Comparison of Numerical Investigations With Measured Heat Transfer Performance of a Film Cooled Turbine Vane

Detailed surface measurements of the heat transfer coefficient and the film cooling effectiveness by application of the transient liquid crystal method were carried out on a heavily film cooled nozzle guide vane (NGV) in a linear cascade wind tunnel at the EPFL as part of the European Research Project TATEF2 (Turbine Aero-Thermal External Flows 2). The external cooling setup included a showerhead cooling scheme and suction and pressure side of the airfoil several rows of fanshaped cooling holes. By testing two different cooling flow rates at a NGV exit Reynolds number of 1.46E+06, detailed aerodynamic and heat transfer measurement data were obtained that can be used for validation of numerical codes and design tools for cooled airfoils. The data include the NGV surface static pressure distribution and wall heat transfer and film cooling effectiveness obtained by application of the transient liquid crystal technique. An engine representative density ratio between the coolant and the external hot gas flow was achieved by using CO2 as coolant gas. For the coupled simulation of internal cooling and external flow the numerical model was composed of the cooling air feeding the internal plenum, the cooling holes, and the outer external flow domain. An unstructured mesh was generated for the simulations by applying two different commercial CFD codes (Fluent and CFX). Identical boundary conditions were chosen in order to allow for a direct comparison of both codes. The computations were carried in two ways, first using a builtin transition model and second by imposing fully turbulent flow starting at the leading edge. For both codes the same built-in turbulence models were applied. The computations were set up to solve for the aerodynamic flow quantities both within and around the test model and for the thermal quantities on the vane surface, i.e. heat transfer coefficient and film cooling effectiveness. The computational results from the two codes are compared and validated against the results from the experiments. The numerical results were able to confirm a suspicion that the cross flow in the feeding plenum causes an observed non-symmetry of the measured film cooling effectiveness at the outlet of some cooling holes.

Experimental investigation of heat transfer and flow characteristics in various geometries of 2-pass internal cooling passages of gas turbine airfoils

Denis Chanteloup

An experimental investigation of the flow and heat transfer characteristics in internal coolant passages of gas turbine airfoils has been conducted. The PIV method was employed for the flow measurements. A stereoscopic PIV system was used and automated. The stereoscopic method allowed measuring the three velocity components in measurement planes. The system was capable of measuring 100 planes in each configuration. Each measurement plane was composed of 30*30 measurement points. The transient liquid crystal technique was adopted for the heat transfer measurements. Full surface Nusselt number distributions were obtained on all the five outer walls of the test models. The transient TLC technique was adapted in large-scale models of internal cooling channels. The gas temperature evolution in location and time was taken into account in the data processing. CFD simulations have been performed in the measured passage configurations. For the calculations, the unstructured flow solver FLUENT/UNS was employed. A test rig was designed and constructed for the experiments to meet the requirements of the intended investigation. For the present investigation, the test section was a large-scale model of a two-pass coolant passage with a sharp 180° bend. Four different coolant passage configurations were tested: A passage with 45° ribs (baseline configuration). A passage similar to the baseline configuration with extraction holes simulating film-cooling extraction. A passage with film-cooling extraction and a turning vane in the bend region. A passage with ribs 50% bigger than in the baseline configuration. This passage also had a turn region back wall angled at 30° to the incoming flow. The four configurations were tested at three Reynolds numbers (25,000, 50,000, 70,000 based on the hydraulic diameter). The configurations with extraction were tested at three extractions (30%, 40%, 50% of the inlet massflow). For the first time, the PIV and TLC techniques were employed simultaneously for a detailed investigation of the turbulent flow and heat transfer characteristics within internal coolant passages connected with 180° turns. The following conclusions can be drawn: 3D-streamlines were extracted from the flow measurements in the various cooling channels. The streamlines underlined the strong influence of both the ribs and the bend geometry on the creation of the secondary flow motion in the channels. The secondary flow motion dominates the heat transfer in the entire cooling channel. In the fully developed region, the rib-induced vortex governs the heat transfer distribution on the ribbed walls behind the ribs, and also on the sidewalls. In the bend region, the bend corner flow cells (with low streamwise velocity motion) deviate the bend incoming flow. They act as if the sharp bend geometry was modified. This geometry modification is very dependant on the studied configuration. The resulting heat transfer distribution is thus strongly configuration dependant. The regions of high heat transfer in the channels have been linked with high impinging flow regions. High gradients of mean impinging velocity components near the walls induce high heat transfer on the wall. In these regions, the Nusselt number showed a good correlation with the Reynolds normal stress corresponding to the impinging velocity. The configurations with extractions have the best thermal performances at all the tested Reynolds numbers. The variation of the extraction ratio does not modify substantially the thermal performances. CFD tools has benefited from the tremendous progress of computing power. 3D simulations of internal cooling configurations are now possible. CFD predictions have gained accuracy in such highly 3D flow problems. The predictions compared well with the measurements, with worst discrepancies of 25%.

