IJRET: International Journal of Research in Engineering and Technology eISSN: 2319-1163 | pISSN: 2321-7308 __________________________________________________________________________________________ Volume: 03 Issue: 03 | Mar-2014, Available @ http://www.ijret.org 603 THERMAL ANALYSIS OF COOLING EFFECT ON GAS TURBINE BLADE Amjed Ahmed Jasim AL-Luhaibi 1 , Mohammad Tariq 2 1 Technical collage Kirkuk, Fuel and Energy Engineering Department, Foundation of Technical Education, Iraq 2 Assistant Professor, Department of Mechanical Engineering, SSET, SHIATS-DU, Allahabad Abstract Performance of a gas turbine is mainly depends on various parameters e.g. ambient temperature, compressor pressure ratio, turbine inlet temperature etc. The most important parameter to increase the life of the turbine blade is the cooling of the blade, which is necessary after reaching a certain temperature of the gases passing through the blades. Various types of cooling models are available for a turbine blade cooling. The power output of a gas turbine depends on the mass flow rate through it. This is precisely the reason why on hot days, when air is less dense, power output falls off. This paper is to analyze the film cooling technique that was developed to cool gases in the initial stages of the turbine blades, where temperature is very high (>1122 K). It is found that the thermal efficiency of a cooled gas turbine is less as compare to the uncooled gas turbine for the same input conditions. The reason is that the temperature at the inlet of the turbine is decreased due to cooling and the work produced by the turbine is slightly decreased. It is also found that the power consumption of the cool inlet air is of considerable concern since it decreases the net power output of gas turbine. In addition, net power decreases on increasing the overall pressure ratio. Furthermore, the reviewed works revealed that the efficiency of the cooled gas turbine largely depends on the inlet temperature of the turbine and previous research said that the temperature above 1123K, require cooling of the blade. Keywords: Gas turbine, Turbine blade cooling, film cooling technique, Thermal Efficiency ----------------------------------------------------------------------***-------------------------------------------------------------------- 1. INTRODUCTION In a bid to remain at the forefront of technological development as well as a technical expert to United States industry, NASA identified the need for an improved design process within the civilian aero engine industry, in hopes of improving their market share, reducing time to market, and minimizing research costs [2]. Areas of interest included, but were not limited to, high temperature materials, advancing turbine analysis techniques, and improving the overall engine design and analysis process. The latter interest called for the impact assessment of engine component technologies from the micro to system levels [2]. A good example of the need for this type of analysis comes from determining the required service life of a turbine blade, which is limited by the exit temperature from the combustor and the material properties that in turn, limits the performance of the gas turbine. Ideally, the engine would operate at a high enough temperature to achieve the highest possible thrust rating [3], while at the same time maintaining an economic service life. Currently, to address the issue of exit temperature, modern turbines utilize a cooled turbine blade to improve the possible rotor inlet temperature, and this necessary cooling flow has a strong impact on the turbine efficiency [4, 5]. By improving cooling technology for a gas turbine blade it is possible to increase the combustor exit temperature sufficiently, therefore achieving good improvement in turbine efficiency and thrust [6]. However, this improvement does not prove viable when considering the complete process. The increase in cooling flow to the turbine blades and vanes takes bleed air away from the compressor, reducing its on efficiency. This detrimentally affects the efficiency of the whole system, such that the improvements in the turbine are eclipsed. Being able to track all these whilst looking at a particular component within an aircraft gas turbine, is obviously very desirable, especially at the early stages of a design. A successful design process incorporates flexibility and freedom at the early conceptual stages and continuing as far into the design as possible, saving time and money in fixing problems that would have arisen had an investigation not taken place. Looking into technologies to improve an engine, one needs to provide useful benchmarks from which comparisons can be made. If a superior product is going to be produced, analysis of the areas affected by the new product needs to be considered. If this is not the case a lot of money could be invested in designs that in the end prove to be impractical. A new technique has been reviewed up to date that was developed to cool inlet air to gas turbine [9]. The techniques including the mechanical chillers, media type evaporative coolers and absorption chillers have been reviewed. It is found that the power consumption of the cool inlet air is of considerable concern since it decreases the net power output of gas turbine. Experimental tests to investigate the film cooling performance of converging slot hole (console) rows on the turbine blade have been performed [10]. Film cooling effectiveness of each single hole row is measured