Analysis of Gas Turbine Blade Cooling System

Gas turbine engines are designed to continuously and efficiently convert the energy of
fuel into useful power. Gas turbines have been developed into very reliable, high
performance engines (high ratio of power output to weight, high efficiency and low
maintenance costs).
As the turbine inlet temperature increases, the heat transferred to the turbine blade also
increases. The operating temperatures are far above the permissible metal temperatures.
Therefore, there is a critical need to cool the blades for safe operation. The cooling
methods currently implemented in the turbine industry can be classified into two types:
internal cooling and external cooling.
In the present work the internal cooling of a gas turbine blade is analyzed. The blade has
a rectangular 9mm x 18mm compressed air channel along the blade span. FiniteDifference
method is used to predict temperature distribution for blade cross section at
different heights from the root. Effect of compressed air mass flow rate, inlet temperature
and the temperature of combustion gasses have been considered. The investigations are
carried out for both smooth and two opposite ribbed-walls channels. The results are
presented and discussed as temperature distribution in various sections of the blade and
the comparison between ribbed and smooth channel based on Nu values. Also, various
ribs configurations have been considered in the analysis. Results at rib angles, a of 90 °,
60', 45 ° and 30' and ribs blockage ratios, e!D, ranging from 0.042 to 0.078 are
compared in terms of Nu and friction factor, f It is found that maximum Nu number
occurs when 60° ribs are introduced in the channel. An enhancement of 149.45% is
achieved with penalty of increase in the friction factor by 114.5%.