Journal of Thermal Science Vol.24, No.1 (2015) 4957
Received: September 2014
Wan Aizon W Ghopa: PhD.
www.springerlink.com
DOI: 10.1007/s11630-015-0755-5 Article ID: 1003-2169(2015)01-0049-09
Aero-Thermal Performances of Leakage Flows Injection from the Endwall Slot
in Linear Cascade of High-Pressure Turbine
Wan Aizon W Ghopa
1
, Zambri Harun
1
, Ken-ichi Funazaki
2
, Takemitsu Miura
2
1. Department of Mechanical and Materials Engineering, Universiti Kebangsaan Malaysia, 43600 Bangi, Malaysia
2. Department of Mechanical Engineering, Iwate University, Morioka 020-8551 Iwate, Japan
© Science Press and Institute of Engineering Thermophysics, CAS and Springer-Verlag Berlin Heidelberg 2015
The existence of a gap between combustor and turbine endwall in the real gas turbine induces to the leakages
phenomenon. However, the leakages could be used as a coolant to protect the endwall surfaces from the hot gas
since it could not be completely prevented. Thus, present study investigated the potential of leakage flows as a
function of film cooling. In present study, the flow field at the downstream of high-pressure turbine blade has
been investigated by 5-holes pitot tube. This is to reveal the aerodynamic performances under the influenced of
leakage flows while the temperature measurement was conducted by thermochromic liquid crystal (TLC). Expe-
rimental has significantly captured theaerodynamics effect of leakage flows near the blade downstream. Further-
more, TLC measurement illustrated that the film cooling effectiveness contours were strongly influenced by the
secondary flows behavior on the endwall region. Aero-thermal results were validated by the numerical simulation
adopted by commercial software, ANSYS CFX 13. Both experimental and numerical simulation indicated almost
similar trendinaero and also thermal behavior as the amount of leakage flows increases.
Keywords: endwall filmcooling, leakage flows,secondary flows,aerodynamic loss, film cooling effectiveness
Introduction
The thermal efficiency of gas turbine need to be in-
creased in order to reduce the fuel burnt. This can be rea-
lized by increasing the turbine inlet temperature. Modern
gas turbine are designed to operate at turbine inlet tem-
peratures in excess of 1600°C, placing high thermal loads
on blade components. Thus, without cooling, the lifetime
of components may be shortened due to thermal stresses.
One of the critical regions that require special thermal
protection intention is the endwall. This region is consid-
erably difficult to be cooled due to complex secondary
flow structure that occurs at the blade passage which
depend on the blade profile. The endwall flow structure
has been revealed by Takeshi et. al [1]. The main flow
structure consists of pressure side and suction side leg
horse-shoe vortex, cross flow, corner vortex and passage
vortex. The earliest study that relates the endwall flow
structure and film cooling has been made by Blair [2].
The work clarifies that, the horse-shoe vortex and pas-
sage vortex hasa dominant impact on the heat transfer of
film-cooled endwall. Graziani [3] documented the en-
hancement of heat transfer on the suction surface of the
bladedue to the passage vortex. Due to the existence of
complex flow structures in this region, detail studies are
required in order to improve the cooling performance.
Several studies have presented the flow field data within
the vane stagnation plane illustrating the formation and
dynamics of the leading edge horse-shoe vortex which
include the works of Goldstein and Spores [4]. Small but