Endwall Flow Research Articles

Effects of upstream slot geometry on the endwall aerothermal performance of a gas turbine blade under different ejection angle conditions

In the current study, effects of upstream slot geometry with different ejection angles on endwall flow and heat transfer characteristics of a gas turbine blade are numerically investigated. Based on validation of the numerical method, the Reynolds-averaged Navier-Stokes (RANS) equations combined with the standard k-ω turbulence model are adopted in this study. The effects of convergent and divergent slot geometries on near-endwall flow structure and endwall heat transfer behavior are studied when ejection angles are 45deg and 90deg respectively. The results indicate that compared with a normal slot, the convergent slot significantly improves the endwall cooling performance when the ejection angle is 45deg, so the thermal loads on the endwall are remarkably reduced. Moreover, the convergent slot can also decrease the aerodynamic losses in blade passage when the coolant to main flow mass flow ratio is higher than 1.0%. However, when the ejection angle is 90deg, the advantage of the convergent slot in endwall cooling does not exist and heat transfer on the fore part of the endwall is increased by the convergent slot. Consequently the thermal loads on the fore part of the endwall are enhanced, especially, when mass flow ratio reaches 1.5%, the thermal loads on the endwall near the leading edge are increased by 28% compared with that for the normal slot. Besides, the convergent slot also increases the aerodynamic losses when the ejection angle is 90deg. As to the divergent slot, it has an adverse effect on cooling protection for the endwall with both ejection angles and its effects on both cooling protection and aerodynamic performance are much less than those for the convergent slot.

Experiment study of the winglet-cavity tip on the aerodynamic performance in a turbine cascade

The effects of a novel winglet-cavity tip on the flow field and aerodynamic performance of a turbine blade with tip clearance have been investigated in a low-speed wind tunnel. A calibrated five-hole probe is used for the measurement of three-dimensional flows downstream of the cascade. The method of oil-flow visualization is used to show the endwall flow field structure. The distribution of endwall static pressure is measured particularly by using the special moveable endwall. The downstream results show that, compared with the flat tip and cavity tip, the winglet-cavity tip reduces aerodynamic loss in the region of tip leakage vortex and passage vortex effectively and gives a 8.5% reduction of total pressure losses at a tip clearance of τ/ H = 1.0%. Meanwhile, a more uniform flow angle is obtained with the winglet-cavity tip. Thus, the winglet-cavity tip provides better aerodynamic performance. It was found that more endwall flow enters the cavity from the front of suction side gap, combines with the flow entering the tip from the pressure side, and then separates upon the cavity. This reduces the loss of passage vortex. The endwall static pressure indicates that the winglet-cavity tip reduces the driving pressure difference and weakens the tip leakage flow. With the tip clearance increasing, the leakage flow is significantly enhanced. This strengthens the interaction between the tip leakage vortex and the passage vortex. With respect to the flat tip and cavity tip, the winglet-cavity tip obtains the lowest total pressure loss at all tested tip clearances.

Blade-end treatment for axial compressors based on optimization method

This paper concentrates on the application of blade-end treatment to axial compressors by means of the optimization algorithm. The blade-end treatment reduces the end wall losses and extends the stable margin by modifying blade shape near the end wall region. It contains three types of passive flow control measures, i.e., the end-bend, end-dihedral and end-sweep treatment. Firstly, the effects of blade-end treatment were reviewed based on the open literature published over the past 30 years. All of these effects essentially influence the compressor performance by changing the blading loading distributions in the streamwise or spanwise directions. There is a trade-off between the improved end wall flows and the deteriorated mid-span flows. It’s difficult to quantitatively apply these measures to achieve an optimal balance according to the traditional engineering experience. Optimization algorithm provides an efficient access to resolve this issue by automatically obtaining the utmost benefit. Secondly, an optimization example of NASA Stage 35 was conducted to validate against the summarized flow mechanisms. The optimal geometry parameters of cantilever stator vane near the end wall region were obtained by employing a surrogate model in conjunction with a genetic algorithm for optimization. Finally, optimization results indicated that the optimal vane blade featured an obvious combination of forward end-sweep, positive end-dihedral and end-bend. The stator total pressure losses were reduced with the blade-end treatment based on optimization method. A significant reduction of loss occurred near the shroud region, from the 80% span to the casing, while the performance was degraded within the mid-span region, approximately 50%–80% span. The resulting mechanisms are consistent with the knowledge obtained from the literature review and this will provide meaningful guidance on the further compressor design process.

