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A two-dimensional numerical model for compressible branching flow through a slot is described for the purpose of predicting the discharge coefficients of film cooling holes in gas turbine blades. The method employs free-streamline theory and the hodograph transformation. It calculates the area ratio of hole to duct and the contraction coefficient from a set of prescribed boundary conditions. An approximate method for calculating the compressible contraction coefficients is also discussed in the thesis. It employs the incompressible theory previously developed by McNown and Hsu (1951) for the free efflux, the 'compressibility factor' and the flow parameter (P_o-P_j)/(P_o-P₁), where P_o, P_j, P₁ represent the stagnation pressure, the static pressure of the jet and the static pressure of the approach flow, respectively. The advantages of using this method are the direct input of the area ratio of hole to duct and its speed of calculation. Experimental tests were performed using a specially designed rig in a supersonic wind tunnel. The investigations included sharp-edged slots with three different widths, a single hole and a row of two holes. The approach velocity in terms of the characteristic Mach number ranged from 0.18 to 0.58 and the pressure ratio P_o/P_j, ranged from 1.10 to 1.97. Agreement between the experimental data and the theoretical values was good to within the experimental accuracy (typically around +/- 5%) for the slots and the 2-hole configuration. For the 1-hole configuration, less bleed flow than predicted was observed, with the discrepancy varying from 7% to 18%. The latter case is a very severe test of a purely two-dimensional theory. The results for the 2-hole plate suggest that the slot theory can in fact be used to predict the flow through a row of holes with small pitch to diameter ratios.

Heat transfer study on gas turbine blades is very important due to the resultant increase in cycle thermal efficiency. This study is focused on the heat transfer effects on a reference nozzle guide vane and test rig component preparation in heat and power technology division at KTH. In order to prepare the current test rig for heat transfer experiments, some feature should be changed in the current layout to give a nearly instant temperature rise for heat transfer measurement. The heater mesh component is the main component to be added to the current test rig. Some preliminary design parameters were set and the necessary power for the heater mesh to achieve required step temperature rise was calculated. For the next step, it is needed to estimate the heat transfer coefficient and the other parameters for study on the reference blade using numerical methods. Boundary layer analysis is very important in heat transfer modeling so to model the reference blade heat transfer and boundary layer properties, a 2D boundary layer code TEXSTAN is used and the velocity distribution around the vane was set to an input dataset file. After elements refinement to ensure the numerical accuracy of TEXSTAN code, various turbulence modeling was check to predict the heat transfer coefficient and boundary layer assessments. It was concluded from TEXTAN calculations that both suction and pressure side have transition flow while for the suction side it was predicted that the flow regime at trailing edge is fully turbulent. Based on the Abu-Ghannam –Shaw Transition model and by the aid of shape factor data, momentum Reynolds number and various boundary layer properties, it was concluded that the pressure side remains in transient region.

The Large Eddy Simulation approach is employed to predict the flow physics and heat transfer characteristics of a film-cooling problem that is formed from the interaction of a coolant jet with a hot mainstream flow. The film-cooling technique is used to protect turbine blades from thermal failure, allowing the gas inlet temperature to be increased beyond the failure temperature of the turbine blade material in order to enhance the efficiency of gas turbine engines. A coolant fluid is injected into the hot mainstream through several rows of injection holes placed on the surface of a gas turbine blade in order to form a protective coolant film layer on the blade surface. However, due to the complex, unsteady and three-dimensional interactions between the coolant and the hot gases, it is difficult to achieve the desired cooling performance. Understanding of this complex flow and heat transfer process will be helpful in designing more efficiently cooled rotor blades. A comprehensive numerical investigation of a rotating film-cooling performance under different conditions is conducted in this thesis, including film-cooling on a flat surface and film-cooling on a rotating gas turbine blade. The flow-governing equations are discretised based on the finite-volumes method and then solved iteratively using the well-known SIMPLE and PISO algorithms. An in-house FORTRAN code has been developed to investigate the flat plate film-cooling configuration, while the gas turbine blade geometry has been simulated using the STAR-CCM+ CFD commercial code. The first goal of the present thesis is to investigate the physics of the flow and heat transfer, which occurs during film-cooling from a standard film hole configuration. Film-cooling performance is analysed by looking at the distribution of flow and thermal fields downstream of the film holes. The predicted mean velocity profiles and spanwise-averaged film-cooling effectiveness are compared with experimental data in order to validate the reliability of the LES technique. Comparison of adiabatic film-cooling effectiveness with experiments shows excellent agreement for the local and spanwise-averaged film-cooling effectiveness, confirming the correct prediction of the film-cooling behaviour. The film coverage and film-cooling effectiveness distributions are presented along with discussions of the influence of blowing ratio and rotation number. Overall, it was found that both rotation number and blowing ratio play significant roles in determining the film-cooling effectiveness distributions. The second goal is to investigate the impact of innovative anti-vortex holes on the film-cooling performance. The anti-vortex hole design counteracts the detrimental kidney vorticity associated with the main hole, allowing coolant to remain attached to the blade surface. Thus, the new design significantly improves the film-cooling performance compared to the standard hole arrangement, particularly at high blowing ratios. The anti-vortex hole technique is unique in that it requires only readily machinable round holes, unlike shaped film-cooling holes and other advanced concepts. The effects of blowing ratio and the positions of the anti-vortex side holes on the physics of the hot mainstream-coolant interaction in a film-cooled turbine blade are also investigated. The results also indicate that the side holes of the anti-vortex design promote the interaction between the vortical structures; therefore, the film coverage contours reveal an improvement in the lateral spreading of the coolant jet.

