专利汇可以提供Turbine blade with blade tip cooling notches专利检索,专利查询,专利分析的服务。并且A turbine rotor blade for use in a gas turbine engine, the rotor blade including a squealer pocket on the blade tip to provide a seal between the blade tip and the outer shroud of the engine. A row of film cooling holes extend along the pressure side wall and the suction side wall of the blade near the tip and discharge film cooling air up and over the blade tip. Associated with each of the film cooling holes is a notch formed on the edge of the tip rail in which each notch is located above a film cooling hole and functions to collect and accelerate the cooling air discharged from the film cooling hole up and into the blade tip. The notches reduce the effective thickness of the blade tip in order to reduce the heat load, and increase the effective convection surface area for the cooling air flow. The notches also function to maintain the film layer of cooling air over the blade tip longer than in the prior art without the notches.,下面是Turbine blade with blade tip cooling notches专利的具体信息内容。
I claim the following:
1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to a turbine blade 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 section includes a plurality of stages of turbine rotor blades with blade tips that from a gap with an outer shroud of the engine in which the hot gas flow passing through the turbine can leak past. The blade tip gap leakage not only reduces the efficiency of the turbine by not impacting all of the gas flow onto the turbine rotor blades, but can cause thermal damage to the blade tips and result in shortened life for the blades.
In a high temperature turbine blade tip section, the 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 must be addressed as a single problem. In the prior art, a turbine blade tip includes a squealer tip rail that extends around the perimeter of the airfoil flush with the airfoil wall and forms an inner squealer pocket. The main purpose of using a squealer tip in a blade design is to reduce the blade tip leakage and also to provide the rubbing capability for the blade.
In the prior art, blade tip cooling is accomplished by drilling holes into the upper extremes of the serpentine coolant passages 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 located along the airfoil pressure side and suction side tip sections and from the leading edge to the trailing edge to provide edge cooling for the blade squealer tip. In addition, convective cooling holes are also located along the tip rail at the inner portion of the squealer pocket to provide for additional cooling for the squealer tip rail. Since the blade tip region is subject to severe secondary flow field, a large quantity of film cooling holes and cooling flow is required in order for adequate cooling of the blade tip periphery.
The blade squealer tip rail is subject to heating from three exposed sides which are heat load form 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 peripheral and conduction through the base region of the squealer tip 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 the tip section convective cooling holes becomes very limited.
U.S. Pat. No. 6,494,678 B1 issued to Bunker on Dec. 17, 2002 and entitled FILM COOLED BLADE TIP discloses a turbine rotor blade with a tip having multi-channel cooling grooves (#50 in the Bunker patent) arranged along the tip edge on the pressure side wall of the blade and discharge cooling air from an internal cooling channel of the blade. These cooling holes and channels do not collect and accelerate the cooling air as in the present invention.
It is therefore an object of the present invention to provide for a turbine rotor blade with film cooling holes for the blade tip edge periphery that greatly reduces the airfoil tip edge metal temperature and therefore reduces the cooling flow requirement and improve turbine efficiency.
It is another object of the present invention to provide for a turbine rotor blade with film cooling holes for the blade tip edge periphery that will provide for a film of cooling air to pass over the blade tip.
It is another object of the present invention to provide for a turbine rotor blade with notches on the pressure side tip edge that will retain the cooling air longer in the notches.
It is another object of the present invention to provide for a turbine rotor blade with increased tip section cooling side convection wedded surface area.
It is another object of the present invention to provide for a turbine rotor blade with a blade tip with reduced hot side convective area.
It is another object of the present invention to provide for a turbine rotor blade with a blade tip that will reduce the effective leakage flow effective area which reduces leakage flow and thus lowers the heat load onto the blade tip.
A turbine rotor blade for use in a gas turbine engine, the turbine blade including a row of pressure side film cooling holes just below the tip edge and a row of suction side film cooling holes just below the tip edge, and a row of notches formed on the edge of the squealer tip on the pressure side and a row of notches formed on the edge of the suction side of the blade tip. The pressure side and suction side film cooling holes discharge film cooling air in an upward direction toward the tip edge. Just above the film cooling holes are the notches which have a concave shape with a narrow downstream portion and a wider upstream portion that opens on the side of the tip edge and the top of the tip edge. One notch is associated with one of the film cooling holes such that film cooling air exiting a hole passes into the notch while maintaining a film layer of cooling air over the tip edge. The film cooling air form the holes are retained within the respective notches.
The present invention is directed to the cooling hole arrangement for rotor blade tips used in a gas turbine engine. The concept of the present invention is represented in
The film cooling holes 15 on the pressure and suction side walls of the blade slant upward as seen in
The notches are shaped and sized for several purposes related to the cooling of the blade tip. The notches 16 provide a larger convective area around the blade tip edge to produce greater heat transfer. The notches allow for the film layer to remain longer and farther over the blade tip. The notches also reduce the heat load area of the tip. The size and shape of the notches are such that the notches will catch the flow from the film cooling holes 15 and accelerate the flow over the tip to reduce the leakage flow area between the outer shroud and the blade tip.
Locating the notches 16 above the film cooling holes 15 will allow for the film cooling air exiting the holes to flow in the same direction of the vortex flow over the blade tip from the pressure side wall to the suction side wall. The notches 16 in the blade tip also increases the tip section cooling side surface area and reduce the hot gas convective surface area from the tip crown and therefore reduces the heat load from the tip crown.
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