Blade Tip Heat Transfer Research Articles

Blade tip leakage flow and heat transfer characteristics over a gas turbine blade at subsonic and transonic exit conditions

The present study numerically investigates the blade tip leakage flow (TLF) over a gas turbine blade cascade at low speed ( M ex = 0.1) and transonic ( M ex = 0.92) blade exit Mach numbers, with and without shroud motion. Various blade tip flow phenomena like flow separation and reattachment, flow choking, shock-boundary layer interaction (SBLI), and tip leakage vortex (TLV) are studied, and their influence on the heat transfer coefficient (HTC) distributions over the blade tip and near tip blade suction surface is investigated. The study has found some new zones of heat transfer due to the interaction of horseshoe vortices (HSV) with the blade tip. In addition to the primary TLV, a secondary TLV is formed, which increases near tip blade suction surface heat transfer. The flow transitions to supersonic speed in the aft portion of the blade tip for the transonic case resulting in complex heat transfer distribution due to SBLI, compared to near-uniform distribution for the low speed case. This supersonic flow results in a smaller flow separation zone near the aft portion of the pressure side tip edge for the transonic case compared to the low speed case. The qualitative nature of HTC near the leading edge of the blade tip resembles for low speed and transonic cases. However, near the trailing edge of the blade tip, the qualitative nature of HTC for the transonic case shows a substantial difference due to supersonic flow compared to the low speed case.

Comprehensive Aerothermal Investigation of Turbine Blade Multicavity Squealer Tip Using a Novel Methodology With Uncertainty Quantification

Abstract Ascertaining the uncertainty in the aerothermal performance of blade tips is crucial, as it represents the most delicate component of modern gas turbines. In this research article, a novel and efficient approach is proposed for quantifying uncertainties in aerothermal performance using a combination of universal kriging, polynomial chaos expansions, and Smolyak sparse grid technology. This method was applied to investigate the aerothermal performance of a high-pressure gas turbine rotor blade tip with high-dimensional robustness. The outcomes of the uncertainty quantification calculation reveal that the downstream total pressure loss coefficient and leakage flowrate increase under normal-speed (subsonic) and high-speed (transonic) conditions. The key uncertainty input that affects the aerodynamic performance of normal-speed and high-speed squealer tip is inlet total pressure fluctuation, with a variance index on the leakage flowrate of normal-speed and high-speed squealer tip of up to 73.92 and 83.85%, respectively. The study suggests that it is more important to control the operating conditions fluctuation than the cavity depth machining accuracy for aerodynamic performance robustness, which applies to both normal-speed and high-speed squealer tips. In line with the aerodynamic performance, the heat flux of normal-speed and high-speed squealer tip increases during operation. Notably, the sensitivity of high-speed squealer tip aerodynamic performance to operating condition fluctuations increases compared to the normal-speed squealer tip, necessitating active intervention for fluctuations in operating conditions at a higher cost for the high-speed squealer tip. The sensitivity analysis results indicate that the inlet total temperature fluctuation is the key parameter that controls the normal-speed and high-speed squealer tip heat flux uncertainty. Finally, it is worth noting that while the use of ribs can effectively enhance the robustness of blade tip heat transfer performance, the heat flux near the root of the ribs fluctuates significantly, which may further increase the thermal fatigue tendency in this region during actual operation.

Development of a Large-Scale High-Speed Linear Cascade Rig for Turbine Blade Tip Heat Transfer Study

