Abstract:
An integrated shroud structure surrounds a circumferential array of stator vanes and a circumferential array of rotor blades of a gas turbine engine. The shroud structure includes a plurality of vane shroud segments and a plurality of blade shroud segments. The blade shroud segments integrally extend downstream from the vane shroud segments and each pair of circumferentially adjacent blade shroud segments defines an inter-segment gap. At least one slot extends axially from a location downstream of the vane shroud segments to an aft end of the blade shroud segment. The inter-segment gaps and slots are sealed by a sealing band mounted around the full circumference of the integrated shroud structure.

Description:
RELATED APPLICATIONS 
     This application is a continuation of U.S. Pat. No. 9,500,095 issued on Nov. 22, 2016, the content of which is hereby incorporated by reference. 
    
    
     TECHNICAL FIELD 
     The application relates generally to the field of gas turbine engines, and more particularly, to shroud segments for surrounding the blades of gas turbine engine rotors. 
     BACKGROUND OF THE ART 
     The turbine shrouds surrounding turbine rotors are normally segmented in the circumferential direction to allow for thermal expansion. Being exposed to very hot combustion gasses, the turbine shrouds usually need to be cooled. Since flowing coolant through a shroud assembly diminishes overall engine efficiency, it is desirable to minimize cooling flow consumption without degrading shroud segment durability. Individual feather seals are typically installed in confronting slots defined in the end walls of circumferentially adjacent turbine shroud segments to prevent undesirable cooling flow leakage at the inter-segment gaps between adjacent shroud segments. While such feather seal arrangements generally provide adequate inter-segment sealing, there is a continued need for alternative sealing and cooling shroud arrangements. 
     SUMMARY 
     In one aspect, there is provided a shroud structure integrated to a circumferential array of stator vanes for surrounding a circumferential array of rotor blades of a gas turbine engine, the circumferential array of stator vanes positioned axially upstream of the circumferential array of rotor blades, the shroud structure comprising: a plurality of blade shroud segments disposed circumferentially one adjacent to another and configured to surround the circumferential array of rotor blades, the blade shroud segments extending integrally from the circumferential array of stator vanes, each pair of circumferentially adjacent blade shroud segments defining an inter-segment gap, at least one of the plurality of blade shroud segments having a radially inner gas path surface and an opposed radially outer surface and at least one slot extending axially from a location downstream of the circumferential array of stator vanes to a downstream end of the at least one of the plurality of the blade shroud segments between the radially inner gas path surface and the opposed radially outer surface thereof; and a sealing band mounted around the radially outer surface of the blade shroud segments and extending across the inter-segment gaps and the at least one slot around the full circumference of the integrated shroud structure. 
     In a second aspect, there is provided a shroud assembly surrounding stator vanes and rotor blades of a gas turbine engine, the shroud assembly comprising: a plurality of integrated shroud structures disposed circumferentially one adjacent to another to form a circumferentially segmented shroud ring, the segmented shroud ring comprising: a plurality of vane shroud segments; and a plurality of blade shroud segments integrally extending from the plurality of vane shroud segments, each one of the blade shroud segments having a body axially defined from a forward end to an aft end in a direction from an upstream position to a downstream position of a gas flow passing through the integral shroud assembly, and being circumferentially defined between opposite first and second lateral sides, said body including a platform having a radially inner gas path surface and an opposed radially outer back surface, and forward and aft arms extending from the back surface of the platform, said forward and aft arms being axially spaced-apart from each other, at least one slot extending axially from the aft arm towards the forward arm and between the radially inner gas path surface and the opposed radially outer surface thereof; and a sealing band mounted between the forward and aft arms on the back surface of the blade shroud segments, the sealing band encircling the segmented blade shroud ring and circumferentially spanning all the inter-segment gaps and at least partially axially covering the at least one slot. 
     In a third aspect, there is provided a method for sealing and cooling a circumferentially segmented integrated shroud structure, the shroud structure including a segmented blade shroud ring integrally extending from a segmented vane shroud ring in a gas turbine engine, the method comprising surrounding the segmented blade shroud ring with a sealing band configured to fully encircle the segmented blade shroud ring; surrounding at least a portion of axially extending slots defined in the segmented blade shroud ring with the sealing band; forming a pressurized air plenum around the sealing band for urging the sealing band in sealing engagement against a radially outer surface of the segmented blade shroud ring; and providing impingement jet holes in the sealing band to allow some of the pressurized air in the plenum to impinge upon a radially outer surface of the segmented blade shroud ring. 
