Squealer Rim Research Articles

Tip gap effects on film cooling performance of a transonic turbine blade with a distinctive coolant supply arrangement

Considered are the influences of tip gap on film cooling performance variations near to and along the tip or end of a turbine blade with a squealer rim contained within a linear cascade with five airfoils. The film coolant is supplied by two plenums which are positioned at forward and aft locations within the blade. The aft blowing ratio varies between 0.88 and 1.48, the forward blowing ratio is set to be approximately constant in the vicinity of 2.75 to 3.0, and a value of about 1.3 is utilized for the ratio of coolant density relative to the main flow density. Measured results include linearly-averaged and spatially-resolved variations of adiabatic effectiveness, which are presented both for the top portion of the pressure surface and end surface of the blade with the squealer rim. Local effectiveness variations along the top pressure side of the blade show significant effectiveness values downstream of aft pressure side hole outlet locations mostly because film lift-off downstream of these holes is less substantial than observed downstream of the forward holes. Spatially-resolved effectiveness data for the end of the blade with a squealer rim show that the largest values of effectiveness are evident near to and immediately downstream of dusting hole outlet locations. Associated linearly-averaged effectiveness data for the tip of the blade recess region with the 1.2 mm tip gap are as much as about two times higher than values for the 2.0 mm tip gap when compared at each surface location. Local and linearly-averaged surface effectiveness variations for the downstream edge region of the blade squealer tip are most strongly influenced by the film cooling holes which are aft and by the dusting hole which is positioned along the tip of the blade rim near the downstream edge. Here, effectiveness magnitudes are also often higher along the downstream edge rim region with the 1.2 mm tip gap, compared to data associated with the 2.0 mm tip gap.

Unsteady numerical study of film cooling and heat transfer on turbine blade squealer tip with coolant jet

Cooling performance of blade squealer tip with film holes in its cavity was investigated numerically under unsteady condition, and the effects of blowing ratio (BR), cavity depth, and tip clearance were comprehensively discussed. Results show that flow inside tip gap periodically fluctuates and affects the outlet pressure, BR, and jet characteristics of film hole, resulting in unsteady tip cooling effectiveness which could not be observed in steady condition. Average cooling effectiveness on 40-70% axial chord of tip changes the most over time. In comparison to steady result, unsteady time-average cooling effectiveness of tip and its region before 40% axial chord is lower while it is relatively higher in the region after 40% axial chord. Coolant from tip hole located in the flow area of rolling vortex is strongly mixed with rolling vortex, leading to high heat transfer coefficient downstream of that hole. Unsteady leakage flow and boundary layer flow cause stronger heat transfer on squealer rim than steady result. Small BR under unsteady condition could result in insufficient coolant supply and high temperature gas intrusion in tip holes located in 40-70% axial chord of tip. Increasing cavity depth brings longer normal jet distance and higher momentum of coolant, leading to coverage deterioration on floor with a drop in tip average cooling effectiveness of 11.6-34.7%. Large tip clearance causes the distribution of BR to become more uneven, and then more coolant flows out from trailing region, resulting in decrease tendency of average cooling efficiency with change in the range of −23.4%∼6.1% over time compared to base design.

Aerodynamics of Tip Leakage Flows Near Partial Squealer Rims In an Axial Flow Turbine With and Without Wall Motion

The objective of present investigation is to study aerodynamics of tip leakage flows near partial squealer rims in an axial flow turbine cascade with and without end wall in motion. The reaction turbine blade used in the present investigations has a blade length of 400 mm and a geometric deflection of 82°. The blade chord is 102 mm with an aspect ratio of 3.9. The tip gap is maintained at 2% (2.04 mm) and the squealer rims were of 1.02 mm height. A total of seven cases were investigated with different squealer geometries including case without squealer rim. The coefficient of secondary loss coefficient (CSKE) is used as objective function. Minimum CSKE indicates minimum losses. From the investigations it is concluded that full suction side squealer rim is the optimum squealer geometry among all cases studied as CSKE was found to be minimum for this geometry. The optimum case was selected for wall motion studies. Three wall motions, namely; 75%,100% and 125% of the inlet meridional velocity were maintained and investigated. The CSKE is reduced with all three end wall motions except very close to the end wall. The exit flow angle is reduced near the end wall as compared to no wall motion.