The high-speed Over-Tip-Leakage (OTL) flow has a significant impact on the aerodynamic performance of the High-Pressure Turbine (HPT) passage and generates high thermal load for the blade tip. Different tip sealing and cooling design strategies are applied to reduce the OTL loss and help turbine survive in a high temperature environment. High-speed linear cascade experimental rigs play an important role in understanding the flow physics and evaluating their performance. Multiple blades and passages are often required to maintain a reasonable flow periodicity. To match the engine representative Reynolds number and Mach number, a high-speed multi-passage cascade design inevitably demands more compressed air supply. A very large amount of heating power is also required if the engine condition wall-to-gas temperature ratio needs to be matched. In this study, a simplified 2-Passage linear cascade rig for high-speed tip heat transfer research was developed. Both the design method and the rig performance are presented. Different from existing design method to match two-dimensional blade loading, this study shows there are other design flexibilities, such as assist blade tip gap, tailboard adjustment, and profiling adjustment, to match the periodic three-dimensional OTL flow structures. The design method was validated by experimental effort. High resolution tip heat transfer coefficient distribution at stationary and rotating conditions (Rotating Mach number = 0.35) are reported. The enlarged test model can offer much more improved resolution of optical measurement near the tip region.

Influence of Gas-to-Wall Temperature Ratio on the Leakage Flow and Cooling Performance of a Turbine Squealer Tip

In high-pressure turbines, there is a large difference in temperature between the mainstream and the turbine blade surface. Most of the turbine blade tip heat transfer studies were conducted under the assumption that the Over-Tip-Leakage (OTL) flow field is independent of the wall thermal condition. Recent numerical and experimental studies have revealed that the two-way coupling effect between aerodynamics and heat transfer should not be neglected. The heat transfer coefficient obtained by the conventional method is not able to match the realistic engine condition accurately. This study investigates the impact of the wall thermal boundary condition on the tip cooling performance of squealer turbine blades. The RANS CFD result was validated against experimental tip heat transfer data obtained from a high-speed test rig with the effect of high-speed relative casing motion. The aerothermal performance for both uncooled and cooled squealer tips was studied at two different gas-to-wall temperature ratios, 1.7 and 1.1; the reference temperature is 204 K. It was found that the location and strength of cavity vortices varied with different wall thermal boundary conditions, leading to different signatures in tip heat transfer and cooling performance. It is recommended that the experimental heat transfer data and film cooling effectiveness obtained at the near-adiabatic wall boundary condition should be corrected before their application to the tip cooling design process. It would be more reliable to match the wall-to-gas temperature ratio during the tip experimental study.

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.

Effect of cavity depth on thermal performance of a cooled blade tip under rotation

In this paper, the effect of cavity geometry of a squealer blade tip on its thermal performance is investigated. Numerical analyses were performed for three depth-to-width ratios for a cooled squealer tip. Also investigated is the effect of cooling introduction into the flow tip and the effect of rotation via the comparison of stationary and rotational domains where the latter was simulated by rotating the blade. The rotational speed and boundary conditions used in the numerical study are chosen to represent the turbine operating environment. The blade tip used had a cooling configuration consisting of seven holes of the same size located on the tip near the blade pressure side. The study presents a discussion on the flow physics within the tip gap and heat transfer on the blade tip. Film-cooling performance of the considered blade tip geometries is demonstrated using distributions of film-cooling effectiveness. It is observed that both the rotation and the addition of cooling at the tip significantly alter the tip gap flow, which is found to have a strong effect on the blade tip heat transfer. The addition of cooling increased the heat transfer coefficient on the blade tip, while an increase in the rim height resulted in higher film effectiveness since the amount of leakage flow entering the cavity floor was reduced.

Experimental Evidence of Temperature Ratio Effect on Turbine Blade Tip Heat Transfer

Determination of a scalable Nusselt number (based on “adiabatic heat transfer coefficient”) has been the primary objective of the most existing heat transfer experimental studies. Based on the assumption that the wall thermal boundary conditions do not affect the flow field, the thermal measurements were mostly carried out at near adiabatic condition without matching the engine realistic wall-to-gas temperature ratio (TR). Recent numerical studies raised a question on the validity of this conventional practice in some applications, especially for turbine blade. Due to the relatively low thermal inertia of the over-tip-leakage (OTL) flow within the thin clearance, the fluids' transport properties vary greatly with different wall thermal boundary conditions and the two-way coupling between OTL aerodynamics and heat transfer cannot be neglected. The issue could become more severe when the gas turbine manufacturers are making effort to achieve much tighter clearance. However, there has been no experimental evidence to back up these numerical findings. In this study, transient thermal measurements were conducted in a high-temperature linear cascade rig for a range of tip clearance ratio (G/S) (0.3%, 0.4%, 0.6%, and 1%). Surface temperature history was captured by infrared thermography at a range of wall-to-gas TRs. Heat transfer coefficient (HTC) distributions were obtained based on a conventional data processing technique. The profound influence of tip surface thermal boundary condition on heat transfer and OTL flow was revealed by the first-of-its-kind experimental data obtained in the present experimental study.

