Patent Document

TECHNICAL FIELD 
     The described subject matter relates generally to gas turbine engines, and more particularly, to a fan and boost joint. 
     BACKGROUND OF THE ART 
     Aircraft gas turbine turbofan engines generally include a low pressure spool assembly having a fan rotor, low pressure compressor and a low pressure turbine connected by a low pressure spool shaft, and a high pressure spool assembly having a high pressure compressor and a high pressure turbine connected by a high pressure spool shaft which is hollow and disposed coaxially around the low pressure spool shaft. Conventionally, the fan rotor and the low pressure compressor, particularly a boost stage which is positioned upstream of the low pressure compressor, are tied together on the low pressure spool shaft, for example by a spline and a spigot arrangement. During flight, a bird strike event and other blade-off loads which create imbalanced loads to the fan rotor, may cause a fan rotor deflection. The fan rotor deflection may be transmitted downstream to the boost stage of the low pressure compressor to cause the boost stage to move with the fan rotor deflection, due to the fact that they are tied together on the low pressure spool shaft. The boost stage deflection affects tip clearance on the boost stage of the low pressure compressor, thereby further affecting the performance of the gas turbine engine. 
     Accordingly, there is a need to provide an improved fan rotor and boost compressor joint in aircraft gas turbine engines. 
     SUMMARY 
     In one aspect, the described subject matter provides a gas turbine engine having at least one spool assembly, the at least one spool assembly comprising a fan rotor, a compressor disposed downstream of the fan rotor, a turbine and a shaft connecting the fan rotor, compressor and turbine, a joint affixed to an upstream end of the shaft, the joint including a first link connecting the fan rotor to the shaft and a second link connecting the compressor to the shaft, the second link being less rigid than the first link wherein the first link comprises an annular front leg extending generally radially outwardly from the shaft, and wherein the second link comprises an annular rear leg extending generally radially outwardly from the shaft. 
     In another aspect, the described subject matter provides a gas turbine engine having at least one spool assembly, the at least one spool assembly comprising a fan rotor, a compressor disposed downstream of the fan rotor, a turbine and a shaft connecting the fan rotor, compressor and turbine, means affixed to an upstream end of the shaft for connecting the fan rotor to the shaft in a first link and for connecting the compressor to the shaft in a second link, the second link being less rigid than the first link. 
     In a further aspect, the described subject matter provides a method for disassociating a fan rotor deflection from a compressor deflection during an undue imbalance event of a fan rotor in a gas turbine engine, the method comprising: a) connecting a fan rotor to an engine shaft by a first link, the link frustoconically extending outwardly of an upstream end of the shaft; and b) connecting a compressor to the engine shaft by a second link, the second link frustoconically extending outwardly of the upstream end of the shaft, the second link being less rigid than the first link. 
     Further details of these and other aspects of the described subject matter will be apparent from the detailed description and drawings included below. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Reference is now made to the accompanying drawings depicting aspects of the described subject matter, in which: 
         FIG. 1  is a schematic cross-sectional view of a turbofan gas turbine engine, showing one embodiment of the described subject matter; and 
         FIG. 2  is a partial cross-sectional view in an enlarged scale, of the circled area  2  of  FIG. 1 , showing a structural arrangement of one embodiment. 
     
    
    