    
    
     
       DESCRIPTION OF THE DRAWINGS 
       Reference is now made to the accompanying figures, in which: 
         FIG. 1  is a schematic cross-section view of a gas turbine engine; 
         FIG. 2  is a cross-section view of a portion of the turbine section of the gas turbine engine shown in  FIG. 1  and illustrating first and second integrated impingement baffle and shroud seals respectively surrounding a circumferentially segmented turbine shroud and a segmented turbine shroud integrated to an upstream segmented vane ring; 
         FIG. 3  is an enlarged cross-section view illustrating the integrated impingement baffle and shroud seal surrounding the full periphery of a circumferentially segmented turbine blade shroud; 
         FIG. 4  is a rear end view of a split turbine shroud segment integrated to a turbine vane segment; 
         FIG. 5  is a schematic end view illustrating a sealing band mounted about a circumferentially segment shroud ring for sealing the inter-segment gaps; 
         FIG. 6  is a isometric view of a portion of the inter-segment sealing band shown in  FIG. 5 . 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  illustrates a gas turbine engine  10  of a type preferably provided for use in subsonic flight, generally comprising in serial flow communication a fan  12  through which ambient air is propelled, a multistage compressor  14  for pressurizing the air, a combustor  16  in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section  18  for extracting energy from the combustion gases. 
     Referring to  FIG. 2 , it can be observed that the turbine section  18  of the engine  10  may include a number of turbine stages. More particularly,  FIG. 2  illustrates a first stage of turbine rotor blades  20  axially followed by a second stage of stationary turbine vanes  22  disposed for channelling the combustion gases to an associated second stage of turbine blades  24  mounted for rotation about the engine centerline. 
     Surrounding the first stage of turbine blades  20  is a stationary shroud ring  26 . The shroud ring  26  is circumferentially segmented to accommodate differential thermal expansion during operation. Accordingly, the shroud ring  26  may be composed of a plurality of circumferentially adjoining shroud segments  25  (see  FIG. 5 ) concentrically arranged around the periphery of the turbine blade tips  27  so as to define a portion of the radially outer boundary of the engine gas path  28 . The shroud segments  25  may be individually supported and located within the engine by an outer housing support structure  30  so as to collectively form a continuous shroud ring about the turbine blades  20 . As shown in  FIG. 2 , each shroud segment  25  comprises an arcuate platform  32  extending axially from a forward end  34  to an aft end  36  and circumferentially between first and second opposed ends. The platform  32  has a radially inner gas path surface  38  and an opposed radially outer back surface  40 . Axially spaced-apart forward and aft arms  42 ,  44  extend radially outwardly from the back surface  40  of each segment. The arms  42 ,  44  are provided with respective axially projecting distal hooks or rail portions  45 ,  47  for engagement with corresponding mounting flange projections  48 ,  50  on the surrounding support structure  30 . A shroud plenum  52  is defined between the arms  42 ,  44  and the radially outer back surface  40  of the platform  32  for receiving pressurized cooling air from a cooling air source, for example bleed air from the compressor  14 . A feed hole  54  may be defined in the support structure  30  for directing the cooling air in the plenum  52 . As well known, once the shroud ring  26  is assembled, small circumferential inter-segment gaps  53  ( FIG. 5 ) exist between the first and second circumferential ends of adjacent shroud segments  25 . As will be seen hereafter, a sealing arrangement is provided to limit cooling air leakage into the engine gas path through the inter-segment gaps. 
     As shown in  FIGS. 2 and 4 , the second stage of turbine vanes  22  is also typically segmented. Each vane segment  60  comprises at least one vane  22  extending radially between inner and outer vane shroud segments  62 ,  64  that defines the radial flow boundaries for the annular stream of hot gases flowing through the vane ring. In the example illustrated in  FIG. 4 , each vane segment  60  is cast or otherwise suitably manufactured with four circumferentially spaced-apart vanes  22 . Typically, for a given turbine stage, the blade shroud segments are separate from the vane segments. However, as shown in  FIG. 2 , it is herein proposed to combine the vane segments  60  and the blade shroud segments into integral parts. More particularly, each vane segment  60  may be cast with a shroud blade portion  66  extending rearwardly from the outer vane shroud  64 . The integrated structure may be provided with a forward support arm  68  extending radially outwardly from the vane shroud  64  and an aft support arm  70  extending radially outwardly from the blade shroud portion  66 . The forward and aft support arms  68 ,  70  are provided with respective axially projecting distal hooks or rail portions  72 ,  74  for engagement with corresponding mounting flange projections  76 ,  78  on the surrounding support structure  30 . An intermediate ridge  80  may project radially outwardly from the integrated vane and blade shroud to allow for the formation of separate cooling air plenums  82 ,  84  for the vane and blade shroud portions  64 ,  66 . The ridge  80  is configured for radially abutting a radially inner surface of the surrounding support structure  30 . Separate feed holes  86 ,  88  may be provided in the support structure  30  for individually feeding the plenums  82 ,  84  with cooling air. 
