The present invention relates to vanes and vane assemblies for use with gas turbine engines.
Known vane (or stator) assemblies, such as low pressure compressor (LPC) exit guide vane assemblies for gas turbine engines, often include an inner shroud ring, and outer shroud ring, and a plurality of vane details having airfoils that bridge an annular gap between the inner and outer shroud rings in a cascade configuration. In some designs, an inner end of each vane detail includes a platform that is riveted to the inner shroud ring. An outer end of each vane detail lacks a platform like the inner end, but instead has a “free” end that is potted within an opening in the outer shroud using a “slug” of conformable material (e.g., rubber, etc.). Potting the outer ends of the vane details facilitates assembly processes, and provides a damping effect during engine operation. Clips or other retainers are sometimes also used to retain the potted ends of the vane details relative to a shroud. The riveted connection is often located at the inner shroud ring and the potted connection at the outer shroud ring, because some engine designs provide a more secure and desirable mounting arrangement relative to the engine structural frame at the inner shroud location.
However, the amount of space available for securing the platforms of the vane details is limited, particularly at the inner shroud. In order to provide large numbers of vane details, that is, to provide a high vane count, the vane detail platforms have been positioned next to each other in close proximity in a nested configuration. Yet, there are still limits on how closely adjacent vane platforms can be positioned before interfering with each other and raising problems with structural integrity. For instance, there are generally minimum requirements for a distance provided between rivets and an adjacent edge of a riveted part to maintain structural integrity during engine assembly and operation. In short, known nested designs are not readily scaled to allow any number of vanes within a given vane assembly in an engine, but rather face maximum vane count limits.
The present invention provides an alternative vane and vane assembly configuration that allows for relatively high vane counts.
A vane for a gas turbine engine includes an airfoil portion, a platform, and a first flange. The airfoil portion has first and second ends spaced apart in a first direction, and the first end of the airfoil portion defines an unshrouded tip. The platform is integrally formed at the second end of the airfoil, and is configured to define a flowpath boundary segment. The first flange extends from the platform away from the airfoil portion. The first flange defines a first circumferential extension and an adjacent second circumferential extension, each defining forward and aft faces. The first and second circumferential extensions are offset in a second direction such that the forward face of the first circumferential extension is substantially aligned with the aft face of the second circumferential extension in the second direction.
In general, the present invention provides a vane (or stator) and an assembly thereof for use in a gas turbine engine. Each vane includes an integrally formed platform with a flange configured for attachment with an adjacent, similarly-configured vane in a shiplap joint.
The airfoil portion 50 has an aerodynamic curvature (e.g., a three-dimensional “bowed” profile) to interact with fluid passing along the primary flowpath 49 through the LPC section 24. The airfoil portion 50 has a free end (or tip) 58, that is, an end without an integral shroud or platform. In the illustrated embodiment, the free end 58 of the airfoil portion 50 is configured to be inserted into a slot in the OD shroud ring 42 and potted with a conformable material (e.g., rubber) in a conventional manner. In that respect, the free end 58 of the airfoil portion 50 is positioned radially outward in the LPC exit guide vane assembly 40 (see
The platform 52 is arranged at an opposite end of the airfoil portion 50 from the free end 58, and can have a parallelogram-shaped profile. The platform 52 can be positioned radially inward in the LPC exit guide vane assembly 40, as shown in
The first and second flanges 54 and 56 both extend from the platform 52 away from the airfoil portion 50, that is, in a radially inward direction. The first and second flanges 54 and 56 can both be configured to be substantially perpendicular to the engine centerline CL when the vane 44 is installed in the LPC exit guide vane assembly 40 of the engine 20.
The first flange 54 is arranged adjacent to the lip 60 at the downstream edge 52A of the platform 52, and can be integrally formed with the platform 52. The first flange 54 includes a first circumferential extension 62 and a second circumferential extension (or lobe) 64. The first and second circumferential extensions 62 and 64 meet at a central portion 66. Openings 68 and 70 are located in the first and second circumferential extensions 62 and 64, respectively, which enable the first flange 54 to be secured to the downstream ring 48 with suitable fasteners, such as rivets (see
In the illustrated embodiment, the first circumferential extension 62 is integrally joined to the platform 52 along an entire radially outward extent of the first circumferential extension 62, and is generally circumferentially aligned with platform 52. The central portion 66 is positioned at a circumferential edge of the platform 52, and the second circumferential extension extends from the central portion 66 beyond the circumferential edge of the platform 52 in a cantilevered configuration. The first and second circumferential extensions 62 and 64 are both substantially planar. However a chamfered edge 72 is provided at a distal end of the cantilevered second circumferential extension 64 at an aft face thereof.
