JP5650181B2 - Gas turbine engine and method for disassembling gas turbine engine front structure - Google Patents

Description

  The present invention relates to a gas turbine engine, and more particularly to a case structure of a gas turbine engine.

  Gas turbine engines typically include one or more rotor shafts that transmit power and rotational motion from the turbine section to the compressor section and the fan section. The rotor shaft is supported within the stationary structure of the engine, generally composed of modules made up of individual case parts connected by bolted flanges. The flange forms a connection that can withstand various loads transmitted through the stationary structure of the engine.

  A continuing challenge for gas turbine engines is the ease and speed of maintenance of certain components of the engine.

  In one exemplary embodiment, the gas turbine engine includes a center body support that provides an inner annular wall of the core flow path. The center body support portion has a first spline. A geared structure connects the spool and a fan rotatable about the shaft. The flexible support portion connects the geared structure to the center body support portion. The flexible support has a second spline that meshes with the first spline to transmit torque.

  In a further embodiment of the gas turbine engine described above, the center body support includes circumferentially spaced vanes extending radially between the inner and outer annular walls and connecting the annular walls. Have.

  In a further embodiment of the gas turbine engine described above, the first spline has teeth that each include a plurality of teeth. These tooth groups are spaced apart in the circumferential direction, and an area without teeth is provided between the tooth groups.

  In a further embodiment of the gas turbine engine described above, the vanes are circumferentially aligned with the toothless region.

  In a further embodiment of the gas turbine engine described above, the second spline has a corresponding tooth group circumferentially aligned and meshed with the tooth group of the first spline and corresponding between these corresponding tooth groups. An area without teeth is provided.

  In a further embodiment of the gas turbine engine described above, the center body support has a plurality of circumferentially spaced fastening bosses. The cluster of fastening bosses is aligned with the teeth.

  In a further embodiment of the gas turbine engine described above, the toothless region is provided by a reinforcing rail that projects radially inward from a center body portion that provides an inner annular wall.

  In a further embodiment of the gas turbine engine described above, the centerbody support has an annular recess and an annular pocket that are axially spaced to provide first and second sides of the reinforcing rail.

  In a further embodiment of the gas turbine engine described above, the tooth group has a tooth root provided at the first tooth radius and extends radially inward to a tooth tip provided at the second tooth radius. Have The reinforcing rail extends radially inward to a rail radius that is smaller than the first tooth radius.

  In a further embodiment of the gas turbine engine described above, the geared structure includes a planetary gear having a sun gear, an internal gear, and an intermediate gear circumferentially disposed around the sun gear and meshing with the sun gear and the internal gear. A gear device is provided.

  In a further embodiment of the gas turbine engine described above, the intermediate gear is a planetary gear supported against rotation about an axis by a flexible support. The sun gear is supported by the spool, and the internal gear is connected to the fan.

  In a further embodiment of the gas turbine engine described above, the center body support has a first inner surface provided in the vicinity of the first spline and the flexible support is relative to the center body support. In order to position the flexible support portion in the radial direction, the flexible support portion has a first outer surface that is interference-fitted to the first inner surface.

  In a further embodiment of the gas turbine engine described above, the center body support has a second inner surface, and the flexible support has a second outer surface that is interference fitted to the second inner surface. Have. The first inner surface and the outer surface are provided in front of the first spline, and the second inner surface and the outer surface are provided in the rear of the first spline. The second outer surface is disposed on the radially inner side of the first outer surface.

  In a further embodiment of the gas turbine engine described above, the flexible support is secured to the center body support by fasteners, and these fasteners have a head portion facing forward.

  In a further embodiment of the gas turbine engine described above, the center body support has circumferentially spaced fastening bosses, and the flexible support has a fastening flange extending radially outward, The fastening flange contacts the fastening boss to position the flexible support portion in the axial direction with respect to the center body support portion.

  In a further embodiment of the gas turbine engine described above, the fastening flange has circumferentially spaced openings that receive the fasteners.

  In another exemplary embodiment, a method for disassembling a gas turbine engine forward structure includes accessing a forward-facing fastener that secures a centerbody support to a flexible support. The flexible support portion includes a geared structure supported by the flexible support portion. The above-described method further includes the step of removing the fastener and separating the first and second splines respectively provided on the center body support portion and the flexible support portion.

  In a further embodiment of the above disassembling method, the accessing step includes removing the fan module from the fan shaft bearing support while the fan shaft bearing support remains fixed to the center body support.

