Abstract:
A turbine shroud apparatus for a gas turbine engine having a centerline axis includes: a shroud segment having: an arcuate body extending axially between forward and aft ends and laterally between opposed end faces, wherein each of the end faces includes seal slots formed therein; and an arcuate stationary seal member mounted to the body; a turbine vane disposed axially aft of the shroud segment; and a casing surrounding the shroud segment and the turbine vane; wherein the turbine vane is mounted to the case so as to bear against the stationary seal member, compressing it and forcing the shroud segment radially outward against the casing.

Description:
BACKGROUND OF THE INVENTION 
     This invention relates generally to gas turbine engine turbines and more particularly to apparatus for sealing turbine sections of such engines. 
     A gas turbine engine includes a turbomachinery core having a high pressure compressor, a combustor, and a high pressure turbine in serial flow relationship. The core is operable in a known manner to generate a primary gas flow. In a turbojet or turbofan engine, the core exhaust gas is directed through an exhaust nozzle to generate thrust. 
     A turbofan engine uses a low pressure turbine downstream of the core to extract energy from the primary flow to drive a fan which generates propulsive thrust. The low pressure turbine includes annular arrays of stationary vanes or nozzles that direct the gases exiting the combustor into rotating blades or buckets. Collectively one row of nozzles and one row of blades make up a “stage”. Typically two or more stages are used in serial flow relationship. 
     These components operate in a high temperature environment. Nearby components outside the gas flow path (such as casings) must be protected from the high temperatures to ensure adequate service life. Leakage of flowpath gases between components, for example between turbine rotor shrouds and adjacent turbine nozzles, is therefore undesirable. Prior art designs have attempted to minimize the leakage gap through the compression of the honeycomb on the shroud. While somewhat effective this does not completely prevent leakage. 
     Accordingly, there is a need for a turbine shroud configuration that prevents leakage between the shroud and adjacent components. 
     BRIEF SUMMARY OF THE INVENTION 
     This need is addressed by the present invention, which provides a turbine shroud which is mounted with a combination of compressed honeycomb seals and spline seals to prevent leakage. 
     According to one aspect of the invention, a turbine shroud apparatus for a gas turbine engine having a centerline axis includes: a shroud segment having: an arcuate body extending axially between forward and aft ends and laterally between opposed end faces, wherein each of the end faces includes seal slots formed therein; and an arcuate stationary seal member mounted to the body; a turbine vane disposed axially aft of the shroud segment; and a casing surrounding the shroud segment and the turbine vane; wherein the turbine vane is mounted to the case so as to bear against the stationary seal member, compressing it and forcing the shroud segment radially outward against the casing. 
     According to another aspect of the invention, a turbine shroud apparatus for a gas turbine engine having a centerline axis includes: an annular array of rotatable turbine blades, each blade having an annular seal tooth projecting radially outward therefrom; a shroud surrounding the turbine blades, the shroud comprising an annular array of side-by-side shroud segments, each shroud segment having: an arcuate body extending axially between forward and aft ends and laterally between opposed end faces, wherein each of the end faces includes seal slots formed therein; and an arcuate stationary seal member mounted to the body, wherein the end faces of adjacent shroud segments abut each other and at least one spline seal is received in the seal slots so as to span the gap between adjacent shroud segments; an annular array of airfoil-shaped turbine vanes disposed axially aft of the shroud; and a casing surrounding the shroud segments and the turbine vanes; wherein each of the turbine vanes is mounted to the case so as to bear against one of the stationary seal members, compressing the seal member and forcing the associated shroud segment radially outward against the casing. 
    
    
     
       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 turbine engine constructed in accordance with the present invention; 
         FIG. 2  is an enlarged view of a portion of a turbine section of the engine shown in  FIG. 1 ; 
         FIG. 3  is a front elevational view of a turbine shroud segment shown in  FIG. 2 ; 
         FIG. 4  is a side view of a portion of the shroud segment shown in  FIG. 2 ; and 
         FIG. 5  is a cross-sectional view of a portion of two side-by-side shroud segments, showing a spline seal installed therein. 
     
    
    
     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 portion of a gas turbine  10  engine having, among other structures, a fan  12 , a low-pressure compressor or “booster”  14 , a high-pressure compressor  16 , a combustor  18 , a high-pressure turbine  20 , and a low-pressure turbine  22 . The high-pressure compressor  16  provides compressed air that passes primarily into the combustor  18  to support combustion and partially around the combustor  18  where it is used to cool both the combustor liners and turbomachinery further downstream. Fuel is introduced into the forward end of the combustor  18  and is mixed with the air in a conventional fashion. The resulting fuel-air mixture is ignited for generating hot combustion gases. The hot combustion gases are discharged to the high pressure turbine  20  where they are expanded so that energy is extracted. The high pressure turbine  20  drives the high-pressure compressor  16  through an outer shaft  24 . The gases exiting the high-pressure turbine  20  are discharged to the low-pressure turbine  22  where they are further expanded and energy is extracted to drive the booster  14  and fan  12  through an inner shaft  26 . 
     In the illustrated example, the engine is a turbofan engine. However, the principles described herein are equally applicable to turboprop, turbojet, and turbofan engines, as well as turbine engines used for other vehicles or in stationary applications. 
     The low pressure turbine  22  includes a rotor carrying a array of airfoil-shaped turbine blades  28  extending outwardly from a disk that rotates about a centerline axis “A” of the engine  10 . As seen in  FIG. 2 , the tip  30  of each blade  28  has one or more annular, flange-like seal teeth  32  extending radially outward therefrom. A plurality of shroud segments  34  are arranged in an annulus so as to closely surround the turbine blades  28  and thereby define the outer radial flowpath boundary for the hot gas stream flowing through the rotor. 
     Each shroud segment  34  includes an arcuate body  36  extending between end faces  38  (see  FIG. 3 ) and having forward and aft ends  40  and  42 . From rear to front the body  36  includes a first leg  44  which extends at an acute angle to the centerline axis A, a second leg  46  which also extends at an acute angle to the centerline axis A, a third leg  48  extending generally radially inward from the second leg  46 , and a fourth leg  50  extending generally axially forward from the third leg  48 . The first leg  44  and the second leg  46  meet in a shallow “V” angle with the apex of the V facing radially outwards. 
     The forward end of the second leg  46  overhangs the third leg  48  in the axial direction so that the two define a forward flange  52 . Also, a boss  54  is disposed adjacent the intersection of the first and second legs  44  and  46  and includes a radially-outward-facing groove  56  formed therein. 
     At the end faces  38 , each of the legs  44 ,  46 ,  48 , and  50  includes a slot  58  sized and shaped to receive a conventional spline seal  59  (seen in  FIG. 5 ). A spline seal takes the form of a thin strip of metal or other suitable material which is inserted in slots  58 . The spline seals span the gaps between shroud segments  34 . 
     A stationary seal member  60  is mounted to the radially inner face of the body  36 . The seal member  60  serves the purpose of forming a non-contact rotating seal in conjunction with the seal teeth  32 . The seal member  60  is configured so as to be sacrificial in the even of contact with the seal tooth  32  during operation, an event known as a “rub”. Various types of sacrificial materials exist, such as nonmetallic abradable materials and honeycomb structures. 
     In the illustrated example, the seal member  60  comprises a known type of metallic honeycomb structure comprising a plurality of side-by-side cells, extending in the radial direction. The seal member  60  has a back surface which conforms to the inner surface of the body  36 . It also includes a flowpath surface  62 . The flowpath surface  62  comprises a plurality of cylindrical sections that define a stepped profile, with the surface of each “step” being selected to provide a desired clearance to the adjacent seal tooth  32 . At the aft end of the body  36 , the seal member  60  extends radially inward beyond the first leg  44  of the body  36 , so as to create a slight interference fit, as described in more detail below. The height “H” of the overhang is shown in  FIG. 4 , greatly exaggerated for illustrative purposes. 
     Referring back to  FIG. 2 , a nozzle is positioned downstream of the rotor, and comprises a plurality of circumferentially spaced airfoil-shaped vanes  64 , each of which terminates at an arcuate tip shroud  66 . Arcuate forward and aft hooks  68  and  70  extend outward from the tip shroud  66 . The forward hook  68  extends axially forward and radially outward, and includes a flange  72  extending axially forward at its distal end. 
     An annular casing  74  surrounds shroud segments  34  and the vanes  64 . The casing  74  includes an annular mounting slot  76  which faces axially aft, and also an annular mounting hook  78  with an L-shaped cross-sectional shape. The forward flange  52  of the shroud segment  34  is received in the mounting slot  76 . The slot  56  of the boss  54  receives the mounting hook  78 . 
     The forward hook  68  of the vane  64  is received in a slot defined by the mounting hook  78 . When assembled, the tip shroud  66  of the vane  64  bears radially outward against the shroud segment  34 . 
     The radial distance between the mounting hook  78  and the tip shroud  66  is selected such that the tip shroud  66  creates a slight interference fit with the stationary seal member  60 . The seal member  60  compresses to accommodate this interference, creating a reliable seal against air leakage and holding the shroud segment  34  firmly against the mounting hook  78 . 
     The addition of spline seals on the first leg  44  of the shroud segment  34  and the interference of the tip shroud  66  allows for very little leakage area through the backside of the shroud segment  34  and into the cavity in front of the forward leg of the nozzle. Additionally, the line of sight leakage from the flow path to the case mounting hook  78  is reduced or eliminated. The configuration as described herein will prevent gas path air from leaking over the forward leg of the tip shroud  66  and into the cavity between the shroud segment  34  and the nozzle. The sealing of this cavity from the hot gas path temperatures will protect the mounting hooks  78 . 
     A technical advantage of this configuration is a reduction in leakage through the gaps and a reduction in air temperature in the cavity. The reduction in leakage and air temperature through the gaps will allow for better performance. Alternatively the reduction of air temperature in the cavity will help protect the case hooks from increased temperature and prevent cracking. 
     The foregoing has described a turbine shroud sealing configuration 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, the invention being defined by the claims.