Patent Publication Number: US-11022144-B2

Title: Diffuser case assembly

Description:
This invention was made with Government support awarded by the United States. The Government has certain rights in this invention. 
    
    
     BACKGROUND 
     1. Technical Field 
     This disclosure relates generally to gas turbine engines, and more particularly to diffuser case assemblies. 
     2. Background Information 
     During operation of a gas turbine engine, heated core gases flow from a compressor section to a combustor section where they are mixed with fuel and ignited. Elevated core gas temperatures may induce large thermal gradients on engine components in the core flowpath. 
     For example, during a transient acceleration from idle to takeoff power, a support structure for an inner diffuser case, forming part of the core flowpath, may rapidly reach takeoff metal temperatures. The resulting thermal gradient may create excessive stress concentrations at intersections of comparatively hotter and colder portions of the diffuser cases and associated support structure. The thermal stress concentrations are exacerbated by the need for the inner diffuser case structure to be stiff enough to support a shaft bearing of the gas turbine engine. 
     SUMMARY 
     According to an embodiment of the present disclosure, a diffuser case assembly for a gas turbine engine includes a fairing disposed circumferentially about a longitudinal axis. The fairing defines a plurality of passages circumferentially spaced apart and forming at least a portion of a fluid path between a compressor and a combustor of the gas turbine engine. A diffuser frame includes a plurality of struts. Each of the plurality of struts is disposed between a pair of adjacent passages of the plurality of passages. The diffuser frame is configured to couple an inner diffuser case to an outer diffuser case. 
     In the alternative or additionally thereto, in the foregoing embodiment, a space between each pair of adjacent passages of the plurality of passages defines a recessed portion of the fairing extending axially from an axial end of the fairing through a portion of the fairing. 
     In the alternative or additionally thereto, in the foregoing embodiment, the diffuser frame and the inner diffuser case form an integral component. 
     In the alternative or additionally thereto, in the foregoing embodiment, at least one of the struts is hollow. 
     In the alternative or additionally thereto, in the foregoing embodiment, the at least one hollow strut defines a channel extending radially through the strut. 
     In the alternative or additionally thereto, in the foregoing embodiment, the diffuser frame is physically independent of the fairing. 
     In the alternative or additionally thereto, in the foregoing embodiment, the diffuser frame is made of a first material and the fairing is made of a second material different than the first material. 
     In the alternative or additionally thereto, in the foregoing embodiment, the diffuser case assembly further includes at least one seal disposed between the fairing and the diffuser frame. 
     In the alternative or additionally thereto, in the foregoing embodiment, the diffuser case assembly further includes a sliding joint forming an interface between the fairing and the diffuser frame. 
     In the alternative or additionally thereto, in the foregoing embodiment, the sliding joint is configured to move radially in response to at least one of thermal expansion and contraction of the fairing in a radial direction. 
     In the alternative or additionally thereto, in the foregoing embodiment, the channel is configured to conduct a flow of fluid between a compartment radially outside the inner diffuser case to a compartment radially inside the inner diffuser case. 
     In the alternative or additionally thereto, in the foregoing embodiment, an auxiliary line extends through the channel. 
     In the alternative or additionally thereto, in the foregoing embodiment, the fairing is a single-piece casting. 
     According to another embodiment of the present disclosure, a diffuser case assembly for a gas turbine engine includes a fairing disposed circumferentially about a longitudinal axis and a diffuser frame including a plurality of hollow struts. The fairing defines a plurality of passages circumferentially spaced apart and forming at least a portion of a fluid path between a compressor and a combustor of the gas turbine engine and a space between each pair of adjacent passages of the plurality of passages. The space defines a recessed portion of the fairing extending axially from an axial end of the fairing through a portion of the fairing. Each strut of the plurality of struts defines a channel extending radially through the strut and each strut of the plurality of struts is disposed between a pair of adjacent passages of the plurality of passages. The diffuser frame is configured to couple an inner diffuser case to an outer diffuser case. 
     In the alternative or additionally thereto, in the foregoing embodiment, the diffuser frame and the inner diffuser case form an integral component. 