A cutaway portion is defined in the first flange 54 at a forward face of the first circumferential extension 62. The cutaway portion at the first circumferential extension 62 has a shape that corresponds to that of the second circumferential extension 64. In the illustrated embodiment, the cutaway portion extends to a radially inward edge of the first circumferential extension 62 but its radially outward extent does not reach the platform 52. A depth of the cutaway portion (measured in the axial direction) at the first circumferential extension 62 can be at least as great as a thickness of the second circumferential extension 64 (measured in the axial direction), with a thickness of the central portion 66 being equal to a total distance between an aft face of the first circumferential extension 62 and a forward face of the second circumferential extension 64.
The first flange 54 is configured to form a shiplap joint when engaged with an adjacent vane 44 of similar configuration, as explained further below. In this respect, the first and second circumferential extensions 62 and 64 are axially offset, such that the forward face of the first circumferential extension 62 within the cutaway portion is substantially axially aligned (i.e., co-planar) with the aft face of the second circumferential extension 64.
The second flange 56 is arranged at an upstream edge 52B of the platform opposite the first flange 54, and in the illustrated embodiment is substantially planar, with a substantially rectangular profile, and axially aligned with the upstream edge 52A. Circumferential edges of the second flange 56 are aligned with the circumferential edges of the platform 52 in the illustrated embodiment. The second flange 56 includes an opening 74, enabling the second flange 56 to be secured to the upstream ring 46 with a suitable fastener, such as a rivet (see
A plurality of vanes 44, as described above with respect to
As best shown in
The configuration of the shiplap joint in the illustrated embodiment, with the first circumferential extension 62 offset so as to be positioned generally aft of the second circumferential extension 64, can help reduce tensile stress in the rivets 78. In the illustrated embodiment, operational loading on the airfoil portion 50 will tend to cause the first circumferential extension 62 to pull away from the downstream ring 48 and the second circumferential extension 64 (located at a suction side of the airfoil portion 50, as best shown in
The OD shroud ring 42 and the downstream ring 48 each include connection features, such as bayonet mount lugs, bolt holes, etc., to enable the LPC exit guide vane assembly 40 to be mounted and secured within the gas turbine engine 20. In the illustrated embodiment, the downstream ring 48 provides the primary structural support attachment between the assembly 40 and the rest of the engine 20 (see
When the LPC exit guide vane assembly 40 is assembled in the engine 20, the lip 60 extends downstream (or aft) of the first flange 54, creating an overhang adjacent to the shiplap joint (see
Should one or more of the vanes 44 of the LPC exit guide vane assembly 40 require repair or replacement, it is possible to remove the rivets 78 (or other fasteners) attaching the selected vane 44 and adjacent vanes 44. The selected vane 44 can be removed or replaced, and then the LPC exit guide vane assembly 40 reassembled in the manner described above with regard to the installation of the last vane in the assembly.
It should be recognized that the present invention provides numerous advantages. For example, vane assemblies having vanes secured at a shiplap joint according to the present invention can be positioned relatively close together, allowing relatively high vane counts. This is particularly advantageous where it is desired to secure vanes with fasteners (e.g., rivets) at ID locations, where space is more limited than at corresponding OD locations. The present invention also places fasteners (e.g., rivets) for securing the vanes away from an engine's primary flowpath, which helps promote aerodynamic efficiency and also helps limit a risk of DOD.
Although the present invention has been described with reference to preferred embodiments, workers skilled in the art will recognize that changes may be made in form and detail without departing from the spirit and scope of the invention. For instance, the present invention can be applied to nearly any vane assembly for a gas turbine engine, and the particular shape and configuration of the airfoil portion, platform, and flanges of each vane can vary as desired for particular applications. Additionally, though the illustrated embodiments depict a shiplap joint at an ID location of a vane assembly, in alternative embodiments of the present invention the shiplap joint can be located at an OD location of the vane assembly.
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Entry |
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Official Search Report of the European Patent Office in counterpart foreign Application No. EP08254064 filed Feb. 19, 2008. |
Number | Date | Country | |
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20090208332 A1 | Aug 2009 | US |