Research On The Aero-Thermal Aspects Of Contoured Endwalls In Gas Turbine Nozzle Guide Vane Cascades

Gas turbine engines are still one of the most important energy and power equipments.They are relatively clean and more efficient.With the development of gas turbine,the turbine endwall regions of the Nozzle Guide Vane(NGV)passage are facing aerodynamic and cooling challenges.Endwall contouring,including axisymmetric endwall contouring and non-axisymmetric endwall contouring,is a practical way for the NGV endwall to reduce passage aerodynamic loss.Meanwhile,endwall cooling is the most challenging part in NGV cooling,and endwall contouring can change the endwall cooling performance.To achieve both better aerodynamic and cooling performance,better understandings of the aerodynamic and cooling characteristics of contoured endwall passages are needed.This thesis studies the aero-thermal aspects of contoured endwall in gas turbine NGV cascades.The axisymmetric contoured endwall varies the endwall profile in the axial direction only,which makes the near-endwall flow accelerate or decelerate.This thesis carries out an aerodynamic optimization on axisymmetric contoured endwall and studies the cooling performance on the optimized endwall with a slot leakage flow located upstream of the passage.For the optimized axisymmetric contoured passage,a proper coolant mass flow ratio range is found.In this range,the cooling effectiveness value is relatively large and the aerodynamic loss coefficient value is relatively low.Larger film cooling effectiveness values are found in the axisymmetric contoured endwall passage than those in the flat endwall passage with the same mass flow ratios.The momentum ratio is found to be the scaling factor of endwall film cooling performance between different axisymmetric contoured endwall passages.The non-axisymmetric contoured endwall varies the endwall profile both in the axial and the circumferential directions,causing raised or sunken part inside of the passage,which influences the flow and heat transfer characteristics locally there.This thesis develops a multi-objective optimization method for the non-axisymmetric contoured endwall.With the flow field analysis and global sensitivity analysis,it is found that the optimized non-axisymmetric contoured endwall changes the acceleration features near the endwall to improve the aerodynamic performance and changes the local velocity to improve the heat transfer characteristics.A non-axisymmetric annular cascade endwall film cooling test rig is designed and built.Measurements for endwall film cooling performance are done using pressure sensitive paint.The effects of the non-axisymmetric endwall on film cooling performance at different locations on the endwall are studied and documented.It is found that the change in film cooling performance by non-axisymmetric endwall is dominated by the change in acceleration features.In a more engine representative contoured endwall test rig which is equipped with a simulated combustor outlet flow field and combustor cooling flow,the passage flow field and endwall film cooling performance were measured with two-row staggered film cooling hole configuration and one-row film cooling hole configuration.Different secondary flow structures and film cooling effectiveness distribution characteristics were found in the current passage from those in the conventional passages.The interactions between the film cooling injection and the mainstream vortex,such as the passage vortex,impingement vortex and momentum gradient vortex,are documented with different mass flow ratios and different film cooling configurations.Better understandings of the aero-thermal aspects of contoured endwall in gas turbine NGV cascade are achieved in this thesis which is a foundation to achieve both better aerodynamic and cooling performance.