  In a further embodiment of the disassembly method described above, the accessing step includes removing the fan shaft bearing support from the center body support without removing the geared structure.

  In a further embodiment of the above disassembling method, the separating step includes removing the geared structure module including the geared structure and the flexible support. The step of separating does not move the bearing that supports the forward portion of the spool that can be connected to the geared structure.

1 is a schematic cross-sectional view of an embodiment of a gas turbine engine. FIG. 2 is an enlarged cross-sectional view of a front center body assembly portion of the embodiment of the gas turbine engine shown in FIG. 1. It is an expanded sectional view of the geared structure of the embodiment of the gas turbine engine shown in FIG. FIG. 2 is an exploded perspective view of the front center body assembly of the embodiment of the turbine engine shown in FIG. 1. FIG. 2 is an enlarged partial cross-sectional perspective view of a front center body support portion of the front center body assembly of the embodiment of the turbine engine shown in FIG. 1. It is an expanded sectional view of the front center body support part of the Example of the turbine engine shown in FIG. It is a perspective view of the center body support part of the Example of the turbine engine shown in FIG. It is an end view of the center body support part of the Example of the turbine engine shown in FIG. It is an exploded view of the front center body support part of the embodiment of the turbine engine shown in FIG. It is a schematic explanatory drawing which shows the Example from which a front gearbox is removed from a gas turbine engine.

  FIG. 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is shown here as a two-shaft turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26, and a turbine section 28. Alternative engines may include augmentors (not shown), among other systems and features. The fan section 22 sends air along the bypass flow path B, and the compressor section 24 sends air along the core flow path C for compression and communicates with the combustor section 26, and subsequently the turbine section 28. Inflate through. Although the disclosed non-limiting example shows a turbofan gas turbine engine, the concepts described herein are not limited to use with turbofans, but other types of turbine engines that include a three-shaft structure. It should be understood that the teachings are also applicable.

  The engine 20 generally includes a low speed spool 30 and a high speed spool 32, which spools 30, 32 with respect to a stationary structure 36 of the engine about a longitudinal central axis A of the engine via a plurality of bearing systems 38. It is attached to rotate. The low speed spool 30 generally includes an inner shaft 40 that connects a fan 42, a low pressure (or first) compressor section 44, and a low pressure (or first) turbine section 46. The inner shaft 40 is connected to the fan 42 via a geared structure 48 in order to drive the fan 42 at a lower speed than the low speed spool 30. The second bearing support portion 38 </ b> A located in the compressor portion 24 supports the front end of the inner shaft 40. Various bearing systems 38 may alternatively or additionally be provided at various locations.

  The high speed spool 32 includes an outer shaft 50 that connects the high pressure (or first) compressor section 52 and the high pressure (or first) turbine section 54. A combustor 56 is provided between the high pressure compressor 52 and the high pressure turbine 54. Here, the “high pressure” compressor or turbine has a higher pressure than the corresponding “low pressure” compressor or turbine.

  The core air stream C is compressed by the low pressure compressor 44 and then compressed by the high pressure compressor 52, mixed with fuel in the combustor 56 and burned, and then expanded through the high pressure turbine 54 and the low pressure turbine 46. The turbines 46 and 54 rotate and drive the corresponding low speed spool 30 and high speed spool 32 according to the expansion.

  In one example, engine 20 is an aircraft engine with a high bypass gear. In a further example, the bypass ratio of engine 20 is greater than about 6, and in the exemplary embodiment is greater than 10. The geared structure 48 is also a planetary gear system or other gear system such as a planetary gear system with a gear reduction ratio greater than about 2.3, and the pressure ratio of the low pressure turbine 46 is greater than about 5. In one example, the geared structure 48 includes a sun gear, an internal gear, and an intermediate gear that is circumferentially disposed about the sun gear and meshes with the sun gear and the internal gear. The intermediate gear is a planetary gear that is supported against rotation about axis A by a flexible support 68 (shown in FIG. 6). The sun gear is supported by the low speed spool 30, and the internal gear is connected to the fan 42.

