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Sequentially stacked impingement channels allow for a more effective use of the cooling fluid and can limit the heat transfer reduction due to the crossflow when the number of jets increases. In this paper, sequential impingement channels are experimentally investigated using the transient liquid crystal technique. Twelve configurations are tested: the baseline consists in a double rectangular channel with 5 jets per channel; other configurations feature different number of jets, variations of the jet-to-jet spacing and of the transition zone length, reduction of the cross-section in the transition zone and addition of a bypass between the channels. Results show that heat transfer coefficients (HTC) in the 2nd channel are similar to the 1st, and that the cross-section reduction in the transition zone allows for an increase of the HTC in this critical zone with only a moderate increase of pressure losses.

High pressure turbines are nowadays designed to a point where most design enhancements only yield marginal efficiency improvements. This challenges research facilities to reliably resolve ever smaller differences in efficiency caused by individual design changes. In recent years, immense efforts towards such highly accurate delta-efficiency measurements have been undertaken at the Large Scale Turbine Rig (LSTR). This paper comprises an overview of the applied methodology and the achievements on the basis of various validation cases. By thoroughly controlling the operation point and accounting for all variables affecting the efficiency η, the rig can resolve efficiency-differences Δη of ±0.1 % for a single day measurement. Four benchmark cases are investigated to validate the rig’s capabilities. First, the influence of tip clearance is investigated for a squealer-type geometry for swirling inflow. It is found that for an increase in tip clearance of 1 %, η is decreased by 2.68 %. Then, it is shown that a winglet-type tip geometry may improve the efficiency by Δη 0.33% in comparison to the squealer tip. Third, it is shown that these trends are similar for plain inflow, however swirl decreases efficiency by up to 1.25 % in comparison to plain inflow. Finally, the clocking-position of the combustor-module relative to the nozzle guide vanes is varied leading to efficiency differences of up to 0.52 %.

The trailing edge region of gas turbine blades is generally subjected to extremely high external heat loads due to the combined effects of high mach numbers and gas temperatures. In order to maintain the metal temperatures of these trailing edges to a level, which fulfils both the part mechanical integrity and turbine performance, highly efficient and reliable cooling of the trailing edges is required without increasing the coolant consumption. In this paper, the heat transfer and pressure drop characteristic of three different turbulator designs in a very high aspect ratio passage have been investigated. The turbulator designs included angled and tapered ribs, broken discrete ribs and V-shaped small chevrons ribs. The heat transfer and pressure drop characteristics of all the turbulator configurations was initially investigated via numerical predictions and subsequently in a scaled experimental perspex model. The experimental study was conducted for a range of operational Reynolds numbers and the TLC (thermochromic liquid crystal) method was used to measure the detailed heat transfer coefficients on all surfaces of the passage. Pressure taps were located at several locations within the perspex model and both the local and average heat transfer coefficients and pressure loss coefficients were determined. The measured and predicted results show, that for all cases investigated, the local internal heat transfer coefficient, which is driven by the highly three dimensional passage flows, is highly non-uniformly within the passage. The highest overall average heat transfer was obtained for the angled and tapered turbulator. Although the average heat transfer coefficient of the discrete broken turbulator and the small chevron turbulator were slightly lower than the baseline case, they had much higher pressure losses. In terms of the overall non-dimensional performance index, which incorporates both the heat transfer and the pressure drop, it was found that the angled and tapered turbulator gave the best overall performance.

The leading edge regions of first stages blades and vanes of heavy-duty gas turbines are subjected to high thermal loads. Efficient cooling allows the reduction of the coolant mass flow required to drive the metal temperatures to a range satisfying mechanical integrity requirements. This paper investigates the heat transfer and pressure loss behavior for the internal cooling channel of a leading edge of a gas turbine blade. The geometrical profile of the blade leading edge and the operating conditions considered are representative of that normally found in a heavy-duty gas turbine. The geometries investigated cover angled turbulators of various angles, pitches and heights. Partial and full length rib coverage as well as broken ribs are also considered. In addition, the impact of including fillets in the geometry is assessed. The experimental and numerical studies are conducted at passage Reynolds numbers ranging from 7.5·104 to 1.3·105. Experiments are performed using Perspex models at atmospheric conditions. The internal heat transfer coefficients on all internal surfaces are measured via thermochromic liquid crystal method and the pressure drop is measured via pressure taps distributed along the channel. The predicted and experimental heat transfer enhancements are compared on the leading-edge, pressure, suction and web surfaces. The overall non dimensional cooling performance numbers are also compared for the various geometries. The results show a large variation of heat transfer enhancement and pressure loss over the various turbulator geometries investigated. Also, the complex flow structures lead to highly differentiated results for leading edge, pressure, suction and web surfaces. Although some configurations with higher ribs lead to increased heat transfer, the associated pressure losses are also shown to increase substantially.