EPFL2002

Detailed heat transfer distributions of narrow impingement channels for cast-in turbine airfoils

Alexandros Terzis

Gas turbine operation at elevated temperatures ensures increased thermal efficiency and useful power output. However, the industrial tendency to push further firing temperatures is limited by the capabilities of the current turbine blade cooling technologies which approach their limits. Nowadays, the capabilities of the foundry industry to produce high integrity and dimensionally accurate castings allow the production of integrally cast turbine airfoils with the so-called double-wall cooling technology. Cast-in turbine airfoils include a more distributed form of cooling, where the coolant can be injected in the form of impingement jets within the wall rather than the hollow of the airfoil increasing dramatically the heat exchanged capabilities. Narrow impingement cooling passages can be therefore generated close to the external hot gas flow. In this study detailed heat transfer distributions for all internal surfaces of narrow impingement channels are evaluated. The test models consist of a single row of five impingement jets investigated over a range of engine representative Reynolds numbers. Effects of jet-to-jet spacing, channel width and height and impingement jet pattern are independently examined composing a test matrix of 22 large-scale geometries. For the evaluation of the heat transfer coefficients, the transient liquid crystal technique is used considering also thermocouple thermal inertia effects. Given that the use of different impingement geometries could play an important role on the distribution of convection coefficients, and hence, on the homogeneity of blade metal temperatures, the experimental results of this study can be used along with the external heat load for the development of thermal design models for integrally cast turbine vanes and blades.

2014

Experimental investigations of showerhead film cooling on the leading edge of a turbine blade

Guillaume Wagner

The turbine components of a gas turbine need extensive cooling to withstand the high heat loads generated by the flow of the hot combustion gases. The leading edge of a turbine blade is one of the areas that faces the hottest gas flow conditions and is thus one of the most critical areas to be cooled. The stagnation region of turbine airfoils is generally cooled with showerhead cooling scheme, which consist in injecting coolant onto the stagnation region through closely spaced rows of film cooling holes. This provides convective cooling inside the holes and produces a layer of coolant film on the external surface. The objective of this study is to investigate experimentally the influence of mainstream incidence on showerhead film cooling. The film cooling effectiveness and heat transfer coefficient were thus measured on a series of symmetric blunt bodies with different showerhead configurations simulating the leading edge of turbine airfoils. Tests have been performed at 0° on-design and 5° off-design mainstream incidences at an approach Mach number of 0.26. Besides the mainstream incidence angle, the investigated parameters were the blowing ratio, the wedge angle, the row location, the spanwise pitch between the holes and the hole shape. In addition to the film cooling measurements, baseline heat transfer tests were performed on test models without film cooling holes in order to evaluate the film cooling performances. To characterize the mainflow conditions and the velocity profile on the test models, aerodynamic measurements have also been carried out. The heat transfer measurements have been performed using the transient liquid crystal technique. The transient experiments were generated by a pre-conditioning system and a rapid insertion mechanism. The data analysis was adapted to measurements on highly curved surfaces using the approximate analytical solution of the heat conduction into a solid cylinder with convective boundary conditions on the surface. Time-varying adiabatic wall temperatures were accounted for using Duhamel's superposition theorem. The first tested showerhead geometry was a triangular blunt body representative of a turbine rotor blade showerhead with a circular leading edge, a 21.1° wedge angle and two film cooling rows. Because of jet lift-off, the film cooling performances of this two-row configuration were relatively low and were almost unaffected by the increase of blowing ratio from 1.5 to 3.9. The heat transfer coefficient was found to be greatly increased at the row location in comparison to the case without film cooling. At 5° off-design incidence angle, the stagnation line moves passed the pressure side row. The two rows were then located on the same side of the stagnation line and were both blowing towards the suction side. On the suction side the film cooling protection was thus very good due to film accumulation but the pressure side was left almost unprotected resulting in highly asymmetric film cooling performances. Two-dimensional FEM calculations of the metal temperature in this second geometry have been performed at engine representative conditions for design and off design mainstream incidences. It was found that the contribution of the film cooling to cool the leading edge region was rather small in comparison to the contribution of the internal convective cooling. Showerhead film cooling was however required to effectively cool the stagnation region and to obtain a uniform surface temperature. The second tested showerhead geometry was similar to the first geometry but with an increased wedge angle of 55° and cooling rows located further downstream. The film cooling performances were found to decrease when the blowing ratio was increased from 1.2 to 3.0 because of jet lift-off. At middle blowing ratio (2.0), the film cooling effectiveness had similar values as for the first geometry. However, the heat transfer increase due to the presence of the film cooling jets was much higher for the second geometry because of higher local mainstream velocities at the injection locations. The resulting film cooling performances were thus lower for the second geometry than for the first geometry. Nevertheless, the larger wedge angle and the cooling rows located further downstream assured that the stagnation line stayed in-between the two cooling rows even at 5° off-design incidence. The second geometry had thus the advantage to be much less sensitive to incidence angle and, unlike the first geometry, provided very similar film cooling performances at 5° and at 0° incidence. The use of shaped holes was found to greatly improve the film cooling performances of the second geometry due to a better lateral coverage and a reduced tendency of jet lift-off.