1. Field of the Invention
The present invention relates generally to a turbine blade tip seal, and more specifically to a turbine blade tip seal with tip cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, especially an industrial gas turbine engine, the turbine includes stages of turbine blades that rotate within a shroud that forms a gap between the rotating blade tip and the stationary shroud. Engine performance and blade tip life can be increased by minimizing the gap so that less hot gas flow leakage occurs.
High temperature turbine blade tip section heat load is a function of the blade tip leakage flow. A high leakage flow will induce a high heat load onto the blade tip section. Thus, blade tip section sealing and cooling have to be addressed as a single problem. A prior art turbine blade tip design is shown in FIGS. 1-3 and includes a squealer tip rail that extends around the perimeter of the airfoil flush with the airfoil wall to form an inner squealer pocket. The main purpose of incorporating the squealer tip in a blade design is to reduce the blade tip leakage and also to provide for improved rubbing capability for the blade. The narrow tip rail provides for a small surface area to rub up against the inner surface of the shroud that forms the tip gap. Thus, less friction and less heat are developed when the tip rubs.
Traditionally, blade tip cooling is accomplished by drilling holes into the upper extremes of the serpentine coolant passages formed within the body of the blade from both the pressure and suction surfaces near the blade tip edge and the top surface of the squealer cavity. In general, film cooling holes are built in along the airfoil pressure side and suction side tip sections and extend from the leading edge to the trailing edge to provide edge cooling for the blade squealer tip. Also, convective cooling holes also built in along the tip rail at the inner portion of the squealer pocket provide additional cooling for the squealer tip rail. Since the blade tip region is subject to severe secondary flow field, this requires a large number of film cooling holes that requires more cooling flow for cooling the blade tip periphery. FIG. 1 shows the prior art squealer tip cooling arrangement and the secondary hot gas flow migration around the blade tip section. FIG. 2 shows a profile view of the pressure side and FIG. 3 shows the suction side each with tip peripheral cooling holes for the prior art turbine blade of FIG. 1.
The blade squealer tip rail is subject to heating from three exposed side: 1) heat load from the airfoil hot gas side surface of the tip rail, 2) heat load from the top portion of the tip rail, and 3) 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 peripheral and conduction through the base region of the squealer pocket becomes 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 film cooling and tip section convective cooling holes become very limited.
One prior art reference, U.S. Pat. No. 5,476,364 issued to Kildea on Dec. 19, 1995 and entitled TIP SEAL AND ANTI-CONTAMINATION FOR TURBINE BLADES shows a turbine blade with a tip region having a plurality of spaced holes or slots extending chordwise from the leading edge to the trailing edge of the blade tip, the holes being connected to an internal cooling passage, the holes opening into a cavity formed on the outer surface of the tip. The cavities are separate from each other along the tip edge and all are connected to the internal cooling air passage to discharge cooling air out onto the corner of the tip on the pressure side. The cavities in the Kildea patent are not diffusion openings and do not allow for a formation of a layer of film cooling air on the side of the tip that flows over the tip edge as does the trenches in the present invention.
This problem associated with turbine airfoil tip edge cooling can be minimized by incorporation of a new and effective blade tip cooling geometry design of the present invention into the prior art airfoil tip section cooling design.