Combustor and turbine aerothermal interactions in gas turbines with can combustors

Thesis

As the research into the improvement of gas turbine performance progresses, the combustor- turbine interface becomes of increasing importance. In new engine designs components come closer together and the study of the combustor and turbine interactions can prove to be valuable for the improvement of the aerothermal performance of the vane.

This thesis presents an experimental and numerical investigation of the aerodynamic and heat transfer aspect of the interactions between the combustor and the nozzle guide vane. In the gas turbine studied the trailing edge of the combustor transition duct wall is found upstream of every second vane. In the experimental measurements carried out in a purpose-built high speed experimental facility, the wake of this wall is shown to increase the aerodynamic loss of the vane. On the other hand, the wall alters secondary flow structures and has a protective effect on the heat transfer in the leading edge-endwall junction, a region that has proven to be detrimental to component life. The effect of different clocking positions of the vane relative to the combustor wall are tested experimentally and shown to alter the aerodynamic field and the heat transfer to the vane. The experimental methods and processing techniques adopted in this work are utilized to highlight the differences between the different cases studied.

A new concept of using the combustor wall to shield the nozzle guide vane leading edge is introduced, followed by a proposed design that is numerically analysed, including a new cooling system. This uses continuous cooling slots on the upstream combustor wall to cool the vane leading edge. Coolant to the endwalls is provided from continuous slots on the combustor-turbine interface. The reduction of secondary flow through the removal of the horseshoe vortex in the new design results in improved cooling of the endwalls, with a higher average adiabatic effectiveness than in the original case, using the same coolant mass flow rate. The vane surface and suction side are also successfully cooled using less air than that required for a showerhead.

The new vane is tested in the experimental facility. The improved aerodynamic and thermal performance of the shielded vane is demonstrated under engine-representative inlet conditions. The new design is shown to have a lower average total pressure loss than the original vane for all inlet conditions. The heat transfer on the vane surface is overall reduced for all inlet conditions and the peak heat transfer on the vane leading edge-endwall junction is moved further upstream, to a region that can be effectively cooled from the upstream cooling slots on the combustor wall trailing edge and the endwalls.

Authors: Aslanidou, I

• Type of award: DPhil
• Level of award: Doctoral
• Awarding institution: University of Oxford
• Language: English
• Keywords: secondary flow; experimental; heat transfer; aerodynamics; numerical; nozzle guide vane
• Subjects: Gas-turbines; Aerodynamics; Heat Transfer