  In one disclosed embodiment, the bypass ratio of the engine 20 is greater than about 10 (10: 1), the fan diameter is significantly greater than the diameter of the low pressure compressor 44, and the pressure ratio of the low pressure turbine 46 is Greater than about 5: 1. The pressure ratio of the low pressure turbine 46 is the pressure measured before the inlet of the low pressure turbine 46 versus the pressure at the outlet of the low pressure turbine 46 before the exhaust nozzle. The geared structure 48 may be a planetary gear set or other gear system, such as a planetary gear system having a gear reduction ratio greater than about 2.5: 1. However, the parameters described above are merely illustrative of one embodiment of a geared structure engine, and the invention is applicable to other gas turbine engines including direct drive turbofans.

  Due to the high bypass ratio, a large thrust is provided by the bypass flow B. The fan section 22 of the engine 20 is designed for specific flight conditions, typically about 0.8 Mach and about 35,000 feet of cruising. 0.8 Mach and 35000 feet are the flight conditions where the fuel consumption of the engine is best. The fuel eradication rate is also called “fuel consumption rate (TSFC) for moving blade cruise”, and is an industry standard that divides the lbm of the burned fuel by the lbf of the current thrust generated by the engine. Parameters. "Low fan pressure ratio" is the pressure ratio over only fan blades that do not include a fan outlet guide vane (FEGV) device. In one non-limiting example disclosed herein, the low fan pressure ratio is less than about 1.45. “Low correction fan tip speed” is the actual fan tip speed of ft / sec divided by the industry standard temperature correction value [(Tamient deg R) /518.7) ^ 0.5]. The “low correction fan tip speed” in the non-limiting example disclosed herein is less than about 1150 ft / sec. The above parameters of engine 20 are exemplary.

  Referring to FIG. 2, the engine stationary structure 36 proximate to the compressor section 24 includes a front center body assembly 60 adjacent to the second bearing support 38A. The front center body assembly 60 generally has a front center body support 62. The second bearing support portion 38 </ b> A generally includes a seal unit 64, a bearing unit 66, and a centering spring 70.

  2 and 3, the flexible support 68 provides a flexible attachment for the geared structure 48 within the front center body support 62 (also shown in FIG. 4). The flexible support 68 reacts to the torsional load from the geared structure 48 so that vibration isolation and other support functions can be easily obtained. Centering spring 70 is a generally cylindrical cage-like structural component with a plurality of beams extending between the end structures of the flange (also shown in FIG. 4). The centering spring 70 elastically positions the bearing unit 66 with respect to the low speed spool 30. In one embodiment, the beam is a double taper beam aligned circumferentially to control the radial spring rate, which includes matters relating to bearing load, bearing life, rotor dynamics, and rotor deflection. Is selected based on a number of considerations that are not limited to these.

  The front center body support portion 62 includes a front center body portion 72 and a bearing portion 74 defined around the axis A, and a frustoconical interface portion 76 is provided therebetween (see FIG. 5). The front center body portion 72 at least partially defines a core flow path leading to the low pressure compressor 44. The front center body portion 72 includes an annular core passage having a front center body vane 71 arranged in the circumferential direction having a front edge 72A and a rear edge 72B shown in cross section in FIG. The bearing portion 74 is defined on the radially inner side of the front center body portion 72. The bearing portion 74 positions the bearing unit 66 and the seal unit 64 with respect to the low speed spool 30. The frustoconical interface portion 76 connects the front center body portion 72 and the bearing portion 74 so that the typical kinks that would otherwise occur in a conventional flange connection portion are substantially eliminated from the bearing unit 66 to the stationary structure 36 of the engine. A unified load path to the outer periphery is formed. The frustoconical interface portion 76 may include a welded portion W (see FIG. 5) or may be an integral part so that the front center body support 62 is an integral part.

  The one-piece construction of the frustoconical interface portion 76 without flanges facilitates a structure that is lightweight, has a reduced number of parts, and is more adjustable in overall stiffness and achieves rotor mechanical requirements. Such a structure further integrates functions such as the supply of lubricating oil and air within the bearing compartment surrounding the bearing unit 66.

  Referring to FIG. 6, the front center body support 62 includes an attachment mechanism that receives the flexible support 68. The flexible support 68 has a conical support 158 that supports an integral flexible member 160 that bends to absorb vibrations. In one disclosed non-limiting example, the attachment mechanism for the front center body support 62 is provided with a first spline 78, each of which is an internal spline in the embodiment, provided in the front center body 72, and in the radial direction. A fastening boss 80 extending inward. The flexible support 68 includes a corresponding second spline 82, which in the embodiment is an external spline, and a fastening flange 84 extending radially outward. The flexible support 68 is received by the front center body support 62 at the spline interface 86 formed by the first and second splines 78 and 82, and the front flange is in contact with the fastening boss 80. It is held in the body support. The spline interface 86 transmits torque between the first and second splines 78 and 82. A set of fasteners 88 such as bolts are threadedly engaged with the fastening boss 80 and the fastening flange 84 to attach the flexible support 68 to the front center body support 62. The fastener 88 has a head portion 89 that faces forward so that it can be accessed from the front of the engine 20.