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.

Loss Mechanisms of Interplatform Steps in a 1.5-Stage Axial Flow Turbine

In this paper, the detailed steady and unsteady numerical investigations of a 1.5-stage axial flow turbine are conducted to determine the specific influence of interplatform steps in the first stator—as caused by deviations in manufacturing or assembly. A basic first stator design and a design consisting of a bow and endwall contours are compared. Apart from step height, the position and geometry of the interplatform border are varied for the basic design. To create the steps, every third stator vane was elevated, together with its platforms at hub and shroud, such that the flow capacity is only little affected. The results show that the effects of steps on the platform borders in front and aft of the first stator can be decoupled from those occurring on the interplatform steps. For the latter, being the main contributor to the additional loss, the intensity of recirculation zones and losses increase substantially when the platform border is located close to the suction side. Using a relative step height of 1.82% span, the entropy production doubles when compared to a position close to the pressure side, which can be explained by differences in local flow velocity level. Regarding a circular-arc-shaped platform, the losses can be more than halved—mainly due to lower included angles between step and endwall flow streamlines. The findings can be explained by a nondimensional relation of the local entropy production using local values for step height and characteristic flow quantities. Furthermore, a reduction in step height leads to an attenuation of the otherwise linear relationship between step height and entropy production, which is mainly due to lower local ratio of step height and boundary layer thickness. In the case of laminar or transitional flow regions on the endwall, typical for turbine rigs with low inlet turbulence and low-pressure turbines under cruise conditions, the steps lead to immediate local flow transition and thus substantially different results.

Numerical Investigation of Three-Dimensional Separation in an Axial Flow Compressor: The Influence of Freestream Turbulence Intensity and Endwall Boundary Layer State

Regions of three-dimensional separations are an inherent flow feature of the suction surface-endwall corner in axial compressors. These corner separations can cause a significant total pressure loss and reduce the compressor's efficiency. This paper uses wall-resolved LES to investigate the loss sources in a corner separation, and examines the influence of the inflow turbulence on these sources. Different subgrid scale (SGS) models are tested and the choice of model is found to be important. The σ SGS model, which performed well, is then used to perform LES of a compressor endwall flow. The time-averaged data are in good agreement with measurements. The viscous and turbulent dissipation are used to highlight the sources of loss, with the latter being dominant. The key loss sources are seen to be the 2D laminar separation bubble and trailing edge wake, and the 3D flow region near the endwall. Increasing the freestream turbulence (FST) intensity changes the suction surface boundary layer transition mode from separation induced to bypass. However, it does not significantly alter the transition location and therefore the corner separation size. Additionally, the FST does not noticeably interact with the corner separation itself, meaning that in this case the corner separation is relatively insensitive to the FST. The endwall boundary layer state is found to be significant. A laminar endwall boundary layer separates much earlier leading to a larger passage vortex. This significantly alters the endwall flow and loss. Hence, the need for accurate boundary measurements is clear.