Thermal performance of blade tip and casing coolant injection on a turbine blade with cavity and winglet-cavity tip

The heat transfer and film cooling effects of cavity tip and winglet-cavity tip have been numerically investigated with blade tip and casing coolant injection with a turbine cascade exit Ma number 0.96. The film-cooling effectiveness, heat transfer coefficient and net heat flux reduction ratio of blade tip and casing at relative tip clearance of 0.8% blade span are compared with blade tip cooling, casing purge cooling and the combination of these two cooling types. The validations and verifications of the numerical methods are performed by contrasting with the experimental results. The results show that the winglet-cavity tip alters the tip leakage flow and affects the heat transfer of blade tip compared to the cavity tip. For these two blade tip types, the blade tip cooling reduces the heat load of blade tip effectively, while it has little cooling effects on the casing. The casing purge cooling obtains high film-cooling effectiveness of the blade tip and the casing at the same time. However, it has no influence on the heat transfer of the pressure side near the tip region and the pressure-side rim. Adopting the blade tip and casing coolant injection simultaneously can further reduce the heat load and improve the film-cooling effectiveness of the blade tip and casing.

Effects of unsteady wakes on heat transfer of blade tip and shroud

An experimental study has been conducted to investigate the heat-transfer characteristics of blade tips and shrouds with and without unsteady wakes. Depending on the presence of unsteady wakes, the local heat/mass-transfer coefficients of the tip and shroud were measured using the naphthalene sublimation method. Wakes from unsteady blades were modeled as wakes generated from moving cylindrical rod bundles. Test conditions were set to a Reynolds number of 100,000, based on an inlet velocity of 11.4 m/s and the axial chord length. The Strouhal number was varied from 0 to 0.22. For the case without unsteady wakes (St = 0), high heat/mass-transfer coefficients appeared in regions where various flow patterns, such as flow reattachment, swirling flow, and vortices, occurred. Unsteady wakes (St = 0.22) made high turbulence intensity of the tip leakage flow, and flow patterns ranging from flow reattachment to tip leakage vortex in the tip and shroud were changed and dispersed in the presence of unsteady wakes, thus changing the heat/mass-transfer distributions in these areas. Due to the high thermal load on the tip and shroud under unsteady wake conditions, more detailed cooling designs must be considered, ranging from film cooling holes to the installation of additional structures on the blade tip, especially in the leading edge region and mid-region I of the tip, and the mid-region I and trailing edge region of the shroud.

