Patent Publication Number: US-10328490-B2

Title: Turbine shroud segment with inter-segment overlap

Description:
RELATED APPLICATIONS 
     The present application is a continuation of U.S. Pat. No. 9,079,245 issued on Jul. 15, 2015, the content of which is incorporated herein by reference. 
    
    
     TECHNICAL FIELD 
     The application relates generally to the field of gas turbine engines, and more particularly, to turbine shroud segments. 
     BACKGROUND OF THE ART 
     Gas turbine engines are operated at extremely high temperatures for the purpose of maximizing engine efficiency. Components of a gas turbine engine, such as turbine shroud segments and their supporting structures, are thus exposed to extremely high temperatures. The shroud is constructed to withstand primary gas flow temperatures, but its supporting structures are not and must be protected therefrom. Therefore, it is desirable to prevent the shroud supporting structure from being directly exposed to heat radiations from the hot gaspath. It is also desirable to achieve the required cooling of the turbine shroud segments and surrounding structure with the minimum use of coolant so as to minimize the negative effect on the overall engine efficiency. 
     There is thus a need to provide an improved turbine shroud arrangement which addresses theses and other limitations of the prior art. 
     SUMMARY 
     In one aspect, there is provided a turbine shroud assembly of a gas turbine engine, comprising a plurality of shroud segments disposed circumferentially one adjacent to another, wherein circumferentially adjacent shroud segments have confronting sides defining an inter-segment gap therebetween, and wherein a flow restrictor integrally projects from a first one of said confronting sides of a first shroud segment through the inter-segment gap and into overlapping relationship with a cooperating joint surface provided at a second one of said confronting sides of an adjacent second shroud segment, said flow restrictor and said joint surface defining a clearance therebetween configured to accommodate thermal expansion during hot operating conditions, said clearance and said inter-segment gap being configured to cooperatively define a tortuous leakage path in a generally radial direction between said first and second shroud segments at said hot operating conditions. 
     In a second aspect, there is provided a turbine shroud assembly of a gas turbine engine, comprising a plurality of shroud segments disposed circumferentially one adjacent to another, each of the shroud segment having a metal injection molded body (MIM) being axially defined from a leading edge to a trailing edge in a direction from an upstream position to a downstream position of a hot gas flow passing through the turbine shroud assembly, and being circumferentially defined between opposite first and second lateral sides, said MIM shroud body including a platform having a hot gas path side surface and a back side surface, and forward and aft arms extending from the back side surface of the platform, said forward and aft arms being axially spaced-apart from each other, said MIM shroud body of each of said shroud segments further comprising an integral flow restrictor projecting from said second lateral side through an inter-segment gap defined between confronting first and second lateral sides of adjacent shroud segments, each of said shroud segments having a groove defined in said first lateral side for receiving the flow restrictor of an adjacent shroud segment, the groove being oversized relative to the flow restrictor to provide for the presence of a clearance between the groove and the flow restrictor, the clearance defining a tortuous leakage path between adjacent shroud segments. 
     In a third aspect, there is provided a method of manufacturing a turbine shroud segment for a gas turbine engine, the method comprising: forming a shroud segment body with a groove defined in a first lateral side thereof and with a flow restrictor projecting integrally from an opposite second lateral side thereof, the groove being oversized relative to the flow restrictor to provide for a clearance fit between the flow restrictor and the groove of adjacent turbine shroud segment when assembled together in a ring formation, and wherein the step of forming comprises metal injection molding (MIM) the flow restrictor together with the shroud segment body, and then subjecting the turbine shroud segment body with the integrated flow restrictor to debinding and sintering operations. 
    
    
     
       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 an isometric view of a turbine shroud segment which may be metal injection molded (MIM) with an integral inter-segment flow restrictor; 
         FIG. 3  is an axial cross-section view illustrating a turbine shroud segment mounted to a turbine support case about a turbine rotor including a circumferential array of turbine blades; and 
         FIG. 4  is an enlarged cross-section view illustrating an overlap interface between two circumferentially adjacent shroud segments in cold assembly and hot operating conditions. 
     