This thesis focuses on the study of flow and heat transfer in porous media in both laminar and turbulent flow regimes, by using Volume Averaged Reynolds Navier Stokes (VARNS) approach. The main concern is to investigate the possibility of using porous media for the gas turbine blade cooling. Very recently, using this technique in blade cooling, particularly with internal cooling, has motivated many researchers due to an effective enhancement in the blade cooling. In this study turbulence is represented by using the Launder-Sharma low-Reynolds-number k-Îμ turbulence model, which is modified via proposals by Nakayama and Kuwahara (2008) and Pedras and de Lemos (2001) for extra source terms in the turbulent transport equations to account for the porous structure, which is treated as rigid and isotropic. Due to the changing of the effective porosity as the clear fluid region is approached, the porosity and additional source term in the macroscopic Reynolds averaged Navier-Stokes equations are relaxed across a thin transitional layer at the edges of the porous media. This is achieved by utilizing exponential damping relations to consider these changes. The Local Thermal Equilibrium (LTE) (one-energy equation) model is used for the thermal analysis in porous media. In order to investigate the validity of the extended model, laminar and turbulent flow in different cases, fully developed and developing flows, have been considered. For laminar flows, fully developed plane channel flows with one and two porous layers, a channel with a single porous block and partially filled porous channel flows have been examined for the purpose of validating the extra drag terms in the momentum equations. For the validation purpose for turbulent flows in porous media, the extended model has been tested in homogeneous porous media, turbulent porous channel flows, turbulent solid/porous rib channel flows, and repeated turbulent porous baffled channel flows. Results of all laminar cases show excellent qualitative agreements with the available numerical calculations and experimental data. Results of all turbulent cases show that the extended model returns generally satisfactory accuracy through the comparisons with the available data, except for some predictive weaknesses in regions of either impingement or adverse pressure gradients, both of which are largely due the underlying eddy-viscosity model formulation employed. Thus, from all results, it can be confirmed that the extended model is promising for engineering applications.

The required useful service lives of gas turbine components and parts are naturally one of the major design constraints limiting the gas turbine design space. For example, the required service life of a turbine blade limits the firing temperature in the combustor, which in turn limits the performance of the gas turbine. For a cooled turbine blade, it also determines the necessary cooling flow, which has a strong impact on the turbine efficiency. In most gas turbine design practices, the life prediction is only emphasized during or after the detailed design has been completed. Limited life prediction efforts have been made in the early design stages, but these efforts capture only a few of the necessary key factors, such as centrifugal stress. Furthermore, the early stage prediction methods are usually hard coded in the gas turbine system design tools and hidden from the system designer s view. The common failure mechanisms affecting the service life, such as creep, fatigue and oxidation, are highly sensitive to the material temperatures and/or stresses. Calculation of these temperatures and stresses requires that the geometry, material properties, and operating conditions be known; information not typically available in early stages of design. Even without awareness of the errors, the resulting inaccuracy in the life prediction may mislead the system designers when examining a design space which is bounded indirectly by the inaccurate required life constraints. Furthermore, because intensive creep lifing analysis is possible only towards the end of the design process, any errors or changes will cost the engine manufacturer significant money; money that could be saved if more comprehensive creep lifing predictions were possible in the early stages of design. A rapid, physics-based life prediction method could address this problem by enabling the system designer to investigate the design space more thoroughly and accurately. Although not meant as a final decision method, the realistic trends will help to reduce risk, by providing greater insight into the bounded space at an earlier stage of the design. The method proposed by this thesis was developed by first identifying the missing pieces in the system design tools. Then, by bringing some key features from later stages of design and analysis forward through 0/1/2Ds dimensional modeling and simulation, the method allows estimation of the geometry, material selection, and the loading stemming from the operating conditions. Finally, after integration with a system design platform, the method provides a rapid and more complete way to allow system designers to better investigate the required life constraints. It also extracts the creep life as a system level metric to allow the designers to see the impact of their design decisions on life. The method is to be first applied to a cooled gas turbine blade and could be further development for other critical parts. These new developments are integrated to allow the system designers to better capture the blade creep life as well as its impact on the overall design.