가스터빈 블레이드 팁의 열전달과 유동 특성에 대한 수치적 해석

본 연구에서는 전산해석을 통해 냉각 터빈 블레이드 팁 간극의 유동 및 열전달 특성을 연구하였다. 1단 고압터빈 노즐 출구에서 획득한 속도, 온도 프로파일을 로터의 입구에 적용하여 로터 도메인을 대상으로 전산해석을 하였다. 스퀼러 팁이 적용된 블레이드의 팁 간극을 스팬의 1%부터 2.5%로 조절하여 팁 간극의 공력 손실, 열전달 계수와 막냉각 효율의 영향을 고찰했다. 팁 간극이 커질수록 출구에서 공력 손실과 블레이드 끝단 표면에서 열전달 계수는 증가하였다. 특히 팁 간극이 스팬의 2%일 때 평균 열전달 계수가 급격히 증가하였다. 팁 영역의 막냉각 효율은 팁 간극이 작을수록 높았고, 캐비티 내부 냉각 홀 근처의 막냉각 효율이 높았다. The heat transfer and flow characteristics of gas turbine blade tip were investigated in this paper by using the conjugate heat transfer analysis. The rotor inlet boundary condition profile which was taken from the first stage nozzle outlet was used to analyse. The profile contained the velocity and temperature information. This study presents the influence of tip clearance about aerodynamic loss, heat transfer coefficient and film cooling effectiveness with the squealer tip designed blade model which tip clearance variation range from 1% to 2.5% of span. Results showed that the aerodynamic loss and the heat transfer coefficient were increased when the tip clearance was increased. Especially when the tip clearance was 2% of the span, the average heat transfer coefficient on the tip region was increased obviously. The film cooling effectiveness of tip region was increasing with decreasing of the tip clearance. There was high film cooling effectiveness at cavity and near tip hole region.

Aerothermal characteristics of a transonic tip flow in a turbine cascade with tip clearance variations

A significant portion of flows over a modern high-pressure turbine blade tip is transonic, and the transonic tip leakage flows lead to significant aerodynamic losses and high heat loads onto the blade tips. This paper aims to develop a deeper understanding of the transonic tip leakage flow physics and its influence on the loss mechanism and blade-tip heat transfer. Three-dimensional (3D) Reynolds-averaged Navier-Stokes (RANS) calculations were performed using the ANSYS CFX 14.5 numerical prediction code, adopting the SST k-ω turbulence model to investigate the sensitivity of aerothermal performance of transonic tip flows to tip clearances in a RT27a turbine cascade. The transonic tip leakage flow pattern within the tip gap, the flowfield downstream of the cascade, and the blade tip heat transfer distribution are studied. The numerical results give a reasonable agreement with the experimental data. The tip aerodynamics and surface heat transfer variations with tip clearances are opposite between leading and trailing edge regions of blade tips, and tip clearance effects on the former are relatively small. As the tip clearance increases, the shock wave reflections are delayed but more evident, and it therefore leads to reduced leakage massflow density and decreased heat loads on the rear part of blade tips. Despite this, since the leakage flow near the leading edge of blade tips remains subsonic resulting in increased leakage mass flowrate and tip heat transfer, the leakage losses and overtip heat loads are increased with the increasing tip clearance.

Ramp Heating in High-Speed Transient Thermal Measurement with Reduced Uncertainty

The uncertainty associated with the convective heat transfer coefficient obtained in transient thermal measurement is often high, especially in high-speed flow. The present study demonstrates that the experimental accuracy could be much improved by an actively controlled ramp heating instead of the conventional step-heating approach. A general design guideline for the proposed ramp-heating method is derived theoretically and further demonstrated by simulation cases. This paper also presents a detailed experimental study for transonic turbine blade-tip heat transfer. A repeatable, high-resolution tip heat transfer coefficient contour is obtained through transient infrared measurement with the proposed ramp-heating method. Detailed uncertainty analysis shows that the resulting heat transfer coefficient uncertainty level is much lower than the experimental data currently available in the open literature. The ramp-heating approach is especially recommended to the high-speed heat transfer experimental research community to improve the accuracy of the transient thermal measurement technique.

An experimental and computational study of tip clearance effects on a transonic turbine stage

This paper describes an experimental and computational investigation into the influence of tip clearance on the blade tip heat load of a high-pressure (HP) turbine stage. Experiments were performed in the Oxford Rotor facility which is a 112 stage, shroudless, transonic, high pressure turbine. The experiments were conducted at an engine representative Mach number and Reynolds number. Rotating frame instrumentation was used to capture both aerodynamic and heat flux data within the rotor blade row. Two rotor blade tip clearances were tested (1.5% and 1.0% of blade span). The experiments were compared with computational fluid dynamics (CFD) predictions made using a steady Reynolds-averaged Navier–Stokes (RANS) solver. The experiments and computational predictions were in good agreement. The blade tip heat transfer was observed to increase with reduced tip gap in both the CFD and the experiment. The augmentation of tip heat load at smaller clearances was found to be due to the ingestion of high relative total temperature fluid near the casing, generated through casing shear.

