International Journal of Heat and Mass Transfer 51 (2008) 3066–3081 www.elsevier.com/locate/ijhmt
Numerical study of ﬁlm cooled rotor leading edge with tip
clearance in 1-1/2 turbine stage
Huitao Yang a, Hamn-Ching Chen a, Je-Chin Han a,*, Hee-Koo Moon b b
Texas A&M University, College Station, TX 77843, United States Solar Turbines Incorporated, 2200 Paciﬁc Highway, San Diego, CA 92101, United States
Received 28 September 2006
Available online 14 November 2007
Numerical simulations were performed to predict the ﬁlm cooling eﬀectiveness and the associated heat transfer coeﬃcient in a 1-1/2 turbine stage. The leading edge of the rotor blade is ﬁlm cooled with three rows of ﬁlm cooling holes. The rotor tip leakage has been investigated for a clearance of 0.8% of blade span. Sliding grid is used for the rotor domain, and interface technique is employed to exchange information between stator and rotor domains. Simulations were carried out for both design and oﬀ-design conditions to investigate the eﬀects of the stator–rotor interaction on the ﬁlm cooling characteristics. The commercial code FLUENT with Reynolds stress model is used in the prediction. It is found that the tilted stagnation line on the rotor leading edge moves from the pressure side to the suction side, and the instantaneous coolant streamlines shift from the suction side to the pressure side with the increasing rotating speed. For the ﬁxed inlet/outlet pressure ratio of turbine stage, the high rpm reduces the heat transfer coeﬃcient on the rotor due to the low rotor relative velocity, and increases the ‘‘sweet spot” on the rotor tip. These trends are well supported by the experimental results. Ó 2007 Elsevier Ltd. All rights reserved.
Keywords: Gas turbine; Film cooling; Rotating blade; Leading edge; CFD
The inlet temperature of a modern turbine has continually increased to achieve high thrust power and high thermal eﬃciency. A turbine stage consists of one row of nozzle guide vanes (stators) and one row of rotating blades
(rotors). Due to the stagnation ﬂow, the leading edge
region of the rotors experiences a high heat load, which
may result in material failure because of the high temperature and thermal stress. To reduce the heat load, several advanced internal cooling technologies have been applied
in this region, including impingement cooling on the inner
wall of the leading edge with diﬀerent kinds of turbulators, such as pins or ribs, to enhance internal cooling eﬀects.
Besides these internal cooling technologies, ﬁlm cooling,
ejecting coolant to the outer surface of the blade, is com*
E-mail address: email@example.com (J.-C. Han).
0017-9310/$ - see front matter Ó 2007 Elsevier Ltd. All rights reserved. doi:10.1016/j.ijheatmasstransfer.2007.09.025
monly used in the leading edge region in the real engine.
Although the ﬁlm cooling slightly increases the heat transfer coeﬃcient, the driving temperature diﬀerence (Taw À Tw) is reduced signiﬁcantly. Therefore, the overall heat load decreases.
Several experiments have been carried out to study the
heat transfer on the leading edge of the rotors in turbine
stages. Dunn et al. [1,2] used a full stage rotating turbine of the Garrett TFE 731-2 engine in a shock-tunnel facility
and thin-ﬁlm heat ﬂux gages to study heat transfer on the vane, end walls and rotors. Abhari and Epstein  studied
the time-resolved heat transfer for cooled and un-cooled
rotors by thin heat ﬂux gages. They found that the heat
transfer was highly unsteady for rotors in a transonic turbine. Takeishi et al.  employed the CO2 mass transfer analogy technique to measure the local ﬁlm cooling eﬀectiveness on a rotor blade, and found higher cooling eﬀectiveness on the suction side compared to the pressure side of the blade. They believed that this phenomenon is caused
H. Yang et al. / International...