Patent Publication Number: US-10767555-B2

Title: Fan drive gear system module and inlet guide vane coupling mechanism

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application is a continuation of U.S. application Ser. No. 14/429,071 filed Mar. 18, 2015 which is a United States National Phase of PCT Application No. PCT/US2013/059506 filed on Sep. 12, 2013 which claims priority to U.S. Provisional Application No. 61/789,207 filed on 15 Mar. 2013 and U.S. Provisional Application No. 61/703,489 filed on 20 Sep. 2012. 
    
    
     BACKGROUND 
     This disclosure relates to a coupling mechanism for removably securing a fan drive gear system module from the rest of the gas turbine engine. 
     A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines. 
     The high pressure turbine drives the high pressure compressor through an outer shaft to form a high spool, and the low pressure turbine drives the low pressure compressor through an inner shaft to form a low spool. The fan section may also be driven by the low inner shaft. A direct drive gas turbine engine includes a fan section driven by the low spool such that the low pressure compressor, low pressure turbine and fan section rotate at a common speed in a common direction. One type of gas turbine engine uses a geared architecture between the turbine section and the fan section, which reduces the rotational speed of the fan section. 
     Turbine engine manufacturers, including those of geared gas turbine engines, continue to seek further improvements to engine performance and assembly including improvements to manufacture, maintainability, thermal, transfer, and propulsive efficiencies. 
     SUMMARY 
     In one exemplary embodiment, a gas turbine engine includes a fan that includes a plurality of fan blades rotatable about an axis. An on-wing portion includes a compressor section and a combustor that is in fluid communication with the compressor section. A turbine section is in fluid communication with the combustor. A core flow path is arranged within a core nacelle. The fan is arranged upstream from the core flow path. A fan drive gear system module is coupled to the turbine section for rotating the fan about the axis. A connector assembly including first and second members respectively is secured to the on-wing portion and the fan drive gear system module. The first and second members are removably secured to one another by radially extending fasteners. The first members are connected to an on-wing portion of the gas turbine engine and the second members are connected to the fan drive gear system module. The fasteners are accessible through the bypass flow path. 
     In a further embodiment of any of the above, the fasteners are captured within the fan drive gear system module to prevent the fasteners from entering the flow path. 
     In a further embodiment of any of the above, the first and second members include complementary shaped mating contoured surfaces configured to align the fan drive gear system module with respect to structure of the gas turbine engine. 
     In a further embodiment of any of the above, the contoured surfaces are chevron-shaped. 
     In a further embodiment of any of the above, the fasteners extend through the contoured surfaces. 
     In a further embodiment of any of the above, the fasteners are captured within the fan drive gear system module to prevent the fasteners from entering the core flow path. 
     In a further embodiment of any of the above, the first members are secured to an inner diameter shroud of an inlet guide vane assembly of the compressor section. The inner diameter shroud includes first and second portions arranged about bearing members of variable inlet guide vanes. A first set of fasteners secure the first member and the first and second portions to one another. 
     In a further embodiment of any of the above, the first and second members include complementary shaped mating contoured surfaces configured to align the fan drive gear system module with respect to the on-wing portion. The fasteners extend through the contoured surfaces. 
     In one exemplary embodiment, a method of servicing a gas turbine engine includes removing a fan section from an on-wing portion of a gas turbine engine to expose a bypass flow path and manipulating fasteners in a radial direction through the bypass flow path to separate a fan drive gear system module from the on-wing portion. 
     In a further embodiment of any of the above, the removing step includes removing a nose cone from a fan hub and removing a fan nut from the fan drive gear system module. 
     In a further embodiment of any of the above, a shaft nut is removed from a shaft to separate the fan drive gear system module from the on-wing portion. 
     In a further embodiment of any of the above, the manipulating step includes loosening the fasteners to detach the fan drive gear system module from an inlet guide vane inner shroud diameter. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: 
         FIG. 1  schematically illustrates a gas turbine engine embodiment. 
         FIG. 2  is an exploded view of a forward portion of the gas turbine engine shown in  FIG. 1 , including a geared architecture and a fan section. 
         FIG. 3  is a cross-sectional view of a portion of the gas turbine engine shown in  FIG. 2 , relating to a connection assembly including an inlet guide vane assembly. 
         FIG. 4A  illustrates a first member of the connection assembly. 
         FIG. 4B  illustrates a second member of the connection assembly. 
         FIG. 5  is a perspective view of the connection assembly shown in  FIG. 3 . 
         FIG. 6  is a cross-sectional view of another example connection assembly. 
         FIG. 7A  is a perspective view of the connection assembly shown in  FIG. 6 . 
         FIG. 7B  is an enlarged perspective view of a bracket of the connection assembly shown in  FIG. 7A . 
         FIG. 8  is a cross-sectional view depicting a fan drive gear system module removal procedure in which the connection assembly is disassembled. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates an example gas turbine engine  20  that includes a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmenter section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B while the compressor section  24  draws air in along a core flow path C where air is compressed and communicated to a combustor section  26 . In the combustor section  26 , air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section  28  where energy is extracted and utilized to drive the fan section  22  and the compressor section  24 . 
     Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, 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; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section. 
     The example engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided. 
     The low speed spool  30  generally includes an inner shaft  40  that connects fan blades  42  and a low pressure (or first) compressor section  44  to a low pressure (or first) turbine section  46 . The inner shaft  40  drives the fan blades  42  through a speed change device, such as a geared architecture  48 , to drive the fan blades  42  at a lower speed than the low speed spool  30 . The high-speed spool  32  includes an outer shaft  50  that interconnects a high pressure (or second) compressor section  52  and a high pressure (or second) turbine section  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via the bearing systems  38  about the engine central longitudinal axis A. 
     A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . In one example, the high pressure turbine  54  includes at least two stages to provide a double stage high pressure turbine  54 . In another example, the high pressure turbine  54  includes only a single stage. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. 
     The example low pressure turbine  46  has a pressure ratio that is greater than about 5. The pressure ratio of the example low pressure turbine  46  is measured prior to an inlet of the low pressure turbine  46  as related to the pressure measured at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. 
     A mid-turbine frame  58  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  58  further supports bearing systems  38  in the turbine section  28  as well as setting airflow entering the low pressure turbine  46 . 
     The core airflow C is compressed by the low pressure compressor  44  then by the high pressure compressor  52  mixed with fuel and ignited in the combustor  56  to produce high speed exhaust gases that are then expanded through the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  58  includes vanes  60 , which are in the core airflow path and function as an inlet guide vane for the low pressure turbine  46 . Utilizing the vane  60  of the mid-turbine frame  58  as the inlet guide vane for low pressure turbine  46  decreases the length of the low pressure turbine  46  without increasing the axial length of the mid-turbine frame  58 . Reducing or eliminating the number of vanes in the low pressure turbine  46  shortens the axial length of the turbine section  28 . Thus, the compactness of the gas turbine engine  20  is increased and a higher power density may be achieved. 
     The disclosed gas turbine engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine  20  includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example geared architecture  48  is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3. 
     In one disclosed embodiment, the gas turbine engine  20  includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor  44 . It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines. 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point. 
     “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.50. In another non-limiting embodiment the low fan pressure ratio is less than about 1.45. 
     “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)] 0.5 . The “Low corrected fan tip speed”, as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second. 
     The example gas turbine engine includes the fan blades  42  that comprises in one non-limiting embodiment less than about 26 fan blades. In another non-limiting embodiment, the fan section  22  includes less than about 20 fan blades. Moreover, in one disclosed embodiment the low pressure turbine  46  includes no more than about 6 turbine rotors schematically indicated at  34 . In another non-limiting example embodiment the low pressure turbine  46  includes about 3 turbine rotors. A ratio between the number of fan blades  42  and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine  46  provides the driving power to rotate the fan section  22  and therefore the relationship between the number of turbine rotors  34  in the low pressure turbine  46  and the number of fan blades  42  in the fan section  22  disclose an example gas turbine engine  20  with increased power transfer efficiency. 
     Bearing and gear serviceability inspection requires removal of the fan drive gear system (FDGS) Assembly. This inspection typically requires four weeks off the aircraft. The disclosed forward engine components include features for simplifying and accommodating this inspection. Assembly component modifications uniquely meet this serviceability requirement. 
     Referring to  FIG. 2 , the disassembly of the geared architecture  48  is schematically depicted in an exploded view. The disassembly permits removal of a fan drive gear system module  62  of the geared architecture  48  with the remainder of the gas turbine engine  20  remaining on-wing. “On-wing” does not necessarily require the engine  20  to be mounted on the aircraft, but means the portion of the engine including the compressor, combustor and turbine sections  24 ,  26  and  28 . 
     The geared architecture  48  includes a fan drive gear system module  62  that is removable as an assembly. During the disassembly procedure, a nose cone  64  is removed from a fan hub  66 , which supports the fan blades  42 . With the nose cone  64  removed, a fan hub nut  86  can be unthreaded from a fan shaft  68  to remove the fan hub  66  and fan blades  42 . A fan exit stator  70  is removed from a fan intermediate case  72 . A shaft nut  88  is unthreaded from the inner shaft  40  permitting the input coupling  74  along with the fan drive gear system module  62  to be detached from the inner shaft  40 . 
     The fan drive gear system module  62  includes a number one bearing  76  supporting the fan shaft  68  relative to the fan intermediate case  72 . A number two bearing  78  is arranged between a center body support  84  of the fan drive gear system module  62  and the input coupling  74 . 
     The fan intermediate case  72  includes inlet struts  71  and aft struts  73  arranged rearward of the inlet struts  71  which are arranged in the core flow path. The fan intermediate case  72  is supported relative to a fan case  90  by flow exit guide vanes  80 . The engine  20  is supported relative to an aircraft  94  by a pylon  92 . 
     In the example, the fan drive gear system module  62  generally includes the fan intermediate case  72 , center body support  84 , number one and two bearings  76 ,  78 , fan shaft  68 , input coupling  74  and gear train. 