    
    
     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. 
     The turbine section  18  generally comprises one or more stages of rotor blades  17  extending radially outwardly from respective rotor disks, with the blade tips being disposed closely adjacent to an annular turbine shroud  19  supported from a turbine shroud support  21  ( FIG. 3 ). The turbine shroud  19  includes a plurality of shroud segments disposed circumferentially one adjacent to another to jointly form an outer radial gaspath boundary for the hot combustion gases flowing through the stage of rotor blades  17 .  FIG. 2  illustrates an example of one such turbine shroud segments  20 . 
     Referring concurrently to  FIGS. 2 and 3 , it can be appreciated that the shroud segment  20  extends axially from a leading edge  29  to a trailing edge  31  in a direction from an upstream position to a downstream position of a hot gas flow (see arrow  23  in  FIG. 3 ) passing through the turbine shroud  19 , and circumferentially between opposite first and second lateral sides  35 ,  37 . The shroud segment  20  has axially spaced-apart forward and aft arms which can be provided in the form of hooks  22  and  24  extending radially outwardly from a back side or cold radially outer surface  26  of an arcuate platform  28 . The hooks  22  and  24  each have a radially extending leg portion  22   a ,  24   a  and an axially extending flange mounting portion  22   b ,  24   b  for engagement with a corresponding hook structure of the turbine shroud support  21 , which may be provided in the form of a shroud hanger as shown in  FIG. 3 . The radially extending leg portions  22   a  and  24   a  define therebetween a cavity  25  which is in fluid flow communication with a source of coolant under pressure (e.g. bleed air from the compressor  14 ). The platform  28  has a radially inner hot gas flow surface  30  adapted to be disposed adjacent to the tip of the turbine blades  17 . Cooling passages (not shown) are typically defined in the platform  28  for receiving cooling air under pressure from the cavity  25  between the forward and aft hooks  22  and  24 . 
     It is desirable to protect the turbine shroud support  21  and the other surrounding turbine structures from the high temperatures of the gas flow  23  flowing through the turbine shroud  19 . It is also desirable to minimize coolant consumption. To that end, it is herein proposed to provide an inter-segment overlap between circumferentially adjacent shroud segments  20 . An example of one such inter-segment overlap is shown in  FIG. 4 . As will be seen hereinafter, the overlap interface at the confronting side faces of each pair of adjacent shroud segments prevents the shroud support structure  21  from being directly exposed to heat radiations from the hot gaspath, while at the same time restricting coolant leakage through the inter-segment gaps, which is advantageous from an engine performance point of view. 
     Referring back to  FIGS. 2 and 3 , the overlap interface between adjacent shroud segments  20  may be provided by forming each shroud segment  20  with a groove  38  in the first lateral side  35  thereof and with a complementary tongue or flow restrictor  40  on its opposite second lateral side  37 . In the embodiment shown in  FIGS. 2 and 3 , the groove  38  and the flow restrictor  40  have both axial and radial components. More particularly, the flow restrictor  40  has a forward leg portion  40   a  projecting from the forward hook  22 , an axially extending base portion  40   b  projecting from the platform  28 , and an aft leg portion  40   c  projecting from the aft hook  24 . The groove  38  has corresponding forward and aft leg portions  38   a  and  38   c  and an axially extending base portion  38   b  respectively defined in the forward and aft hooks  22  and  24  and in the platform  28 . In the illustrated embodiment, the forward and aft leg portions  40   a  and  40   c  of the flow restrictor  40  and associated groove  38  both have a radially outer axially extending component defined on the flanges  22   b  and  24   b  of the forward and aft hooks  22  and  24 . However, it is understood that the flow restrictor  40  and the groove  38  could adopt various other configurations. For instance, they could be provided on the platform  28  only. According to another non-illustrated embodiment, the flow restrictor  40  and the groove  38  could have a U-shaped configuration corresponding to the forward and aft hooks  22  and  24  and the portion of the platform  28  extending between the forward and aft hooks  22  and  24 . 
       FIG. 4  illustrates an example of an inter-segment gap W between the first lateral side  35  of a first shroud segment  20  and the opposed facing second lateral side  37 ′ of a second adjacent shroud segment  20 ′ at a cold assembly condition (i.e. room temperature). The stippled lines in  FIG. 4  illustrate the inter-segment gap W′ at a representative hot engine operating condition. 
     It can be appreciated from  FIG. 4 , that the flow restrictor  40 ′ of shroud segment  20 ′ projects through the inter-segment gap W and partly into the opposed facing groove  38  of shroud segment  20  so as to provide an overlap L between the adjacent segments  20  and  20 ′. It can also be appreciated that the groove  38  is oversized relative to the flow restrictor  40 ′ to provide a clearance fit therebetween. More particularly, the groove  38  and the flow restrictor  40 ′ are sized to provide a clearance C at the cold assembly condition. The clearance C is selected to ensure that a clearance C′ will remain under hot operating conditions. For illustration purposes, during hot operation conditions, the clearance C′ and the inter-segment gap W′ may be of about 0.005 inches and the overlap L′ between the segments  20  and  20 ′ may be of about 0.05 inches. During engine operation, the clearance C′ and the inter-segment gap W′ define a tortuous path which will prevent the shroud support structure  21  from being directly exposed to hot radiations H from the gaspath while allowing a controlled or restricted amount of coolant to flow over the lateral side edges of the shroud segments to properly cool same and avoid hot spots to occur thereat. 
