Patent Publication Number: US-9410447-B2

Title: Forward compartment service system for a geared architecture gas turbine engine

Description:
BACKGROUND 
     The present disclosure claims priority to U.S. Provisional Patent Disclosure Ser. No. 61/677,284, filed Jul. 30, 2012. 
    
    
     The present disclosure relates to a gas turbine engine, and in particular, to a case structure that provides a service pathway around a geared architecture. 
     Gas turbine engines with geared architectures may utilize epicyclic reduction gearbox for their compact design and efficient high gear reduction capabilities. The reduction gearbox of the geared architecture isolates and de-couples the fan and low spool, which may result in isolation of the forwardmost bearing compartment from service pathways. 
     SUMMARY 
     A gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a first component that defines a first passage and a jumper tube that extends through the first passage. 
     In a further embodiment of the foregoing embodiment, the first component is an engine case. 
     In a further embodiment of any of the foregoing embodiments, the jumper tube extends from a second component. In the alternative or additionally thereto, the foregoing embodiment includes the second component is a bearing support. In the alternative or additionally thereto, the foregoing embodiment includes the first component is a fan inlet case and the second component is a #1/#1.5 bearing support. 
     In a further embodiment of any of the foregoing embodiments, the jumper tube is resiliently mounted within the first passage. 
     In a further embodiment of any of the foregoing embodiments, further comprising a flange that mounts the jumper tube to the first component. In the alternative or additionally thereto, the foregoing embodiment includes the flange defines an opening in communication with a bore through the jumper tube. 
     In a further embodiment of any of the foregoing embodiments, the jumper tube includes a lateral opening. In the alternative or additionally thereto, the foregoing embodiment includes the lateral opening communicates with one of the first component and the second component. 
     In a further embodiment of any of the foregoing embodiments, the jumper tube communicates with a hollow strut. 
     A gas turbine engine according to another disclosed non-limiting embodiment of the present disclosure includes a fan inlet case that defines a first passage, the fan inlet case includes a hollow strut, a bearing support that defines a second passage, and a jumper tube that extends through the first passage and the second passage to communicate with the hollow strut. 
     In a further embodiment of the foregoing embodiment, the jumper tube includes a lateral opening. In the alternative or additionally thereto, the foregoing embodiment includes the lateral opening communicates with the bearing support. 
     In a further embodiment of any of the foregoing embodiments, further comprising a flange that mounts the jumper tube to one of the first component and the second component. In the alternative or additionally thereto, the foregoing embodiment includes the flange defines an opening in communication with a bore through the jumper tube. 
     A method of assembling a gas turbine engine, according to another disclosed non-limiting embodiment of the present disclosure includes assembling a first component that defines a first passage to a second component that defines a second passage, and inserting a jumper tube through the first passage and the second passage. 
     In a further embodiment of the foregoing embodiment, comprising resiliently mounting the jumper tube with a multiple of seals. 
     In a further embodiment of any of the foregoing embodiments, further comprising providing a service pathway through the jumper tube. In the alternative or additionally thereto, the foregoing embodiment includes directing the service pathway through a lateral opening in the jumper tube. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiment. The drawings that accompany the detailed description can be briefly described as follows: 
         FIG. 1  is a schematic cross-section of a gas turbine engine; 
         FIG. 2  is an enlarged schematic cross-section of the gas turbine engine; 
         FIG. 3  is an enlarged schematic cross-section of a forward section of the gas turbine engine; 
         FIG. 4  is a side perspective exploded view of a #1/1.5 bearing support structure with a multiple of jumper tubes mounted therein; 
         FIG. 5  is a perspective view of a jumper tube according to one disclosed non-limiting embodiment; and 
         FIG. 6  is a perspective view of a jumper tube according to another disclosed non-limiting embodiment. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flowpath while the compressor section  24  drives air along a core flowpath for compression and communication into the combustor section  26  then expansion through the turbine section  28 . 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 such as a three-spool (plus fan) engine wherein an intermediate spool includes an intermediate pressure compressor (IPC) between the LPC and HPC and an intermediate pressure turbine (IPT) between the HPT and LPT. 
