Martin Schnieder

Design Considerations and Validation of Trailing Edge Pressure Side Bleed Cooling

One important design measure, which allows achieving a higher efficiency in heavy-duty gas turbines, is reducing the trailing edge thickness of the turbine vanes and blades. A common approach to achieve an efficient cooling of thin trailing edges is pressure side coolant ejection. As the trailing edge is one of the areas of the vanes and blades with the highest heat load, a careful design and validation approach is required. First, the most important design parameters were identified for the design of pressure side bleed. In the present investigation, focus is given to the blockage of the internal features, slot width, overhang length, pressure side lip thickness, effect of rotation and blowing rate. Flat plate test results from a low speed rig are used to choose suitable parameters, which fulfill requirements of a specific design. Previous investigations have shown that contrary to CFD using steady RaNS, unsteady detached eddy simulations can predict film effectiveness of pressure side bleed with good accuracy. Therefore, this approach is used to complete experimental data. The effect of going from low speed rig conditions to engine conditions is modeled with this approach. The final design is investigated in a high-speed cascade rig. Film effectiveness is measured using thermocromic liquid crystals. The cascade results confirm results from the low speed rig. High levels of film effectiveness allow effective cooling of the trailing edge overhang.Copyright

Experimental Study of Heat Transfer in Gas Turbine Blades Using a Transient Inverse Technique

To enhance specific power output and thermal efficiency of gas turbine engines, industry searches for ways to increase the turbine inlet temperatures. Therefore, temperatures of turbine blades increase as well and necessitate active cooling of these components. Experimental design work on such internal cooling schemes is carried out to find acceptable compromises between heat transfer and pressure losses. It is often carried out by using transient thermochromic liquid crystal techniques in combination with Plexiglas models. However, for real turbine blades this experimental technique is inappropriate due to the lack of optical access. Therefore, to study actual turbine blades there is need for development of noninvasive, nondestructive methodologies. This article describes a measurement technique that allows determination of internal heat transfer coefficients of real turbine blades experimentally. Thus, a test rig with a rapidly responding heater was designed to fulfill the requirement of a sudden increase in the air temperature within the cooling passages. The outer surface temperatures were measured using infrared thermography. To estimate the spatial distribution of internal heat transfer coefficients from transient surface temperatures the inverse heat transfer problem was solved. As optimization algorithm the Levenberg–Marquardt method was chosen. Outer surface temperature data was measured for a rectangular reference model with rib turbulators and compared with simultaneously acquired data using the thermochromic liquid crystal technique. It is concluded that the new experimental measurement technique could be used to quantitatively determine internal heat transfer coefficients.

The Influence of Including a Partially Smooth Section in the 2nd Leg of an Internally Ribbed Two Pass Cooling Channel

Internal cooling schemes for blades in a gas turbine engine often are subject to compromises between increased pressure losses in return for greater levels of heat transfer required to maintain durability levels in the engine’s harsh environment. Rib configurations have been the subject of much study in past years, however these configurations are normally presumed to be used in “full-coverage” mode, meaning that the ribs are placed in the channel in a continuous and uniform manner. This study investigates the interaction between the bend effects downstream of a 180° bend, which cause higher local heat transfer, and the effect of ribs. Some of the ribs directly downstream of the 180° bend in the 2nd leg of a two pass high aspect ratio (4:1) channel were removed and the effect on heat transfer was assessed. Experimental results showed that the heat transfer level recovered quickly once ribs were encountered. As expected, some decrease in heat transfer was observed in the region where ribs were removed; however total pressure losses in the channel were also much lower. Results include detailed two-dimensional heat transfer distributions determined by the transient liquid crystal method as well as an analysis of the balance between pressure recovery and local heat transfer levels. Generally, for the accuracy of the transient liquid crystal technique in complex three-dimensional flows, strongly varying fluid temperatures present in and downstream of the bend region must be taken into account. For this study, time and position dependent fluid temperature distributions were measured to account for these effects, making it possible to obtain high quality heat transfer results in those regions.Copyright