     In the alternative or additionally thereto, in the foregoing embodiment, the diffuser frame is physically independent of the fairing. 
     According to another embodiment of the present disclosure, a gas turbine engine includes an inner diffuser case, an outer diffuser case, and a diffuser case assembly coupling the inner diffuser case to the outer diffuser case. The diffuser case assembly includes a fairing disposed circumferentially about a longitudinal axis. The fairing defines a plurality of passages circumferentially spaced apart and forming at least a portion of a fluid path between a compressor and a combustor of the gas turbine engine. A diffuser frame includes a plurality of struts. Each of the plurality of struts is disposed between a pair of adjacent passages of the plurality of passages. 
     In the alternative or additionally thereto, in the foregoing embodiment, the diffuser frame and the inner diffuser case form an integral component. 
     In the alternative or additionally thereto, in the foregoing embodiment, the diffuser frame is physically independent of the fairing. 
     In the alternative or additionally thereto, in the foregoing embodiment, the diffuser frame is made of a first material and the fairing is made of a second material different than the first material. 
     The present disclosure, and all its aspects, embodiments and advantages associated therewith will become more readily apparent in view of the detailed description provided below, including the accompanying drawings. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  illustrates a schematic cross-sectional view of a gas turbine engine. 
         FIG. 2  illustrates a cross-sectional side view of a diffuser case assembly of a gas turbine engine. 
         FIG. 3  illustrates a cross-sectional perspective view of a portion of the diffuser case assembly of  FIG. 2 . 
         FIG. 4  illustrates a cross-sectional perspective view of a portion of the diffuser case assembly of  FIG. 2 . 
         FIG. 5  illustrates an exploded view of the diffuser case assembly of  FIG. 2 . 
     
    
    
     DETAILED DESCRIPTION 
     It is noted that various connections are set forth between elements in the following description and in the drawings. It is noted that these connections are general and, unless specified otherwise, may be direct or indirect and that this specification is not intended to be limiting in this respect. A coupling between two or more entities may refer to a direct connection or an indirect connection. An indirect connection may incorporate one or more intervening entities. 
       FIG. 1  schematically illustrates a gas turbine engine  10 . The gas turbine engine  10  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  12 , a compressor section  14 , a combustor section  16 , and a turbine section  18 . The fan section  12  drives air along a bypass flowpath B while the compressor section  14  drives air along a core flowpath C for compression and communication into the combustor section  16  then expansion through the turbine section  18 . Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
     The gas turbine engine  10  generally includes a low-speed spool  20  and a high-speed spool  22  mounted for rotation about an engine central longitudinal axis  24  relative to an engine static structure  26 . It should be understood that various bearing systems at various locations may alternatively or additionally be provided. 
     The low-speed spool  20  generally includes an inner shaft  28  that interconnects a fan  30 , a low-pressure compressor  32  and a low-pressure turbine  34 . The inner shaft  28  is connected to the fan  30  through a geared architecture  36  to drive the fan  30  at a lower speed than the low-speed spool  20 . The high-speed spool  22  includes an outer shaft  38  that interconnects a high-pressure compressor  40  and high-pressure turbine  42 . A combustor  44  is arranged between the high-pressure compressor  40  and high-pressure turbine  42 . 
     The core airflow is compressed by the low-pressure compressor  32  then the high-pressure compressor  40 , passed through a diffuser case assembly  60 , mixed and burned with fuel in the combustor  44 , and then expanded over the high-pressure turbine  42  and the low-pressure turbine  34 . The turbines rotationally drive the respective low-speed spool  20  and high-speed spool  22  in response to the expansion. 
       FIG. 2  illustrates a cross-sectional view of the diffuser case assembly  60  of the gas turbine engine  10  illustrating the high-pressure compressor  40 , the combustor  44 , and the core flowpath C therebetween. An exit guide vane  46  is positioned within the core flowpath C immediately aft of the high-pressure compressor  40  and alters flow characteristics of core gases exiting the high-pressure compressor  40 , prior to the gas flow entering the combustor  44 . 
