Patent Abstract:
A shroud sealing arrangement for a gas turbine engine, which comprises a static shroud assembly mounted to an engine case and having a platform surrounding a rotatable airfoil array. The platform has an inner side and an outer side and extends from a leading edge to a trailing edge. A shroud support structure mounts the shroud platform to the case. A circumferential groove is defined on the outer side of the shroud platform proximal to one of the leading edge and the trailing edge. A sealing ring is set in the groove and adapted to seal cooling air from escaping directly to the gas path.

Full Description:
TECHNICAL FIELD 
     The application relates generally to gas turbine engines and, more particularly, to static shroud assemblies for rotor blade arrays. 
     BACKGROUND OF THE ART 
     Typically, an axial gap is provided between a turbine shroud and the outer wall of a gas path duct at ambient temperatures, to allow for thermal expansion of the duct and/or the turbine shroud at engine operating temperatures. The magnitude of such thermal expansion can be predicted, and the gap sized, so that thermal expansion generally seals the gap to prevent leakage through the gap. 
     However, the seal is not perfect and it must be ensured to adequately purge the adjacent cavity with sufficient cooling air to avoid hot gas ingestion. Reducing such uses of secondary air can increase gas turbine engine efficiency. 
     Accordingly, there is a need for an improved turbine shroud sealing arrangement. 
     SUMMARY 
     In one aspect, there is provided a shroud sealing arrangement for a gas turbine engine, the arrangement comprising: a static shroud assembly mounted to an engine case and having a circumferential array of shroud segments surrounding a rotatable blade array, the shroud segments each having a platform, the platform having a radially inner side and a radially outer side and extending axially from a leading edge to a trailing edge, and a forward leg and an aft leg extending radially outwardly from the radially outer side of the platform; a shroud support structure engaged with the forward and aft legs of the shroud segments for mounting the shroud segments to the engine case; a circumferentially extending groove defined on the radially outer side of the shroud segments proximal to one of the leading edge and the trailing edge; and a sealing ring mounted in the circumferentially extending groove, the sealing ring cooperating with the shroud support structure to define a cooling air plenum with one of said forward and aft legs. 
     In another aspect, a gas turbine engine has a circumferential array of shroud segments surrounding a rotatable blade array in a gas path whereby the shroud segments are secured to an engine case by a shroud support structure. An adjacent stator vane assembly forms a gap with the array of shroud segments. An annular slot is defined in the shroud segments near the gap and a radial sealing ring is set in the slot for sealing cooling air to the array of shroud segments. 
     In accordance with another aspect, there is provided a method for cooling the shroud segments of a circumferential array of shroud segments surrounding a rotatable turbine blade array in a gas path, the shroud segments each having forward and aft legs extending radially outwardly from a radially outer surface of a platform, the method comprising: capturing cooling air leaking from between the forward or aft legs in a cooling air plenum closing a leading edge or trailing edge cavity of the shroud segments, and reusing said cooling air to provide impingement cooling on an adjacent component. 
     In accordance with a still further general aspect, there is provided a method for cooling the shroud segments of a circumferential array of shroud segments surrounding a rotatable turbine blade array in a gas path, the method including: supplying cooling air to the array of shroud segments, sealing the cooling air in the area of the shroud segments by defining a radially outwardly facing annular slot near an edge of the shroud segments; providing a sealing ring in the slot and providing discharge ports in the sealing ring 
    
    
     
       DESCRIPTION OF THE DRAWINGS 
       Reference is now made to the accompanying figures, in which: 
         FIG. 1  is a schematic, cross-sectional view of a turbofan engine having a reverse flow annular combustor; 
         FIG. 2  is a schematic, fragmentary view in axial cross-section of the turbine shroud area of the engine shown in  FIG. 1 ; 
         FIG. 3  is a schematic, fragmentary view in axial cross-section of the turbine shroud area similar to  FIG. 2 , but showing the cooling air flow and 
         FIG. 4  is a schematic, fragmentary view in axial cross-section of the turbine shroud area of the engine in accordance with another embodiment. 
     
    
    
