1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine rotor blade tip section with cooling and sealing.
2. Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98
A gas turbine engine, especially an industrial gas turbine (IGT) engine, includes a turbine section with a number of rows or stages or rotor blades and stator vanes to react with a hot gas flow to power the engine. The efficiency of the engine can be increased by passing a higher temperature gas flow into the turbine. However, the highest turbine inlet temperature is limited to the airfoil materials and the cooling capability of the first stage blades and vanes. An improvement in the material properties or to provide better cooling to allow for higher temperatures will allow for higher engine efficiency.
Another problem with high temperature exposure to the turbine air foils is from erosion due to the hot gas flow acting on a section of the blade tip that is not adequately cooled. A high temperature turbine blade tip section heat load is a function of the blade tip leakage flow. A high leakage flow will induce high heat load onto the blade tip section, leading to increased cooling ability. Thus, blade tip section sealing and cooling must be addressed as a single problem. In the prior art, a turbine blade tip includes a squealer tip rail which extends around the perimeter of the airfoil flush with the airfoil wall to form an inner squealer pocket. The main purpose of incorporating a squealer tip into a blade design is to reduce the blade tip leakage and also to provide rubbing capability for the blade tip against an inner surface of the engine shroud that forms a blade outer air seal or BOAS. The tip rail provides for a minimum amount of material that contacts the shroud surface while minimizing the gap.
In general, film cooling holes are positioned along the airfoil pressure side wall near the tip section and extend from the leading edge to the trailing edge to provide edge cooling for the tip rail at the inner portion of the squealer pocket to provide for additional cooling for the squealer tip rail. Secondary hot gas leakage flow (shown by the arrows over the tip in FIG. 1) migrates around the blade tip section. The vortex flow from the blade suction side is shown by the spiraling arrow flow in FIG. 1. The film holes in the prior art of FIGS. 2 and 4 are drilled into the airfoil surface just below the tip edges on the pressure side wall and the suction side wall and extend from the leading edge to the trailing edge. Also, convection cooling holes are formed along the tip rail at an inner portion of a squealer pocket to provide additional cooling for the squealer tip rail. Since the blade tip region is subject to severe secondary flow, a large number of film cooling holes and cooling flow is required for the cooling of the blade tip periphery.
FIG. 2 shows a prior art blade with the pressure side tip section having a row of pressure side film cooling holes that open just below the tip edge and extend from the leading edge region to the trailing edge region. FIG. 3 shows the film hole breakout pattern of these film holes.
FIG. 4 shows the prior art turbine blade with cooling for the blade suction side tip rail arrangement. The suction side blade tip rail is subject to heating from three exposed sides. Cooling of the suction side squealer tip rail by means off a row of discharge film cooling holes located along the blade suction side peripheral and at the bottom of the squealer floor becomes insufficient. This is primarily due to the combination of tip rail geometry and the interaction of the hot gas secondary flow mixing. The effectiveness induced by the suction side film cooling and the tip section convective cooling holes is very limited.
The blade squealer tip rail is subject to heating from the three exposed sides which includes heat load from the airfoil hot gas side surface of the tip rail, heat load from the top portion of the tip rail, and heat load from the back side of the tip rail. Cooling of the squealer tip rail by means of discharge row of film cooling holes along the blade pressure side and suction side periphery and conduction through the base region of the squealer is insufficient. This is primarily due to the combination of squealer pocket geometry and the interaction of hot gas secondary flow mixing. The effectiveness induced by the pressure side film cooling and the tip section convection cooling holes is very limited. Also, a TBC is normally used ion the industrial gas turbine airfoil for the reduction of blade metal temperature. However, applying the TBC around the blade tip rail without effective backside convection cooling may not reduce the blade tip rail metal temperature. FIG. 5 shows the state of the art prior art blade tip section cooling design. This blade includes the pressure side wall 11 with a pressure side tip film cooling hole 12 and a TBC applied over the wall 11, a pressure side tip rail 15, a suction side wall 21 with a suction side film cooling hole 22, a TBC 13 applied on the suction side wall 21, and a suction side tip rail 25. The blade tip is formed by a tip floor 16 with a TBC 13 on it as well that even extends up along the inner surfaces of the two tip rails 15 and 25. The blade forms a cooling air supply channel 17 that is connected to two rows of tip cooling holes with one row next 18 to the pressure side tip rail 15 and the second row 19 next to the suction side tip rail 25. The two tip rails 15 and 25 include tip caps that form a seal with a BOAS 27 on the stationary casing of the engine.