TY - GEN
T1 - Contoured endwall flow and heat transfer experiments with combustor coolant and gap leakage flows for a turbine nozzle guide vane
AU - Saxena, Reema
AU - Alqefl, Mahmood H.
AU - Liu, Zhao
AU - Moon, Hee Koo
AU - Zhang, Luzeng
AU - Simon, Terrence W.
PY - 2016/1/1
Y1 - 2016/1/1
N2 - Flow in a high pressure gas turbine passage is complex, involving systems of secondary vortex flows and strong transverse pressure gradients. This complexity causes difficulty in providing film cooling coverage to the hub endwall region, which is subjected to high thermal loading due to combustor exit hot core gases. Therefore, an improved understanding of these flow features and their effects on endwall film cooling is needed to assist designers in developing efficient cooling schemes. The experimental study presented in this paper is performed on a linear, stationary, two-passage cascade representing the first stage nozzle guide vane of a high-pressure gas turbine. The sources of film cooling flows are the upstream combustor liner coolant and the leakage flow from the combustor-nozzle guide vane interfacial gap. Measurements are performed on an axisymmetrically-contoured endwall passage under conditions of various leakage mass flow rates to mainstream flow ratios (MFR= 0.5%, 1.0%, 1.5%). Flow migration and mixing are documented by measuring passage thermal fields and adiabatic effectiveness values over the endwall. It is found that, compared to our previous studies with a rotor inlet leakage slot geometry, the thin slot geometry of the nozzle leakage path gives a more uniform coolant spread over the endwall with significant coverage reaching the downstream and pressure-side regions of the passage. Interestingly, the coverage is seen to be only weakly dependent on the leakage mass low ratio and even reduce slightly with an increase in mass flow ratio above 1%, as indicated by lowered endwall adiabatic effectiveness values.
AB - Flow in a high pressure gas turbine passage is complex, involving systems of secondary vortex flows and strong transverse pressure gradients. This complexity causes difficulty in providing film cooling coverage to the hub endwall region, which is subjected to high thermal loading due to combustor exit hot core gases. Therefore, an improved understanding of these flow features and their effects on endwall film cooling is needed to assist designers in developing efficient cooling schemes. The experimental study presented in this paper is performed on a linear, stationary, two-passage cascade representing the first stage nozzle guide vane of a high-pressure gas turbine. The sources of film cooling flows are the upstream combustor liner coolant and the leakage flow from the combustor-nozzle guide vane interfacial gap. Measurements are performed on an axisymmetrically-contoured endwall passage under conditions of various leakage mass flow rates to mainstream flow ratios (MFR= 0.5%, 1.0%, 1.5%). Flow migration and mixing are documented by measuring passage thermal fields and adiabatic effectiveness values over the endwall. It is found that, compared to our previous studies with a rotor inlet leakage slot geometry, the thin slot geometry of the nozzle leakage path gives a more uniform coolant spread over the endwall with significant coverage reaching the downstream and pressure-side regions of the passage. Interestingly, the coverage is seen to be only weakly dependent on the leakage mass low ratio and even reduce slightly with an increase in mass flow ratio above 1%, as indicated by lowered endwall adiabatic effectiveness values.
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U2 - 10.1115/GT2016-56675
DO - 10.1115/GT2016-56675
M3 - Conference contribution
AN - SCOPUS:84991720600
T3 - Proceedings of the ASME Turbo Expo
BT - Heat Transfer
PB - American Society of Mechanical Engineers (ASME)
T2 - ASME Turbo Expo 2016: Turbomachinery Technical Conference and Exposition, GT 2016
Y2 - 13 June 2016 through 17 June 2016
ER -