Numerical investigation of unsteady film cooling on turbine blade squealer tip with pressure side coolant

Unsteady film cooling characteristic from pressure side coolant in squealer tip in turbine stage was investigated by numerical simulation. Due to unsteady flow in turbine stage, flow characteristics and blowing ratio of coolant from various holes fluctuate and change cooling effectiveness on squealer rim, cavity floor and side wall. In unsteady simulation, deterioration of coolant coverage causes unsatisfied cooling performance on squealer tip, which results in the maximum reduction of 18.3% and 20% of cooling effectiveness on squealer rim and cavity floor at specific instants compared to steady simulation. In different moments, pressure side coolant increases average heat transfer coefficient on pressure side squealer rim by 15% ∼ 47% and forms discrete distribution of high heat transfer coefficient there. Unsteady heat transfer on squealer rim is much intense compared to steady condition. Raising blowing ratio and decreasing cavity depth promote cooling effectiveness in tip cavity. Too small blowing ratio might lead to the risk of high temperature mainstream invasion into some pressure side film holes. Cooling effectiveness in tip cavity increases with cavity depth decreasing. Due to relative motion between blade tip and shroud, scraping vortex pushed rolling vortex more closely to pressure side and thus facilitates coolant to touch cavity floor.

Augmented cooling performance in gas turbine blade tip with slot cooling

The blade tip is one of the most susceptible parts of gas turbine blade due to the thermally induced failure caused by a hot gas tip leakage flow through tip clearance. A highly sophisticated and efficient blade tip cooling scheme plays a crucial role in reliable operation of a gas turbine blade. In the present study, we adopted a special slot cooling scheme in turbine blade tip. The film cooling effectiveness (FCE) and flow characteristics of turbine blade tip were investigated according to the cooling flow through the slot gap in blade tip rim. A linear cascade was fabricated to measure and evaluate the cooling performance of the blade tip with the film cooling through a slotted rim versus the blade tip with cooling holes. The local FCE values of the blade tip floor and rim upper surface were measured using the pressure-sensitive paint (PSP) method. The Reynolds number based on the axial chord length and the inlet velocity in the experiment was set at 192,000. The tip clearance was varied from 1 to 3% of the blade span and the coolant mass flow rate ranged from 0.3 to 0.7% of the main flow. The slot cooling in the blade tip, which can be incorporated through an additive manufacturing, showed highly enhanced cooling performance ranging from 53 to 116% compared to that of the blade tip with film cooling holes for various tip gaps and coolant mass flow rates. The slot cooling in the blade tip is expected to improve the thermal durability and reliability of blade tip with better coolant coverage and additional convective cooling through the internal cooling passage in the narrow squealer rim.

Dusting Hole Film Cooling Heat Transfer on a Transonic Turbine Blade Tip

Investigated is a transonic turbine blade tip with a squealer rim and a squealer recess, with a single dusting film cooling hole contained within the leading edge region of the squealer recess. Data are provided for transonic flow conditions for a range of film cooling blowing ratios for two tip gap values, using a linear cascade, with no relative motion between the blade and the casing. Surface heat transfer characteristics are measured using the transient impulse-response measurement approach, employed with infrared thermography. Line-averaged adiabatic film cooling effectiveness values, for the 1.4 mm tip gap, are generally very small along the pressure side rim, with only small, locally increased values along the suction side rim. For the 0.8 mm tip gap, line-averaged adiabatic film cooling effectiveness values are generally somewhat higher along the pressure side rim and along the suction side rim. In general, effectiveness values for both tip gap values, for these locations, and for the recess region, increase as the blowing ratio increases. As the tip gap decreases from 1.4 mm to 0.8 mm, line-averaged adiabatic film cooling effectiveness generally increases on the rims and downstream regions of the recess, with increased magnitudes which are spread over larger spatial surface areas. For tip gaps of 0.8 mm and 1.4 mm, for regions where the line-averaged heat transfer coefficient ratio deviates significantly from 1.00, values generally decrease as the blowing ratio increases. Across every region of the blade, line-averaged heat transfer coefficient ratios either decrease or remain approximately invariant, as the tip gap value decreases from 1.4 mm to 0.8 mm.