Heat Transfer Measurements in an Internal Cooling System Using a Transient Technique With Infrared Thermography

In this paper, a transient method for measuring heat transfer coefficients in internal cooling systems using infrared thermography is applied. The experiments are performed with a two-pass internal cooling channel connected by a 180 deg bend. The leading edge and the trailing edge consist of trapezoidal and nearly rectangular cross-sections, respectively, to achieve an engine-similar configuration. Within the channels, rib arrangements are considered for heat transfer enhancement. The test model is made of metallic material. During the experiment, the cooling channels are heated by the internal flow. The surface temperature response of the cooling channel walls is measured on the outer surface by infrared thermography. Additionally, fluid temperatures as well as fluid and solid properties are determined for the data analysis. The method for determining the distribution of internal heat transfer coefficients is based on a lumped capacitance approach, which considers lateral conduction in the cooling system walls as well as natural convection and radiation heat transfer on the outer surface. Because of time-dependent effects, a sensitivity analysis is performed to identify optimal time periods for data analysis. Results are compared with available literature data.

Validation of a First Vane Platform Cooling Design

Low emission requirements for large industrial 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 vane platforms increases and platform film cooling is an important requirement. Furthermore, high lift airfoils generate stronger secondary flows including complex vortex flows over the platforms, which impacts heat transfer coefficients and film cooling. Cascade tests have been performed on a high lift profile with a platform film configuration and will be presented. The linear cascade was operated at engine representative Mach numbers. Pressure measurements are compared to design data to ensure correct operating conditions and periodicity of the cascade. The thermochromic liquid crystal measurement technique is used to obtain adiabatic film cooling effectiveness. The upstream gap (corresponding to the gap between the combustor and turbine) and the purge air exiting this gap are included in the investigations. The effect of the purge air on the recovery temperature is very strong and needs to be taken into account for the layout of the cooling scheme. The heat transfer coefficient distribution on the platform is obtained for an uncooled configuration using a transient infrared imaging technique with heat flux reconstruction. Computational fluid dynamics (CFD) assessments are used to support the validation results. Heat transfer coefficients and the effect of the purge air on adiabatic wall temperatures are compared with experimental results.Copyright

Thermal Validation of a Heat Shield Surface for a High Lift Blade Profile

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

Validation and Analysis of Numerical Results for a Varying Aspect Ratio Two-Pass Internal Cooling Channel

Numerical results for an internal ribbed cooling channel including a 180° bend with a 2:1 inlet and 1:1 aspect ratio outlet channel were validated against experimental results in terms of spatially resolved heat transfer distributions, pressure losses, and velocity distributions. The numerical domain consisted of one rib segment in the inlet channel and three ribs segments in the outlet channel to reduce the overall numerical effort and allow for an extensive parametric study. The results showed good agreement for both heat transfer magnitudes and spatial distributions and the numerical results captured the predominate flow physics resulting from the 180° bend. The production of Dean vortices and acceleration of the flow in the bend produced strongly increased heat transfer on both the ribbed and unribbed walls in the outlet channel in addition to increases due to the ribs. Numerical simulations were performed for a wide range of divider wall-to-tip wall distances, which influenced the position of the highest heat transfer levels on the outlet walls and changed the shape of the heat transfer distribution on the tip wall. Analysis of section averages of heat transfer in the bend and outlet channel showed a strong influence of the tip wall distance while no effect was seen upstream of the bend. A similarly large effect on pressure losses in the bend was observed with varying tip wall position. Trends in averaged heat transfer varied linearly with tip wall distance while pressure losses followed a non-linear trend, resulting in an optimum tip wall distance with respect to heat transfer efficiency.Copyright

The Effect of Ribs and Tip Wall Distance on Heat Transfer for a Varying Aspect Ratio Two-Pass Ribbed Internal Cooling Channel