  With reference to FIGS. 5-6A, the center body support 62 provides an inner annular wall 128 for the core airflow C. A vane 71 connects the inner annular wall 128 to the outer annular wall 129 to provide a unitary structure. The first spline 78 has a group of teeth 146 each including a plurality of teeth. The tooth groups 146 are spaced apart in the circumferential direction, and an area without teeth is provided between the tooth groups 146. The vane 71 is circumferentially aligned with the toothless region to structurally reinforce the interface between the first and second splines 78,82. The second spline 82 has a corresponding tooth group that meshes with the tooth group 146 of the first spline 78 in the circumferential direction. Corresponding toothless regions are provided between the teeth of the second spline 82.

  In this example, the fastening boss 80 is provided as a cluster separated in the circumferential direction as shown in FIG. 6A. The fastening boss 80 is aligned with the tooth group 146. However, the fastening boss 80 may be arranged in other forms. The fastening flange 84 extends radially outward from an annular flange 127 that extends in the axial direction from the second spline 82. The fastening flange 84 has a rear surface 142 that contacts the surface 144 of the fastening boss 80 and positions the flexible support portion 68 in the axial direction with respect to the center body support portion 62. The fastening flange 84 is provided as a cluster spaced apart in the circumferential direction and includes an opening 132 for receiving a fastener 88 fixed to the hole 130 of the fastening boss 80. The fastening flange 84 may include interruptions or recesses that allow components to pass through the flexible support 68 at the periphery of the fastening flange 84.

  The toothless region is provided by a reinforcing rail 148 that projects radially inward from the center body portion 72 that provides the inner annular wall 128. The center body support 62 includes an annular recess 150 and an annular pocket 152 that are axially spaced to provide first and second side surfaces 154, 156 of the reinforcing rail 148. The teeth of the tooth group 146 have a tooth bottom provided at the first tooth radius T1, and extend radially inward to a tooth tip provided at the second tooth radius T2. As shown in FIG. 6B, the reinforcing rail 148 is smaller than the first tooth radius T1, and in one example extends radially inward to a rail radius R equal to the second tooth radius T2. By the circumferential alignment of the reinforcing rail 148 and the reinforcing rail 148 and the vane 71, the cylindricity of the center body portion 72 during the operation of the engine is surely improved.

  The center body support 62 has a first inner surface 134 provided in the vicinity of the first spline 78 and provided by the annular recess 150. The flexible support 68 is a first outer surface that is interference fit with the first inner surface 134 at room temperature to radially position the flexible support 68 relative to the center body support 62. 138. A second inner surface 136 is provided on the center body support 62, and the flexible support 68 has a second outer surface 140 that is interference fit with the second inner surface 136 at room temperature. . The first inner and outer surfaces 134 and 138 are disposed in front of the first spline 78, and the second inner and outer surfaces 136 and 140 are disposed in the rear of the first spline 78. In order to facilitate attachment / detachment of the flexible support portion 68 from the front of the engine 20, the second outer surface 140 is provided smaller than the first outer surface 138.

  Referring to FIG. 7, the head portion 89 of the fastener 88 is provided forward so as to provide access from the front portion of the front center body assembly 60 on the opposite side of the bearing unit 66 of the second bearing system 38A. . Thereby, the fastener 88 can be easily removed and the gearbox 90 of the geared structure 48 can be accessed.

  In order to provide access to the geared structure 48 from the front of the engine 20, a fan shaft bearing support front wall 102 located behind the fan 42 is attached to the front of the front center body support 62. The front wall 102 includes a flange 103 that can be attached to the front center body support 62 at the flange 61 by a plurality of fasteners 105, which can be bolts in one non-limiting example. The front wall 102 and the front center body support 62 define a bearing compartment 100 (also shown in FIG. 2) that is attached to the bearing unit 66. The front wall 102 is removable so that the gearbox 90 is accessible as a module. This allows the gearbox 90 to be accessed to facilitate rapid on-wing maintenance.