     The blade shroud portion  66  of each integrated segment will be classified for different rotor tip diameters. For enhance tip clearance control, multiple blades shroud segments may be incorporated in the same cast vane segment. The integrated approach has several benefits including: less part count, cost and weight reduction, reduced secondary air leakage and smoother gas path, and durability improvement as the TSC is not directly exposed to gas path conditions. Also the vane and shroud segment parts are designed to the same life target, so they should be replaced at overhaul. 
     Referring concurrently to  FIGS. 2 and 4 , it can be observed that the blade shroud portion  66  of each integrated segment may be slotted either mechanically (i.e. EDM, grinding, etc.) or cast-in, to minimize thermal stress and blade shroud uncurling. The number of slots  90  depends on static structures requirements (uncurling, thermal stress, etc.). In the embodiment illustrated in  FIG. 4 , five circumferentially spaced-apart slots  90  are defined in the blade shroud portion  66  of an integrated quad vane segment. As shown in  FIG. 2 , each slot  90  may extend axially from the aft end of the integrated blade shroud portion to a location upstream of the blades  24  relative to the flow of gases flowing through the engine gas path  28 . 
     As shown in  FIG. 2 , a sealing band  92   a,    92   b  may be disposed in each of the plenums  52 ,  84  to seal all the inter-segment gaps (such as the ones shown at  53  in  FIG. 5 ) around the segmented shroud rings and, thus, limit cooling air leakage from the plenums  52 ,  84  into the engine gas path  28 . Each sealing band  92   a,    92   b  is configured to be fitted in sealing engagement with the boundary surfaces of the associated plenum. The pressurized air directed in the plenums  52 ,  84  may be used to urge the sealing bands  92   a,    92   b  in proper sealing engagement with the plenum boundary surfaces. The first sealing band  92   a  has a generally C-shaped cross-section including an annular base  94   a  and forward and aft radially outwardly extending annular sealing faces  96   a,    98   a.  The forward and aft sealing faces  96   a,    98   a  are urged by the pressurized air in uniform sealing contact with the forward and aft arms  42 ,  44 . Likewise, the annular base  94   a  is urged in sealing contact with the radially outer surface of the circumferentially segmented shroud ring  26 . Similarly, the second sealing band  92   b  has an annular base  94   b  and forward and aft annular sealing faces  96   b,    98   b.  The aft sealing face  98   b  may have an axially forwardly bent end portion  100  for engagement with a radially inner surface of the support structure  30  for sealing the aft hook interface between the shroud and support structure. The forward annular face  96   b  of the sealing band  92   b  is urged in sealing engagement against a corresponding axially facing surface of the support structure  30 . The aft annular sealing face  98   b  is urged in sealing engagement with the aft arm  70 . The annular base  94   b  is urged in sealing engagement with the radially outer surface of the blade shroud portions  66  of the segmented blade shroud ring. 
     Each sealing band  92   a,    92   b  covers 360 degrees and, thus, extends across the inter-segment gaps around the full circumference of the associated segmented shroud. The second sealing band  92   b  also seals the portion of the slots  90  extending forwardly from the aft support arm  74 . Each sealing band  92   a,    92   b  may be provided in the form of a full ring, a single split ring with overlapping end portions ( FIG. 3 ) or a single split ring with a butt joint. Sheet metal may be used to form the sealing bands. Impingement jet holes  106  ( FIGS. 2 and 6 ) may be defined in the sealing bands  92   a,    92   b  to allow the same to also act as impingement baffles for cooling the shroud segments. A portion of the air directed in the plenums  52 ,  84  can thus flow through the impingement jet holes  106  for impinging upon the underlying radially outer surface of the segmented shroud rings. 
     As shown in  FIG. 3 , if the sealing bands  92   a,    92   b  are provided with overlapping end portions, a window opening  108  may be defined in the radially outer base layer  110  in order not to block the underlying impingement jets  106  defined in the radially inner base layer  112 . The window opening  108  may be oversized to ensure proper registry between the window opening  108  and the underlying impingement jet holes  106  when the overlapping end portions of the sealing band  92   a,    92   b  slide relative to each other to accommodate thermal growth during engine operation. The use of sealing bands  92   a,    92   b  to seal the inter-segment gaps instead of conventional feather seals result in less part count. It also provides cost reduction (eliminate feather seal slots and feather seals). It also contributes to reduce the assembly time. Finally, it may result in reduced secondary air leakage. 
     It is noted that conventional feather seals  110  ( FIG. 2 ) may still be used to prevent the air directed into the plenum  82  surrounding the second stage of vanes  22  to leak into the engine gas path  28  via the inter-segment gaps in the shroud vane portion  64  of the integrated vane-blade shroud segments. 
     The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without departing from the scope of the invention disclosed. Modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims.