Patent Publication Number: US-8123473-B2

Title: Shroud hanger with diffused cooling passage

Description:
BACKGROUND OF THE INVENTION 
     This invention relates generally to gas turbine engine turbines and more particularly to methods for cooling turbine sections of such engines. 
     A gas turbine engine includes a turbomachinery core having a high pressure compressor, combustor, and high pressure or gas generator turbine in serial flow relationship. The core is operable in a known manner to generate a primary gas flow. 
     The gas generator turbine includes one or more rotors which extract energy from the primary gas flow. Each rotor comprises an annular array of blades or buckets carried by a rotating disk. The flowpath through the rotor is defined in part Typically two or more stages are used in serial flow relationship. These components operate in an extremely high temperature environment, and must be cooled by air flow to ensure adequate service life. Typically, the air used for cooling is extracted from one or more points in the compressor. 
     Conventional cooled turbine shrouds are supported by segmented hangers through which the shroud cooling air is supplied. This air is typically supplied through holes in the main body of the hanger. Once through the hanger, the air enters a plenum formed by the hanger and a sheet metal impingement baffle. The air then passed through the baffle and impinges on the shroud. In order to not damage the sheet metal baffle, it is preferable that the hanger holes be angled such that the air does not directly impinge on the baffle, or that the air is diffused before entering the plenum. 
     Current turbine shroud hangers either use straight holes which impinge directly on the baffle, or holes with partially cast diffusers. Turbine shroud hangers utilizing the direct impingement have experienced sheet metal baffle cracking due to excitation from the high velocity air coming from the hanger holes. Conventional cast diffusers require substantial space to be incorporated in and may require the use of quartz rods in the casting process. 
     BRIEF SUMMARY OF THE INVENTION 
     These and other shortcomings of the prior art are addressed by the present invention, which provides a turbine shroud hanger which incorporates a simple, compact impingement air diffuser. 
     According to one aspect of the invention, shroud hanger for a gas turbine engine has an arcuate body with opposed inner and outer faces and opposed forward and aft ends, the channel having at least one cooling passage therein which includes: (a) a generally axially-aligned channel extending through the body, the channel having one end open to an exterior of the body; and (b) a generally radially-aligned diffuser extending through the inner face and intersecting the channel. 
     According to another aspect of the invention a method of making a shroud hanger for a gas turbine engine includes: (a) casting an arcuate body with opposed inner and outer faces and opposed forward and aft ends; (b) forming a generally radially-aligned diffuser extending through the inner face; and (c) forming a generally axially-aligned channel extending through the body, the channel having one end open to an exterior of the body and intersecting the diffuser. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The invention may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which: 
         FIG. 1  a schematic cross-sectional view of a gas generator core of a turbine engine constructed in accordance with an aspect of the present invention; 
         FIG. 2  is a cross-sectional view of a turbine shroud hanger shown in  FIG. 1 ; 
         FIG. 3  is a view taken along lines  3 - 3  of  FIG. 2 ; 
         FIG. 4  is a view taken along lines  4 - 4  of  FIG. 2 ; 
         FIG. 5  is a schematic cross-sectional view of a mold for casting a turbine shroud hanger; 
         FIG. 6  is a schematic cross-sectional view of a shroud hanger cast using the mold of  FIG. 5 ; 
         FIG. 7  is a view of the shroud hanger of  FIG. 9  after a cooling passage has been machined therein; 
         FIG. 8  is a cross-sectional view of an alternative turbine shroud hanger constructed in accordance with an aspect of the present invention; 
         FIG. 9  is a view taken along lines  9 - 9  of  FIG. 8 ; and 
         FIG. 10  is a view taken along lines  10 - 10  of  FIG. 8 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views,  FIGS. 1 and 2  depict a gas generator turbine  10  which forms a portion of a gas turbine. It includes a first stage nozzle  12  which comprises a plurality of circumferentially spaced airfoil-shaped hollow first stage vanes  14  that are supported between an arcuate, segmented first stage outer band  16  and an arcuate, segmented first stage inner band  18 . The first stage vanes  14 , first stage outer band  16  and first stage inner band  18  are arranged into a plurality of circumferentially adjoining nozzle segments that collectively form a complete 360° assembly. The first stage outer and inner bands  16  and  18  define the outer and inner radial flowpath boundaries, respectively, for the hot gas stream flowing through the first stage nozzle  12 . The first stage vanes  14  are configured so as to optimally direct the combustion gases to a first stage rotor  20 . 
     The first stage rotor  20  includes a array of airfoil-shaped first stage turbine blades  22  extending outwardly from a first stage disk  24  that rotates about the centerline axis of the engine. A segmented, arcuate first stage shroud  26  is arranged so as to closely surround the first stage turbine blades  22  and thereby define the outer radial flowpath boundary for the hot gas stream flowing through the first stage rotor  20 . 
     A second stage nozzle  28  is positioned downstream of the first stage rotor  20 , and comprises a plurality of circumferentially spaced airfoil-shaped hollow second stage vanes  30  that are supported between an arcuate, segmented second stage outer band  32  and an arcuate, segmented second stage inner band  34 . The second stage vanes  30 , second stage outer band  32  and second stage inner band  34  are arranged into a plurality of circumferentially adjoining nozzle segments that collectively form a complete 360° assembly. The second stage outer and inner bands  32  and  34  define the outer and inner radial flowpath boundaries, respectively, for the hot gas stream flowing through the second stage turbine nozzle  28 . The second stage vanes  30  are configured so as to optimally direct the combustion gases to a second stage rotor  36 . 
