Recently, film cooling has been continuously studied to increase the efficiency of gas turbines. A turbine inlet temperature increase occurs as a way to improve the efficiency. However, it is essential to improve the cooling performance of the blade surface because of the melting point of the part. In this paper, a side hole shape wherein a general cylinder hole and two auxiliary holes are combined, is proposed to improve the film cooling efficiency, and the blowing ratio was set to 0.4, 0.8, 1.2, and 2.0. When side hole was applied, the vortex interference at the hole entrance occurred less than that of the cylinder hole. That is, the flow rate of the coolant adsorbed to the surface increased to improve the cooling performance. In conclusion, compared to the cylinder hole, the cooling efficiency of the shape to which the side hole was applied was excellent, and in particular, the average area cooling efficiency with spanwisely designed side holes improved by 83%.

In the gas turbine, the clearance between the blade tip of the rotor and the inside of the stationary casing varies depending on the rotation of the rotor and the heat output of the combustor. Accordingly, the assembly clearance is determined, and the leakage of the gas occurs because of the gap during operation, affecting the efficiency of the system. Thus, designers use a variety of techniques to optimize this clearance, a typical method that reduces the relative variation of the clearance using heating and cooling mechanisms. In this study, we developed a method to control the blade tip clearance through the axial movement of the inclined blade without using heating and cooling mechanisms. Recently, we designed an advanced blade tip clearance control system that can control multi-step, not on-off control, to apply to large gas turbines developed by Doosan. The designed system is hydraulic and can be used with a maximum thrust of 100 tons, and the desired displacement can be moved in multiple stages as required. We have completed the reliability verification of the entire lifecycle level and applied it to the newly developed gas turbine.

The large gas turbine rotor used for power generation has a structural characteristic comprising a shaft, disk, and blade assembled to the disk. Because the start/stop is repeated, the tightening force may be reduced in the process of repeating the tightening force between the tie rod and the disk. When the tightening force falls below the threshold, changing the critical speed, increasing the vibration, or in extreme cases, the rotor may loosen and cause a major accident. Also, it is imperative to continuously maintain the proper tightening force because the thread of the tie rod is damaged when the tightening force exceeds the yield stress condition of the tie rod. In this paper, the gas turbine rotor system is modeled and simplified to identify the control variable of the tightening force of the tie rod bolts of the rotor. For verification, a simplified model of the gas turbine rotor was designed, manufactured, and verification tests were conducted to confirm the adequacy of the calculation method. As a result, the tightening force decreased as the stiffness of the pressing disk decreased, so the stiffness of the pressing disk should have a stiffness range similar to that of the tie rod.

Research on advanced cooling system design is significant in achieving a high turbine inlet temperature in the gas turbine industry. The higher turbine inlet temperature of the gas turbine increases thermal efficiency. However, it also aggravates the gas turbine deterioration, lifespan, and efficiency. In this study, a numerical model is developed for simulating the cooling performance of the gas turbine vane with the turbine inlet temperature of 1528 K. The impact of the coolant air flow rate and hole-shape were investigated. The expanded hole shape had better cooling performance than the general cylindrical shape, and showed higher cooling efficiency. We suggest that there is a relationship between the shape of the film cooling holes and the cooling air flow rate that achieves the desired cooling effectiveness.

The purpose of this study was to investigate the flow characteristics and cooling performance for the heavy turbine blade with different shapes. Research was focused on the numerical study on forced convective heat transfer coefficients for three different blades with base, tip, and hole. Thus, selected local locations for various temperature distributions were shown in the flow domain. Final temperature on the local surface of blades was compared with three different blades. According to the results of velocity and temperature distributions in the fluid domain, the blade with holes had the best convective cooling performance with higher 13-16% average heat transfer coefficient than the other two blades. Apparent vortex at the tip of tip and hole blade caused the stable temperature drop. According to the calculations of local convective heat transfer coefficient between blade surface and atmosphere in the blade, approximately 18% of heat transfer coefficient at hole was higher than the base blade and 7% at hole blade was higher than the base blade. Lowest cooling performance existed at the center position of all three blades.

Gas turbine blades are important parts of a power plant, and thus, it is necessary to be able to predict the low-cycle fatigue life of the blades. In this study, a low-cycle fatigue test of In738LC, which is used primarily in gas turbine blade manufacture, was performed at various high temperatures (750oC, 800oC, and 850oC). From the test results, the stressstrain curve and the stress-strain hysteresis loop were obtained. It was established that In738LC has no strain hardening or softening. The life prediction equations for low-cycle fatigue were derived using the Coffin-Manson equation and the energy model. In conclusion, one equation for predicting the life low-cycle fatigue was obtained using the energy level with temperature as the varying factor.

A gas turbine is a power plant unit that converts thermal energy into rotational energy by rotating a blade using hightemperature and high-pressure combustion gas. A gas turbine blade is directly exposed to a high-temperature flame. Various studies have aimed to improve the durability of the blade in harsh conditions. One proposes coating the blade with a thermal barrier to protect it from the flame, using a ceramic material with better thermal insulation. Another proposes using internal cooling, by creating an air flow path inside the blade to lower its temperature. Because both these techniques create a thermal gradient in the cross section of the blade, they amplify the difference in thermal expansion, thereby producing thermal stress in the blade and the thermal barrier coating. This study investigates the internal cooling effect on thermal gradient fatigue by using finite element analysis.