Effect of Blade Profile Contouring on Endwall Flow Structure in a High-Lift Low-Pressure Turbine Cascade

Previous work has shown that low-stagger contouring near the endwall of a nominally high-lift and high-stagger angle front-loaded low-pressure turbine (LPT) airfoil is successful in reducing endwall loss by limiting the development and migration of low momentum fluid associated with secondary flow structures. The design modification that leads to loss reduction in that study was determined from an intuitive approach based on the premise that reducing flow separation near the endwall will lead to reduced loss production. Those authors also relied heavily upon Reynolds-averaged Navier–Stokes (RANS) based computational tools. Due to uncertainties inherent in computational fluid dynamics (CFD) predictions, there is little confidence that the authors actually achieved true minimum loss. Despite recent advances in computing capability, turbulence modeling remains a shortcoming of modern design tools. As a contribution to overcoming this problem, this paper offers a three-dimensional (3D) view of the developing mean flow, total pressure, and turbulence fields that gave rise to the loss reduction of the airfoil mentioned above. Experiments are conducted in a linear cascade with aspect ratio of 3.5 and Re = 100,000. The results are derived from stereoscopic particle image velocimetry (PIV) and total pressure measurements inside the passage. Overall, the loss reduction correlates strongly with reduced turbulence production. The aim of this paper is to provide readers with a realistic view of mean flow and turbulence development that include all the components of the Reynolds stress tensor to assess, at least qualitatively, the validity of high fidelity computational tools used to calculate turbine flows.

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 on endwall flow and heat transfer characteristics of a gas turbine blade

This paper presents a numerical investigation on the influence of purge flow on endwall flow and heat transfer characteristics of a gas turbine blade. Upon the numerical validation with experiment data, the Reynolds-averaged Navier-Stokes equations coupled with standard k-ω turbulence model are utilized in this study. Five mass flow ratios (MFR) of the purge flow (MFR=0.5%, 0.75%, 1.0%, 1.25%, 1.5%) and four ejection angles α of the upstream slot (α=30°, 45°, 60°, 90°) are selected to investigate the effects of purge flow on endwall flow structure and their thermal behaviors. The results indicate that the purge flow provides some cooling effectiveness and increases the heat transfer on the endwall. The reduction of the ejection angle α improves the film cooling effectiveness and increases the heat transfer coefficient of the endwall. The averaged film cooling effectiveness of the endwall is reduced by 53.4% and the heat transfer coefficient at the leading edge is increased by 18.89% when the ejection angle α is increased from 30° to 90° at MFR=1.5%. Comparing to another case without purge flow, the purge flow increases the aerodynamic losses, and as the increasing of MFR, the aerodynamic losses is increased first and reduced afterwards, obtaining the largest aerodynamic losses at MFR=1.0% for the ejection angle α=30°.

Numerical investigation of secondary flows in a high-lift low pressure turbine

In turbomachines, secondary flows (or endwall flows) typically originate at the junction between endwalls and the blade surface. Within the blade passage, the strength of the secondary flows is amplified by the crossflow from the pressure to the suction surface of the blade. The enhanced mixing due to secondary flows induce additional losses into the system. This decreases the overall work output and also changes the flow incidence onto the downstream blade rows. Using a series of high-fidelity eddy resolving simulations, the current study attempts to provide an improved understanding for the complex flow physics over the endwalls of a high-lift Low Pressure Turbine (LPT) blade. The effect of three different inflow conditions has been studied. These include a laminar boundary layer (LBL), a turbulent boundary layer (TBL) and wakes with secondary flow (W&S) from an upstream blade row. For the simulations with TBL and W&S, precursor eddy resolving simulations were used to prescribe realistic inflows. The loss generation mechanisms were subsequently studied both at the endwall and the midspan, which includes evaluating the mass-averaged total pressure loss coefficient (Yp) and the loss generation rate.When compared to LBL, additional disturbances from an incoming TBL and wakes with secondary flows enhanced the mixing within the blade passage resulting in a substantial increase in the total pressure loss. Prior to flow transition, incoming wakes with secondary flows increased the local loss generation rate at both the endwall and the midspan in the front portion of the blade passage (x/Cx < 0.84). In contrast, in the aft portion of the passage (x/Cx > 0.8), the incoming wakes effectively suppressed the separation bubble at the midspan thereby decreasing the local loss generation rate. It is also demonstrated that the wakes shed from the trailing edge at the mid-span mix out rapidly when compared to the passage vortex at the endwall.

