, but the spanwise averaged values of η do not show discernible variation.

The Effects of Inlet Boundary Layer Condition and Periodically Incoming Wakes on Secondary Flow in a Low Pressure Turbine Cascade

AbstractA particular turbine cascade design is presented with the goal of providing a basis for high quality investigations of endwall flow under high-speed conditions with unsteady inflow. The key feature of the design is an integrated two-part flat plate serving as a cascade endwall at part-span, which enables a variation of the inlet endwall boundary layer conditions. The new design is applied to the T106A low pressure turbine cascade for endwall flow investigations in the High-Speed Cascade Wind Tunnel of the Institute of Jet Propulsion at the Bundeswehr University Munich. Measurements are conducted under realistic flow conditions (M2th = 0.59, Re2th = 2 · 105) in three cases of varying endwall boundary layer conditions with and without periodically incoming wakes. The endwall boundary layer is characterized by 1D-CTA measurements upstream of the blade passage. Secondary flow is evaluated by five-hole-probe measurements in the turbine exit flow. A strong similarity is found between the time-averaged effects of unsteady inflow conditions and the effects of changing inlet endwall boundary layer conditions regarding the attenuation of secondary flow. Furthermore, the experimental investigations show that all design goals for the improved T106A cascade are met.

Experimental and numerical study on film cooling effectiveness of an annular cascade endwall with different slot configuration

To obtain higher thermal efficiency, advanced gas turbines operate at a turbine inlet temperature (TIT) up to 2000K, which causes an extremely high thermal load on the first stage vane endwall. Therefore, it is critical to understand the complex thermal environment near the endwall and develop efficient cooling schemes. In this study, adiabatic cooling effectiveness measurements on an annular cascade endwall with purge flow were conducted using a high-resolution infrared thermography technique. Effects of profiles and locations of an upstream slot on endwall film cooling behaviors were investigated at three coolant-to-mainstream mass flow ratios and an engine-like density ratio. Using numerical simulation, the flow details were illustrated to support the measured cooling effectiveness. Results indicate that moving the slot to a further upstream location will reduce film cooling effectiveness on the endwall. However, a further upstream slot can improve the uniformity of the film cooling effectiveness distribution. Compared with the normal slot, the convergent slot accelerates the coolant ejection to a higher momentum and thus inhibits the influence of secondary flows on endwall film cooling. Therefore, the convergent slot significantly improves both the coolant coverage and local film cooling effectiveness for all mass flow ratios and slot locations investigated. Quantitatively, an enhancement of approximately 25% in laterally averaged cooling effectiveness can be achieved by the convergent slot.

Investigation of improving the hydraulic turbine cascade performance using non-axisymmetric endwall contouring

Non-axisymmetric endwall contouring techniques have been widely applied in gas turbines; numerous papers and experimental studies have shown that it can significantly improve the overall performance of the turbine. This article first presents the non-axisymmetric endwall profile construction and optimization both for the rotor hub and shroud of a turbine drilling hydraulic cascade in the presence of an axisymmetric rotor, in order to improve the performance and investigate non-axisymmetric endwall contouring’s influence laws on a hydraulic cascade. Rotor cascade endwall design is studied by confining the axisymmetric variation near hub, shroud, and both endwalls. Endwall surface is parameterized with control points of Bezier curve, and control points are considered as design variables having a height constraint of 10%, 12%, or 15% span to move in radial coordinate. This methodology also combines Latin hypercube sampling with NSGA-II algorithm to obtain optimum solution. Computational fluid dynamics simulation results show that the non-axisymmetric endwall contouring technology applied to hub and shroud all realize an improvement in efficiency, while the NEW_S_15% has the best comprehensive performance and it causes an improvement of 1.52% in efficiency and 5.1% in torque. Off-design experiment shows that NEW_S_15% improves the output torque and efficiency as well even it deviates from design working condition, which proves valuableness of the non-axisymmetric endwall contouring technology.