The aerothermal performance of turbine blade squealer tip at various Reynolds numbers and Mach numbers with moving endwall

Abstract The changes of operating conditions can lead to variations in the aerodynamic conditions of the turbine blades. Then numerical studies are conducted to study the aerothermal performance of the turbine blade with a squealer tip at various Reynolds numbers (Re) and exit Mach numbers (Ma) with moving endwall. Besides, the effect of the rim height on the squealer tip is studied. The increasing of Reynolds numbers at Mach number = 0.78 enhances the heat-transfer at the tip of blade entirely, however, it has little impact on the leakage flow of the blade tip. The total heat flux increases with Mach number increasing from 0.78 to 0.96 and changes slightly with a further increasing of Mach number at a design Reynolds numbers. The height of squealer rim affects the tip leakage flow in the cavity and tip leakage vortex, and then affects the heat transfer distribution and the heat load of the blade tip.

Effects of inclined squealer rims on tip leakage vortex and loss in a transonic axial turbine

In this paper, to investigate the effects of inclined squealer rims on tip leakage flow and loss, numerical simulations have been performed on a transonic high-pressure turbine. Based on the geometry of prototype with conventional cavity tip, six modifications with different inclined rims are constructed. The effects of inclined suction side squealer rim (SSSR) on tip leakage flow (TLF) is analyzed emphatically. The results show that it is advantageous to control the TLF and loss if the internal and external surface of SSSR are inclined in a proper direction. When the internal surface of SSSR is inclined toward the cavity, the scraping vortex and its pneumatic labyrinth sealing effect are enhanced, which is beneficial to blocking the TLF and reducing the mixing loss. As the external surface of SSSR is inclined toward the passage, the pressure gradient near the tip changes and the intensity of adverse pressure gradient decreases, which is conducive to suppressing the breakdown of tip leakage vortex (TLV). In addition, the inclined external surface could make the TLV away from the blade and reduce the viscous dissipation. Regarding the aerodynamic performance of the turbine, the inclined pressure side squealer rim (PSSR) could improve the stage efficiency of the turbine by 0.08% relative to the prototype, while the proper inclined SSSR could further improve that by 0.18%.

Investigation of the film-cooling and flow characteristics of trapezoidal slot scheme on the blade tip with partial pressure side squealer in transonic flow

The present work first introduces and studies the blade tip with a discrete trapezoidal slot scheme. The film-cooling characteristics of the discrete trapezoidal slot scheme on the blade tip with the partial pressure side squealer and full suction side squealer is experimentally investigated by the Pressure Sensitive Paint technique in transonic flow. The effects of the mass-flow-ratio and tip-clearance-gap are investigated, and the experiment is carried out at four mass-flow-ratios of 0.1% + 0.05%, 0.14% + 0.07%, 0.17% + 0.085%, and 0.21% + 0.105% and three tip-clearance-gaps of 0.7%, 1.1%, and 1.5%. The cascade exit Mach numbers is 1.05. The flow characteristics of the blade tip is studied by the calculation. The results indicate that the coolant can cover the whole blade tip bottom surface at the tip clearance of 1.5% and high mass-flow-ratios; With the tip-clearance-gap increasing, the adiabatic film effectiveness near the squealer rim of the suction side is significantly improved; The adiabatic film effectiveness near the leading edge at the small tip-clearance-gap is higher than that at the large tip-clearance-gap under the mass-flow-ratio of 0.1% + 0.05% condition due to the mainstream-blockage-effect; With the tip-clearance-gap increasing, the high aerodynamic loss coefficient region is significantly enlarged. At three tip-clearance-gaps, changing the mass-flow-ratio shows a slight effect on the aerodynamic loss coefficient distribution.