  To seal the lubricant and to support the rotation of the output shaft 108, various bearing structures 104 (schematically illustrated and also illustrated in FIG. 2) and (schematically illustrated and also illustrated in FIG. 2) A seal 106 (shown) can be supported by the front wall 102. The output shaft 108 is connected to the geared structure 48 for driving the fan 42. A fan blade 42B extends from the fan hub 110, which is attached to the output shaft 108 and rotates with the output shaft 108. In the disclosed non-limiting example, after the fan hub 110 is removed, the bearing structure 104 and seal 106 can be removed from the front wall 102 as a unit.

  The gear box 90 is driven by the low speed spool 30 (see FIG. 1) via the connecting shaft 112. The connecting shaft 112 transmits torque to the gearbox via the bearing unit 66 and facilitates isolation of vibrations and other transient phenomena. The connection shaft 112 generally includes a front connection shaft portion 114 and a rear connection shaft portion 116 extending from the bearing unit 66, but a greater or lesser number of parts are used to provide the connection shaft 112. Is possible. The front connection shaft portion 114 includes an interface spline 118 that meshes with the rear spline 120 of the rear connection shaft portion 116. In this non-limiting example, the interface spline 122 of the rear connecting shaft portion 116 connects the connecting shaft 112 to the low speed spool 30 by spline engagement with a spline 124 provided on the low pressure compressor hub 126 of the low pressure compressor 44. To do.

  In summary, the forward structure of engine 20 is disassembled by removing the fan module from the fan shaft bearing support. The fan shaft bearing support (front wall 102) remains fixed to the center body support 62 so as to cover the gear box 90. The fan shaft bearing support (front wall 102) is removed from the center body support 62 without removing the gear box 90. The front facing fastener 88 is accessed and removed. The first and second splines 78 and 82 are separated, and the gear box 90 is removed together with the fan shaft bearing support (front wall 102) and the flexible support 68. The bearing 38A is not moved.

  To disengage and separate the gear box 90, the fan hub 110 is removed from the output shaft. Next, the plurality of fasteners 105 are removed, and the front wall 102 is separated from the front center body support portion 62. The front wall 102 is then removed from the engine. Subsequently, the plurality of fasteners 88 are removed from the front of the engine 20. Thereafter, the geared structure 48 is slid forward and removed from the front center body support portion 62, whereby the interface spline 118 is slid out of the rear spline 120 and the external spline 82 is connected to the internal spline. Slide off 78. As a result, the geared structure 48 is removable from the engine 20 as a module (shown schematically in FIG. 8). Other components may need to be removed to remove the geared structure 48 from the engine 20, but such removal is not critical and does not require detailed description. Further, other components such as the bearing unit 66 and the seal 64 are also easily accessible from the front of the engine 20.

  By removing the gearbox 90 from the front of the engine 20 as described above, time and costs are greatly reduced. The geared structure 48 can be removed as a module from the engine 20 and need not be disassembled further. Further, the geared structure 48 needs to be removed from the engine to access the bearing unit 66 and the seal 64, but the geared structure 48 need not be removed to access the engine core itself.

  Throughout the drawings, corresponding parts and corresponding parts are given the same reference numerals. While the illustrated embodiment shows a particular arrangement of components, other arrangements are possible.

  Although specific step sequences are described in the disclosure, description and claims, unless otherwise stated, the steps can be performed in any order and can be separated or combined and still benefit from the present invention.

  Each different embodiment has the specific components illustrated, but embodiments of the invention are not limited to these specific combinations. Several components and features of one embodiment can be used in combination with components and features of other embodiments.

  The above description is illustrative and not restrictive. While various non-limiting embodiments have been disclosed, those skilled in the art will appreciate that various modifications and variations in light of the above teachings are within the scope of the appended claims. Thus, within the scope of the appended claims, the present invention may be practiced otherwise than as specifically described. For this reason, the following claims should be studied to determine the true scope and content of this invention.