Influences of large fillets on endwall flows in a vane cascade with upstream slot film-cooling

Investigations in cascade employ filleted blades to influence the near endwall secondary flows and total-pressure losses. The secondary flows aggravate the aerodynamic losses and endwall thermal stresses in the gas turbine passages. Investigations of different configurations of the slot-bleed flow from the endwall of cascade upstream show significant influences on the near endwall flows. In the present paper, the near endwall flow-field in a linear cascade employing filleted vanes and bleed flow from the upstream endwall slots is measured experimentally. Two fillet profiles are tested, one is larger than the other. The objectives are to quantify the combined effects of endwall fillet, fillet profiles, and film-cooling flow on the endwall region flow-field. The fillet covers the junction of vane and endwall upstream of the cascade throat region. The discontinuous bleed-slots near the cascade entrance provide the film-cooling flow on endwall and simulate the gaps between combustor/nozzle-vane discs or stator/rotor discs in the gas turbine. The inlet Reynolds number of 2.0E+05 is based on the chord length of vane profile. The density and temperature ratios of the coolant flow to mainstream are both about 1.0. The inlet blowing ratio of the film-cooling flow varies between 1.0 and 2.8. The flow-field is measured through the distributions of flow temperature, yaw angle, axial vorticity, and total-pressure losses along the vane passage. The effects on the flow-field are then presented by comparing the cases of filleted vanes with the cases of un-filleted vanes. The results of flow yaw angle and axial vorticity in the filleted passage are smaller than those in the passage without the fillet. The yaw angles responsible for the endwall secondary flows are the smallest for the smaller fillet (Fillet-2). The temperature field indicates the pitchwise distributions of the coolant on endwall specially near the pressure side are better when the smaller fillet is employed. Also, the weakened passage vortex of the endwall secondary flows in the filleted passage reduces the total-pressure losses. Although the total-pressure losses decrease at very high coolant mass flux with and without the fillet, the losses are always smaller for the filleted passage than for the un-filleted passage. The present investigation is beneficial for improving and optimizing the endwall film-cooling in the gas turbine passage.

Effects of Stator Platform Geometry Features on Blade Row Performance

This paper details an experimental investigation, using a linear cascade, into the effects of real geometry features on the aerodynamic performance of stator blade rows within axial flow compressors. The specific geometric features investigated include shroud cavities, inter-platform gaps, vane-pack gaps and the effects of misalignment of the platform endwalls due to manufacturing tolerances. A computational investigation into these effects is also included. To ensure that the linear cascade measurements are representative of a multi-stage compressor environment a novel experimental technique was developed to generate a hub endwall boundary layer which had skew. The boundary layer skew generation method involves injecting flow along the cascade endwall in such a manner as to control both the displacement thickness and tangential momentum thickness of the resulting boundary layer. Without the presence of the endwall boundary layer skew the linear cascade could not reproduce the flow features typically observed in a multi-stage compressor. The investigation reveals that real geometry features can have a significant impact on the flowfield within a blade passage. For a shrouded stator, increasing the leakage flow rate increases the stagnation pressure loss coefficient. However, high levels of whirl pickup of the leakage flow as it passes through the stator-shroud cavity can offset the natural secondary flow within the stator passage and thus reduce the stagnation pressure loss. All of the steps and gaps that were observed to be present in real compressors were found to increase the stagnation pressure loss relative to that of a smooth endwall. It is also shown that the computational method is capable of capturing the trends observed in the experiments.

The sensitivity of turbine cascade endwall loss to inlet boundary layer thickness

The development of hub and casing boundary layers through a turbomachine is difficult to predict, giving rise to uncertainty in the boundary conditions experienced by each blade row. Previous studies in turbine cascades disagree on the sensitivity of endwall loss to such inlet conditions. This paper explores the problem computationally, by examining a large number of turbine cascades and varying the inlet boundary layer thickness. It is demonstrated that the sensitivity of endwall loss to inlet conditions is design dependent, and determined by the component of endwall loss associated with the secondary flow. This Secondary-Flow-Induced loss is characterised by a vorticity factor based on classical secondary flow theory. Designs that produce high levels of secondary vorticity tend to generate more loss and are more sensitive to inlet conditions. This sensitivity is largely driven by the dissipation of Secondary Kinetic Energy (SKE): thickening the inlet boundary layer causes the secondary vorticity at the cascade exit to be more dispersed within the passage, resulting in larger secondary flow structures with higher SKE. The effects are captured using a simple streamfunction model based on classical secondary flow theory, which has potential for preliminary design and sensitivity assessment.

Enhancing film cooling effectiveness in a gas turbine end-wall with a passive semi cylindrical trench

Computational studies were carried out in the end-wall of a linear cascade, of chosen blade profile, which is provided with one row of cylindrical film cooling holes inclined at 30o to the end wall. The CO gas was used as the coolant supplied through the film holes, 2 maintaining a blowing ratio of 0.6. The film cooling hole row was positioned at the leading edge of the cascade. The mainstream fluid was air and based on its properties at the cascade inlet, the flow was found turbulent. A semi cylindrical trench was placed at two positions upstream of the cascade leading edge and three positions downstream of it. ANSYS FLUENT 15.0 was used to compute the film cooling effectiveness of the cascade endwall. Trench positioned at a distance of twice that of film hole diameter, was found to show a highest increase of area averaged effectiveness value by 30.4% over the baseline. Further to this, the influence of the trench diameter was carried out where the trench with diameter twice that of film hole diameter was found to show a 31.3% increase of cooling effectiveness over the baseline. Studies on the influence of blowing ratio showed a highest increment of cooling effectiveness value by 43.5% over the baseline a blowing ratio of 1.2.