     The second stage rotor  36  includes a radially array of airfoil-shaped second stage turbine blades  38  extending radially outwardly from a second stage disk  40  that rotates about the centerline axis of the engine. A segmented arcuate second stage shroud  42  is arranged so as to closely surround the second stage turbine blades  38  and thereby define the outer radial flowpath boundary for the hot gas stream flowing through the second stage rotor  36 . 
     The segments of the first stage shroud  26  are supported by an array of arcuate first stage shroud hangers  44  that are in turn carried by an arcuate shroud support  46 , for example using the illustrated hooks, rails, and C-clips in a known manner. A shroud plenum  48  is defined between the first stage shroud hangers  44  and the first stage shroud  26 . The shroud plenum  48  contains a baffle  50  that is pierced with impingement cooling holes in a known manner. 
       FIGS. 2 ,  3 , and  4  show one of the first stage shroud hangers  44  in more detail. It is noted that the first stage shroud hanger  44  is used merely as an example to illustrate the principles of the present invention, which are equally applicable to other similar components, for example the hangers supporting the second stage shrouds  42 . The first stage shroud hanger  44  is a unitary casting and has an arcuate body  52  with opposed inner and outer faces  54  and  56 , and opposed forward and aft ends  58  and  60 . A forward hook  62  having a generally L-shaped cross-section extends radially inward from the inner face  54 , at the forward end  58 . An aft hook  64  having a generally L-shaped cross-section extends radially inward from the inner face  54 , at the aft end  60 . 
     A forward mounting rail  66  having a generally L-shaped cross-section with axial and radial legs  68  and  70  extends from the outer face  56 , at the forward end  58 . An aft mounting rail  72  having a generally L-shaped cross-section extends from the outer face  56 , at the aft end  60 . 
     An annular array of cooling passages  74  are formed in the body  52 . Each cooling passage  74  has a generally axially-aligned channel  76  and a generally radially-aligned diffuser  78 . The channel  76  passes through the radial leg  70  of the forward mounting rail  66  and extends through the body  52 . In the illustrated example each of the channels  76  passes through an optional boss  80  which protrudes radially outward from the outer face  56  of the body  52 . The aft end of the channel  76  joins the diffuser  78 . The diffuser  78  passes through the inner face  54  and extends through the body  52  into the boss  80 . The cross-sectional flow area of the diffuser  78  is significantly greater than that of the channel  76 . In this example the angle θ 1  between a back wall  82  of the diffuser  78  and the centerline of the channel  76  is about 90 degrees. 
     In operation, cooling air from a source within the engine, for example compressor bleed air, is supplied to the channel  76 . The high velocity air coming through the channel  76  will lose some of its velocity head when it impinges on the back wall  82  of the diffuser  78 . As this is a part of a relatively thick casting, it can be made to have sufficient thickness such that there is no risk of damage due to excitation from the cooling air. The air, with lower velocity, then turns radially inward as shown by the arrow in  FIG. 2 , and diffuses. It subsequently flows into the shroud plenum  48  (see  FIG. 1 ) where is it used for impingement cooling in a known manner. Based on analysis, the axial position of the diffuser  78  can be preferentially located for each specific application, to ensure a uniform distribution of air in the shroud plenum  48 , which results in uniform impingement cooling for the first stage shroud  26 . 
     The shroud hanger  44  may be manufactured using a known investment casting process, in which a ceramic mold is created (shown schematically at “M” in  FIG. 5 ) which has a cavity “C” that defines the form of the shroud hanger  44  and its interior features. The mold cavity C includes an integral positive feature or plug “P” in the shape of the diffuser  78 . The mold M is placed in a furnace, and liquid metal, for example a known cobalt- or nickel-based “superalloy”, is poured into an opening therein (not shown). After the metal is allowed to cool and solidify, the external shell is broken and removed, exposing the casting which has taken the shape of the shroud hanger  44  including the diffuser  78 , as shown in  FIG. 6 . Optionally, the diffuser  78  could be formed by machining after casting. 
     After the casting process is complete, the channel  76  is formed by machining (e.g. by drilling, ECM, EDM, or a similar process) through the radial leg  70  and the boss  80  to intersect the diffuser  78 , as shown in  FIG. 7 . Optionally, the channel  76  could be formed during casting by incorporating a quartz rod or other refractory core element into the mold M in a known manner. 
     The dimensions and shapes of the cooling passages  74  may be varied to suit a particular application. For example,  FIGS. 8-10  illustrate an alternative shroud hanger  144  similar in construction to the shroud hanger  44  described above. It includes a cooling passage  174  comprising a channel  176  and a diffuser  178 . In this example the angle θ 2  between a back wall  182  of the diffuser  178  and the centerline of the channel  176  is about 45 degrees. This design produces a lower pressure drop in the flow exiting the cooling passage  174  than the design shown in  FIGS. 2-4 , which may be desirable in some applications. 
     The shroud hanger described herein has several advantages over a conventional design. By targeting the channel  74  at a cast surface, baffle distress caused by high velocity impingement air is avoided. This configuration is also optimized to work in areas of limited space where there is not enough room for a typical in-line diffuser configuration. Finally, the cast features are relatively simple to create, reducing the cost and complexity of the manufacturing process. 
     The foregoing has described a shroud hanger for a gas turbine engine. While specific embodiments of the present invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be made without departing from the spirit and scope of the invention. Accordingly, the foregoing description of the preferred embodiment of the invention and the best mode for practicing the invention are provided for the purpose of illustration only and not for the purpose of limitation.