1. Technical Field
The disclosure generally relates to gas turbine engines.
2. Description of the Related Art
Turbines of gas turbine engines typically incorporate alternating sets of rotating blades and stationary vanes. In this regard, it is commonplace to incorporate seals between the adjacent sets of blades and vanes. Such seals tend to prevent cooling air leakage from the inner cavities to the gas flow path along which the vanes and blades are located. Oftentimes, such a seal is provided by a coverplate that is secured to a turbine disk, which mounts a set of rotating blades. These coverplates are also often used to provide blade retention.
A bayonet type coverplate is typically characterized by having slotted appendages that interface with corresponding slotted appendages located radially inboard of the live rim of the disk on which the coverplate is mounted. This interface provides axial retention for the coverplate. Radial retention for the coverplate is typically created by a surface located radially inboard of the live rim of the disk. When cooling air for the blades needs to pass through the coverplate, holes are often used. These holes can create high stress concentrations and can limit the operational life of the coverplate.
Additionally, coverplate installation and removal typically involves high tool forces, heating of the turbine disk and/or cooling of the coverplate to relieve interference fits. Unfortunately, these techniques can often be complex and difficult.
Gas turbine engine systems involving rotor bayonet coverplates and tools for installing such coverplates are provided. In this regard, an exemplary embodiment of a turbine assembly for a gas turbine engine comprises: a turbine disk operative to mount a set of turbine blades; and a coverplate having an annular main body portion and a spaced annular arrangement of tabs extending radially inwardly from the main body portion with open-ended gaps being located between the tabs, the tabs being operative to secure an inner diameter of the coverplate to the turbine disk.
An exemplary embodiment of a coverplate for a turbine disk of a gas turbine engine comprises: a main body portion defining a downstream, annular cavity; and a spaced annular arrangement of tabs extending radially inwardly from the main body portion, the tabs being operative to secure an inner diameter of the coverplate to the turbine disk.
An exemplary embodiment of a tool for installing a coverplate on and removing a coverplate from a turbine disk of a gas turbine engine comprises: a body portion; upstream and downstream axial compression surfaces operative to be positioned along a range of axial positions relative to each other such that engagement of the axial compression surfaces with a coverplate applies an axial compression load to the coverplate; and a radial compression surface operative to be positioned along a range of radial positions with respect to the body portion such that engagement of the radial compression surface with the coverplate applies a radial load to the coverplate.
Other systems, methods, features and/or advantages of this disclosure will be or may become apparent to one with skill in the art upon examination of the following drawings and detailed description. It is intended that all such additional systems, methods, features and/or advantages be included within this description and be within the scope of the present disclosure.
Many aspects of the disclosure can be better understood with reference to the following drawings. The components in the drawings are not necessarily to scale. Moreover, in the drawings, like reference numerals designate corresponding parts throughout the several views.
Gas turbine engine systems involving rotor bayonet coverplates and tools for installing such coverplates are provided, several exemplary embodiments of which will be described in detail. In some embodiments, the coverplate extends radially outwardly beyond the live rim (i.e., into the dead rim) of the turbine disk to which the coverplate is installed. Additionally or alternatively, some embodiments incorporate a spaced annular arrangement of tabs that interlock with corresponding annularly spaced locking features of the turbine disk. In addition to securing the coverplate to the turbine disk, locations between the tabs provide open passages that permit the flow of cooling air.
With reference to the partially cut-away, schematic diagram of
As shown in
The radial interference between the coverplate and disk is located radially outboard of the disk live rim. Notably, the live rim is defined by continuous material capable of carrying hoops stresses. This configuration tends to reduce coverplate weight significantly compared to conventional configurations. Because of the weight savings, there is potentially a weight savings for the host turbine disk as well.
As shown more clearly in
A spaced set of locking tabs (e.g., locking tab 160) extend radially inwardly from main body portion 132 of the coverplate. Notably, in the embodiment of
As shown more clearly in
As best shown in
An embodiment of a tool for installing a coverplate to a turbine disk is depicted schematically in
The radial compression jaws are received at least partially within an annular cavity 220 of the base. Each of the jaws is movable between a radial outboard position (not shown) and a radial inboard position. In the embodiment of
Radial compression jaw 206 incorporates dual compression surfaces 234, 236 that are spaced from each other to facilitate radial compression of the coverplate. Each of the compression surfaces is aligned with a corresponding surface of the coverplate. In the embodiment of
Positioning of a radial compression jaw is facilitated by a radial adjustment mechanism (e.g., mechanism 240). In the embodiment of
Axial compression of the coverplate is facilitated by axial compression ring 204, which also is moveably attached to the base. In the embodiment of
An adjustment mechanism 250 that incorporates an annular arrangement of bolts (e.g., bolt 252) facilitates axial positioning of the axial compression ring with respect to the base. In contrast to the compression jaws, which can be moved between radial outboard and inboard positions, the axial compression ring can be moved between axial upstream and downstream positions. In the upstream position, the compression surface 244 is positioned away from corresponding locking tabs of the coverplate. In the downstream position, the compression surface urges the locking tabs toward the turbine disk to provide clearance between the locking tabs of the coverplate and corresponding flange segments of the turbine disk. The compression force is reacted out by the fingers on the downstream side of the main body.
The combined axial and radial compression from the tool releases the interference fits between the coverplate and disk. This allows the coverplate to be positioned onto the disk or taken off the disk with little additional force and no heating or cooling of components.
For installation, the coverplate is positioned inside the tool, which compresses the coverplate radially and axially. The coverplate and tool are then brought towards the disk so that the coverplate locking tabs fit between corresponding tabs of the disk. The coverplate and tool are then rotated so that the coverplate tabs are positioned behind the disk tabs and coverplate cooling air openings are aligned properly with the disk. The axial and radial compression is then removed from the coverplate. Blades are installed surrounding the coverplate antirotation tabs, thus providing positive antirotation. Removal of the coverplate is the opposite of installation.
It should be emphasized that the above-described embodiments are merely possible examples of implementations set forth for a clear understanding of the principles of this disclosure. Many variations and modifications may be made to the above-described embodiments without departing substantially from the spirit and principles of the disclosure. All such modifications and variations are intended to be included herein within the scope of this disclosure and protected by the accompanying claims.
This application is a divisional of U.S. patent application Ser. No. 11/952,367 filed Dec. 7, 2007, which is hereby incorporated by reference in its entirety.
The U.S. Government may have an interest in the subject matter of this disclosure as provided for by the terms of contract number N00421-99-C-1270 awarded by the United States Navy.
Number | Date | Country | |
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Parent | 11952367 | Dec 2007 | US |
Child | 13544668 | US |