Effective end wall profiling rules for a highly loaded compressor cascade

This paper presents a numerical study of a linear compressor cascade to investigate the effective end wall profiling rules for highly loaded axial compressors. The first step in the research applies a correlation analysis for the different flow field parameters by a data mining over 600 profiling samples to quantify how variations of loss, secondary flow and passage vortex interact with each other under the influence of a profiled end wall. The result identifies the dominant role of corner separation for control of total pressure loss, providing a principle that only in the flow field with serious corner separation does the does the profiled end wall change total pressure loss, secondary flow and passage vortex in the same direction. Then in the second step, a multiobjective optimization of a profiled end wall is performed to reduce loss at design point and near stall point. The development of effective end wall profiling rules is based on the manner of secondary flow control rather than the geometry features of the end wall. Using the optimum end wall cases from the Pareto front, a quantitative tool for analyzing secondary flow control is employed. The driving force induced by a profiled end wall on different regions of end wall flow are subjected to a detailed analysis and identified for their positive/negative influences in relieving corner separation, from which the effective profiling rules are further confirmed. It is found that the profiling rules on a cascade show distinct differences at design point and near-stall point, thus loss control of different operating points is generally independent.

Reduction of end wall loss in axial compressor by using non-axisymmetric profiled end wall: A new design approach based on end wall velocity modification

Aircraft axial compressors are commonly designed with high load that generates strong secondary flow and produces high secondary loss. To deal with this problem, profiled end wall has been introduced as a promising technique to reconstruct the end wall flow during the last decade. In this study, we put forward a physics-based approach to design non-axisymmetric profiled end wall which enables researchers to generate the end wall geometry for an arbitrary compressor. The approach is less relying on design experience, much more time-saving than numerical optimization design, but comes to a considerably effective and universal geometry of profiled end wall for performance improvement. A theoretical study is applied to the end wall region at first, during which we proposed a simplified model and established a 6-equation system for the quantitative correlation between the end wall geometry and secondary flow. Then, the new physics-based profiling approach is developed from the theoretical study. By defining a more suitable distribution of end wall velocity as the design target, the corresponding end wall geometry can be solved from the equation system and is expected to modify the end wall flow as the design target. This new approach is used on a high-load compressor cascade to testify its performance, accompanied by two other profiling using the conventional approach and the numerical optimization with the same profiling limitation. A detailed analysis of the cascade, including loss generation and secondary flow, shows excellent influence of the new approach in the front end wall region while negative influence in rear region. The combined effect of the two influences finally reduces the loss by 1.65%, much higher than the conventional approach, which is reduced by only 0.22%, and reaches 73% of the numerical optimization. A further discussion on the relation between the modified end wall velocity and the positive/negative influence of PEW is given and it generally confirmed the physical rationality as well as the effectiveness of the new approach.

Role of tip injection in desensitizing the compressor to the tip clearance size

Numerical simulations are performed to investigate the role of tip injection in desensitizing a transonic compressor rotor to the detrimental effects of tip leakage flow. First, the effects of tip clearance size on the compressor stability and performance are investigated. Endwall injection of high-speed fluid upstream of the blade is then applied to the compressor with different tip clearance sizes. Results reveal that tip injection can effectively desensitize the compressor stability to the tip clearance size and improve the total pressure ratio, at the expense of some efficiency loss. It is further found that the larger the tip clearance size, the greater the stability enhancement and the smaller the efficiency penalty due to injection. The near-stall endwall flow structures for different tip clearance sizes, with and without injection, are also investigated in the current study. Results suggest that tip clearance size does not influence the compressor stalling mode, whereas endwall injection has the potential to change it (i.e., from “wall-stall” to “blade-stall” with the injection applied in the current study).

