Joerg Krueckels

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

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

Probabilistic Lifetime Prediction of Thermal Barrier Coating Systems Depending on Manufacturing Scatter

The assessment of Bondcoat/Thermal barrier coating systems is an inherent part of the lifing process of gas turbine component. On the one hand, coatings are considered in the constitutive modelling — e.g. in the thermal model and for the prediction of eigenfrequencies of gas turbine blades. On the other hand, the influence of the coating system on the lifetime of the part (target cyclic life and target operation hours) needs to be assessed. This paper addresses the prediction of coating lifetime. Lifing models of Bondcoat/Thermal barrier coating systems (BC/TBC) are commonly built using isothermal furnace cyclic tests (FCT). The lifetime of the BC/TBC under such test conditions has been shown to depend on multiple coating parameters like TBC thickness, TBC porosity, BC thickness, BC roughness, and also on testing temperature. For example, the TBC life (defined as time to partial TBC spallation) is reduced with increasing temperature, with increasing TBC thickness and decreasing porosity and BC roughness. When operating in a gas turbine (GT), the TBC surface temperature and the BC temperature depend on engine operating conditions, heat transfer of combustion gas and cooling air, coating microstructure and thickness. For instance, a TBC with high porosity typically demonstrates a lower thermal conductivity than that with low porosity. For otherwise same boundary conditions, the BC temperature will decrease with increasing TBC porosity and increasing TBC thickness. The benefit of having a high coating porosity observed in FCT is further amplified by its impact on reducing the BC temperature in GT operation. To the contrary, the positive impact of a reduced TBC thickness observed in FCT is reduced by its negative impact on an increased BC temperature during GT operation. Taking these effects into account a probabilistic lifing model is proposed based on Monte Carlo simulations. Using this model the impact of the manufacturing scatter on the BC/TBC life can be assessed, and enables improved manufacturing by focusing on those parameters that are most critical for coating lifetime.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

Aerodynamic and Heat Transfer Characterization of a Nozzle Vane Cascade With and Without Platform Cooling

In the present paper, aerodynamic and thermal performance of a linear nozzle vane cascade is fully assessed. Tests have been carried out with and without platform cooling, with coolant ejected through a slot located upstream of the leading edge. Cooling air is also ejected through a row of cylindrical holes located upstream of the slot, simulating a combustor cooling system. The cascade was tested at a high inlet turbulence intensity level (Tu1 = 9%) and at variable cooling injection conditions. Aero-thermal characterization of vane platform was obtained through 5-hole probe measurements, oil & dye surface flow visualizations, measurements of end wall adiabatic film cooling effectiveness and heat transfer coefficient. The platform cooling scheme operated at nominal injection rate was shown to effectively reduce the heat load over most of the platform surface, with only a small increase in secondary flows loss. Combustor holes injection resulted beneficial in controlling momentum of coolant approaching the cascade, thus limiting the secondary flows growth and resulting in an increase of the coolant film length inside of the passage.Copyright