Exploring Topology Optimization for High Pressure Turbine Blade Tips

Abstract This work presents the aerodynamic topology optimization of high-pressure turbine rotor blade tips. Before carrying out the topology optimization on the blade tip, some initial tip design studies were carried out. A winglet tip was optimized first and it was found that the optimum winglet design features a combination of small and largest overhangs possible that increase the aerodynamic efficiency by 1.40% compared to the datum design. Second, a radial basis function-based parametrization was set up to allow the creation of a single squealer rim on the datum blade’s tip that could move position in the circumferential direction. The optimum case proved to increase efficiency by 0.46% compared to the flat datum tip of the same tip gap. After that, a combination of winglet and topology free squealer tips was investigated via topology optimization. The winglet tip was created as in the winglet-only optimization cases and topology free squealer walls were parametrized and created using mapping of a radial basis function surface. It was shown that the radial basis function surface-based parametrization creates a very flexible design space containing novel squealer topologies. Combining both winglet and novel squealer topology optimization, better designs than the flat tip winglet can be achieved. However, because of the flexibility of the design space, gradient-based methods were found to struggle to reach an optimum solution. This was resolved by optimizing the most promising design subspace.

Measurement and determination of local film cooling performance along a transonic turbine blade tip with viscous dissipation

Presented are experimental and analytic procedures for the measurement and determination of film cooling performance parameters for a transonic flow environment along a turbine blade tip, which account for the influences of viscous dissipation. Such viscous dissipation magnitudes are vital to ascertain appropriate driving temperatures for convective heat transfer within high velocity, compressible environments. A key step in this procedure is the separation of adiabatic surface temperature magnitudes due to the thermal field from adiabatic surface temperature values associated with flow effects related to viscous dissipation. These latter adiabatic surface temperature magnitudes, from flow effects only, are determined with no film cooling, which, when considered relative to flow stagnation temperature, are directly related to local Mach number values within the tip gap flow along the blade tip surface. After the separation is implemented, adiabatic surface temperature variations from film cooling thermal effects only are provided relative to local baseline values with no film cooling. Resulting local adiabatic film cooling effectiveness and local heat transfer coefficients are subsequently determined which are based entirely upon locally measured temperatures (without the use of global temperature parameters), which is the most appropriate perspective for presentation of spatially-resolved surface heat transfer characteristics. Associated data are provided for the surface of a blade tip with a squealer rim and recess, with and without film cooling. An array of five film cooling holes, located along the upper pressure side of a two-dimensional turbine blade, are employed to provide the film coolant, with a local blowing ratio BR of 3.02. Ratios of tip gap to true blade span are 0.66%, and 1.16%. Experimental procedures employed to obtain the spatially-resolved surface data include infrared thermography, transient testing techniques, and the impulse response method for transient data analysis.

Measurement and determination of local film cooling performance along a transonic turbine blade tip with viscous dissipation