Experimental and numerical investigations of overall cooling effectiveness on a vane endwall with jet impingement and film cooling

In this paper, a conjugate heat transfer model for the endwall of a turbine four-vane linear cascade is developed to examine the conjugate cooling effects generated by internal jet impingement and external film cooling as well as heat conduction through the metal endwall. Aerodynamic and geometrical parameters are appropriately scaled to match engine conditions. The conjugate model with a maximum Biot number of 1.5 is tested in engine-like oncoming flows with a turbulence intensity of 9.8% and an integral length scale of 10 mm. The effects of varying passage inlet Reynolds numbers from 1.40 × 105 to 4.20 × 105 and coolant-to-mainstream mass flow ratios from 1.5% to 3.8% are investigated by using experimental measurements and numerical simulations in the presence of an upstream slot. Both experimental and numerical results reveal that overall cooling effectiveness on the endwall increases with the increase of coolant mass flow rate. The effects of passage flow inlet Reynolds number on endwall overall cooling performance are more complicated, that depends on competing effects of internal and external heat transfer. Overall cooling effectiveness is found to be significantly enhanced in the vicinity of the film cooling holes due to higher in-hole convective heat transfer levels. Computational results, which show good agreement with measurements, provide additional information of thermal behavior in the endwall and explain why there is improvement with coolant mass flow ratio.

Effect of discrete-hole arrangement on film-cooling effectiveness for the endwall of a turbine blade cascade

This paper investigates film cooling in a turbine cascade endwall for two discrete-hole configurations using liquid crystal thermography. The discrete holes, arranged in rows aligned in the pitchwise direction, gave rise to relative maxima of film-cooling effectiveness downstream of each row, followed by a marked decrease of effectiveness along the gap between adjacent rows of holes. This resulted in a not efficient coverage of the endwall surface, with succession of over- and under-cooled regions. The re-design of discrete-hole configuration, based on knowledge of the heat transfer coefficient map on the endwall without film cooling, enabled the redistribution of the coolant to provide a better coverage of the endwall and a significant increase of the area-averaged film-cooling effectiveness.

Vortex control and aerodynamic performance improvement of a highly loaded compressor cascade via inlet boundary layer suction

Effects of inlet boundary layer suction on the vortex structure and cascade loss in a highly loaded compressor cascade were investigated experimentally. Ink-track visualization was undertaken on cascade endwall and the blade surface. Ten traverse planes from upstream to downstream of the cascade in a rectangular wind tunnel were measured by an L-shaped five-hole probe. These tested planes revealed the process of emergence, development and decline of several principal vortices as well as the corresponding additional losses. Details of ink-track visualization displaying the secondary flow behavior of boundary layer upon endwall and blade surface assist to make judgment on vortex evolution. Inspection of the vortex structure revealed that highly loaded compressor was characterized by large-scale vortices in the endwall region. After suction, these vortices are all well organized and under control. Among all of them, passage vortex is most sensitive to the variation of the inlet boundary layer, and its main function is to spread low-energy fluid rather than to produce loss. On the other hand, a wall vortex and a concentrated shedding vortex take place inside and after the cascade, respectively, and engender considerable accompanying loss as they dissipate. The effects of inlet boundary layer suction on them are correspondingly weaker. About one forth of the total loss in the baseline cascade was eliminated when boundary layer suction flow rate reaches 2.5 % of the inlet mass flow. The feasibility of simplifying the suction system is also verified through this work.

Large-eddy simulation analysis of mechanisms for viscous losses in a turbomachinery tip-clearance flow

The tip-leakage flow in a turbomachinery cascade is studied using large-eddy simulation with particular emphasis on understanding the underlying mechanisms for viscous losses in the vicinity of the tip gap. Systematic and detailed analysis of the mean flow field and turbulence statistics has been made in a linear cascade with a moving endwall. Gross features of the tip-leakage vortex, tip-separation vortices, and blade wake have been revealed by investigating their revolutionary trajectories and mean velocity fields. The tip-leakage vortex is identified by regions of significant streamwise velocity deficit and high streamwise and pitchwise vorticity magnitudes. The tip-leakage vortex and the tip-leakage jet which is generated by the pressure difference between the pressure and suction sides of the blade tip are found to produce significant mean velocity gradients along the spanwise direction, leading to the production of vorticity and turbulent kinetic energy. The velocity gradients are the major causes for viscous losses in the cascade endwall region. The present analysis suggests that the endwall viscous losses can be alleviated by changing the direction of the tip-leakage flow such that the associated spanwise derivatives of the mean streamwise and pitchwise velocity components are reduced.