DESCRIPTION OF SYMBOLS 24 ... Compressor part 38A ... 2nd bearing support part 48 ... Structure with gears 60 ... Front center body assembly 62 ... Front center body support part 64 ... Seal unit 66 ... Bearing unit 68 ... Flexible support part 70 ... Centering spring 90 ... Gearbox 100 ... Bearing compartment 102 ... Front wall 104 ... Bearing structure 106 ... Seal

Claims (19)

A center body support that provides an inner annular wall of the core flow path and has a first spline;
A geared structure that connects the spool and a fan that can rotate about an axis; and
A flexible support portion having a second spline that connects the geared structure to a center body support portion and meshes with the first spline to transmit torque ;
The first spline has tooth groups each including a plurality of teeth, these tooth groups are spaced apart in the circumferential direction, and a region without teeth is provided between the tooth groups. Gas turbine engine.   The gas according to claim 1, wherein the center body support portion includes circumferentially spaced vanes extending between the inner annular wall and the outer annular wall to connect the annular walls. Turbine engine. The gas turbine engine according to claim 2 , wherein the vane is circumferentially aligned with the toothless region. The second spline has a corresponding tooth group that meshes with the tooth group of the first spline in a circumferential direction, and a corresponding toothless region is provided between the corresponding tooth groups. The gas turbine engine according to claim 1 . Centerbody support portion includes a plurality of fastening bosses that are circumferentially spaced, these fastening bosses, the gas turbine engine according to claim 1, characterized in that aligned with the teeth. Region without the tooth, a gas turbine engine according to claim 1, characterized in that it is provided by the reinforcing rails projecting from the center body portion to provide said inner annular wall radially inwardly. The gas turbine according to claim 6 , wherein the center body support portion includes an annular recess and an annular pocket that are axially spaced to provide first and second side surfaces of the reinforcing rail. engine. The tooth group has a tooth bottom provided at a first tooth radius and inner teeth extending radially inward to a tooth tip provided at a second tooth radius. The gas turbine engine according to claim 6 , wherein the gas turbine engine extends radially inward to a rail radius smaller than one tooth radius.   The geared structure includes a planetary gear device having a sun gear, an internal gear, and an intermediate gear arranged circumferentially around the sun gear and meshing with the sun gear and the internal gear. Item 4. The gas turbine engine according to Item 1. The intermediate gear is a planetary gear supported by the flexible support portion with respect to rotation about an axis, the sun gear is supported by a spool, and the internal gear is connected to a fan. The gas turbine engine according to claim 9 .   The center body support portion has a first inner surface provided in the vicinity of the first spline, and the flexible support portion is configured such that the flexible support portion is arranged in a radial direction relative to the center body support portion. The gas turbine engine of claim 1, further comprising a first outer surface that is interference fitted to the first inner surface for positioning. The center body support portion has a second inner side surface, and the flexible support portion has a second outer side surface that is interference-fitted to the second inner side surface, and the first inner side surface and the outer side surface The side surface is provided in front of the first spline, the second inner surface and the outer surface are provided behind the first spline, and the second outer surface is radially inward of the first outer surface. The gas turbine engine according to claim 11 , wherein the gas turbine engine is arranged.   The gas turbine engine according to claim 1, further comprising a fastener that fixes the flexible support portion to the center body support portion, and each of the fasteners includes a head portion facing forward. The center body support portion has a fastening boss spaced apart in the circumferential direction, and the flexible support portion has a fastening flange extending radially outward, and the fastening flange contacts the fastening boss. The gas turbine engine according to claim 13, wherein the flexible support portion is axially positioned with respect to the center body support portion. The gas turbine engine according to claim 14 , wherein the fastening flange has a circumferentially spaced opening that receives the fastener. A method for disassembling a front structure of a gas turbine engine according to claim 1,
Accesses the centerbody support portion fastener facing forward is fixed to the flexible support, said flexible support part supports a structure with the gear,
Remove the fastener,
A method for disassembling a front structure of a gas turbine engine, comprising the steps of separating a first spline and a second spline respectively provided on a center body support and the flexible support. 17. The gas turbine according to claim 16 , wherein the step of accessing includes the step of removing the fan module from the fan shaft bearing support portion while the fan shaft bearing support portion is fixed to the center body support portion. A method for disassembling the engine front structure. The disassembling of the front structure of a gas turbine engine according to claim 17 , wherein the step of accessing includes the step of removing the fan shaft bearing support from the center body support without removing the geared structure. Method. The separating step includes removing the geared structure module including the geared structure and the flexible support, and the separating step allows the spool to be operatively connected to the geared structure. The method for disassembling a front structure of a gas turbine engine according to claim 18, wherein the bearing supporting the front of the gas turbine engine is not moved.

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