Patent Application
20130052037
This disclosure relates to a cooling passage for an airfoil.
Turbine blades are utilized in gas turbine engines. As known, a turbine blade typically includes a platform having a root on one side and an airfoil extending from the platform opposite the root. The root is secured to a turbine rotor. Cooling circuits are formed within the airfoil to circulate cooling fluid, such as compressor bleed air. Typically, multiple relatively large cooling channels extend radially from the root toward a tip of the airfoil. Air flows through the channels and cools the airfoil, which is relatively hot during operation of the gas turbine engine.
Some advanced cooling designs use one or more radial cooling passages arranged between the cooling channels and an airfoil exterior surface that extend from the root toward the tip. The cooling passages provide high convective cooling.
Other current airfoil tooling designs make use of some cooling holes drilled through airfoil walls and into internal cooling passages. In this type of configuration, the geometry of the holes is limited to a straight hole with the possibility for some flow diffusing feature added near the exit of the hole. As holes must be drilled in a straight line, minimal angles with the airfoil exterior surface must be observed. The length of holes is dictated by manufacturing constraints.
An example method of manufacturing an airfoil includes providing a ceramic core corresponding to an interior cooling channel. A refractory metal core is provided that corresponds to a cooling passage. The cores are arranged in a mold. An airfoil structure is cast about the cores to provide a turbine engine airfoil.
The turbine engine airfoil includes a wall providing the interior cooling channel and an exterior airfoil surface. The cooling passage is provided in the wall and fluidly connects the interior cooling channel to the exterior airfoil surface. The cooling passage includes multiple inlets and multiple outlets respectively adjoining the interior cooling channel and the exterior airfoil surface. At least one of a first inlet and outlet has a different structural flow characteristic than at least one of a second inlet and outlet.
These and other features of the disclosure can be best understood from the following specification and drawings, the following of which is a brief description.
A gas turbine engine (GTE) 10 is illustrated schematically in
The turbine section 22 includes turbine blades 24 rotatable about the axis A and arranged in a circumferential direction C, shown in
Referring to
In the example, a cooling passage 46 fluidly interconnects the interior cooling channel 42 to the exterior airfoil surface 34 and is arranged on the pressure side of the turbine blade 24. The cooling passage 46 includes multiple inlets 48 adjoining a radially extending intermediate passage 50. Multiple outlets 52 adjoin the intermediate passage 50, which enables the pressure to be better equalized across the outlets 52. The inlets 48 each provide an entrance 54 at the interior cooling channel 42. The extended intermediate passage 50 provide exits 56 arranged at the end of the airfoil structure near the tip 32. The cooling passage 46 has a generally S-shaped cross-section. The flow path from the entrance 54 to the exit 56 can replace the straight, drilled holes previously used. Trip strips 58, schematically shown in
In the example, the interior cooling channel 42 and cooling passage 46 are provided by one or more ceramic cores arranged within a mold. Referring to
The RMC 66 is formed to provide a desired core shape. Typically, the RMC can be stamped out of a flat sheet metal. Subsequently, this stamped RMC shape is bent to a desired shape to provide a correspondingly shaped cooling passage 46, an example of which is illustrated in
The RMC 66 can be configured provide different structural flow characteristics with any desired geometry to produce holes of any desired length, path and exit shape, for example. For example, by utilizing different cross-sectional areas along the length of the RMC 66 (for example in along the flow path from the entrance 54 to the exit 56), each hole may be designed to provide desired pressure drop control across the radial length of the cooling passage 46 rather than over pressurizing many of the drilled holes with only a few holes optimized. The cooling passage 46 may include any heat transfer augmentation features such as trip strips to improve heat transfer characteristics and control pressure drops through the holes. Diffuser features 90 may also be provided in the cooling passage 46 and in the exits 56 (see, e.g.,
Although example embodiments have been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. As another example, the method disclosed above can be applied to manufacturing blade outer air seals (BOAS). For that reason, the following claims should be studied to determine their true scope and content.
