Abstract:
A ring seal for sealing an annular gap defined between opposed surfaces of concentric inner and outer annular engine parts in a gas turbine engine. The ring seal being wholly disposed within the annular gap and having a wave pattern with alternating peaks. The ring seal radially inwardly loads the inner engine part and provides a fluid seal between the inner and outer engine parts.

Description:
TECHNICAL FIELD 
   The invention relates generally to a seal for a gas turbine engine, and more particularly, to an improved turbine shroud segment seal. 
   BACKGROUND OF THE ART 
   Seals provided between turbine shroud segments and outer supporting housings in gas turbine engines are well-known. Such seals reduce gas flow between inner cooling air cavities, defined within the turbine shroud segments, and the main engine hot gas path defined radially within the turbine shroud segments. In many engine designs, relatively cool secondary air flow is fed from the compressor to the cooling cavities defined within the turbine shroud segments to provide cooling thereof. In order to prevent leakage of this cooling air into the main gas path, seals are preferably provided between the upstream and downstream edges of the turbine shroud segments and the outer supporting shroud housing. 
   In order to achieve a tight clearance gap between the turbine blade tips and the surrounding shroud segments, it is common to grind the shroud segments, once assembled, until the desired tip clearance is achieved. Most known turbine shroud segment seals, however, require a special fixture in order to load the segments radially inward during this grinding operation in order to prevent the grinding wheel from pushing the shroud segments outward as a result of deflections in the shroud seals. This results in increased and unacceptable tolerances between the turbine blade tips and the surrounding turbine shroud segments. 
   A simplified turbine shroud seal which is economical to manufacture and which obviates the need for special retaining fixtures of the shroud segment during assembly grinding operations is accordingly desired. 
   SUMMARY OF THE INVENTION 
   It is therefore an object of the present invention to provide an improved turbine shroud segment seal. 
   In a first aspect, the present invention provides a ring seal for sealing an annular gap defined between opposed surfaces of concentric inner and outer annular engine parts in a gas turbine engine, said opposed surfaces being radially spaced apart by a first distance, said ring seal being wholly disposed within said annular gap and having a wave pattern with alternating peaks defining an uncompressed radial height therebetween greater than said first distance, said ring seal disposed within said annular gap exerts solely a radial force on said inner engine part relative to said outer engine part which is fixed, thereby radially inwardly loading said inner engine part and providing a fluid seal between said inner and outer engine parts. 
   In a second aspect, the present invention provides a ring seal for sealing an annular gap radially defined between an axially extending outer surface of at least one turbine shroud segment and an axially extending inner surface of a surrounding shroud housing in a gas turbine engine, said ring seal being wholly disposed within said annular gap and having a wave pattern with alternating peaks, said ring seal exerting a radially inward force on said turbine shroud segment such that said turbine shroud segment is radially inwardly loaded. 
   In a third aspect, the present invention provides a ring seal for a turbine blade tip shroud in a gas turbine engine having a hot main gas flow passage and an outer casing, the shroud being located between the main gas flow passage and a cooling passage formed between the shroud and at least a portion of the outer casing, the shroud having at least one axially extending mounting platform for engagement with corresponding mounting structure of the outer casing such that an annular gap is defined between radially spaced walls formed on the mounting platform of the shroud and the mounting structure of the outer casing, wherein the ring seal is wholly disposed between the radially spaced walls within the annular gap and provides fluid sealing between said main gas flow passage and said cooling passage, said ring seal having a wave pattern with alternating peaks abutting and radially extending between said radially spaced walls, and said ring seal acting on said shroud to exert a radially inward force thereon such that said shroud is inwardly loaded. 
   There is also provided, in accordance with the present invention, a method of assembling a gas turbine engine, comprising the steps of: mounting a turbine shroud segment to a casing; providing a seal between the shroud segment and the casing; and grinding the shroud segment to provide accurate tip clearance, wherein radial support for the shroud segment during grinding is provided substantially by the seal. 
   Further details of these and other aspects of the present invention will be apparent from the detailed description and figures included below. 

   
     DESCRIPTION OF THE DRAWINGS 
     Reference is now made to the accompanying figures depicting aspects of the present invention, in which: 
       FIG. 1  is a schematic cross-section of a gas turbine engine; 
       FIG. 2  is a partial cross-sectional view of a turbine shroud segment seal of the prior art; 
       FIG. 3  is a partial cross-sectional view of turbine shroud segment seals in accordance with the present invention; and 
       FIG. 4  is a partial cross-sectional view of the turbine shroud segment seal of  FIG. 3  shown in greater detail. 
   

   DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
     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 multi-stage 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  may comprise several turbine stages, each of which generally includes a rotatable turbine rotor having a plurality of blades extending therefrom within a surrounding turbine shroud. A plurality of vanes, arranged in an annular configuration, are provided immediately upstream of each turbine rotor. 
   Referring to prior art  FIG. 2 , a turbine stage  20  of a gas turbine engine includes generally a turbine rotor  22  having a plurality of radially extending blades  23  and a turbine stator vane assembly  24  comprising a plurality of vanes  25  extending between inner and outer vane platforms  27 . Surrounding the turbine blades  23 , and downstream of the stator vane assembly  24 , is an annular turbine shroud  26 , which typically comprises a plurality of individual shroud segments  28  concentrically arranged around the periphery of the turbine blade tips. The shroud segments  28  are supported and located within the engine by an outer housing support structure  34 . Spring seals  32  provide sealing between the shroud segments  28  and the surrounding support structure  34 . While the spring seals  32  having a configuration as depicted in  FIG. 2  provide adequate sealing properties, their relatively complex shape makes them expensive to manufacture. Further, additional support is required during the shroud grinding operations to prevent unwanted radial outward movement of the shroud and seals. Other shapes of ring seals are also employed elsewhere in gas turbine engines. For example, M-shaped seals  29  are known to be employed between the main platforms  27  and their surrounding support structure  31  for sealing purposes only. 
   Referring now to the present invention as depicted in  FIGS. 3 and 4 , a plurality of shroud segments  40  surrounding turbine rotor blades  23  include axially extending mounting platforms  41  which engage, and are preferably received within, corresponding mounting flange projections  45  of the surrounding shroud housing  44  in order to help locate the shroud segments  40  in position within the turbine section of the gas turbine engine. Internal cavities  42  are defined within the turbine shroud segments  40  and are generally provided with secondary cooling air via apertures  46  defined in the surrounding housing  44 . According to the present invention, M-shaped sealing rings  50  are wholly disposed within annular gaps  48 , defined between the opposed surfaces of the mounting platforms  41  of the shroud segments  40  and the mounting flanges  45  of the surrounding housing  44 . Particularly, the sealing rings  50  are disposed between an axially extending outer surface  43  of the mounting platforms  41  and an axially extending inner surface  49  of the surrounding shroud housing  44 . The M-shaped sealing rings  50  act to seal the shroud cavity  42  such that leakage of cooling air from the cavity  42  into the main gas path is at least limited, if not prevented. The sealing is particularly achieved by radially pinching the M-shaped sealing rings  50  between the axial projections  41  of the turbine shroud segments  40  and the surrounding outer housing  44 . 
   M-shaped sealing rings  50  are preferably provided both at the upstream and downstream engagement points of the shroud segments  40 , between the shroud segments and the surrounding housing. The M-shaped sealing rings  50  further act to load the shroud segments  40  radially inward, and thereby obviate the need for additional shroud supports at least during the shroud grinding operation performed following assembly of the turbine section of the gas turbine engine. This shroud grinding step is performed in order to achieve the precise tip clearance gap desired between the turbine blade tips and the surrounding shroud segments, thereby minimizing tip clearance losses. The M-shaped sealing rings  50  also have a simplified shape in comparison with the more complex sealing rings  32  known in the prior art for sealing turbine shroud segments, and are therefore less costly to produce than the turbine shroud seals  32  of the prior art, which have a significantly more complex shape necessitating more forming operations. 
   The M-shaped sealing rings  50  may be constructed as an annular ring to be fitted within the annular gap  48 . The M-shaped sealing rings may also be provided with a split therein to allow circumferential expansion if necessary. 
   Referring now to  FIG. 4 , the M-shaped sealing rings  50  comprise preferably three substantially evenly spaced peaks, namely two outer peaks  52  which abut the inner surface  49  of the outer shroud housing  44 , and a central inner peak  56  which abuts the outer surface  43  of the axial projections  41  of the shroud segments. Thus, the M-shaped sealing rings  50  provide radially inward loading of the shroud segments  40  within the fixed outer housing  44 . This configuration of the M-shaped sealing rings permits sufficient radial inward force to be exerted on the shroud segments  40 , thus preventing unwanted outward movement thereof during the above described shroud grinding operation. The M-shaped sealing rings  50  therefore provide a simple and cost effective seal which acts to load the component, particularly to load the turbine shroud segments  40  radially inward, while maintaining a sealing ability sufficient to prevent leakage of cooling air. 
   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 department from the scope of the invention disclosed. For example, the sealing rings of the present invention may be made of any material suitable for the given application. Further, while the ring seals of the present invention are preferably M-shaped with three alternating peaks, they nevertheless comprise a wave pattern having at least two alternating peaks and may include more than three while maintaining a configuration which is cost effective to manufacture. 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.