Numerical Investigation of Contrasting Flow Physics in Different Zones of a High-Lift Low-Pressure Turbine Blade

Using a range of high-fidelity large eddy simulations (LES), the contrasting flow physics on the suction surface, pressure surface, and endwalls of a low-pressure turbine (LPT) blade (T106A) was studied. The current paper attempts to provide an improved understanding of the flow physics over these three zones under the influence of different inflow boundary conditions. These include: (a) the effect of wakes at low and high turbulence intensity on the flow at midspan and (b) the impact of the state of the incoming boundary layer on endwall flow features. On the suction surface, the pressure fluctuations on the aft portion significantly reduced at high freestream turbulence (FST). The instantaneous flow features revealed that this reduction at high FST (HF) is due to the dominance of “streak-based” transition over the “Kelvin–Helmholtz” (KH) based transition. Also, the transition mechanisms observed over the turbine blade were largely similar to those on a flat plate subjected to pressure gradients. On pressure surface, elongated vortices were observed at low FST (LF). The possibility of the coexistence of both the Görtler instability and the severe straining of the wakes in the formation of these elongated vortices was suggested. While this was true for the cases under low turbulence levels, the elongated vortices vanished at higher levels of background turbulence. At endwalls, the effect of the state of the incoming boundary layer on flow features has been demonstrated. The loss cores corresponding to the passage vortex and trailing shed vortex were moved farther from the endwall with a turbulent boundary layer (TBL) when compared to an incoming laminar boundary layer (LBL). Multiple horse-shoe vortices, which constantly moved toward the leading edge due to a low-frequency unstable mechanism, were captured.

Numerical investigation of flow upstream of a circular cylinder mounted vertically over a flat plate with gap

Flow field in the plane of symmetry upstream of circular cylinder mounted perpendicular to a flat plate with varying gap between the cylinder tip and the flat plate is numerically investigated. The results were compared with an experimental study of the configuration. The existence of a multiple vortex system with characteristics similar to that of end-wall flows of surface-mounted obstacles with zero gap was captured by the numerical method. The influence of gap size and Reynolds numbers was also investigated. The results were in good comparison with the experimental data. It has been found that the vortex system structure depends on the free stream Reynolds number; and an increase in the tip gap size caused a shift in the location of vortex system and bifurcation streamline due to increased tip gap flow.

Study on grooved wall flow under rarefied condition using the Lattice Boltzmann Method

Decreasing energy consumption is a key issue for stratospheric airship to maintain its high altitude long-endurance feature. The airship envelope covered micro-grooves is considered as a feasible way to reduce resistance in order to save energy. The key to using grooves as the drag reduction technique is to evaluate its drag reduction ability under the condition of rarefied gas which is different from the original continuum atmospheric state. In this paper, the drag reduction ability of V-shape and rectangular micro-grooves under both continuum and rarefied condition is evaluated using the Lattice Boltzmann Method (LBM). This method modified for rarefied flow has been validated by the analytical solution in this paper. The local vortices are obviously observed within the grooves in the continuum and slippery flow regimes. The mechanism of drag reduction with grooves is elucidated at first and the ability of drag reduction under the rarefied condition is then evaluated by comparison of that under continuum condition. In the end, both the V-shape and rectangular grooved wall flow with different height and width are simulated. The relationships of drag reduction ability and the groove geometry are also determined and evaluated. Our results reveal that the grooves can reduce the resistance with a drag reduction rate of about 6% in the continuum. The drag reduction ability of the same grooves, however, decreases to 2.5% or less in the slippery regime. The grooves with appropriate size which make the vortices full within grooves can reduce shear force between fluid and wall. The results obtained in this paper are good references to the design of stratospheric airship envelope and the drag reduction technique.

