Session: 13-04: Heat transfer modelling

Paper Number: 152700

Potential of a Nozzle Guide Vane Cooling Concept Adapted to Circumferential Temperature Variations 

The turbine inlet temperature of jet engines is constantly being increased to improve thermal efficiency. In addition to the development of high-temperature materials, these temperature increases are possible through the use of active cooling using compressed air derived from the compressor. However, excessive cooling air consumption has a negative impact on overall efficiency. It is therefore important to develop efficient cooling systems for turbine vanes and blades that can withstand higher turbine inlet temperatures while minimizing cooling air consumption. The turbine inlet temperature is also not constant at the turbine inlet, but varies in both time and space. In particular, can-annular combustors exhibit significant circumferential temperature variations at the combustor outlet, with peak temperatures occurring in the center of each can. This spatial temperature distribution is typically characterized by a combustor pattern factor. Currently, nozzle guide vane (NGV) cooling configurations are designed based on the peak gas temperature occurring within the hot streaks at the center of the cans. This leads to excessive cooling in regions downstream of lower temperature zones, thereby reducing turbine efficiency.

The objective of this study is to establish a correlation between the turbine inlet temperature profile and potential cooling air savings by using a circumferentially alternating NGV cooling configuration. A semi-empirical cooling design tool is used to develop a state-of-the-art cooling configuration for a high-pressure turbine NGV as a baseline. For several variations of the combustor outlet profile, the cooling configuration is tailored to the lower thermal load in the area between two adjacent hot streaks, resulting in several designs in the circumferential direction. Selected configurations are further evaluated for mixing losses due to film cooling using computational fluid dynamics simulations and a reduced-order film cooling model.

The findings suggest that a significant reduction in cooling air requirements can be achieved, which also reduces the mixing losses due to the decreased interaction of film cooling jets and the main gas flow. These results allow a designer to identify potential improvements and weigh them against the cost of the increased number of cooling designs required.

Presenting Author: Robin Schöffler German Aerospace Center (DLR)

Presenting Author Biography: I studied mechanical engineering at Baden-Wuerttemberg Cooperative State University (DHBW) from 2014 to 2017. Afterwards I continued studying mechanical engineering at the Ruhr-University Bochum (RUB), where I specialized in turbomachinery. After completing my Master of Science degree in 2019, I started my research work at the German Aerospace Center (DLR) in the area of turbine cooling. My work focuses on the design of cooled turbine vanes and blades and the development of the tools used for this purpose.

Authors:

Robin Schöffler German Aerospace Center (DLR)
Michael Müller German Aerospace Center (DLR)
Clemens Grunwitz German Aerospace Center (DLR)
Robin G. Brakmann German Aerospace Center (DLR)
Robert Krewinkel Graz University of Technology