Analytical-Solution Based Corner Correction for Transient Thermal Measurement

The one-dimensional (1D) conduction analytical approaches for a semi-infinite domain, widely adopted in the data processing of transient thermal experiments, can lead to large errors, especially near a corner of solid domain. The problems could be addressed by adopting 2D/3D numerical solutions (finite element analysis (FEA) or computational fluid dynamics (CFD)) of the solid field. In addition to needing the access to a conduction solver and extra computing effort, the numerical field solution based processing methods often require extra experimental efforts to obtain full thermal boundary conditions around corners. On a more fundamental note, it would be highly preferable that the experimental data processing is completely free of any numerical solutions and associated discretization errors, not least because it is often the case that the main purposes of many experimental measurements are exactly to validate the numerical solution methods themselves. In the present work, an analytical-solution based method is developed to enable the correction of the 2D conduction errors in a corner region without using any conduction solvers. The new approach is based on the recognition that a temperature time trace in a 2D corner situation is the result of the accumulated heat conductions in both the normal and lateral directions. An equivalent semi-infinite 1D conduction temperature trace for a correct heat transfer coefficient (HTC) can then be generated by reconstructing and removing the lateral conduction component at each time step. It is demonstrated that this simple correction technique enables the use of the standard 1D conduction analysis to get the correct HTC completely analytically without any aid of CFD or FEA solutions. In addition to a transient infrared (IR) thermal measurement case, two numerical test cases of practical interest with turbine blade tip heat transfer and film cooling are used for validation and demonstration. It has been consistently shown that the errors of the conventional 1D conduction analysis in the near corner regions can be greatly reduced by the new corner correction method.

Aerothermal Performance of Shroudless Turbine Blade Tips with Relative Casing Movement Effects

Qualitatively different heat transfer characteristics between a transonic blade tip and a subsonic one have recently been discovered. High-resolution experimental data can be acquired for blade-tip heat transfer research using a high-speed linear cascade. A combined experimental and computational fluid dynamics study on several high-pressure turbine blade-tip configurations is conducted to understand the flow physics in both stationary and moving casing setups. Extensive tests measuring aerodynamic loss and heat transfer have been performed on a stationary squealer tip at engine-representative aerodynamic conditions. A systematic validation of the computational fluid dynamics solver (Rolls–Royce, plc. HYDRA code) is introduced, showing good agreement with the experimental data obtained. Relative casing movement effects are then evaluated for two tip configurations at three different tip gaps. The moving casing is shown to affect the aerothermal performance considerably; the trends are consistently captured for the large and medium tip gaps, both in the stationary and moving casing instances. Presented results confirm that, even with a moving casing, the blade tips remain transonic. It is also shown that the heat transfer is not only dependent on the tip gap size but also the tip-geometry configuration. The squealer cavity is subsonic regardless of the tip gap size, whereas the local flow state over a flat tip is much more responsive to tip gap size.

Investigation on heat transfer of a rotor blade tip with various film cooling holes arrangements and groove depths

In this article, the effects of film cooling holes arrangement and groove depth on the heat transfer and film cooling performance of blade tip are investigated. For this numerical research, a high-pressure turbine blade with the squealer tip is applied, and the tip clearance is given to be 0.8 mm (1% of the blade span). Simultaneously, a typical tip cooling technology of film holes in the groove floor is used. The number of film holes is fixed at 10, and two kinds of holes arrangements are considered: (1) equidistance distribution and (2) dense distribution near the leading edge. Three groove depths are studied with the values of 1.5, 2.0, and 2.5 mm (1.875%, 2.5%, and 3.125% of the blade span, respectively). The results show that the area-averaged film cooling effectiveness is higher when the holes distribute densely near the leading edge, and the cooling effect of the groove depth with 2.0 mm is obviously high compared with the other two depths.