     Before the fan intermediate case  72  can be separated from the inner shaft  40 , disassembly of the fan intermediate case  72  occurs at a connection assembly  96 , which removably secures an inlet guide vane assembly  82  relative to the fan intermediate case. In particular, an inner diameter shroud  98  of the inlet guide vane assembly  82  is detached from the fan intermediate case  72  in the area of the aft struts  73 . In the example, the inlet guide vane assembly vanes are variable in that individual guide vanes rotate about a radial axis. 
     The connection assembly  96  is shown in more detail in  FIGS. 3, 4A and 4B . First and second members are removably secured to one another by radially extending fasteners to decouple the fan drive gear system module  62 . 
     The inner diameter shroud  98  includes first and second portions  100 ,  102  secured to one another by a first fastener  108  that is oriented in an axial direction in the example. A bearing member  104  is retained between the first and second portions  100 ,  102  that is used to specifically support the inlet guide vanes. A first bracket  106  is also secured at the inner diameter shroud  98  by the first fastener  108 . 
     A second bracket  110  is supported by the fan intermediate case  72 . In one example, the bracket  110  is secured to the center body support  84  and structure providing the aft struts  73  by second fasteners  112 , which are oriented in an axial direction in the example shown. 
     Third fasteners  114  are supported by the second bracket  110  and secure the first bracket  106  to the second bracket  110 . In the example, the third fasteners  114  are oriented in a radial direction, which permits access through the core flow path to the connection assembly  96  during disassembly of the fan drive gear system module  62 . 
     A head  116  of the third fastener  114  is larger than an access hole  118  that is configured to enable a tool to be inserted through the access hole  118  to manipulate the third fasteners  114 . The third fastener  114  is sized such that when the third fastener  114  is disengaged from the first bracket  106 , the head  116  abuts a stop surface  124  of the fan intermediate case  72  to prevent complete removal and accidental loss of the third fasteners  114 . The second bracket  110  also includes a threaded hole  122  that may be provided by a helicoil, which retains the third fastener  114  in the disengaged position shown in  FIG. 4A . A radial lock  120 , which may be constructed from plastic, engages the threads of the third fastener  114  to provide a slight clamping load to the threads of the third fastener  114  during assembly and disassembly. The radial lock  120  includes a lever portion  119  that exerts a biasing force against the fastener to maintain it in the desired radially outward position. The example lever portion  119  comprises a polyimide spring ratcheting material with RTV filler  121  that prevents fracture and aids in maintaining the desired contact with the radially extending fastener. 
     The second and first brackets  110 ,  106  respectively include first and second contoured surfaces  126 ,  128  that are of a complimentary shape to one another. The complimentary shaped first and second contoured surfaces  126 ,  128 , which are chevron shapes in the example, ensure desired alignment of the first and second brackets  106 ,  110  with respect to one another during assembly as the third fasteners  114  are tightened. 
     In this example, four second brackets  110  are shown disposed about an inner circumference of the aft struts  73 . As appreciated, other numbers of bracket assemblies could be utilized as is required to provide the desired mounting and support for the fan drive gear system. Each of the brackets includes a structural cross member to attach the brackets such that the number of individual separate parts is limited. 
     Referring to  FIG. 4B , the first fasteners  108  each include a capture feature  134 , which ensures that the first fasteners  108  are retained with respect to the second portion  102 . A washer  136  and tab washer  138  are arranged beneath a head of the first fasteners  108  in the example. 
     Referring to  FIG. 5 , the apertures  140  in the second brackets  110  are shown. The second fasteners  112 , shown in  FIG. 4A , are inserted through the apertures  140  in an assembled condition. 
     Another example connection assembly  196  is shown in  FIGS. 6 and 7A . In this example, the first bracket  142  is secured on a forward side of the first portion  200 , which is arranged between the first bracket  142  and the second portion  202 . The first fastener  208  secures the first bracket  142  and the first and second portions  200 ,  202  of the inner diameter shroud  198  to one another. 
     The first bracket  142  carries a nut  144  having a flange  146 . The bushing  144  provides an elongated opening  148 , as shown in  FIG. 7B . Returning to  FIG. 6 , the second bracket  210  carries the third fastener  214  which is removably secured with respect to the bushing  144 . 
     The radially extending fasteners  214  include a length with partial threads. An end of each fastener  214  is smooth and defines a pin  217  that is received within the bushing  144  inserted near the fastener such that upon threading the fastener  214  through a helicoil  215  in the second bracket  210 , the pin  217  of the fastener  214  will engage the bushing  144  supported and thereby provide a non-threaded connection. In this way, the need is eliminated for the threaded alignment between the mated parts of the connection assembly  96 . 
     A tool  150  used during an example removal procedure is shown in  FIG. 8 . The tool  150  is a ratchet wrench  152  arranged within the core flow path of the fan intermediate case  72  circumferentially between the aft struts  73 . The ratchet wrench  152  drives a socket head tool  154 , which may have an allen, torx, ribe or other profile to drive the head  116  of the third fastener  114 . 
     Accordingly, the example connector assembly provides for access and unfastening of the fan drive gear system through a forward portion of the engine. Openings within the flow path allow access for a tool to a radially extending fastener to decouple and remove the fan drive gear system during maintenance and other inspection operations. 
     Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this disclosure.