The durability of gas turbine engines is strongly dependent on the component temperatures. For the combustor and turbine airfoils and endwalls, film cooling is used extensively to reduce component temperatures. Film cooling is a cooling method used in virtually all of today's aircraft turbine engines and in many power-generation turbine engines and yet has very difficult phenomena to predict. The interaction of jets-in-crossflow, which is representative of film cooling, results in a shear layer that leads to mixing and a decay in the cooling performance along a surface. This interaction is highly dependent on the jet-to-crossflow mass and momentum flux ratios. Film-cooling performance is difficult to predict because of the inherent complex flowfields along the airfoil component surfaces in turbine engines. Film cooling is applied to nearly all of the external surfaces associated with the airfoils that are exposed to the hot combustion gasses such as the leading edges, main bodies, blade tips, and endwalls. In a review of the literature, it was found that there are strong effects of freestream turbulence, surface curvature, and hole shape on the performance of film cooling. Film cooling is reviewed through a discussion of the analyses methodologies, a physical description, and the various influences on film-cooling performance.

Gas Turbine Film Cooling

A review of mass transfer measurements using naphthalene sublimation

Publisher Summary The decade of the 1950s saw great advances in heat-transfer measurement techniques. Optical methods became popular, along with several new heat-flux gages that are still in wide use today, as evidenced by the commercial heat-flux-gage manufacturers. In the last 10 to 20 years, a number of new techniques have been developed and applied, which have greatly increased the resolution and operating range of heat-flux instrumentation. Although the older methods are discussed as the background, the focus of this chapter is on the newer techniques and novel uses of the older techniques. In addition, an effort is made to bring together information from a variety of fields that deal with heat transfer, but that often isn't communicated. Three areas of new capability with important applications are time-resolved heat-flux measurements, simultaneous measurement of spatially distributed heat flux, and heat-flux measurement at high-temperature conditions. Examples of recent advances in these areas are discussed. Such new capabilities, when applied to real-world problems, make the field of heat transfer exciting. From these reviews, it is clear that no one gage or method is good for every application. The limitations of previous heat-flux-gage performance are highlighted and the need for better gage characteristics is expressed. Recent advances have been able to overcome some of these limitations and provide measurements with improved accuracy under conditions previously not possible. The discussion of measurement methods is organized by the first three categories of heat-transfer measurement categories, given in the chapter. Recent measurements using the methods discussed are showcased to demonstrate the advances made in measurement capability. Calibration is briefly addressed in the chapter.

Advances in Heat Flux Measurements

This paper presents in detail the transient liquid crystal technique for convective heat transfer measurements. A historical perspective on the active development of liquid crystal techniques for convective heat transfer measurement is also presented. The experimental technique involves using a thermochromic liquid crystal coating on the test surface. The colour change time of the coating at every pixel location on the heat transfer surface during a transient test is measured using an image processing system. The heat transfer coefficients are calculated from the measured time responses of these thermochromic coatings. This technique has been used for turbine blade internal coolant passage heat transfer measurements as well as turbine blade film cooling heat transfer measurements. Results can be obtained on complex geometry surfaces if visually accessible. Some heat transfer results for experiments with jet impingement, internal cooling channels with ribs, flow over simulated TBC spallation, flat plate film cooling, cylindrical leading edge and turbine blade film cooling are presented for demonstration.

https://iopscience.iop.org/article/10.1088/0957-0233/11/7/312/pdf

A transient liquid crystal thermography technique for gas turbine heat transfer measurements

A review of jet-impingement heat transfer in gas turbine systems is presented. Characteristics of the different flow regions for submerged jets--free jet, stagnation flow, and wall jet--are reviewed. Heat transfer characteristics of both single and multiple jets are discussed with consideration of the effects of important parameters relevant to gas turbine systems including curvature of surfaces, crossflow, angle of impact, and rotation.

Jet-impingement heat transfer in gas turbine systems.

