Flow and Thermal Field Measurements in a Combustor Simulator Relevant to a Gas Turbine Aero-Engine

Abstract

The highly competitive gas turbine industry has been motivated by consumer demands for higher power-to-weight ratios, increased thermal efficiencies, and reliability while maintaining affordability. In its continual quest, the industry must continually try to raise the turbine inlet temperature, which according to the well-known Brayton cycle is key to higher engine efficiencies. The desire for increased turbine inlet temperatures creates an extremely harsh environment for the combustor liner in addition to the components downstream of the combustor. Shear layers between the dilution jets and the mainstream, as well as combustor liner film-cooling interactions create a complex mean flow field within the combustor, which is not easy to model. A completely uniform temperature and velocity profile at the combustor exit is desirable from the standpoint of reducing the secondary flows in the turbine. However, this seldom occurs due to a lack of thorough mixing within the combustor. Poor mixing results in non-uniformities, such as hot streaks, and allow non-combusted fuel to exit the combustor.This investigation developed a database documenting the thermal and flow characteristics within a combustor simulator representative of the flowfield within a gas turbine aero-engine. Three-and two-component laser Doppler velocimeter measurements were completed to quantify the flow and turbulence fields, while a thermocouple rake was used to quantify the thermal fields.The measured results show very high turbulence levels due to the dilution flow injection. Directly downstream of the dilution jets, an increased thickness in the filmcooling was noted with a fairly non-homogeneous temperature field across the combustor width. A highly turbulent shear layer was found at the leading edge of the dilution jets. Measurements also showed that a relatively extensive recirculation region existed downstream of the dilution jets. Despite the lack of film-cooling injection at the trailing edge of the dilution hole, there existed coolant flow indicative of a horse-shoe vortex wrapping around the jet. As a result of the dilution jet interaction with the mainstream flow, kidney-shaped thermal fields and counter-rotating vortices developed. These vortices serve to enhance combustor mixing.iii

A Detailed Study of Effects of Row Trenched Holes at the Combustor Exit on Film Cooling Effectiveness

“…For plane 1p, the deviation was equal to 9.76% compared to experimental measurements (Vakil et al 2005) and 8.34% compared to CFD results (Stitzel et al 2004). For observation plane of 2p, the deviation was 13.36% and 11.96% respectively (Figure 3).…”

“…The findings of the current research were compared with the experimental collected results and numerical findings gathered by other researchers (Vakil et al 2005, Stitzel et al 2004. The effectiveness of film-cooling was compared in plane 1p and 2p at y/W=0.4.…”

“…In the present research, the combustor simulator contained two arrangements of cooling holes. For the verification of findings, the first arrangement (baseline or case 1) was designed similar to the combustor modelled by (Vakil et al 2005). The length of these cooling holes was 2.5cm and they drilled at an angle of 30deg from the horizontal surface.…”

“…Researchers presented experimental results of the study of temperature distribution inside a combustor simulator (Vakil et al 2005). In this study, a real large scale of combustor was modelled.…”

“…This approach of cooling holes can also be utilized by engine designers at the fore side of the turbine vanes. Moreover, in order to measure the validity of the results, a comparison between the data gained from this study and the other project was carried out (Vakil et al 2005, Stitzel et al 2004). …”