     The engine  20  generally includes a low spool  30  and a high spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing compartments  38 - 1 - 38 - 4 . The low spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  (“LPC”) and a low pressure turbine  46  (“LPT”). The inner shaft  40  drives the fan  42  through a geared architecture  48  to drive the fan  42  at a lower speed than the low spool  30 . An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system. 
     The high spool  32  includes an outer shaft  50  that interconnects a high pressure compressor  52  (“HPC”) and high pressure turbine  54  (“HPT”). A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
     Core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed with the fuel and burned in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The turbines  54 ,  46  rotationally drive the respective low spool  30  and high spool  32  in response to the expansion. 
     The main engine shafts  40 ,  50  are supported within the static structure  36  at a plurality of points by bearing compartments  38 - 1 - 38 - 4 . In one non-limiting embodiment, a # 1  bearing compartment  38 - 1  located radially inboard of the fan section  22 . 
     In one non-limiting example, the gas turbine engine  20  is a high-bypass geared aircraft engine. In a further example, the gas turbine engine  20  bypass ratio is greater than about six (6:1). The geared architecture  48  can include an epicyclic gear train, such as a planetary gear system or other gear system. The example epicyclic gear train has a gear reduction ratio of greater than about 2.3, and in another example is greater than about 2.5:1. The geared turbofan enables operation of the low spool  30  at higher speeds which can increase the operational efficiency of the low pressure compressor  44  and low pressure turbine  46  and render increased pressure in a fewer number of stages. 
     A pressure ratio associated with the low pressure turbine  46  is pressure measured prior to the inlet of the low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle of the gas turbine engine  20 . In one non-limiting embodiment, the bypass ratio of the gas turbine engine  20  is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about five (5:1). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans. 
     In one embodiment, a significant amount of thrust is provided by the bypass flow path B due to the high bypass ratio. The fan section  22  of the gas turbine engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with the gas turbine engine  20  at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust. 
     Fan Pressure Ratio is the pressure ratio across a blade of the fan section  22  without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine  20  is less than 1.45. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of (“T”/518.7) 0.5 . in which “T” represents the ambient temperature in degrees Rankine. The 
     Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine  20  is less than about 1150 fps (351 m/s). 
     With reference to  FIG. 2 , the engine case structure  36  proximate the compressor section  24  generally includes a fan inlet case  60  with a multiple of hollow struts  62 . The multiple of hollow fan struts  62  may also be referred to as “wet struts” that provide services pathways across a primary airflow path  64 . The services pathways may terminate at a rear bulkhead  65  radially outward of the primary airflow path  64  where services may be readily connected. 
     The fan inlet case  60  defines the annular primary airflow path  64  to direct core airflow into the LPC  44 . The fan inlet case  60  mounts a #1/1.5 bearing support structure  66  therein to define a front bearing compartment  38 - 1 . The frustro-conical shaped #1/1.5 bearing support structure  66  beneficially mounts closely within a frustro-conical fan hub to facilitate a more compact arrangement. It should be appreciated that various case structures may alternatively or additionally be provided, yet benefit from the architecture described herein. The #1/1.5 bearing support structure  66  supports a # 1  bearing  68 , a #1.5 bearing  70 , one or more seals  72  and the geared architecture  48 . The # 1  bearing  68  and the #1.5 bearing  70  rotationally support rotation of a fan output shaft  74  that connects the LPC  44  with the geared architecture  48  to drive the fan  42 . The seals  72  contain oil to define a “wet” front bearing compartment  38 - 1 . For ease of reference, regions or volumes that contain oil may be referred to as a “wet” zone and an oil-free region may be referred to as a “dry” zone. So, for example, the interior of each bearing compartment  38 - 1  may be referred to as a wet zone that ultimately communicates with an oil sump while the regions external thereto may be referred to as a dry zone. 