Local Heat Transfer in Internally Cooled Turbine Airfoil Leading Edge Regions: Part II—Impingement Cooling With Film Coolant Extraction

An experimental technique has been developed for measurement of local convection heat transfer characteristics on rotating surfaces, utilizing thin liquid crystal surface coatings in a thermal transient test procedure. The encapsulated liquid crystal coatings used are sprayed directly on the test surface and their response is observed and processed during the transient with automated computer vision and data acquisition system

Transient Liquid Crystal Measurement of Local Heat Transfer on a Rotating Disk With Jet Impingement

Heat Transfer in the Tip Region of Grooved Turbine Blades

It unshrouded axial turbine stages, a small but generally unavoilable clearance between the blade tips and the stationary outer seal allows a clearance gap leakage flow to be driven across the blade tip by the pressure-to-suction side pressure difference In modern high-temperature machines, the turbine blade tips are often a region prone to early failure because of the presence of hot gases in the gap and the resultant added convection heating that must be counteracted by active blade cooling The blade tip region, particularly near the trailing edge, is often very difficult to cool adequately with blade internal coolant flow, and film cooling injection directly onto the blade tip region can be used in an attempt to reduce the heat transfer rates directly from the hot clearance flow to the blade tip An experimental program has been designed and conducted to model and measure the effects of film coolant injection of convection heat transfer to turbine blade tips The modeling approach follows earlier work that found the leakage flow to be mainly a pressure-driven flow related strongly to the airfoil pressure loading distribution and only weakly, if at all, to the relative motion between blade tip and shroud In the present work the clearance gap and blade tip region is thus modeled in stationary form with primary flow supplied to a narrow channel simulating the clearance gap above a plane blade tip Secondary film flow is supplied to the tip surface through a line away of discrete normal injection holes near the upstream or pressure side Both heat transfer and effectiveness are determined locally over the test surface downstream of injection through the use of this liquid crystal coatings and a computer vision system over an extensive test matrix of clearance heights, clearance flow Reynolds numbers, and film flow rates The results of the study indicate that film injection near the pressure-side corner of plane turbine blade tips can provide significant protection from convection heat transfer to the tip from the hot clearance gap leakage flow

Heat Transfer and Effectiveness on Film Cooled Turbine Blade Tip Models

Unshrouded blades of axial turbine stages move in close proximity to the stationary outer seal, or shroud, of the turbine housing. The pressure difference between the concave and convex sides of the blade drives a leakage flow through the gap between the moving blade tip and adjacent wall. This clearance leakage flow and accompanying heat-transfer are of interest because of long obvious effects on aerodynamic performance and structural durability, but understanding of its nature and influences has been elusive. Previous studies indicate that the leakage through the gap is mainly a pressure-driven flow whose magnitude is related strongly to the airfoil pressure loading distribution and only weakly, if at all, to the relative motion between blade tip and shroud. A simple flow and heat-transfer model incorporating these features can be used to estimate both tip and shroud heat transfer provided that reasonable estimates of the clearance gap size and clearance leakage flow can be made. The present work uses a numerical computation of the leakage flow to link the model to a specific turbine geometry and operating point for which a unique set of measured local tip and shroud heat fluxes are available. The resulting comparisons between the model estimates and measured heat-transfer are good. The model should thus prove useful in the understanding and interpretation of future measurements, and should additionally prove useful for providing early design estimates of the levels of tip and shroud heat transfer that need to be compensated for by active turbine cooling.Copyright © 1990 by ASME

D. E. Metzger

Papers

Local Heat Transfer in Internally Cooled Turbine Airfoil Leading Edge Regions: Part II—Impingement Cooling With Film Coolant Extraction

Transient Liquid Crystal Measurement of Local Heat Transfer on a Rotating Disk With Jet Impingement

Heat Transfer in the Tip Region of Grooved Turbine Blades

Heat Transfer and Effectiveness on Film Cooled Turbine Blade Tip Models

Turbine Tip and Shroud Heat Transfer