EP2565383A2

EP2565383A2 - Airfoil with nonlinear cooling passage

Info

Publication number: EP2565383A2
Application number: EP12182433A
Authority: EP
Inventors: William Abdel-Messeh; Justin D. Piggush
Original assignee: United Technologies Corp
Current assignee: RTX Corp
Priority date: 2011-08-31
Filing date: 2012-08-30
Publication date: 2013-03-06
Anticipated expiration: 2032-08-30
Also published as: EP2565383A3; US20130052037A1; EP2565383B1

Abstract

An example method of manufacturing an airfoil (30) includes providing a ceramic core (64) corresponding to an interior cooling channel (42). A refractory metal core (66) is provided that corresponds to a cooling passage (46). The cores are arranged in a mold (60). An airfoil structure (30) is cast about the cores to provide a turbine engine airfoil. The turbine engine airfoil (30) includes a wall (44) providing the interior cooling channel (42) and an exterior airfoil surface (34). The cooling passage (46) is provided in the wall (44) and fluidly connects the interior cooling channel (42) to the exterior airfoil surface (34). The cooling passage (46) includes multiple inlets (48) and multiple outlets (50) respectively adjoining the interior cooling channel (42) and the exterior airfoil surface (34). At least one of a first inlet (48) and outlet (50) has a different structural flow characteristic than at least one of a second inlet and outlet (48,50).

Description

BACKGROUND

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.

SUMMARY

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.

BRIEF DESCRIPTION OF THE DRAWINGS

Figure 1 is a schematic view of an example gas turbine engine incorporating the disclosed airfoil.
Figure 2 is a perspective view of an example turbine blade.
Figure 3 is a cross-sectional view of a portion of the turbine blade illustrated in Figure 2.
Figure 4 is an airfoil tip cross-sectional view through a cooling passage in a wall of the airfoil structure shown in Figure 3.
Figure 5 is a partial cross-sectional view of a core assembly arranged in a mold prior to casting the airfoil structure.
Figure 6 is a perspective view of a portion of a refractory metal core used to form the cooling passage shown in Figures 3 and 4.
Figures 6A-6C are cross-sectional views of portions of cooling passage outlets illustrated in Figure 6.
Figure 7 is an enlarged top view of a portion of a cooling passage outlet illustrated in Figure 6.

DETAILED DESCRIPTION

A gas turbine engine (GTE) 10 is illustrated schematically in Figure 1. The GTE 10 includes a core section downstream from a fan section 14. The core section 12 includes a compressor section 18 supplying compressed air to a combustor 20. The combusted air expands over a turbine section 22 that rotationally drives a fan 16 within the fan section 14 about an axis A.
The turbine section 22 includes turbine blades 24 rotatable about the axis A and arranged in a circumferential direction C, shown in Figure 2. One example turbine blade is illustrated in Figure 2. The turbine blade 24 has a root 26 that supports a platform 28. An airfoil structure 30 extends in a radial direction R from the platform 28 to a tip 32. The airfoil structure 30 provides an exterior airfoil surface 34 having leading and trailing edges 36, 38 with adjoining spaced apart sides 40.
Referring to Figures 3 and 4, the example turbine blade 24 includes a wall 44 that provides the exterior airfoil surface 34. One or more interior cooling channels 42 are provided by the wall 44 and supply cooling air, for example, compressor bleed air, for cooling the turbine blade 24. This cooling fluid is supplied to various cooling features that ultimately flow through the wall 44 to provide internal convective cooling and a cooling film to the exterior airfoil surface 34.
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 Figure 4, are arranged in the cooling passage 46 as desired, for example, along portions of the outlets 52 to improve cooling. Cross-section of the trip strips can be any shapes such as block (as shown), semi-circular, triangular, semi-elliptic, and alike. Pedestals may also be provided.
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 Figure 5, a ceramic core 64 provides the interior cooling channel 42. A refractory metal core (RMC) 66 provides the cooling passage 46. The ceramic core and the refractory metal core are provided using different materials than one another. One or more locating features 68, such as interlocking protrusions and recesses, locate the RMC 66 relative to the ceramic core 64. The cores 64, 66 are arranged within a cavity 62 of the mold 60. The airfoil structure 30 is typically cast into the mold 60 to provide a structure, such as a single-crystal nickel alloy structure.
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 Figure 6. The RMC 66 includes a first and second ends (generally, 70 and 72), which correspond to the inlets and outlets 48 and 52, joined by a radially extending intermediate portion 74. The first ends 70A, 70B respectively include a first and second inlet area 76, 78 that can be different in shape and size than one another. The outlets 72A, 72B, 72C include first, second and third outlet areas 80, 82, 84 (shown in Figures 6A-6C and respectively represented by cross-sectional lines A-A, B-B, C-C in Figure 6) that can be different than one another. Notches 86 are provided in the RMC 66 to provide corresponding trip strips 58.
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., Figure 4).
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.