     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. A gas path  20  carries the hot combustion gases from the combustor through the turbine section  18  for powering the turbines. 
     The combustor  16  is housed in a plenum  17  supplied with compressed air from compressor  14 . The turbine section  18  is also surrounded by the plenum  17 , defined within the engine case  22 , for supplying cooling air to a turbine shroud surrounding the turbine blades  26  (see  FIG. 2 ). The turbine section  18  generally comprises one or more stages of turbine blades  26  extending radially outwardly from respective rotor disks, with the blade tips  26   a  being disposed closely adjacent to an annular turbine shroud  24  supported from the engine case  22 . The shroud  24  is typically circumferentially segmented.  FIGS. 2 and 3  illustrate an example of one such turbine shroud segments  30 . The various stages of turbine blades  26  are arranged in the gas path  20  with alternating stator vanes  28 . 
     As seen in  FIG. 2 , each shroud segment  30  comprises axially spaced-apart forward and aft hooks or legs  32  and  40  extending radially outwardly from a back side or cold radially outer surface  33   a  of an arcuate platform  33 . The platform  33  has an opposite radially inner hot gas flow surface  33   b  adapted to be disposed adjacent to the tip  26   a  of the turbine blades  26 . The platform  33  is axially defined from a leading edge  34  to a trailing edge  42  in a direction from an upstream position to a downstream position of a hot gas flow passing through gas path  20 , and being circumferentially and longitudinally defined between opposite lateral sides. 
     The forward leg  32  is disposed just downstream of the leading edge  34  of the platform  33 . The leg  32  includes a fastener device  36 , extending, axially downstream of the leg  32 . The fastener device  36  engages a shroud support housing  38  mounted to the engine case  22 . 
     The aft leg  40  is disposed upstream of the trailing edge  42  of the platform  33 . A projection  44  extends downstream and axially from the leg  40 . The projection  44  engages a corresponding axial recess  46  defined in the shroud support housing  38 . A cooling air chamber  48  is defined between the shroud support housing  38  and the forward and aft legs  32 ,  40  of the shrouds segments  30 . Bores  50  traverse the shroud support housing  38  and communicate the plenum  17  with the cooling air chamber  48 . 
     Axial gaps  52  are typically provided between the stator shroud  54  and the leading edge  34  of the shroud segments  30  to provide for thermal expansion. Cooling air can escape through the gaps  52  to exhaust into the gas path  20 . 
     A circumferentially extending slot or groove  58  is defined in the radially outer surface  33   a  of the platform  33  of the shroud segments  30  axially between the leading edge  34  and the forward leg  32 . The grooves  58  of the shroud segments  30  collectively form a full or 360 degrees groove. A 360 degrees sealing ring  56  is mounted in the full circumferential groove  58  formed by the shroud segments  30 . The sealing ring  56  may be provided in the form of a lightweight, annular metal plate. 
     As shown in  FIG. 2 , the outer portion  56   a , of sealing ring  56 , may axially contact the sealing surface  38   a  of the shroud support housing  38 . A circumferential W seal  68  is also resilient and adds pressure to the annular ring  56  to engage the seal surface  38   a . An axial, contact sealing surface  60  is defined on a short axial stub  62  which projects upstream from the annular ring  56  radially inwardly from the outer or peripheral portion  56   a . Part of the stator shroud aft support leg  55  includes a contact surface  64  defined on a short axial stub  66  opposed to the contact surface  60 . Surfaces  60  and  64  form contact sealing faces in running conditions. 
     Referring now to  FIG. 3 , which is identical to  FIG. 2 , there is shown by way of arrows the movement of the cooling air emanating from the plenum  17 . The cooling air enters the shroud array  24  through the bores  50  in the shroud support housing  38  to the cooling air chamber  48 . As there is no feather seal on the forward legs  32  of the shroud segments  30 , the air, under pressure, within the cooling air chamber  48  will leak through the interface between adjacent forward legs  32  of the shroud segments  30 . This leakage air is received in a cooling air plenum  72  defined between the annular ring  56  and the forward leg  32  of the shroud segments  30 . The air in plenum  72  provides cooling along all the length of the forward leg  32 . It also provides for a better cooling of the leading edge region of the platform. This contributes to improve shroud durability. It also eliminates the need for multiple feather seals between the forward legs of the shroud segments. Air also passes by the aft legs  40  in order to enter the plenum  49  where the cooling air can impinge on the downstream portion of the platform of the shroud segments  30 . Along the axial length of the platform  33  of the shroud segments  30  are feather seals  76  and cooling air impinges on the shroud segment  30 , between the feather seals  76 . 
     Cooling air passes from the plenum  72  through impingement holes  70  defined in the sealing ring  56 . The holes  70  may be evenly distributed on a circumferential row and oriented so as to aim at the back face of the adjacent stator shroud  54 . The size and number of discharge ports or holes will be determined by design criteria for a given engine. As depicted by the arrows in  FIG. 3 , the air passing through the holes  70  impinges on the back face of the stator shroud  54 . The air may then be used to purge the gap  52  formed between the stator shroud  54  and the annular ring  56  as well as the leading edge  34  of the shroud segments  30 . 
     Reusing the cooling air to cool the adjacent component (the stator shroud) and to purge the gap between the shroud segments and the adjacent component allows to reduce the amount of cooling air and, thus contributes to the engine efficiency. The 360 degrees sealing plate architecture also provides better control of cooling air leakage as compared to individual feather seals. 
     During operation, the hot environment of the gas path  20  causes the shroud segments  30  and the stator vane shroud  54  as well as shroud support  55  to expand axially towards each other so that the contact surfaces  60  and  64  of the stubs  62  and  66  respectively sealingly engage each other, thus providing a seal against the loss of the cooling air into the gas path  20 . At the same time, the W seal  68  is compressed so that the outer portion  56   a  of the sealing ring  56  abuts the contact surface  38   a  in a sealing arrangement. However a nominal amount of cooling air loss is acceptable. The spent cooling air once into the gas path  20  may form a cooling film along the outer surface of the shroud segments  30 . 
     The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiment described without departing from the scope of the invention disclosed. For example, the sealing ring  56  can be provided with different configurations, and is not limited to application in turbofan engines. Furthermore the spring shown in the drawings can have different configurations and need only be resilient. Also, as shown in  FIG. 4 , the sealing ring could be mounted in an associated groove defined in the radially outer surface of the platform axially between the aft leg and the trailing edge of the platform to provide sealing along the aft leg and ensure proper cooling thereof. 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.

Technology Classification (CPC): 5