Internal cooling channels with differing aspect ratios are typically found in gas turbine blades due to the varying thickness of the blade from the leading to trailing edge. These serpentine passages often contain several 180° bends, which are sharp edged in the region of the blade tip. The 180° bend has a pronounced effect on the heat transfer characteristics in the outlet channel and tip wall, where a strong influence is seen due to the divider wall-to-tip wall distance in the bend. The present study investigates the effect of the divider wall-to-tip wall distance for a ribbed two-pass cooling channel with a 2:1 inlet and 1:1 outlet channel. Spatially resolved heat transfer measurements were made using the transient thermochromic liquid crystal technique for a smooth and a ribbed configuration using parallel 45° ribs. Effects of the 180° bend on heat transfer and rib-induced enhancements were identified separately and bend effects were found to dominate the heat transfer increase in the outlet channel near the bend. Pressure losses due to the bend and ribs were also independently evaluated for a range of tip wall distances. Results show that the smaller tip wall distances increase heat transfer on the tip wall and outlet channel, but at the cost of an increased pressure loss. An optimum tip wall position is suggested, forming a compromise between heat transfer improvement and increased pressure losses.Copyright

Multi-Row Film Cooling Performances of a High Lift Blade and Vane

This paper investigates the aerodynamic and film cooling effectiveness characteristics of a first stage turbine high lift guide vane and its corresponding downstream blade. The vane and blade geometrical profiles and operating conditions are representative of that normally found in a heavy-duty gas turbine. Both the vane and the blade airfoils consist of multi-row film cooling holes located at various axial positions along the airfoil chord. The film cooling holes are geometrically three-dimensional in shape and depending on the location on the airfoil; they can be either symmetrically fan shaped or non-symmetrically fan shaped. Additionally the film cooling holes can be either compounded or in-line with the external flow direction.Numerical studies and experimental investigations in a linear cascade have been conducted at vane and blade exit isentropic Mach number of 0.8. The influence of the coolant flow ejected from the film cooling holes has been investigated for both the vane and the blade profiles.For the nozzle guide vane, the measured film cooling effectiveness compared well with the predictions, especially on the pressure side. The suction side film cooling effectiveness, which consisted of two pre-throat film rows, proved very effective up-to the suction side trailing edge. For the blade, there was a reasonable comparison between the measured and predicted film cooling effectiveness. Again the blade pre-throat fan shaped cooling holes proved very effective up-to the suction side trailing edge. For the vane, the impact of varying the blowing ratios showed a strong variation in the film cooling effectiveness on the pressure side. However, on the blade, the effect of varying the blowing ratio had a greater impact on the suction side film effectiveness compared to the pressure side.Copyright

Effect of the Biot Number on Metal Temperature of Thermal-Barrier-Coated Turbine Parts: Real Engine Measurements

For numerous hot gas parts (e.g. blades or vanes) of a gas turbine, thermal barrier coating (TBC) is used to reduce the metal temperature to a limit that is acceptable for the component and the required lifetime. However, the ability of the TBC to reduce the metal temperature is not constant, it is a function of Biot and Reynolds number. This behavior might lead to a vane’s or blade’s metal temperature increase at lower load relative to a reference load condition of the gas turbine (i.e. at lower operating Reynolds number).A measurement campaign has been performed, to evaluate metal temperature measurements on uncoated and coated turbine parts in Alstom’s GT26 test power plant in Switzerland. Herewith the impact of varying Reynolds number on the ability of the TBC to protect the turbine components was evaluated.This paper reports on engine-run validation, including details on the application of temperature sensors on thermal-barrier-coated parts. Different methods for the application of thermocouples that were taken into account during the development of the application process are shown.Measurement results for a range of Reynolds number are given and compared to model predictions. Focus of the evaluation is on the measurements underneath the TBC. The impact of different Reynolds number on the ability of the TBC to protect the parts against the hot gas is shown. TBC coated components show under certain circumstances higher metal temperatures at lower load compared to a reference load condition. The measurement values obtained from real engine tests can be confirmed by 1D-model predictions that explain the dependency of the TBC effect on Biot and Reynolds number.Copyright