Abstract Presented are experimental and analytic procedures for the measurement and determination of film cooling performance parameters for a transonic flow environment along a turbine blade tip, which account for the influences of viscous dissipation. Such viscous dissipation magnitudes are vital to ascertain appropriate driving temperatures for convective heat transfer within high velocity, compressible environments. A key step in this procedure is the separation of adiabatic surface temperature magnitudes due to the thermal field from adiabatic surface temperature values associated with flow effects related to viscous dissipation. These latter adiabatic surface temperature magnitudes, from flow effects only, are determined with no film cooling, which, when considered relative to flow stagnation temperature, are directly related to local Mach number values within the tip gap flow along the blade tip surface. After the separation is implemented, adiabatic surface temperature variations from film cooling thermal effects only are provided relative to local baseline values with no film cooling. Resulting local adiabatic film cooling effectiveness and local heat transfer coefficients are subsequently determined which are based entirely upon locally measured temperatures (without the use of global temperature parameters), which is the most appropriate perspective for presentation of spatially-resolved surface heat transfer characteristics. Associated data are provided for the surface of a blade tip with a squealer rim and recess, with and without film cooling. An array of five film cooling holes, located along the upper pressure side of a two-dimensional turbine blade, are employed to provide the film coolant, with a local blowing ratio BR of 3.02. Ratios of tip gap to true blade span are 0.66 percent, and 1.16 percent. Experimental procedures employed to obtain the spatially-resolved surface data include infrared thermography, transient testing techniques, and the impulse response method for transient data analysis.

Transonic Turbine Blade Tip Heat Transfer with Pressure Side Film Cooling

Investigated are spatially resolved distributions of surface adiabatic film cooling effectiveness and surface heat transfer coefficients for a transonic turbine blade tip, with a unique and innovative film cooling arrangement, wherein a single row of five film cooling holes, located on the pressure side of the blade very near to the blade tip, provides thermal protection to the blade tip. The blade tip contains a squealer rim, and a squealer recess region. The present tip gap magnitude is 1.2% of the true blade span. When blade upper pressure side data are considered, heat transfer coefficient ratio variations are due, in part, to a horseshoe-shaped vortex structure, which forms around each film cooling jet, and then advects downstream. Spatially resolved distributions of surface adiabatic film cooling effectiveness show that the coolant advects to the upper edge of the blade pressure surface, along the pressure side rim, over the squealer recess region, and then sometimes collects in a substantial manner along portions of the suction side rim. As a result, local effectiveness values are generally higher on the suction side rim relative to pressure side rim values when compared at a particular value of blowing ratio.

Effects of tip gap on transonic turbine blade heat transfer characteristics with pressure side film cooling

Investigated are local, spatially-resolved and line-averaged distributions of surface adiabatic film cooling effectiveness and surface heat transfer coefficient ratios for a transonic turbine blade tip. The tip contains a squealer rim, and a single row of film cooling holes is located on the pressure-side of the blade very near to the blade tip. A two-dimensional linear cascade, with four flow passages and five complete blades, is utilized for the investigation. Associated data are provided for different film cooling flow conditions for two different tip gaps, which correspond to 1.16 and 0.66% of true blade span. Experimentally-measured, spatially-resolved surface data demonstrate that improved surface thermal protection on the squealer tip surface, from pressure side film cooling, is often present with a smaller tip gap. Such a conclusion is illustrated by heat transfer coefficient values at many surface locations which are generally lower, and by adiabatic film cooling effectiveness magnitudes which are often substantially larger for the smaller tip gap, relative to values associated with the larger tip gap. Associated effectiveness values, along the upper pressure side, vary in a significant manner as blowing ratio changes. Along the tip region of the blade, for both tip gap values, line-averaged adiabatic film cooling effectiveness generally increases with increasing blowing ratio for the pressure side rim. Effectiveness data for the suction side rim with the smaller tip gap, show more complicated variations as blowing ratio changes. Heat transfer coefficient data along the squealer tip surface and on the upper pressure side of the blade surface generally have variations which are similar to the baseline blade with no film cooling for many surface locations. The local deviations, which are present, are due to locally augmented mixing and shear, and increased turbulent transport from the advective presence of the film coolant.