     With reference to  FIG. 3 , the #1/1.5 bearing support structure  66  mounts to the fan inlet case  60  with fasteners  76  and to a # 1  seal support  78  with fasteners  80  such as a respective ring of bolts. The #-1/#1.5 bearing support structure  66  and the fan inlet case  60  may be manufactured as cast components with respective passages  82 ,  84  that are integrally cast therein. The cast passages  82 ,  84  provide for cooling, lubrication or other service pathways, but, being cast, may not be air or even fluid tight. 
     A multiple of jumper tubes  88  are mounted within the #1/1.5 bearing support structure  66  ( FIG. 4 ) to provide a sealed services pathway between the passages  82 ,  84  across an interface F 2  and the hollow struts  62 . That is, each jumper tube  88  provides an air or fluid tight services pathway within the the passages  82 ,  84  across the interface F 2  to supply or remove various gaseous or liquid fluids. The jumper tubes  88  may also be utilized to guide wire harnesses or other conduits to and from the relatively remote front bearing compartment  38 - 1 . The jumper tubes  88 , although illustrated as independent components in the disclosed non-limiting embodiment, may alternatively be integral to other structure such as the #1/1.5 bearing support structure  66 . The jumper tubes  88  may also facilitate “blind” assembly. 
     Furthermore, the jumper tubes  88  may provide service communication for needs other than the bearing compartment. For example, de-icing air for a fan nosecone  42 N may be routed in the same way—but is not used by the bearing compartment. 
     With reference to  FIG. 5 , each jumper tube  88 , in one disclosed non-limiting embodiment, includes a multiple of seal grooves  90  each of which may receive a seal  92  such as an O-ring to seal with the passages  82 ,  84  as well as accommodate relative motion and manufacturing tolerances there between. That is, the interfaces provided by the seals  92  between the jumper tube  88  and the passages  82 ,  84  are essentially resilient. 
     A lateral opening  94  through the wall of the jumper tube  88  provides for communication there through (illustrated schematically by arrow C). The jumper tube  88  may have particular applicability, but not be limited to, fluid transfer for communication of, for example, oil “wet” or buffer air “dry”. 
     A flange  96  defines a distal end of the jumper tube  88  to mount the jumper tube  88  to the #1/1.5 bearing support structure  66  with fasteners  98  such as bolts. The flange  96  may include a tab, an oval shape or other shape to receive the fastener  98  generally parallel to the jumper tube  88 . The fasteners  98  readily thread and thereby mount the jumper tube  88  into the #1/1.5 bearing support structure  66 . It should be appreciated that various fasteners and mount arrangements may alternatively or additionally be provided. 
     The jumper tube  88  facilitates assembly of the gas turbine engine  20  and formation of sealed services pathways in communication with the forward bearing compartment  38 - 1 . That is, the jumper tube  88  may be assembled after the #1/1.5 bearing support structure  66  and # 1  bearing compartment  38 - 1  are mounted within the fan inlet case  60 . The jumper tubes  88  provide a continuous sealed services pathway through a multiple engine components, e.g., the #1/1.5 bearing support structure  66  and the fan inlet case  60  to provide service around the geared architecture  48  to and from the hollow strut  62 . The jumper tubes  88  also facilitate the assembly of the geared architecture  48  without resort to “blind assembly”. 
     With reference to  FIG. 6 , a jumper tube  88 ′ in another disclosed non-limiting embodiment includes an open distal end  100  through the flange  96 ′ to define an axial services pathway along a through bore  102  defined along a jumper tube axis T′. The jumper tube  88 ′ may have, but not be limited to, particular applicability for conduit, wire harnesses, cable, etc. 
     It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit here from. 
     Although the different examples have specific components shown in the illustrations, embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples. 
     The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be understood that within the scope of the appended claims, the invention may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content.