Comparative investigation of unsteady flow interactions in endwall regions of shrouded and unshrouded turbines

The flow in turbomachinery is inherently unsteady, and the endwall losses are major sources of lost efficiency in turbine cascades. Therefore, the investigation of unsteady endwall flow interactions and the consideration of the effects into turbine design are valuable to improve the turbine performance. Comparative investigation into the physical mechanisms of unsteady endwall flow interactions of 1.5-stage shrouded and unshrouded turbines are performed by using a three-dimensional Navier–Stokes viscous solver. Emphasis is placed on how unsteady stator–rotor interactions affect turbine endwall secondary flows, and the feasibility of incorporating the unsteady endwall flow effects in turbine design is also discussed. The results show that unsteady interactions between upstream wake, tip leakage vortex/mixing zone and downstream passage vortex are the main factor affecting turbine endwall secondary flows. Unsteady interactions can reduce the radial vorticity of turbine endwall secondary flows, and the effects of these interactions on the streamwise vorticity of endwall secondary flows depend on upstream wake characteristics. The properly controlled unsteady interactions can reduce the size and intensity of endwall secondary flows, and thus improve the turbine performance. Because of the difference of turbine tip architectures, the periodic fluctuations of the flow in the shrouded turbine have smaller amplitude than those in the unshrouded turbine, and the shrouded turbine is of better unsteady performance than the unshrouded turbine.

Performance improvement of shrouded turbines with the management of casing endwall interaction flows

The flow in turbomachinery is inherently unsteady, and the understanding of the flow interaction in the casing endwall region of shrouded turbine stages is a key requirement in achieving further improvements in performances. Thus, comparative numerical investigations of casing endwall flow interactions of 1.5-stage shrouded turbines with straight and bowed vanes are performed, in an attempt to incorporate flow unsteadiness in the control of casing endwall secondary flows and losses. The investigation focuses on a comparison of the flow interaction mechanisms of casing secondary vortices with upstream straight and bowed wakes, and the influences on the performance of casing endwall region are also discussed and quantified. The results show that the bowed wake at different radial sections passes different distances to reach rotor leading edge, which affects their consequent interactions with following blades. And, this is quite similar to clocking effects but with a limited extent. The interaction with upstream bowed wakes has a beneficial effect on casing endwall flow behavior resulting in weakened secondary flows. The combined unsteady effects on turbine performance depend on the balance of unsteady endwall mixing and endwall secondary flows, and the negative impact of these effects should be alleviated to achieve enhanced performance.

Three-Dimensional Flow Field in Highly Loaded Compressor Cascade

Within the present paper, a detailed experimental investigation is presented. The influence of blade loading on the development and interaction of secondary flow structures within an annular compressor stator cascade (CSC) is examined. Experimental results at 3% chord hub clearance were obtained at four different blade loadings. Included are blade and endwall flow visualization, time resolved measurements of the static pressure on the endwall and radial–circumferential hot-wire traverse measurements within the passage as well as five-hole probe traverse measurements at the inlet and the outlet of the passage. The experimentally obtained results give detailed insight on the effect of the incidence on the development and interaction of the clearance vortex, horseshoe vortex, and the passage vortex. Furthermore, it will be shown that a vortex breakdown of the clearance vortex occurs at higher loadings.

Winglets for Improved Aerothermal Performance of High Pressure Turbines

This paper investigates the design of winglet tips for unshrouded high pressure turbine rotors considering aerodynamic and thermal performance simultaneously. A novel parameterization method has been developed to alter the tip geometry of a rotor blade. A design survey of uncooled, flat-tipped winglets is performed using Reynolds-averaged Navier–Stokes (RANS) calculations for a single rotor at engine representative operating conditions. Compared to a plain tip, large efficiency gains can be realized by employing an overhang around the full perimeter of the blade, but the overall heat load rises significantly. By employing an overhang on only the early suction surface, significant efficiency improvements can be obtained without increasing the overall heat transfer to the blade. The flow physics are explored in detail to explain the results. For a plain tip, the leakage and passage vortices interact to create a three-dimensional impingement onto the blade suction surface, causing high heat transfer. The addition of an overhang on the early suction surface displaces the tip leakage vortex away from the blade, weakening the impingement effect and reducing the heat transfer on the blade. The winglets reduce the aerodynamic losses by unloading the tip section, reducing the leakage flow rate, turning the leakage flow in a more streamwise direction, and reducing the interaction between the leakage fluid and end wall flows. Generally, these effects are most effective close to the leading edge of the tip where the leakage flow is subsonic.