An increased demand for lower emission of stationary gas turbines as well as civil aircraft engines has led to new, low emission combustor designs with less liner cooling and a ﬂattened temperature proﬁle at the outlet. As a consequence, the heat load on the endwall of the ﬁrst nozzle guide vane is increased. The secondary ﬂow ﬁeld dominates the endwall heat transfer, which also contributes to aerodynamic losses. A promising approach to reduce these losses is non-axisymmetric endwall contouring. The effects of non-axisymmetric endwall contouring on heat transfer and ﬁlm cooling are yet to be investigated. Therefore, a new cascade test rig has been set up in order to investigate endwall heat transfer and ﬁlm cooling on both a ﬂat and a non-axisymmetric contoured endwall. Aerodynamic measurements that have been made prior to the upcoming heat transfer investigation are shown. Periodicity and detailed vane Mach number distributions ranging from 0 to 50% span together with the static pressure distribution on the endwall give detailed information about the aerodynamic behavior and inﬂuence of the endwall contouring. The aerodynamic study is backed by an oil paint study, which reveals qualitative information on the effect of the contouring on the endwall ﬂow ﬁeld. Results show that the contouring has a pronounced effect on vane and endwall pressure distribution and on the endwall ﬂow ﬁeld. The local increase and decrease of velocity and the reduced blade loading towards the endwall is the expected behavior of the 3d contouring. So are the results of the oil paint visualization, which show a strong change of ﬂow ﬁeld in the leading edge region as well as that the contouring delays the horse shoe vortex hitting the suction side.
A New Test Facility to Investigate Film Cooling on a Non-Axisymmetric Contoured Turbine Endwall, Part I: Introduction and Aerodynamic
ASME Turbo Expo 2015
15.06.2015 - 19.06.2015
Paper No.: GT2015-42272