A number of studies have shown that the flow field exiting a combustor of a gas turbine cycle is highly non-uniform in pressure, velocity and, most importantly, temperature. Much research has been dedicated to the cooling of gas turbine blades using internal, film cooling, impingement jets, and pin/fin cooling technologies. Such designs allow for heated blades to be cooled from the inside out. While advancements in this type of blade cooling technology provide effective means to reduce the occurrence of blade failure due to material overheat conditions, the effect of externally reducing or eliminating the temperature non-uniformities in the exit flow from the combustor would assist in the solution. The goal of this study is to optimize the mixing of primary and dilution air in the dilution zone of the combustor using guide vanes. This improvement in mixing would lead to increase in the degree of temperature uniformity with respect to the radial position at the exit nozzle. To achieve this objective, both experimental and computational studies were performed to investigate the heat and flow behaviors with 45° spherically swept guide vanes attached to the dilution holes. These guide vanes were intended to direct the dilution jets into the primary flow and enhance mixing. A parameter was defined in terms of the temperatures of the dilution and primary flow streams at the inlet and the exit plane, called the mixture fraction. Based on the mixture fraction value, it was found that the guide vanes produce a more uniform exit temperature flow field as compared to the case when there were no guide vanes used. Also, the design was modified for different alignment orientations of the guide vanes — 0°, 30°, 60° and 90° with respect to the primary flow — with the 60° orientation fostering the best results.
Experimental measurements of the flow field around a film-cooled blade leading edge using particle image velocimetry
