Patent Application
20130280081
This disclosure relates to an airfoil for a gas turbine engine. In particular, the disclosure relates to airfoil geometry and cooling features and an example core manufacturing process that produces a core providing such features.
Airfoils, particularly those used in a hot section of a gas turbine engine, incorporate internal cooling features. Current airfoil manufacturing techniques limit possible cooling configurations. Typically, the airfoil is cast within a mold having at least first and second portions secured to one another to define an exterior airfoil surface. The core structure used to form the impingement holes and cooling passages must be retained between the mold portions, which limits the location and configuration of the core, which is quite fragile. The core is typically assembled from multiple elements constructed from different material. The elements are glued to one another through a painstaking assembly process, which may result in scrapped cores.
Core elements of 0.012-0.030 inch (0.30-0.76 mm) may be used to produce microcircuits. These core elements must be glued to other core elements of at least 0.060 inch (1.52 mm). A channel in the thicker element must be provided to receive an end of the thinner element, which also results in an abrupt transition with no taper, bell mouth or smooth transition.
In one exemplary embodiment, a core for an airfoil includes a refractory metal structure having a variable thickness.
In a further embodiment of any of the above, the structure includes molybdenum.
In a further embodiment of any of the above, the structure is defined by an exterior surface providing a contour, the exterior surface including a uniform surface finish.
In a further embodiment of any of the above, the structure includes a radiused edge.
In a further embodiment of any of the above, exterior surface is without machining.
In a further embodiment of any of the above, a coating is adhered to the exterior surface.
In a further embodiment of any of the above, the structure is defined by an exterior surface including a perimeter, and includes a recess inboard of the perimeter.
In a further embodiment of any of the above, the structure is defined by an exterior surface including a perimeter, and includes a protrusion inboard of the perimeter.
In a further embodiment of any of the above, the thickness is less than 0.060 inch (1.52 mm) at a location.
In a further embodiment of any of the above, the thickness is less than 0.020 inch (0.51 mm) at the location.
In a further embodiment of any of the above, the structure is tapered between first and second portions, with the second portion less than 0.020 inch (0.51 mm).
In a further embodiment of any of the above, the first portion is greater than 0.020 inch (0.51 mm).
In a further embodiment of any of the above, the structure is a unitary body having uniform material properties throughout the structure. The structure has at least one portion with a thickness of less than 0.060 inch (1.52 mm), and the structure includes first and second portions in spaced apart and overlapping relationship with one another.
In a further embodiment of any of the above, the structure includes a standoff interconnecting the first and second portions.
In a further embodiment of any of the above, one of the first and second portions has a thickness of less than 0.060 inch (1.52 mm).
In another example embodiment, an airfoil for a gas turbine engine includes a body having leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface defined by a perimeter wall. An interior wall is arranged interiorly and adjacent to the perimeter wall to provide a cooling passage there between. A cooling passage with first and second portions is tapered and respectively has first and second thicknesses. The first thickness is greater than the second thickness, and the second thickness is less than 0.060 inch (1.52 mm).
In a further embodiment of any of the above, the second portion includes a radiused edge.
In a further embodiment of any of the above, the second portion includes a recess providing a thickness less than the second thickness.
In a further embodiment of any of the above, the second portion includes a protrusion providing a thickness greater than the second thickness.
In a further embodiment of any of the above, the second portion includes third and forth portions extending generally in a chordwise direction and overlapping one another.
In a further embodiment of any of the above, the third and fourth portions are joined to one another by interconnecting holes.
In another example embodiment, an airfoil for a gas turbine engine includes a body including leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface defined by a perimeter wall. An interior wall is arranged interiorly and adjacent to the perimeter wall to provide first, second and third cooling passages there between. The first and second cooling passages have a thickness of greater than 0.060 inch (1.52 mm), and the third cooling passage interconnects the first and second cooling passages and has a thickness of less than 0.060 inch (1.52 mm).
In a further embodiment of any of the above, the first cooling passage is arranged adjacent to the perimeter wall at the leading edge.
In another embodiment, a method of manufacturing a refractory metal core for an airfoil includes the steps of depositing multiple layers of powdered metal onto one another, joining the layers to one another with reference to CAD data relating to a particular cross-section of a refractory metal core, and producing the core having a variable thickness.
In a further embodiment of any of the above, the core includes molybdenum.
In a further embodiment of any of the above, the thickness is less than 0.060 inch (1.52 mm) at a location.
In a further embodiment of any of the above, the thickness is less than 0.020 inch (0.51 mm) at the location.
In a further embodiment of any of the above, the core is a unitary body having uniform material properties throughout the core. The core has at least one portion with a thickness of less than 0.060 inch (1.52 mm). The core includes first and second portions in spaced apart and overlapping relationship with one another.
In a further embodiment of any of the above, the core includes a standoff interconnecting the first and second portions.
The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein:
The turbine section 11 includes alternating rows of blades 20 and static airfoils or vanes 19. It should be understood that
An example blade 20 is shown in
Referring to
An example core 40 and resultant airfoil 26 is shown in
The core 40 is provided by a refractory metal structure, constructed from molybdenum, for example, having a variable thickness. The core 40 is defined by an exterior surface 60 providing a contour. The exterior surface 60 including a uniform surface finish from the core manufacturing process, described in connection with
In the example, first and second thicknesses 42, 44 are different than one another. In one example, the first thickness 42 is less than 0.060 inch (1.52 mm). The core 40 may include an aperture 46 with a radiused edge 48 providing the second thickness 44 of less 0.020 inch (0.51 mm), for example. The cast airfoil 26 provides a cooling passage 53 and standoff 50 corresponding to the aperture 46. The standoff 50 illustrated in
Referring to
In another example shown in
Referring to
The core 340 may include first and second portions 77, 79 overlapping one another. The first and second portions 77, 79 are less than 0.060 inch (1.52 mm) thick, in one example, and of varying thickness. At least one of the first and second portions 77, 79 may provide a film cooling hole 74 in the exterior airfoil surface 360. The first and second portions 77, 79 may be joined to one another by a standoff 72 that produces a hole interconnecting the resultant overlapping cooling passages. Standoffs 72 can be used to integrally connect and join all passages 66, 68, 70 to eliminate the need for core assembly and better stabilize the core during casting. However, directly manufacturing the airfoil, as shown in
Similarly, a unitary body having uniform material properties throughout the structure provides the cores 440, 540 shown in
Referring to the airfoil 526 of
The core geometries and associated airfoil cooling passages disclosed in
The coating 90 may be applied to the exterior surface of the core 40, which enables the core 40 to be more easily removed subsequently. The core 40 is arranged in a multi-piece mold and held in a desired orientation by features on the mold, as indicated at 92. The core 40 is more robust and can better withstand handling as it is positioned within the mold. The airfoil 26 is cast about the core 40, as indicated at 94. The core 40 is then removed from the airfoil 26, as indicated at 96, to provide desired cooling passage features.
An additive manufacturing process 180 may be used to produce an airfoil, as schematically illustrated in
Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that reason, the following claims should be studied to determine their true scope and content.