Tip Leakage Flow and Heat Transfer on Turbine Blade Tip and Casing, Part 1: Effect of Tip Clearance Height and Rotation Speed

Steady simulations were performed to investigate tip leakage flow and heat transfer characteristics on the rotor blade tip and casing in a single-stage gas turbine engine. A typical high-pressure gas turbine stage was modeled with a pressure ratio of 3.2. The predicted isentropic Mach number and adiabatic wall temperature on the casing showed good agreement with available experimental data under similar operating condition. The present numerical study focuses extensively on the effects of tip clearance heights and rotor rotational speeds on the blade tip and casing heat transfer characteristics. It was observed that the tip leakage flow structure is highly dependent on the height of the tip gap and the speed of the rotor. In all cases, the tip leakage flow was seen to separate and recirculate just around the corner of the pressure side of the blade tip. This region of re-circulating flow enlarges with increasing clearance heights. The separated leakage flow reattaches afterwards on the tip surface. Leakage flow reattachment was shown to enhance surface heat transfer at the tip. The interaction between tip leakage flow and secondary flows that is induced by the relative casing motion is found to significantly influence the blade tip and casing heat transfer distribution. A region of critical heat transfer exists on the casing near the blade tip leading edge and along the pressure-side edge for all the clearance heights that were investigated. At high rotation speed, the region of critical heat transfer tends to move towards the trailing edge due to the change in inflow angle.

Flowfield and heat transfer past an unshrouded gas turbine blade tip with different shapes

This paper describes the numerical investigations of flow and heat transfer in an unshrouded turbine rotor blade of a heavy duty gas turbine with four tip configurations. By comparing the calculated contours of heat transfer coefficients on the flat tip of the HP turbine rotor blade in the GE-E 3 aircraft engine with the corresponding experimental data, the κ-ω turbulence model was chosen for the present numerical simulations. The inlet and outlet boundary conditions for the turbine rotor blade are specified as the real gas turbine, which were obtained from the 3D full stage simulations. The rotor blade and the hub endwall are rotary and the casing is stationary. The influences of tip configurations on the tip leakage flow and blade tip heat transfer were discussed. It’s showed that the different tip configurations changed the leakage flow patterns and the pressure distributions on the suction surface near the blade tip. Compared with the flat tip, the total pressure loss caused by the leakage flow was decreased for the full squealer tip and pressure side squealer tip, while increased for the suction side squealer tip. The suction side squealer tip results in the lowest averaged heat transfer coefficient on the blade tip compared to the other tip configurations.

Unsteady Effects on Transonic Turbine Blade-Tip Heat Transfer

In a gas turbine engine the blade tips of the high-pressure turbine are exposed to high levels of convective heat transfer, because of the so-called tip-leakage phenomenon. The blade-lift distribution is known to control the flow distribution in the blade–tip gap. However, the interaction between upstream nozzle guide vanes and the rotor blades produces a time-varying flow field that induces varying flow conditions around the blade and within the tip gap. Extensive measurements of the unsteady blade-tip heat transfer have been made in an engine representative transonic turbine. These include measurements along the mean camber line of the blade tip, which have revealed significant variation in both time-mean and time-varying heat flux. The influences of potential interaction and the vane trailing edge have been observed. Numerical calculations of the turbine stage using a Reynolds-averaged-Navier-Stokes-based computational fluid dynamics code have also been conducted. In combination with the experimental results, these have enabled the time-varying flow field to be probed in the blade-relative frame of reference. This has allowed a deeper analysis of the unsteady heat-transfer data, and the quantification of the impact of vane potential field and vane trailing edge interaction on the tip-region flow and heat transfer. In particular, the separate effects of time-varying flow temperature and heat-transfer coefficient have been established.

Effects of Tip Shape on the Gas Turbine Blade Tip Heat Transfer

Effects of Tip Shape on the Gas Turbine Blade Tip Heat Transfer