     Referring to  FIGS. 2-5 , a fairing  48  is disposed immediately aft of the exit guide vane  46  and forms at least a portion of the core flowpath C (i.e., providing fluid communication) between the high-pressure compressor  40  and the combustor  44 . The fairing  48  is disposed circumferentially (e.g., annularly) about the longitudinal axis  24  of  FIG. 1 . The fairing  48  includes a plurality of passages  52  extending (e.g., generally axially) through the fairing  48  and configured to form the core flowpath C through the fairing  48  between the high-pressure compressor  40  and the combustor  44 . The fairing  48  further includes a plurality of recessed portions  50  defined between adjacent passages  52  of the fairing  48 . For example, the recessed portions  50  may extend axially from an aft axial end (i.e., an end proximate the combustor  44 ) of the fairing  48  through a portion of the fairing  48 . In some embodiments, each recessed portion of the plurality of recessed portions  50  may be disposed between each respective pair of circumferentially adjacent passages of the plurality of passages  52 . In some embodiments, the fairing  48  may be configured as a single piece, for example a single-piece casting or a fully machined component. In some other embodiments, the fairing  48  may be configured as a plurality of circumferential segments subsequently assembled (e.g., welded or otherwise attached together) to form the fairing  48 . 
     Annular inner and outer diffuser cases  54 ,  56  radially house the fairing  48 . The outer diffuser case  56  is disposed radially outward of the fairing  48 . The inner diffuser case  54  is disposed radially inward of the fairing  48 . In some embodiments, the inner and outer diffuser cases  54 ,  56  may extend generally axially through all or part of the compressor section  14  and/or the combustor section  16 . The inner and outer diffuser cases  54 ,  56  mechanically support structures of the gas turbine engine  10 , for example, the inner diffuser case  54  may support a shaft bearing of the gas turbine engine  10 . 
     The inner diffuser case  54  includes a diffuser frame  58  which extends between and couples the inner diffuser case  54  and the outer diffuser case  56 . The inner diffuser case  54 , outer diffuser case  56 , and diffuser frame  58  form a diffuser case assembly  60  (i.e., a “cold structure” in contrast to the “hot” fairing  48 ). In some embodiments, the diffuser frame  58  and the inner diffuser case  54  may form a single integral component. 
     The diffuser frame  58  includes a plurality of circumferentially spaced-apart struts  62  with each strut of the plurality of struts  62  configured to radially extend through the fairing  48  between a pair of adjacent passages of the plurality of passages  52 . For example, each strut of the plurality of struts  62  may be disposed within a respective recessed portion of the plurality of recessed portions  50 . In some embodiments, each pair of adjacent passages of the plurality of passages  52  may correspond to a respective strut of the plurality of struts  62 , i.e., a strut of the plurality of struts  62  may radially extend through the fairing  48  between each pair of adjacent passages of the plurality of passages  52 . In other embodiments, a quantity of the plurality of struts  62  may be less than a quantity of the plurality of passages  52 . For example, each strut of the plurality of struts  62  may radially extend through the fairing  48  between each other pair, each third pair, etc. of adjacent passages of the plurality of passages  52  or any other suitable configuration of the plurality of struts  62  and the plurality of passages  52 . This configuration may provide for simpler assembly by allowing the diffuser case assembly  60  to be installed and then allowing the fairing  48  to be installed between the plurality of struts  62  from the forward end (see, e.g.,  FIG. 5 ). In some embodiments, the diffuser frame  58  may be physically independent of the fairing  48  (i.e., there is no physical contact between the diffuser frame  58  and the fairing  48 ). 
     As shown in  FIGS. 4 and 5 , in some embodiments, at least one strut of the plurality of struts  62  may be hollow, thereby defining a channel  86  extending radially through the at least one hollow strut. A hollow configuration of the plurality of struts  62  may provide a reduction in the weight of the diffuser case assembly  60 . The channel  86  may be configured to conduct a flow of fluid (e.g., cooling air), for example, between a compartment radially outside the inner diffuser case  54  to a compartment radially inside the inner diffuser case  54 . 