Aerothermal Optimization of Turbine Squealer Tip Geometries With Arbitrary Cooling Injection

Abstract The optimization of the turbine rotor tip geometry remains a vital opportunity to create more efficient and durable engines. Balancing the aerodynamic and thermal aspects, while maintaining the mechanical integrity, is key to reshape one of the most vulnerable and life-determining parts of the entire turbine. The ever-increasing turbine gas temperatures, combined with the difficulty of cooling the tip and the aerodynamically penalizing nature of the overtip leakage vortex, make the design of the tip a truly multidisciplinary challenge. While many earlier efforts focused on uncooled geometries or studied the aerothermal impact with a fixed cooling configuration, the current paper presents the outcome of a multi-objective optimization where both the squealer rim geometry and the cooling injection pattern were allowed to vary simultaneously. This study explores a significantly wider design space, seeking a further synergistic aerothermal benefit through the combination of a quasi-fully arbitrary cooling arrangement, with mutating squealer rim structures. Specifically, the current manuscript presents the results of over 330 cooled and uncooled squealer tip geometries. The high-pressure turbine tip was automatically altered using a novel parametrization strategy adopting a maximum of 40 design variables to vary the squealer rim structures, as well as the size and location of the various cooling holes. The aerodynamic and thermal characteristics of every design were evaluated through Reynolds-averaged Navier–Stokes computational fluid dynamics (CFD) simulations with the k–ω shear-stress transport (SST) model for the turbulence closure, adopting an unstructured hexahedral grid typically containing more than 8 × 106 cells. A multi-objective differential evolution algorithm was used to obtain a Pareto front of designs which maximize the aerodynamic efficiency, while minimizing the overtip thermal loads. Eventually, a detailed investigation and robustness study was performed on a set of prime squealer geometries, to further investigate the aerodynamic flow topology and the effect of various cooling injection schemes on the heat transfer patterns.

Influence of tip clearance and cavity depth on heat transfer in a cutback squealer tip

Numerical simulations are conducted to present the aerothermal performance of a turbine blade tip with cutback squealer rim. Two different tip clearance heights (0.5%, 1.0% of the blade span) and three different cavity depths (2.0%, 3.0%, and 6.0% of the blade span) are investigated. The results show that a high heat transfer coefficient (HTC) strip on the cavity floor appears near the suction side. It extends with the increase of tip clearance height and moves towards the suction side with the increase of cavity depth. The cutback region near the trailing edge has a high HTC value due to the flush of over-tip leakage flow. High HTC region shrinks to the trailing edge with the increase of cavity depth since there is more accumulated flow in the cavity for larger cavity depth. For small tip clearance cases, high HTC distribution appears on the pressure side rim. However, high HTC distribution is observed on suction side rim for large tip clearance height. This is mainly caused by the flow separation and reattachment on the squealer rims.

Performance of Partial and Cavity Type Squealer Tip of a HP Turbine Blade in a Linear Cascade

Three-dimensional highly complex flow structure in tip gap between blade tip and casing leads to inefficient turbine performance due to aerothermal loss. Interaction between leakage vortex and secondary flow structures is the substantial source of that loss. Different types of squealer tip geometries were tried in the past, in order to improve turbine efficiency. The current research deals with comparison of partial and cavity type squealer tip concepts for higher aerothermal performance. Effects of squealer tip have been examined comprehensively for an unshrouded HP turbine blade tip geometry in a linear cascade. In the present paper, flow structure through the tip gap was comprehensively investigated by computational fluid dynamic (CFD) methods. Numerical calculations were obtained by solving three-dimensional, incompressible, steady, and turbulent form of the Reynolds-averaged Navier-Stokes (RANS) equations using a general purpose and three-dimensional viscous flow solver. The two-equation turbulence model, shear stress transport (SST), has been used. The tip profile belonging to the Pennsylvania State University Axial Flow Turbine Research Facility (AFTRF) was used to create an extruded solid model of the axial turbine blade. For identifying optimal dimensions of squealer rim in terms of squealer height and squealer width, our previous studies about aerothermal investigation of cavity type squealer tip were utilized. In order to obtain the mesh, an effective parametric generation has been utilized using a multizone structured mesh. Numerical calculations indicate that partial and cavity squealer designs can be effective to reduce the aerodynamic loss and heat transfer to the blade tip. Future efforts will include novel squealer shapes for higher aerothermal performance.