The film cooling performance on a convex surface subjected to zero and favourable pressure gradient free-stream flow was investigated. Adiabatic film cooling effectiveness values were obtained for five different injection geometries, three with cylindrical holes and two with shaped holes. Heat transfer coefficients were derived for selected injection configurations. CO2 was used as coolant to simulate density ratios between coolant and free-stream close to gas turbine engine conditions. The film cooling effectiveness results indicate a strong dependency on the free-stream Mach number level. Results obtained at the higher free-stream Mach number show for cylindrical holes generally and for shaped holes at moderate blowing rates significant higher film cooling effectiveness values compared to the lower free-stream Mach number data. Free-stream acceleration generally reduced adiabatic film cooling effectiveness relative to constant free-stream flow conditions. The different free-stream conditions investigated indicate no significant effects on the corresponding heat transfer increase due to film injection. The determined heat flux ratios or film cooling performance indicated that coolant injection with shaped film cooling holes is much more efficient than with cylindrical holes especially at higher blowing rates. Heat flux penalties can occur at high blowing rates when using cylindrical holes.

This paper investigates the heat transfer and pressure loss characteristic in the internal cooling system of the trailing edge of a gas turbine blade. The geometrical profile of the blade trailing edge and the operating conditions considered are representative of that normally found in a heavy-duty gas turbine. The trailing edge geometry consists of two radial passages with inclined turbulators which are connected with a bend. The trailing edge section consists of pins rows and a flow ejection cut-out slot. The impact of a cross-over hole in the web connecting the serpentine passages is also investigated. Both numerical and experimental studies were conducted at several passage Reynolds numbers ranging from 104 to 106. Experiments were conducted in a Perspex model at atmospheric conditions. The internal heat transfer coefficients were measured via the transient liquid crystal method and the pressure drop was measured via pressure taps. The impact of blade rotation on the heat transfer and pressure drop was also assessed numerically. Comparison of the measured and predicted heat transfer coefficients and pressure drops shows a good agreement for several flow conditions. The three-dimensional flow field in the passage and in the downstream pin banks was well captured numerically, with and without coolant injection via cross-over hole.

Antioxidant enzyme profile of two clinical isolates of Entamoeba histolytica varying in sensitivity to antiamoebic drugs

This paper investigates the flow, heat transfer and film cooling effectiveness of advanced high-pressure turbine blade tips and endwall. Two blade tip configurations have been studied, including a full rim squealer and a partial squealer with a leading edge and trailing edge cut-out. Both blade tip configurations have pressure side film cooling, and cooling air extraction through dust holes which are positioned along the airfoil camber line on the tip cavity floor. The investigated clearance gap and the blade tip geometry are typical of that commonly found in the high pressure turbine blades of heavy-duty gas turbines. Numerical studies and experimental investigations in a linear cascade have been conducted at a blade exit isentropic Mach number of 0.8 and a Reynolds number of 9 × 105 . The influence of the coolant flow ejected from the tip dust holes and the tip pressure side film holes has also been investigated. Both the numerical and experimental results showed that there is a complex aero-thermal interaction within the tip cavity and along the endwall. This was evident for both tip configurations. Although, the global heat transfer and film cooling characteristics of both blade tip configurations were similar, there were distinct local differences. The partial squealer exhibited higher local film cooling effectiveness at the trailing edge but also low values at the leading edge. For both tip configurations, the highest heat transfer coefficients were located on the suction side rim within the mid-chord region. However on the endwall, the highest heat transfer rates were located close to the pressure side rim and along most of the blade chord. Additionally, the numerical results also showed that the coolant ejected from the blade tip dust holes partially impinges onto the endwall.Copyright © 2010 by Alstom Technology, Ltd.

Heat transfer measurements of a confined impingement cooling configuration with ribs on the target surfaces are presented. The assembly consists of four non-perpendicular walls of which one holds two rows of staggered inclined jets, each impinging on a different adjacent wall. The ribs are aligned with the inclined jet axes, have the same pitch and are staggered to the impinging jets. The flow exhausts through two staggered rows of holes opposing the impingement wall. The passage geometry is related to a modern gas turbine blade cooling configuration. A transient liquid crystal technique was used to take spatially resolved surface heat transfer measurements for the ground area between the ribs. A comparison with the smooth baseline configuration reveals local differences and a generally reduced heat transfer for the rib-roughened case. Furthermore, lumped heat capacity measurements of the ribs yielded area averaged heat transfer information for the ribs. From the combination of ground and rib heat transfer measurements it is concluded that the overall performance of the ribbed configuration depends on the Reynolds number. Of the five investigated jet Reynolds numbers from 10,000 up to 75,000, only for the highest Re the averaged Nusselt numbers increase slightly compared to the smooth baseline configuration.Copyright © 2010 by ASME

Low emission requirements for heavy-duty gas turbines can be achieved with flat combustor temperature profiles, reducing the combustor peak temperature. As a result, the heat load on the first stage heat shield above the first stage blade increases. High lift airfoils cause increased thermal loading on the heat shield above the blade tip and impact the unavoidable secondary flows, including complex vortex flows. Cascade tests have been performed on a blade with a generic high lift profile and the results on the heat shield are presented. A transient thermochromic liquid crystal measurement technique was used to obtain heat transfer coefficients on the heat shield surface. Several variations of blade tip clearance were investigated, and the impact on heat transfer coefficients is shown. Computational fluid dynamics predictions are compared to the experimental data to interpret the data and validate the CFD.Copyright © 2011 by Alstom Technology Ltd.

