Failure Analysis of Turbine Disc of an Aero Engine

Topics: Turbine, Fluid dynamics, Gas turbine Pages: 32 (8608 words) Published: February 25, 2013
Available online at

International Journal of Heat and Mass Transfer 51 (2008) 3066–3081

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

1. Introduction
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*

Corresponding author.
E-mail address: (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 [3] 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. [4] 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...

References: Part I: Time averaged results, ASME J. Turbomach. 108 (1) (1986)
turbine. Part II: Description of analysis technique and typical timeresolved measurements, ASME J. Turbomach. 108 (1) (1986) 98–107.
blade, ASME J. Eng. Gas Turbines Power 107 (4) (1985) 991–997.
ASME J. Eng. Gas Turbines Power 107 (4) (1985) 1016–1021.
blade with air and CO2 film injection, International J. Heat Mass
Transfer 37 (10) (1994) 2707–2714.
Turbomach. 116 (4) (1994) 730–737.
turbine blade, ASME J. Turbomach. 120 (4) (1998) 808–817.
behavior for a transonic turbine vane, Numer. Heat Transfer Part A
49 (2006) 237–256.
Numer. Heat Transfer Part A 32 (4) (1997) 347–355.
Heat Transfer 20 (3) (2006) 558–568.
Continue Reading

Please join StudyMode to read the full document

You May Also Find These Documents Helpful

  • How Gas Turbine Engines Work Essay
  • Reaction Turbine Essay
  • Tesla Turbine Essay
  • Essay on engines
  • Thermodynamic Analysis and Performance Characteristics of a Turbofan Jet Engine Essay
  • Steam Turbine Essay
  • Hydraulic Turbines Essay
  • Analysis on Big It Project Failure Essay

Become a StudyMode Member

Sign Up - It's Free