     It will be noted that throughout the appended drawings, like features are identified by like reference numerals. 
     DETAILED DESCRIPTION 
       FIG. 1  illustrates a turbofan gas turbine engine according to one embodiment. The engine includes a housing or nacelle  10 , a core casing  13 , a low pressure spool assembly (not numbered) which includes a fan rotor  14 , a low pressure compressor assembly having a boost compressor  16  and a low pressure turbine assembly  18  connected by a shaft  12 , and a high pressure spool assembly (not numbered) which includes a high pressure compressor assembly  22  and a high pressure turbine assembly  24  connected by a turbine shaft  20 . The housing or nacelle  10  surrounds the core casing  13  and in combination the housing  10  and the core casing  13  define an annular bypass duct  28  for directing a bypass airflow. The core casing  13  surrounds the low and high pressure spool assemblies to define a core fluid path  30  therethrough. In the core fluid path  30  there is provided a combustor  26  to form a combustion gas generator assembly which generates combustion gases to power the high pressure turbine assembly  24  and the low pressure turbine assembly  20 . The boost compressor  16  is disposed downstream of the fan rotor  14  and together with the fan rotor  14 , is connected to the shaft  12  via a joint  32 , as schematically shown in the circled area  2  and will be further described hereinafter. 
     The terms “upstream” and “downstream” mentioned in the description below generally refer to the airflow direction through the engine and are indicated by an arrow in  FIG. 1 . The terms “front” and “rear” generally refer to a position sequence from the front to the rear of the engine in a direction as indicated by the arrow in  FIG. 1 . The terms “axial”, “radial” and “circumferential” used for various components below are defined with respect to the main engine axis shown but not numbered in  FIG. 1 . 
     According to one embodiment illustrated in  FIGS. 1 and 2 , the shaft  12  is supported by a bearing assembly  34  disposed around the shaft  12  adjacent to an upstream end  36  of the shaft  12 . The bearing assembly  34  is supported by a stationary structure (not shown) of the engine. The upstream end  36  of the shaft  12  is integrated with the joint  32 . The joint  32  according to this embodiment may have an annular joint body  38  extending generally radially outwardly from the upstream end  36  of the shaft  12 . An annular front leg  40  extends generally radially and outwardly, from the annular joint body  38  to form a first link for connection with the fan rotor  14 . An annular rear leg  42  disposed downstream of the annular front leg  40  and extends generally radially and outwardly from the annular joint body  38  to form a second link for connection with the boost compressor  16 . The joint  32  with the annular front and rear legs  40 ,  42  may expand frustoconically forwardly and rearwardly, respectively, from the annular joint body  38  to form a substantial Y-shaped configuration in a cross-section thereof, as shown in the  FIGS. 1 and 2 . 
     The annular front leg  40  may have a thickness greater than the thickness of the annular rear leg  42 . The annular front leg  40  may also be shorter than the annular rear leg  42 . The annular joint body  38  may have a thickness greater than the thickness of the respective annular front and rear legs  40 ,  42 . Therefore, the joint  32  provides the second link connecting the boost compressor  16  to the shaft  12 , less rigid than the first link connecting the fan rotor  14  to the shaft  12 . The less rigidity and thus relative flexability of the second link provided by the annular rear leg  42  with respect to the first link provided by the annular front leg  40 , reduces transmissibility of deflection through the joint  32  from the fan rotor  14  to the boost compressor  16 , thereby substantially maintaining the tip clearance of the boost compressor  16  during a bird ingestion or other blade detachment event occurring to the fan rotor  14 . 
     According to one embodiment, the fan rotor  14  may include a rearwardly and inwardly extending annular web  44  and an annular flange  46  extending radially and inwardly from a rear end (not numbered) of the annular web  44 . A plurality of holes  48  may be provided in the flange  46  of the of the fan rotor  14 , circumferentially spaced apart one from another. A plurality of holes  50  may be provided in the annular front leg  40 , circumferentially spaced apart one from another and aligning with the respective holes  48  in the flange  46  of the fan rotor  14 , to receive fasteners or fastener assemblies  52  which extend axially therethrough for securing the fan rotor  14  to the annular front leg  40  of the joint  38 . Each of the fastener assemblies  52  may include a fastener, washer, nut, lock element, etc. 
     According to one embodiment, the boost compressor  16  may include a forwardly and inwardly extending annular web  54  and an annular flange  56 , extending radially and inwardly from a front end (not numbered) of the annular web  54 . A plurality of holes  58  may be provided in the annular flange  56  of the boost compressor  16 , circumferentially spaced apart one from another. A plurality of holes  60  may also be provided in the annular leg  42  adjacent an outer periphery of the annular rear leg  42 , circumferentially spaced apart one from another and aligning with the respective holes  58 , in order to receive respective fasteners or fastener assemblies  62  which extend axially therethrough for securing the boost compressor  16  to the annular rear leg  42  of the joint  32 . Each of the fastener assemblies  62  may include a fastener, washer, nut, lock element, etc. 
     Optionally, the annular web  44  of the fan rotor  14  may have a thickness greater than the thickness of the annular web  54  of the boost compressor  16 , in order to further reduce deflection transmissibility from the fan rotor  14  to the boost compressor  16 . 
     Alternatively, the joint  32  need not necessarily be integrated with the upstream end of  36  of the shaft  12 . The joint  32  may be removably connected to the shaft  12  by any known or unknown suitable mechanism. 
     Alternatively, the annular front leg  40  of the joint  32  may be replaced by three or more front legs extending radially and outwardly from the annular joint body  38 , circumferentially spaced apart one from another. 
     Similarly, the annular rear leg  42  of the joint  32  may be alternatively replaced with three or more rear legs radially and outwardly extending from the annular joint body  38 , circumferentially spaced apart one from another. 
     Also alternatively, the annular webs  44 ,  54  of the respective fan rotor  14  and boost compressor  16  may be replaced by any suitable mounting apparatus of the respective fan rotor  14  and boost compressor  16 . 
     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 departure from the scope of the described subject matter. For example, the schematically illustrated turbofan gas turbine engine is an exemplary application of the described subject matter and the described subject matter may also be applicable in gas turbine engines of various types. Still other modifications which fall within the scope of the described subject matter 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 Category: 2