Experimental and numerical heat transfer results in a trapezoidal duct with two staggered rows of inclined impingement jets are presented. The influence of changes in the jet bore geometry on the wall heat transfer is examined. The goal of this project is to minimize the thermal load in an internal gas turbine blade channel and to provide sufficient cooling for local hot spots. The dimensionless pitch is varied between p/djet = 3 –6. For p/djet = 3 , cylindrical as well as conically narrowing bores with a cross section reduction of 25% and 50%, respectively, are investigated. The studies are conducted at 10,000 ≤ Re ≤ 75,000 . Experimental results are obtained using a transient thermochromic liquid crystal technique. The numerical simulations are performed solving the RANS equations with FLUENT using the low-Re k-ω-SST turbulence model. The results show that for greater pitch, the decreasing interaction between the jets leads to diminished local wall heat transfer. The area averaged Nusselt numbers decrease by up to 15% for p/djet = 4.5 , and up to 30% for p/djet = 6 , respectively, if compared to the baseline pitch of p/djet = 3 . The conical bore design accelerates the jets, thus increasing the area-averaged heat transfer for identical mass-flow by up to 15% and 30% for the moderately and strongly narrowing jets, respectively. A dependency of the displacement between the Nu maximum and the geometric stagnation point from the jet shear layer is shown.Copyright © 2011 by ASME and Alstom Technology, Ltd.

Impingement cooling is widely used in cooling configurations for gas turbine components. Relatively high local heat transfer rates and the possibility of jet adjustment to specific geometries are advantageous for internal turbine blade cooling designs. In this paper a confined impingement cooling configuration is investigated. The assembly consists of four non-perpendicular walls of which one holds two rows of staggered inclined jets, each impinging on a different adjacent wall. The flow extraction is realized through two staggered rows of holes opposing the impingement holes wall. Heat transfer experiments were carried out using a transient liquid crystal technique for a series of jet Reynolds numbers between 10,000 and 75,000. The obtained local heat transfer data was evaluated regarding spatially resolved Nusselt numbers as well as line and area averaged values. The results include measurements of the discharge coefficients for the flow through the impingement holes. Numerical simulations of the flow field were carried out, complementary to the experiments. The simulations yield information for a better understanding of the main flow structures. The jets cause high heat transfer in the flow impinging regions with Nusselt numbers up to 180 for Re = 45,000. By contrast, for the same Reynolds number the Nusselt number drops below 20 in flow recirculation regions. For area-averaged Nusselt numbers, the correlation Nu ∝ Rex was found to be valid with slightly modified exponents for each passage wall.Copyright © 2009 by Alstom

Transient liquid crystal experiments have been carried out to measure the effectiveness and heat transfer characteristics of leading-edge film cooling for three different film cooling holes configurations at design and off-design incidence angle. The three configurations are based on the same representative leading edge model of a turbine blade, consisting of a symmetrical blunt body with a specific leading edge wedge angle. Film cooling is introduced from two rows of cooling holes, representative of a pressure-side row and a suction-side row. At design incidence, film cooling performances are symmetric. There is a jet lift-off situation and shaped holes significantly improve the film cooling performances because of a better lateral coverage and a reduced coolant momentum at the hole exit. At 5° off-design incidence angle, on the suction side, the situation is similar to that of a 0° incidence but with higher film cooling performances due to a reduced local blowing ratio. At 5° incidence on the pressure side, a beneficial interaction between the jets of the pressure side row appears. For middle and high blowing ratio, the film cooling performances are also better than 0° incidence. At 5° incidence on the pressure side, shaped holes also improve the film cooling performances in comparison to cylindrical holes.

Sequential impingement channels can reduce the adverse effect of crossflow in narrow impingement channels, as well as increase the cooling efficiency. In this work, sequential impingement channels are experimentally investigated using the transient liquid crystal technique to assess their thermal performances. A low heat transfer region is identified in the downstream part of the first channel where the flow is discharged into the second plenum. Various means of increasing the heat transfer at this location are investigated. Ribs on the target plate allow for an increase of the average heat transfer coefficient with small losses in pressure. Reducing the channel cross-section increases the mean flow velocity and, combined with the ribs, allows for a further increase of the heat transfer. Additionally, the geometrical changes of the channel caused by the addition of a ramp with a rounded corner, allow to decrease the pressure losses associated with the discharge into the second plenum, which is not optimal in the baseline configuration due to the sharp corner of the purge hole. Further reducing the cross-section to increase the heat transfer, however, increases the pressure losses due to the small open area in the transition zone.