     In the embodiment shown in  FIG. 4 , the groove  38  and the flow restrictor  40 ′ have corresponding tapering cross-sectional profiles. The flow restrictor  40 ′ tapers in a direction away from the lateral side  37 ′ of the shroud segment  20 ′. The groove  38  tapers in a depthwise direction. 
     By so overlapping the adjacent shroud segments, it is also possible for a given shroud segment to provide support to an adjacent damaged shroud segment. Indeed, the flow restrictor  40  may be provided in the form of a rigid tongue integrally projecting from one lateral side of each shroud segments, thereby offering a strong arresting surface against which a damaged segment may rest. The overlap joint between the segments may thus also be used to prevent unacceptable deflection and/or collapsing at the shroud segment sides when exposed to excessive temperatures. This contributes to maintaining tip clearance integrity and, thus, engine performances. 
     The shroud segment overlap design may be implemented by using a metal injection molding (MIM) processes. By metal injection molding the flow restrictor together with the body of the shroud segment, the flow restrictor may be incorporated in the shroud segment design at virtually no extra cost and without additional manufacturing operations. That would not be possible with a conventional casting process. The manufacturing process of an exemplary turbine shroud segment may be described as follows. First, an injection mold (not shown) having a plurality of mold details adapted to be assembled together to define a mold cavity having a shape corresponding to the shape of the desired turbine shroud segment  20  is produced. The mold may have a flow restrictor forming feature as well as a groove forming feature. In this way, the flow restrictor  40  and associated groove  38  can be both conveniently formed at the MIM stage. It is noted that the mold cavity is larger than that of the desired finished part to account for the shrinkage that will occur during debinding and sintering of the green shroud segment. Pins or the like may be inserted in the mold cavity to create cooling holes in the MIM shroud body. 
     A MIM feedstock comprising a mixture of metal powder and a binder is injected into the mold to fill the mold cavity. The MIM feedstock may be a mixture of Nickel alloy powder and a wax binder. The metal powder can be selected from among a wide variety of metal powder, including, but not limited to Nickel alloys, Cobalt alloy, equiax single crystal. The binder can be selected from among a wide variety of binders, including, but not limited to waxes, polyolefins such as polyethylenes and polypropylenes, polystyrenes, polyvinyl chloride etc. The maximum operating temperature will influence the choice of metal type selection for the powder. Binder type remains relatively constant. 
     The MIM feedstock is injected at a low temperature (e.g. at temperatures equal or inferior to 250 degrees Fahrenheit (121 deg. Celsius)) and at low pressure (e.g. at pressures equal or inferior to 100 psi (689 kPa)). It is understood that the injection temperature is function of the composition of the feedstock. Typically, the feedstock is heated to temperatures slightly higher than the melting point of the binder. However, depending of the viscosity of the mixture, the feedstock may be heated to temperatures that could be below or above melting point. 
     Once the feedstock is injected into the mold, it is allowed to solidify in the mold to form a green compact. After it has cooled down and solidified, the mold details are disassembled and the green shroud segment with its integral flow restrictor  40  is removed from the mold. The term “green” is used herein to generally refer to the state of a formed body made of sinterable powder or particulate material that has not yet been heat treated to the sintered state. 
     Next, the green shroud segment body is debinded using solvent, thermal furnaces, catalytic process, a combination of these know methods or any other suitable methods. The resulting debinded part (commonly referred to as the “brown” part) is then sintered in a sintering furnace. The sintering temperature of the various metal powders is well-known in the art and can be determined by an artisan familiar with the powder metallurgy concept. 
     Thereafter, the resulting sintered shroud segment body may be subjected to any appropriate metal conditioning or finishing treatments, such as grinding and/or coating. Cooling passages may be drilled in the MIM shroud body if not already formed therein during molding. This also applies to groove  38  if not formed at the MIM stage. 
     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. For example, a wide variety of material combinations could be used for the MIM shroud body and the integrated flow restrictor. Also, the groove  38  could be replaced by a stepped surface formed in the first lateral side of each shroud segment. For instance, the flow restrictor could be positioned to overly a stepped surface formed on the cold radially outer surface of an adjacent shroud segment. Still other 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.