The increase of new gas turbine’s efficiency is connected with further rise of turbine inlet temperature and sometimes as well pressure. In these conditions, first cooled turbine stages of a gas turbine engine usually consist of freestanding airfoils, which do not use an integrated shroud, to avoid risk of shroud overheating. In order to better control the radial gap leakage flow between the rotating blade tip and turbine casing, special design features of the airfoil tip need to be considered in the design process to meet the best possible stage performance. In the general engineering practice, a blade tip squealer provides opportunities to control tip clearance loss. In this paper several simplified types of the tip squealer design are investigated to determine the most effective loss control. At this stage of the investigation, blade tip cooling was not taken into account, but aerodynamic effects were analysed in detail. Based on the most common designs of the blade tip in the literature, four geometry types were investigated: (i) a flat tip design as the reference baseline solution, (ii) full tip squealer, (iii) partial squealer along the pressure side (PS) wall with a cut-out at the pressure side near the trailing edge (TE) and (iv) partial squealer along the suction side (SS) wall with a cut-out at the suction side near TE. All these cases have been compared among each other for two relative radial gaps (gap to blade height) of 0.6% and 1.36%. The flow calculations were done with a full 3D Navier-Stokes CFD code. For the flat tip and for full squealer designs, numerical results were validated against well-known experimental data measured on the GE-E3 blade cascade test rig found in the open literature. By using the 3D numerical data, the special attention was considered to confirm reliability and predictive credibility of the blade tip flow obtained from the analytical model. The obtained loss values and flow details were compared for all studied cases, providing insight into turbine stage aerodynamics with respect to minimal and maximal radial clearance.

Modern gas turbines operate at high temperature, which exceed the endurance limit of material, and therefore the turbines components have to be cooled by the air taken from the compressor. The cooling providing positive impact on lifetime of GT has negative impact on its performance. Firstly because the cooling air bypasses combustor and its capacity is not fully utilized. This effect is usually accounted in thermodynamic calculations of gas turbine. Secondly the injection of cooling air in the turbine disturbs the main flow, and may lead to increased losses. In addition cooling requirements lead to limitation on the blade shape (e.g. limiting the minimal size of trailing edge) and thereby negatively affect the losses. These effects were already discussed in the literature, but further investigations for better understanding of flow physics and design improvement are still useful. There is also additional impact of cooling - impact of heat transfer on near wall boundary layer and coolant properties. This effect was not sufficiently discussed in the open literature, where quite often the walls are considered as adiabatic. The paper consists of two main parts. In the first part the results of experimental investigations of several linear cascades with and without trailing edge injection are presented and discussed. In the second part the results of detailed numerical investigations of one of these cascades are presented. One set of calculations were done at the test rig conditions for comparison with measured data. These calculations were used for validation of CFD model. The next sets of calculations were done for engine typical conditions, including the simulation of blade material temperature. The calculations were performed for adiabatic wall and for surface with heat transfer, including the impact of heat transfer on coolant injection. This analysis provides split of losses caused by different factors, and offers the opportunities for efficiency and lifetime improvements of real engine designs/upgrades.

State-of-the-art gas turbines (GT) operate at high temperatures that exceed the endurance limit of the material, and therefore the turbine components are cooled by the air taken from the compressor. The cooling provides a positive impact on the lifetime of GT but has a negative impact on its performance. In convection-cooled turbine blades the coolant is usually discharged through the trailing edge and leads to limitations on the minimal size of the trailing edge, thereby negatively affecting the losses. Moreover, the injection of cooling air in the turbine disturbs the main flow, and may lead to an additional increase in loss. Trailing edge loss is a significant part of the overall loss in modern gas turbines. This study comprises investigations of the unguided flow angle, the trailing edge shape, and cooling air injection through the trailing edge on the base pressure and profile losses in cooled blades. Some vane and blade cascades with different unguided turning angle and two shapes of trailing edges with and without coolant injection were studied both experimentally and numerically. This analysis provides a split of losses caused by different factors, and offers opportunities for efficiency and lifetime improvements of real engine designs/upgrades. In particular, it is shown that an increase in the unguided turning angle and the use of a round trailing edge result in a reduction of loss in case of a relatively thick trailing edge. Numerical investigation showed that an increase in the unguided turning angle at the initial transonic vane with a thick and blunt trailing edge, without a change in other basic geometric parameters, allowed for a significant reduction of the profile loss by about 3-4% at the exit Mach number M 2is = 0.7-1.0. Experimental investigation of four cascades with cooling air injection into the base flow through the trailing edge allowed us to validate the fact that in blades with a low level of base pressure C pb < −0.1 at m = 0 a non-monotonic dependence of the change of losses against relative cooling air mass flow m is observed. Firstly, the cooling air injection into wake increases base pressure and decreases losses; then the losses start to increase with increasing cooling mass flow due to the interaction between the main flow and the cooling air (mixing losses) and, finally, due to the cooling mass flow increase and momentum increase losses are decreased. In blades with an increased level of the base pressure coefficient C pb ≥ −0.1 at m = 0 the cooling air injection results in an increase in losses right from the beginning of the injection and then, according to the cooling mass flow increase and momentum rise, losses decrease. It is also shown that injection through the trailing edge slot parallel to the main flow leads to a neutral loss impact and even a loss reduction in the subsonic range and a loss increase in the supersonic range of exit Mach numbers.