Claims

A turbine engine airfoil (30) comprising an airfoil structure having a wall (44) providing an interior cooling channel (42) and an exterior airfoil surface (34), a cooling passage (46) provided in the wall (44) fluidly connecting the interior cooling channel (42) to the exterior airfoil surface (34), the cooling passage (46) including at least one inlet (48) and multiple outlets (52) respectively adjoining the interior cooling channel (42) and the exterior airfoil surface (34), at least one of a first inlet (48) and outlet (52) having a different structural flow characteristic than at least one of a second inlet (48) and outlet (52).
The turbine engine airfoil (30) according to claim 1, wherein the structural flow characteristic includes at least one of length, path and shape.
The turbine engine airfoil (30) according to claim 2, wherein the shape includes a cross-sectional area, the first outlet (52) having a different cross-sectional area than the second outlet (52).
The turbine engine airfoil (30) according to any of claims 1 to 3, wherein the cooling passage (46) extends generally axially within the wall (44), and including a generally axially extending intermediate passage (50) fluidly connecting the inlets (48) to the outlets (52).
The turbine engine airfoil (30) according to any preceding claim, wherein multiple inlets (48) each include an entrance (54) at the interior cooling channel (42), and the outlets (52) each include an exit (56) at the exterior airfoil surface (34), the entrances (54) having a greater cross-sectional area than that of the exits (56).
The turbine engine airfoil (30) according to claim 5, wherein a first entrance (54) includes an area that is greater than a second entrance (54).
The turbine engine airfoil (30) according to claim 5 or 6, wherein a first exit (56) has an area that is greater than a second exit (56).
The turbine engine airfoil (30) according to any preceding claim, wherein the cooling passage (46) includes trip strips (58).
The turbine engine airfoil (30) according to any preceding claim, wherein the cooling passage (46) is nonlinear.
A method of manufacturing the airfoil (30) of claim 1, comprising the steps of:
providing a ceramic core (64) corresponding to an interior cooling channel (42);

providing a refractory metal core (66) corresponding to a cooling passage (46);

arranging the cores (64,66) in a mold (60); and

casting an airfoil structure (30) around the cores (64,66), wherein the airfoil structure (30) includes a wall (44) separating the interior cooling channel (42) from an exterior airfoil surface (34), the cooling passage (46) provided in the wall (44) fluidly connects the interior cooling channel (42) to the exterior airfoil surface (36), the cooling passage (46) including at least one inlet (48) and multiple outlets (52) respectively adjoining the interior cooling channel (42) and the exterior airfoil surface (34), at least one of a first inlet (48) and outlet (52) having a different structural flow characteristic than at least one of a second inlet (48) and outlet (52).
The method according to claim 10, wherein the refractory metal core (66) providing step includes forming a desired core shape and bending the formed desired core shape to correspond to the cooling passage (46).
The method according to claim 11, wherein the bending step includes the bending the cooling passage (46) into generally an S-shape in a lateral direction.
The method according to any of claims 10 to 12, wherein the refractory metal core (66) providing step includes providing notches (86) in the cooling passage (46) corresponding to trip strips (58).
The method according to any of claims 10 to 13, wherein the arranging step includes locating the refractory metal core (66) relative to the ceramic core (64).
The method according to any of claims 10 to 14, wherein the casting step includes forming a diffuser feature in the cooling passage (46).