Heat/mass transfer over the cavity squealer tip equipped with a full coverage winglet in a turbine cascade: Part 1 – Data on the winglet top surface

Heat/mass transfer characteristics on the winglet top surface for the cavity squealer tip equipped with a full coverage winglet has been investigated experimentally with the variation of h/s (tip gap height-to-span ratio) by employing the naphthalene sublimation technique. For a squealer rim height-to-span ratio of hst/s=3.75% and a winglet width-to-pitch ratio of w/p=10.55%, the tip gap is changed to be h/s=0.34%, 0.68%, 1.02%, 1.36%, and 1.70%. On the winglet top surface, high local heat/mass transfer rate is found (i) at the sites of the leading edge tip gap vortices, (ii) in the flow separation/reattachment area on the pressure-side winglet, (iii) in the upstream outflow area on the suction-side winglet, (iv) in the downstream outflow area on the suction-side winglet, and (v) in the area where the cavity fluid is discharged toward the trailing edge. On the other hand, very low local heat/mass transfer rate is observed in two separate areas on the suction-side winglet upstream of the mid-chord. In the case of very low h/s, the two separate low heat/mass transfer areas cannot be identified. Instead, a new high heat/mass transfer area is found near the mid-chord on the suction-side winglet. Average heat/mass transfer rate on the winglet top surface tends to increase consistently with increasing h/s, but it is always lower than that on the tip surface of the plane tip with no winglet.

Heat/mass transfer over the cavity squealer tip equipped with a full coverage winglet in a turbine cascade: Part 2 – Data on the cavity floor

Heat/mass transfer characteristics on the floor of the cavity squealer tip equipped with a full coverage (FC) winglet has been investigated experimentally with the variation of h/s (tip gap height-to-span ratio) by employing the naphthalene sublimation technique. For a squealer rim height-to-span ratio of hst/s=3.75% and a winglet width-to-pitch ratio of w/p=10.55%, the tip gap is changed to be h/s=0.34%, 0.68%, 1.02%, 1.36%, and 1.70%. In this study, the cavity floor is divided into four specific regions based on local heat/mass transfer rate, and the effects of h/s on local heat/mass transfer rate are fully discussed. The results show that average heat/mass transfer rate on the floor of the cavity squealer tip with the FC winglet increases steeply with the increment of h/s. This average value is found lower in the case of low h/s but higher in the case of high h/s, in comparison with that for the cavity squealer tip with no winglet. However, it is always lower than that on the plane tip surface with no winglet. For the cavity squealer tip with the FC winglet, average heat/mass transfer rate on the cavity floor is lower than that on the winglet top surface regardless of h/s, but the discrepancy between them tends to decrease rapidly as h/s increases.

Investigations of film cooling and heat transfer on a turbine blade squealer tip

At three tip gap sizes, two kinds of cooling-hole distributions and three blowing ratios, heat transfer coefficient and adiabatic film cooling effectiveness on a blade squealer tip with cooling holes were numerically computed and also validated with the experimental data. The results showed that, for the squealer tip configured with only tip cooling holes, a lower heat transfer coefficient area can be achieved near the pressure side of cavity floor as increasing the blowing ratio. As tip clearance increases, the area-averaged heat transfer coefficient on the squealer tip is increased. While the film cooling effectiveness near the pressure side of the cavity floor is decreased due to less coverage of coolant. For the squealer tip configured with both tip and pressure-side holes, coolant ejection from the pressure side holes can effectively cool the squealer rim, tip trailing edge, and cavity floor near the pressure side. As the blowing ratio or tip clearance increases, heat transfer coefficients on the cavity floor near pressure side, squealer rim and tip trailing edge decrease significantly, while the adiabatic film cooling effect on these areas are enhanced, significantly.