Robert Krewinkel

This work experimentally studied a rotating, blade-shaped, two-pass cooling channel with various 45-deg rib turbulators and a tip-turning vane. The first passage of the cooling channel has an aspect ratio (AR) = 4:1 with the coolant moving radially outward. After a 180-deg blade-shaped tip turn, the coolant flows radially inward into the second passage with an AR = 2:1. The first and second passages are oriented at 50 deg and 105 deg from the rotation direction, respectively. In the tip-turn portion, a tip-turning vane with an oval cross section connects the mid-lines of the two passes. Two orientations of 45 deg angled ribs were considered in this work: unusual and criss-cross. The ribs have a profiled cross section and were placed in-line (P/e = 10, e/H = 0.16) on the leading and trailing surfaces of the channel. Five inlet Reynolds numbers (10,000–45,000) and five rotational speeds (0−400 rpm) were considered in this study, with a maximum rotation number of Ro = 0.38 achieved in the first pass. The results showed that the effect of the turning vane on heat transfer and pressure loss for the 45-deg unusual rib case was insignificant. The overall heat transfer for the criss-cross rib case was slightly increased by the turning vane. For pressure loss, the turning vane provided approximately an 8% reduction in the 45-deg criss-cross rib case, which was the highest in the rib cases. The 45-deg criss-cross ribs presented the highest thermal performance with the turning vane.

Computational fluid dynamics (CFD) is the standard tool in the turbomachinery industry to analyze and optimize internal cooling systems of turbine components, but the code applied has to be validated. This paper presents a combined experimental and numerical study on the flow field and heat transfer in a cooling system consisting of a three-pass serpentine with rib turbulators and trailing edge ejection. The cooling geometry is taken from a stator vane currently used in an industrial gas turbine and operates at a coolant inlet Reynolds number of 45,000. As an experimental technique, magnetic resonance velocimetry (MRV) was used to obtain the three-dimensional time-averaged velocity field of the isothermal flow. The measurements were conducted in a large-scale model and resulted in 3.2 million velocity vectors and measurement uncertainty of 6.1% of the bulk inlet velocity. The local wall heat transfer was measured in a separate experiment using thermochromic liquid crystals (TLC). These measurements yielded the distribution of the heat transfer coefficient on both the pressure and the suction side internal walls with a measurement uncertainty of 12%. The experimental data are used as a reference for the numerical study. In total, eight turbulence models are evaluated here, including one-equation, two-equation, algebraic and differential Reynolds stress models, and a scale adaptive simulation. The results show the differences between the velocity fields and the heat transfer coefficient distribution, allowing for the identification of the optimum turbulence model for this particular type of flow.

The requirement for reduced emissions and the growing demand for gas turbine efficiency are in part met through increasing firing temperatures. However, development budgets leave only limited time for dedicated thermal testing. Consequently, manufacturers are seeking novel temperature measurement technologies to validate new engine designs. This paper will demonstrate how a new temperature mapping technology can be utilized for non-dedicated (multi-cycling) testing while still delivering high-resolution temperature data in a non-dedicated test on a combustor of an industrial gas turbine. Typically, thermocouples are used to monitor the temperature during tests, but they only provide one data point. Color-changing thermal paints are used to deliver measurements over complete surfaces, but they require dedicated testing with short-duration exposure, necessitating dismantling and re-assembling the engine for further testing. Thermal History Coatings (THC) present an alternative solution to providing high-density temperature information. This coating permanently changes consistent with the maximum temperature of exposure during the test. A laser-based instrumentation technique is then used to obtain temperatures. The maximum temperature profile of the surface can be determined through a customized calibration. Given the complex cooling system of a combustor, the high temperatures, and the long-time exposure, this case offers a unique possibility for the testing of the coating under real engine conditions. The coated region covered the external surface of the can. Highly significant is the number of measurement points above 7000 (2 × 2 mm resolution), which enables advanced analysis. This provides insight into the impact of local features, e.g., the region adjacent to a cooling hole. The temperature profile is compared to a computational fluid dynamics-conjugate heat transfer model and thermocouple measurements for the calibration of cooling pre-design methods.

