Numerical simulation on ﬁlm cooling with compound angle of blade leading edge model for gas turbine Wen-jing Gao, Zhu-feng Yue ⇑ , Lei Li ⇑ , Zhe-nan Zhao, Fu-juan Tong School of Mechanics, Civil Engineering and Architecture, Northwestern Polytechnical University, Xi’an 710072, China article info Article history: Received 8 November 2016 Received in revised form 28 June 2017 Accepted 24 July 2017 Keywords: Turbine blade Leading edge Film cooling Numerical simulations Compound angle abstract Film cooling performances of the cylindrical ﬁlm cooling holes with different compound angles on the turbine blade leading edge model are investigated in this paper. Several numerical simulation results are compared with available experimental data, under different blowing ratios. Three rows of holes are arranged in a semi-cylinder model which is used to model the blade leading edge. These three rows of holes have a compound angle of 90° in the ﬂow direction, 30° along the spanwise direction. Besides, the two rows on either side of the stagnation row have an additional angle in the transverse direction. Five different ﬁlm cooling hole compound angles in the transverse direction and four different blowing ratios are studied in detail. The results show that as the blowing ratio increases, the trajectory of the ﬁlm jets in the leading edge region deviates gradually from the mainstream direction to the spanwise direc- tion, for all cases studied. And ﬁlm cooling effectiveness increases with the increasing blowing ratio while a slight decrease appears as the blowing ratio approaches 2.0. In this study, the optimal value of M is around 1.4. For the Baseline Case, the overall averaged cooling effectiveness increases by more than 0.1, compared with M = 0.7. The holes with negative additional compound angle have better performance of cooling. On the one hand, the improvement of ﬁlm cooling effectiveness increases with the increasing negative compound angle, before it reaches -30°. On the other hand, with the increasing blowing ratio, the improvement of the cooling performance due to negative additional compound angle is more signif- icant. For c = 30°, the increase of overall averaged cooling effectiveness varies from 1.75% to almost 20%, with the increase of M. Ó 2017 Elsevier Ltd. All rights reserved. 1. Introduction As the needs arises for higher overall efﬁciency and higher speciﬁc power output, modern gas turbine systems are required to operate at higher turbine inlet temperature, which has already been far beyond the material acceptable level. As a result, more effective cooling schemes must be applied in the turbine vane or blade to protect it from thermal stresses. Film cooling is one of the major cooling methods. In practice, relatively cool air from the compressor stages is injected through holes in the walls of hol- low turbine airfoils in order to protect the metal surface from the hot mainstream. Film cooling is applied to nearly all of the external surfaces associated with the airfoils that are exposed to the hot combustion gasses such as the leading edges, main bodies, blade tips, and endwalls. And there have been several studies around it. In Zhang et al. [1] and Li et al. [2], experiments were conducted to consider the effects of leading edge airfoil geometry and ﬁlm cooling arrangements on endwall ﬁlm cooling, Becchi et al. [3], studied the ﬁlm cooling adiabatic effectiveness measurements of pressure side trailing edge cooling conﬁgurations. Al-Zurﬁ and Turan [4] investigated the effect of rotation on ﬁlm cooling effec- tiveness and heat transfer coefﬁcient distribution on the suction and pressure sides of a gas turbine blade. Ke and Wang [5] further presented a numerical investigation of pulsed ﬁlm cooling on a modiﬁed NASA C3X, with ﬁve rows of cooling hole at leading edge, pressure side and suction side. More information on ﬁlm cooling can be found in Bogard and Thole [6]. However, in a gas turbine, the leading edge of a turbine airfoil often withstands the highest thermal load since it is upwind to high temperature inﬂows, and the airfoil failure often occurs in this region due to material dam- age. According to the study of Bogard and Thole [6], ﬁlm cooling on vanes and blades generally involves a dense array of coolant holes around the leading edge, referred to as a showerhead. And ﬁlm cooling yet has very difﬁcult phenomena to predict. The ﬂow environment around the leading edge is extremely complex with a stagnation mainstream, the curvature of surface, strong pressure gradients and turbulence, as well as interaction http://dx.doi.org/10.1016/j.ijheatmasstransfer.2017.07.105 0017-9310/Ó 2017 Elsevier Ltd. All rights reserved. ⇑ Corresponding authors. E-mail addresses: zfyue@nwpu.edu.cn (Z.-f. Yue), lileinpu@nwpu.edu.cn (L. Li). International Journal of Heat and Mass Transfer 115 (2017) 839–855 Contents lists available at ScienceDirect International Journal of Heat and Mass Transfer journal homepage: www.elsevier.com/locate/ijhmt