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Multimedia in Learning

The Structure of Turbulent Shear Flow By Dr. A. A. Townsend. Pp. xii + 315. 8¾ in. × 5½ in. (Cambridge: At the University Press.) 40s.

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The Structure of Turbulent Shear Flow

Introduction to heat transfer

Film cooling represents one of the few game-changing technologies that has allowed the achievement of today's high firing temperature, high-efficiency gas turbine engines. Over the last 30 years, only one major advancement has been realized in this technology, that being the incorporation of exit shaping to the film holes to result in lower momentum coolant injection jets with greater surface coverage. This review examines the origins of shaped film cooling and summarizes the extant literature knowledge concerning the performance of such film holes. A catalog of the current literature data is presented, showing the basic shaping geometries, parameter ranges, and types of data obtained. Specific discussions are provided for the flow field and aerodynamic losses of shaped film hole coolant injection. The major fundamental effects due to coolant-to-gas blowing ratio, compound angle injection, cooling hole entry flow character, and mainstream turbulence intensity are each reviewed with respect to the resulting adiabatic film effectiveness and heat transfer coefficients for shaped holes. A specific example of shaped film effectiveness is provided for a production turbine inlet vane with comparison to other data. Several recent unconventional forms of film hole shaping are also presented as a look to future potential improvements

A review of shaped hole turbine film-cooling technology

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

One viable option to improve cooling methods used for gas turbine blades is to optimize the geometry of the film-cooling hole. To optimize that geometry, effects of the hole geometry on the complex jet-in-crossflow interaction need to be understood. This paper presents a comparison of detailed flowfield measurements for three different single, scaled-up hole geometries, all at a blowing ratio and density ratio of unity. The hole geometries include a round hole, a hole with a laterally expanded exit, and a hole with a forward-laterally expanded exit. In addition to the flowfield measurements for expanded cooling hole geometries being unique to the literature, the testing facility used for these measurements was also unique in that both the external mainstream Mach number (Ma{sub {infinity}} = 0.25) and internal coolant supply Mach number (Ma{sub c} = 0.3) were nearly matched. Results show that by expanding the exit of the cooling holes, both the penetration of the cooling jet and the intense shear regions are significantly reduced relative to a round hole. Although the peak turbulence level for all three hole geometries was nominally the same, the source of that turbulence was different. The peak turbulence level for both expanded holes was locatedmore » at the exit of the cooling hole resulting from the expansion angle being too large. The peak turbulence level for the round hole was located downstream of the hole exit where the velocity gradients were very large.« less

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Flowfield Measurements for Film-Cooling Holes With Expanded Exits

Recent technological advances in the field of additive manufacturing (AM), particularly with direct metal laser sintering (DMLS), have increased the potential for building gas turbine components with AM. Using DMLS for turbine components broadens the design space and allows for increasingly small and complex geometries to be fabricated with little increase in time or cost. Challenges arise when attempting to evaluate the advantages of DMLS for specific applications, particularly because of how little is known regarding the effects of surface roughness. This paper presents pressure drop and heat transfer results of flow through small, as produced channels that have been manufactured using DMLS in an effort to better understand roughness. Ten different coupons made with DMLS all having multiple rectangular channels were evaluated in this study. Measurements were collected at various flow conditions and reduced to a friction factor and a Nusselt number. Results showed significant augmentation of these parameters compared to smooth channels, particularly with the friction factor for mini-channels with small hydraulic diameters. However, augmentation of Nusselt number did not increase proportionally with the augmentation of the friction factor.Copyright © 2015 by ASME

Roughness Effects on Flow and Heat Transfer for Additively Manufactured Channels

Film cooling on airfoils is a crucial cooling method as the gas turbine industry seeks higher turbine inlet temperatures. Shaped film cooling holes are widely used in many designs given the improved performance over that of cylindrical holes. Although there have been numerous studies of shaped holes, there is no established baseline shaped hole to which new cooling hole designs can be compared. The goal of this study is to offer the community a shaped hole design, representative of proprietary and open literature holes that serves as a baseline for comparison purposes. The baseline shaped cooling hole design includes the following features: hole inclination angle of 30° with a 7° expansion in the forward and lateral directions; hole length of 6 diameters; hole exit-to-inlet area ratio of 2.5; and lateral hole spacing of 6 diameters. Adiabatic effectiveness was measured with this new shaped hole and was found to peak near a blowing ratio of 1.5 at density ratios of 1.2 and 1.5 as well as at both low and moderate freestream turbulence of 5%. Reductions in area-averaged effectiveness due to freestream turbulence at low blowing ratios were as high as 10%.Copyright © 2014 by ASME

Adiabatic Effectiveness Measurements for a Baseline Shaped Film Cooling Hole

With the desire for increased power output for a gas turbine engine comes the continual push to achieve higher turbine inlet temperatures. Higher temperatures result in large thermal and mechanical stresses particularly along the nozzle guide vane. One critical region along a vane is the leading edge-endwall juncture. Based on the assumption that the approaching flow to this juncture is similar to a two-dimensional boundary layer, previous studies have shown that a horseshoe vortex forms. This vortex forms because of a radial total pressure gradient from the approaching boundary layer.This paper documents the computational design and experimental validation of a fillet placed at the leading edge-endwall juncture of a guide vane to eliminate the horseshoe vortex. The fillet design effectively accelerated the incoming boundary layer thereby mitigating the effect of the total pressure gradient. To verify the CFD studies used to design the leading edge fillet, flow field measurements were performed in a large-scale, linear, vane cascade. The flow field measurements were performed with a laser Doppler velocimeter in four planes orientated orthogonal to the vane. Good agreement between the CFD predictions and the experimental measurements verified the effectiveness of the leading edge fillet at eliminating the horseshoe vortex. The flowfield results showed that the turbulent kinetic energy levels were significantly reduced in the endwall region because of the absence of the unsteady horseshoe vortex.Copyright © 2001 by ASME

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Karen A. Thole

Papers

Gas Turbine Film Cooling

Flowfield Measurements for Film-Cooling Holes With Expanded Exits

Roughness Effects on Flow and Heat Transfer for Additively Manufactured Channels

Adiabatic Effectiveness Measurements for a Baseline Shaped Film Cooling Hole

Computational Design and Experimental Evaluation of Using a Leading Edge Fillet on a Gas Turbine Vane