Turbine Blade Tip Heat Transfer Research Articles

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.

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.

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.

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.

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

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

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

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

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

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

Aerothermal Performance of a Winglet at Engine Representative Mach and Reynolds Numbers

This paper presents an experimental and numerical investigation of the aerothermal performance of an uncooled winglet tip, under transonic conditions. Spatially resolved heat transfer data, including winglet tip surface and near-tip side-walls, are obtained using the transient infrared thermography technique within the Oxford high speed linear cascade test facility. Computational fluid dynamics (CFD) predictions are also conducted using the Rolls-Royce HYDRA suite. Most of the spatial heat transfer variations on the tip surface are well-captured by the CFD solver. The transonic flow pattern and its influence on heat transfer are analyzed, which shows that the turbine blade tip heat transfer is greatly influenced by the shock wave structure inside the tip gap. The effect of the casing relative motion is also numerically investigated. The CFD results indicate that the local heat transfer distribution on the tip is affected by the relative casing motion but the tip flow choking and shock wave structure within the tip gap still exist in the aft region of the blade.

Overtip Shock Wave Structure and Its Impact on Turbine Blade Tip Heat Transfer

In this paper, the transonic flow pattern and its influence on heat transfer on a high-pressure turbine blade tip are investigated using experimental and computational methods. Spatially resolved heat transfer data are obtained at conditions representative of a single-stage high-pressure turbine blade (Mexit=1.0, Reexit=1.27×106, gap=1.5% chord) using the transient infrared thermography technique within the Oxford high speed linear cascade research facility. Computational fluid dynamics (CFD) predictions are conducted using the Rolls-Royce HYDRA/PADRAM suite. The CFD solver is able to capture most of the spatial heat flux variations and gives prediction results, which compare well with the experimental data. The results show that the majority of the blade tip experiences a supersonic flow with peak Mach number reaching 1.8. Unlike other low-speed data in the open literature, the turbine blade tip heat transfer is greatly influenced by the shock wave structure inside the tip gap. Oblique shock waves are initiated near the pressure-side edge of the tip, prior to being reflected multiple times between the casing and the tip. Supersonic flow within the tip gap is generally terminated by a normal shock near the exit of the gap. Both measured and calculated heat transfer spatial distributions illustrate very clear stripes as the signature of the multiple shock structure. Overall, the supersonic part of tip experiences noticeably lower heat transfer than that near the leading-edge where the flow inside the tip gap remains subsonic.

Gas Turbine Blade Tip Heat Transfer and Cooling: A Literature Survey

Gas turbines are widely used for aircraft propulsion, land-base power generation, and other industrial applications like trains, marines, automobiles, etc. To satisfy the fast development of advanced gas turbines, the operating temperature must be increased to improve the thermal efficiency and output work of the gas turbine engine. However, the heat transferred to the turbine blade is substantially increased as the turbine inlet temperature is continuously increased. Thus, it is very important to cool the turbine blades for a long durability and safe operation. Cooling the blade must include cooling of the key regions being exposed to the hot gas. The blade tip region is such a critical area and is indeed difficult to cool. This results from the tip clearance gap where the complex tip leakage flow occurs and thereby local high heat loads prevail. This paper presents a literature survey of blade tip leakage flow and heat transfer, as well as research of external and internal cooling technologies. The present paper does not intend to review all published results in this field, nor review all papers from the past to now. This paper is limited to a review of recently available published works by several researchers, especially from 2001 to present, concerning blade tip leakage flow associated with heat transfer, and external or/and internal tip cooling technologies.

A review of turbine blade tip heat transfer.

This paper presents a review of the publicly available knowledge base concerning turbine blade tip heat transfer, from the early fundamental research which laid the foundations of our knowledge, to current experimental and numerical studies utilizing engine-scaled blade cascades and turbine rigs. Focus is placed on high-pressure, high-temperature axial-turbine blade tips, which are prevalent in the majority of today's aircraft engines and power generating turbines. The state of our current understanding of turbine blade tip heat transfer is in the transitional phase between fundamentals supported by engine-based experience, and the ability to a priori correctly predict and efficiently design blade tips for engine service.

Darryl E. Metzger Memorial Session Paper: A Summary of the Cooled Turbine Blade Tip Heat Transfer and Film Effectiveness Investigations Performed by Dr. D. E. Metzger

The clearance gap between the stationary outer air seal and blade tips of an axial turbine allows a clearance gap leakage flow to be driven through the gap by the pressure-to-suction side pressure difference. The presence of strong secondary flows on the pressure side of the airfoil tends to deliver air from the hottest regions of the mainstream to the clearance gap. The blade tip region, particularly near the trailing edge, is very difficult to cool adequately with blade internal coolant flow. In this case, film cooling injection directly onto the blade tip region can be used in an attempt to directly reduce the heat transfer rates from the hot gases in the clearance gap to the blade tip. The present paper is intended as a memorial tribute to the late Professor Darryl E. Metzger, who made significant contributions in this particular area over the past decade. A summary of this work is made to present the results of his more recent experimental work, which was performed to investigate the effects of film coolant injection on convection heat transfer to the turbine blade tip for a variety of tip shapes and coolant injection configurations. Experiments are conducted with blade tip models that are stationary relative to the simulated outer air seal based on the result of earlier works that found the leakage flow to be mainly a pressure-driven flow, which is related strongly to the airfoil pressure loading distribution, and only weakly, if at all, to the relative motion between blade tip and shroud. Both heat transfer and film effectiveness are measured locally over the test surface using a transient thermal liquid crystal test technique with a computer vision data acquisition and reduction system for various combinations of clearance heights, clearance flow Reynolds numbers, and film flow rates with different coolant injection configurations. The present results reveal a strong dependency of film cooling performance on the choice of the coolant supply hole shapes and injection locations for a given tip geometry.