Measurement and Calculation of Turbine Cascade Endwall Pressure and Shear Stress

The complex three-dimensional fluid flow on the endwall in an axial flow turbine blade or vane passage has been extensively investigated and reported on in turbomachinery literature. The aerodynamic loss producing mechanisms associated with the endwall flow are still not fully understood or quantitatively predictable. To better quantify wall friction contributions to endwall aerodynamic loss, low Mach number wind tunnel measurement of skin friction coefficients have been made on one endwall of a large scale cascade of high pressure turbine airfoils, at engine operating Reynolds numbers. Concurrently, predictive calculations of the endwall flow shear stress have been made using a computational fluid dynamics (CFD) code. Use of the oil film interferometry skin friction technique is described and applied to the endwall, to measure local skin friction coefficients and shear stress directions on the endwall. These are correlated with previously reported measured local endwall pressure gradients. The experimental results are discussed and compared to the CFD calculations, to answer questions concerning endwall aerodynamic loss predictive ability.

Flow visualization on the linear compressor cascade endwall using oil flows and laser Doppler anemometry

Surface oil flow visualizations were conducted in the VPI&SU Department of AOE Low Speed Linear Cascade Wind Tunnel on the endwall. Oil flow visualizations of the endwall flow features (passage flow, cross flow and tip leakage vortex) for tip gap to chord ratios (t/c) of 1.65% and 3.3% are discussed in detail. For the first time, quantitative comparisons are made between the oil-flow direction and the wall shear stress direction obtained from laser Doppler anemometer experimental results. The oil flow streaks are aligned with the wall shear stress direction, except near separation, which is proved theoretically and experimentally. Wall shear stress directions in the oil-flow images are measured and the separation line in the oil flow is calibrated based on the skin friction results from the LDA measurement.

1632 リブ付きタービン翼列後流壁面の静圧分布特性

1632 リブ付きタービン翼列後流壁面の静圧分布特性

Heat transfer near turbine nozzle endwall.

This paper gives an overview and reviews recent findings concerning turbine endwall cooling in the literature. The text below begins with a brief discussion of the secondary flows and heat transfer around cascade endwall. This will be followed by a review of recent developments in cooling concepts and related heat transfer results. Key topics include: film cooling, upstream bleeding, endwall contouring, and leakage through component interfaces.

Counter-rotating streamwise vortex formation in the turbine cascade with endwall fence

The three-dimensional turbulent flows in the turbine cascade with and without endwall fences are numerically investigated by solving the incompressible Navier–Stokes equations with a high-Reynolds number k– ϵ turbulence closure model. The limiting streamline patterns and the static pressure contours at the suction surface of the blade as well as on the cascade endwall are employed to visualize the effectiveness of the endwall fence for the secondary flow control. The analysis on the streamwise vorticity contour maps along the cascade with the three-dimensional representation of their iso-surfaces reveals the complicated structures of the vortical flows in the turbine cascade with endwall fence, and also leads to an understanding on the formation of a counter-rotating streamwise vortex over the fence. The mechanisms of controlling the secondary flow and the proper selection of an optimal fence height are also explained.

Measurements of Endwall Heat(Mass) Transfer Coefficient in a Linear Turbine Cascade Using Naphthalene Sublimation Technique

Heat (mass) transfer characteristics have been investigated on the endwall of a large-scale linear turbine cascade. Its profile is based on the mid-span of the first-stage rotor blade in a industrial gas turbine. By using the naphthalene sublimation technique, local heat (mass) transfer coefficients are measured for two different free-stream turbulence intensities of 1.3% and 4.7%. The results show that local heat (mass) transfer Stanton number is widely varied on the endwall, and its distribution depends strongly on the three-dimensional vortical flows such as horseshoe vortices, passage vortex, and corner vortices. From this experiment, severe heat loads are found on the endwall near the blade suction side as well as near the leading and trailing edges of the blade. In addition, the effect of the free-stream turbulence on the heat (mass) transfer is also discussed in detail.