This study features a rotating, blade-shaped, two-pass cooling channel, which follows the contour and shape of realistic airfoils. Effects of two different 45-deg rib orientations on the heat transfer and pressure loss were investigated and discussed. The cross section and orientation with respect to rotation vary in the serpentine cooling channel. The first passage of the channel is oriented at 50 deg from the direction of rotation, and the second passage is angled at 105 deg. The radially outward first passage has an aspect ratio (AR) = 4:1. After a 180-deg blade-shaped tip turn, the coolant flows radially inward into the AR = 2:1 s passage. The copper plate method was applied to calculate average heat transfer coefficients in each region of the cooling channel. The 45 deg angled ribs with a profiled cross section are placed on the leading and trailing surfaces in two ways: unusual and criss-cross. The rib spacing is P/e = 10, and rib height is e/H = 0.16. In this study, the Reynolds number ranges from 10,000 to 45,000 in the first passage, resulting in Re = 16,000 to 73,000 in the second passage. The rotational speed varies from 0 to 400 rpm, corresponding to maximum rotation numbers of 0.38 and 0.15 in the first and second passages, respectively. From the results, the 45-deg usual and unusual ribs generally have higher heat transfer than the criss-cross ribs. However, the criss-cross ribs have the lowest pressure loss penalty among the three cases. Taking both heat transfer and pressure loss into account, the 45-deg unusual ribs have a higher thermal performance under the stationary condition while the criss-cross ribs are better than other rib orientations under rotating conditions. Thus, the 45-deg unusual and criss-cross ribs can be considered for applications in real engines. The heat transfer and pressure loss data in this study provide important information for internal cooling of gas turbine blades.

This comparative study is concerned with the advances in nozzle guide vane (NGV) design developments and their influence on endwall film cooling performance by injecting coolant through the purge slot. This experimental study compares the film cooling effectiveness and the aerodynamic effects for different purge slot configurations on both a flat and an axisymmetrically contoured endwall of a NGV. While the flat endwall cascade was equipped with cylindrical vanes, the contoured endwall cascade consisted of modern NGVs, which represent state-of-the-art high-pressure turbine design standards. Geometric variations, e.g., the slot width and injection angle, as well as different blowing ratios were realized. The mainstream flow parameters were set to meet real engine conditions with regard to Reynolds and Mach numbers. Pressure-sensitive paint was used to determine the adiabatic film cooling effectiveness. Five-hole probe measurements were performed to measure the flow field in the vane wake region. For a more profound insight into the origin of the secondary flows, oil dye visualizations were carried out. The results show that the advances in NGV design have a significantly positive influence on the distribution of the coolant. This has to be attributed to lesser disturbance of the coolant propagation by secondary flow for the optimized NGV design, since the design features are intended to suppress the formation of secondary flow. Therefore, it is advisable to take these effects into account when designing the film cooling system of a modern high-pressure turbine.

This study features a rotating, blade-shaped, two-pass cooling channel with a variable aspect ratio (AR). Internal cooling passages of modern gas turbine blades closely follow the shape and contour of the airfoils. Therefore, the cross section and the orientation with respect to rotation varies for each cooling channel. The effect of passage orientation on the heat transfer and pressure loss is investigated by comparing to a planar channel design with a similar geometry. Following the blade cross section, the first pass of the serpentine channel is angled at 50 deg from the direction of rotation while the second pass has an orientation angle of 105 deg. The coolant flows radially outward in the first passage with an AR = 4:1. After a 180-deg tip turn, the coolant travels radially inward into the second passage with AR = 2:1. The copper plate method is applied to obtain the regionally averaged heat transfer coefficients on all the interior walls of the cooling channel. In addition to the smooth surface case, 45 deg angled ribs with a profiled cross section are also placed on the leading and trailing surfaces in both the passages. The ribs are placed such that P/e = 10 and e/H = 0.16. The Reynolds number varies from 10,000 to 45,000 in the first passage and 16,000 to 73,000 in the second passage. The rotational speed ranges from 0 to 400 rpm, which corresponds to maximum rotation numbers of 0.38 and 0.15 in the first and second passes, respectively. The blade-shaped feature affects the heat transfer and pressure loss in the cooling channels. In the second passage, the heat transfer on the outer wall and trailing surface is higher than the inner wall and leading surface due to flow impingement and the swirling motion induced by the blade-shaped tip turn. The rotational effect on the heat transfer and pressure loss is lower in the blade-shaped design than the planar design due to the feature of angled rotation. The tip wall heat transfer is significantly enhanced by rotation in this study. The overall heat transfer and pressure loss in this study is higher than the planar geometry due to the blade-shaped feature. The heat transfer and pressure loss characteristics from this study provide important information for the gas turbine blade internal cooling designs.

