Numerical study on the influence of coolant temperature, pressure, and thermal barrier coating thickness on heat transfer in high-pressure blades

• Published in: International journal of heat and mass transfer Vol. 189; p. 122715
• Main Authors: Vo, Duy-Tan, Mai, Thanh Dam, Kim, Byungwook, Ryu, Jaiyoung
• Format: Journal Article
• Language: English
• Published: Oxford Elsevier Ltd 15.06.2022; Elsevier BV
• Subjects: Actual high-pressure blade; CFD; Compressibility effects; Conjugate heat transfer; Coolants; Cooling; Cooling rate; Film cooling; Gas pressure; Gas turbine; Heat transfer; Heat transfer coefficients; Mathematical analysis; Overheating; Poly-hexcore mesh; Pressure ratio; Temperature; Temperature ratio; Thermal barrier coating; Thermal barrier coatings; Thermal stress; Thickness; Topology; Turbine blades; CFD; Conjugate heat transfer; Gas turbine; Poly-hexcore mesh; Actual high-pressure blade; Thermal barrier coating
• ISSN: 0017-9310; 1879-2189
• DOI: 10.1016/j.ijheatmasstransfer.2022.122715

•Influence of various factors on the actual gas turbine blade was investigated.•High thermal stress is at platform, leading edge, and leading and trailing tip.•Low-pressure leads to lack of supplying coolant for trailing passage, while at high-pressure, film cooling lift-off at leading tip reduces cooling performance.•Nonlinear compressible air effects make higher-temperature coolants less effective.•Thermal barrier coating efficiently reduces heat transfer to the blade surface. Several studies have focused on investigating the heat transfer in blades; however, these studies have primarily adopted numerical approaches, and the heat transfer in an actual turbine blade remains unclear thus far, owing to the complex topology of the blade. To address this gap in literature, this study investigates the influences of the thermal barrier coating (TBC) thickness and coolant on the mainstream hot gas pressure ratio (pc/pt) and temperature ratio (Tc/Tg) in blade film cooling, respectively, by simulating the conjugate heat transfer (CHT) in the first-stage cooling blade of an W501F engine. A complete blade design is used to obtain a realistic heat transfer profile. The results demonstrate equivalent high-thermal-stress regions between the simulation and industrial patterns. The leading tip (LT), trailing tip (TT), platform, and leading edge (LE) regions have higher thermal stresses than the other regions. Coolants with higher temperature are less effective in cooling the blade due to the nonlinear effects of compressible air. Although the coolant reduces the temperature across the blade with similar efficacy, the temperature variation is still significant; a decrease of 100 K in the coolant temperature results in a reduction of 58 K in the average blade temperature. At a fixed coolant temperature and standard pressure ratio pc/pt (approximately 1.07), the coolant film uniformly envelopes the entire LT surface. However, for the same blade design, the trailing passage lacks a coolant film at this pressure ratio. As the pressure ratio exceeds 1.50, the coolant flows to all regions in the trailing passage cooling the TT at an enhanced rate. However, as the ratio continues to increase, liftoff jet occurs and reduces the cooling performance at the LT, resulting in overheating in this region. The application of a 0.8 mm-thick TBC layer on the blade surface reduces the heat transfer coefficient by 35%, thus decreasing the heat transfer.