The present invention relates to gas turbine engines, and more particularly to cooled turbine blades for a gas turbine engine.
Gas turbine engines, such as an axial-flow turbine engine, include rotor assemblies with a rotating disk and a number of rotor blades circumferentially disposed around the disk. Rotor blades include a root portion for engaging with the disk and an airfoil portion for positioning within the gas path of the engine. The temperature within this gas path is very high, resulting in the heating of the rotor blades. If the rotor blades become too hot, the durability of the airfoil will be adversely affected. Therefore, various methods of cooling turbine blades have been used to improve the longevity and durability of the turbine blades.
One method of cooling rotor blades is known as simple radial flow film cooling. In this method, internal passages are formed in the rotor blade by drilling straight holes from the base to the tip. The internal passages are cylindrical passages having a single diameter from base to tip, due to the fact that they are formed by drilling. Film holes are then drilled through the side walls of the rotor blade to adjoin with the internal passages. Cooling air is then forced through the passages, which flows out of the film holes. Cooling occurs both as the cooling air passes through the internal passages, and also as it flows out and around the turbine blade. Simple radial flow film cooling was largely abandoned because the method resulted in the release of too much cooling air into the engine. It is undesirable to overcool the engine, because cooling air reduces the pressure within the engine, thereby decreasing the efficiency and thrust of the engine.
To alleviate this problem, airfoil designs began to include complex internal passageways such as serpentine designs. In these airfoils, internal passages were formed to direct cooling air first from the base to the tip of the airfoil. The passage would then turn sharply to direct the cooling air from the tip back to the base. The passage would then turn again directing the cooling air back toward the tip, and so on until it had gone through the entire serpentine passageway.
Serpentine designs present numerous problems. One of the challenges is the difficulty of predicting and modeling the air flow and pressure distribution through a serpentine system while it is rotating. In particular, rotational forces resist the flow of cooling air in the direction from the tip to the base of the airfoil. Predicting and modeling the precise effect of the rotational forces on the cooling air is very difficult. In order to overcome the rotational forces resisting the flow of cooling air from the tip to the base of the airfoil, a sufficient pressure must be present at the tip of an internal passage to force the air back down to the base of the airfoil. If sufficient pressure is not present, cooling air will not flow through the desired serpentine path, resulting in inadequate cooling of the turbine blade.
Furthermore, insufficient pressure within the cooling channel can result in backflow. Backflow occurs when hot gases from the gas path flow into film holes of the airfoil, rather than cooling air flowing out. This leads to undesired heating of the turbine blade. On the other hand, if too much cooling air is present in the cooling channels, too much cooling air will escape from the film holes, resulting in overcooling of the engine.