     During operational transients of the gas turbine engine  10 , the fairing  48  may experience an increased flow of hot gases along the core flowpath C. For example, during a transient acceleration from idle to takeoff power, the increase flow of hot gases through the fairing  48  may cause the fairing  48  to rapidly increase in temperature. Separation of the core flowpath C from the diffuser case assembly  60  (i.e., the “cold structure”) by the fairing  48  may prevent the development of large thermal gradients across one or both of the diffuser case assembly  60  and the fairing  48 . As a result, the temperature of the fairing  48  may increase while the diffuser case assembly  60  remains at a more constant, lower temperature compared to the fairing  48 . Similarly, the fairing  48  may achieve a more constant, higher temperature compared to the diffuser case assembly  60 . Thus, thermal stress concentrations, for example, between the diffuser frame  58  and the inner diffuser case  54  or across the fairing  60  may be reduced as a result of the minimized thermal gradients. 
     The fairing  48  may include one or more seals  68  between the fairing  48  and the diffuser case assembly  60 . In the illustrated embodiment, the fairing  48  includes a seal  68  between the fairing  48  and the inner diffuser case  54 . The fairing  48  includes an additional seal  68  between the fairing  48  and a seal carrier  84 . The seals  68  may be configured to maintain the seal between the diffuser case assembly  60  and the fairing  48  as the fairing  48  expands and contracts (e.g., in a radial, axial, etc. direction), independent of the diffuser case assembly  60 , as a result of changes in the temperature of the fairing  48 . The seals  68  may be configured, for example, as piston seals or any other suitable type of seal. In other embodiments, the number and location of the seals  68  may vary according to diffuser case assembly  60  configuration. In some embodiments, the seal carrier  84  may include a retaining ring  88  configured to maintain the sealing function of the seal carrier  84  in response to radial movement of the fairing  48 . In some embodiments, the diffuser case assembly  60  may include a mixing seal  70  configured to provide a seal between an aft portion of the diffuser frame  58  and the outer diffuser case  56 . 
     The diffuser case assembly  60  may include at least one sliding joint  72  to provide a support interface between the fairing  48  and the diffuser case assembly  60 , while still allowing the fairing  48  to thermally expand and contract. In the illustrated embodiment, the at least one sliding joint  72  includes an alignment pin  74  extending radially inward from the diffuser frame  58 . The alignment pin  74  mates with a pin bushing  76  disposed on the fairing  48  (i.e., a pin boss configuration), thereby movably supporting the fairing  48  by allowing relative radial movement between the fairing  48  and the alignment pin  74 . For example, the alignment pin  74  may move radially within the pin bushing  76  in response to at least one of thermal expansion and contraction of the fairing  48  in a radial direction. 
     As discussed above, the gas turbine engine  10  transients may cause the fairing  48  to thermally expand or contract while the diffuser case assembly  60  maintains a more consistent and cooler temperature. Accordingly, in some embodiments, the diffuser frame  58  may be made from a first material while the fairing  48  is made from a second material, different than the first material. For example, the fairing  48  may be made from a high-temperature resistant material (e.g., waspaloy, nickel-based alloys, ceramics, ceramic matrix composites, etc.) while the diffuser frame  58  is made from a comparatively stronger material (e.g., Inconel  718 , titanium, etc.) for improved support and structural stiffness of the diffuser case assembly  60 . 
     In some embodiments, one or more auxiliary lines  78  may extend through one or both of an aperture  64  of the outer diffuser case  56  and a channel  86  of the plurality of struts  62 . For example, the at least one auxiliary line  78  may be a bearing service line configured to convey oil to or from a bearing of the gas turbine engine  10 . 
     While various aspects of the present disclosure have been disclosed, it will be apparent to those of ordinary skill in the art that many more embodiments and implementations are possible within the scope of the present disclosure. For example, the present disclosure as described herein includes several aspects and embodiments that include particular features. Although these particular features may be described individually, it is within the scope of the present disclosure that some or all of these features may be combined with any one of the aspects and remain within the scope of the present disclosure. Accordingly, the present disclosure is not to be restricted except in light of the attached claims and their equivalents.