With the growing importance of regenerative power generation and especially of a hydrogen-based economy, the full potential of gas turbines of the smaller output class (<10 MW) can be ideally exploited to provide peak coverage of the energy need whilst stabilizing the electric grids in the mid- and low-voltage range. Such machines can be typically started in a relatively short time (similarly to aero engines) and are capable, at the same time, of delivering dispatchable power-on-demand. A safe, stable and profitable operation under highly unsteady conditions poses renewed challenges for an optimal thermal management (especially in the HP stages) as well as control and surveillance of the machines. The understanding and hence predictability of the propagation of the temperature inhomogeneities originating at the combustor outlet remains hence a primary objective of current research, as persistent distortion patterns could be adopted at the turbine exhaust as diagnostic indications of a malfunction of the combustor, for example. In the present study, low-frequency disturbances introduced by a periodic load variation have been simulated and superimposed to the inhomogeneous, unsteady flow entering a three-stage, high-pressure industrial gas turbine fed by a can-type combustion chamber comprising six silo-burners. The effects of the unsteadiness realized at the combustor exit have been investigated by means of Detached Eddy Simulations, whereby a density-based solution approach with detailed thermodynamics has been employed. The periodic disturbances at the turbine inlet have been obtained by means of an artificially generated, unsteady field, resulting from a two-dimensional snapshot of the flow field at the combustor exit. Also, a combustor failure has been mimicked by reducing (respectively increasing) the mean temperature in some of the turbine inlet regions corresponding to the outlet of two burners. The propagation and amplitude changes of temperature fluctuations have been analyzed in the frequency domain. Tracking of the temperature fluctuations' maxima at the lowest frequencies revealed characteristic migration patterns indicating that the corresponding fluctuations persist with a non-negligible amplitude up to the last rows. A distinct footprint could also be observed at the same locations when a combustor failure was simulated, showing that, in principle, the early detection of combustor failures is indeed possible.

The thermal performance of two different V-shaped rib configurations is measured in a rotating, two-pass rectangular cooling channel. Modeling modern, high pressure turbine blades, the cross-sectional area of the cooling channel varies from the first passage to the second passage. The coolant air travels radially outward in the rectangular first passage with an aspect ratio (AR) of

The thermal performance of two V-type rib configurations is measured in a rotating, two-pass cooling channel. The coolant travels radially outward in the rectangular first pass (aspect ratio (AR—channel width:height) = 4:1) and radially inward in the second pass (AR = 2:1). Both the passages are oriented 90 deg to the direction of rotation. The leading surface (LS) and trailing surface (TS) of the channel are roughened with V-type ribs. The first V-shaped configuration has a narrow gap at the apex of the V. The configuration is modified by off-setting one leg of the V to create a staggered discrete, V-shaped configuration. The ribs are oriented 45 deg relative to the streamwise coolant direction. The heat transfer enhancement and frictional losses are measured with varying Reynolds and rotation numbers. The Reynolds number varies from 10,000 to 45,000 in the AR = 4:1 first pass; this corresponds to 16,000 to 73,500 in the AR = 2:1 second pass. The maximum rotation numbers are 0.39 and 0.16 in the first and second passes, respectively. The heat transfer enhancement on both the leading and trailing surfaces of the first pass of the 45 deg V-shaped channel is slightly reduced with rotation. In the second pass, the heat transfer increases on the leading surface, while it decreases on the trailing surface. The 45 deg staggered, discrete V-shaped ribs provide increased heat transfer and thermal performance compared to the traditional V-shaped and standard, 45 deg angled rib turbulators.