US20150098816A1

US20150098816A1 - System and method for controlling backbone bending in a gas turbine engine

Info

Publication number: US20150098816A1
Application number: US14/045,552
Authority: US
Inventors: Randy Scott Longtin
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2013-10-03
Filing date: 2013-10-03
Publication date: 2015-04-09

Abstract

A flexible coupling for a gas turbine engine includes a fan hub frame having a first flange and a core engine having a second flange. The first flange includes a first plurality of fingers extending from a face of the flange and the second flange includes a second plurality of fingers extending from a face of the second flange. The second plurality of fingers of the second flange is complementary to the first plurality of fingers extending from the first flange. The first flange and the second flange are coupled together in form-fitting engagement such that the first plurality of fingers and the second plurality of fingers form an interdigitated configuration.

Description

BACKGROUND

The subject matter described herein relates generally to gas turbine engines and, more particularly, to a flexible coupling system between a turbine engine fan frame and an engine core to facilitate reducing engine backbone bending.
At least some known turbine engines direct a portion of a process fluid, i.e., air, from an engine inlet through the compression, combustion, and turbine sections of the engine core. Generally, a fan assembly pushes air into the compression section where a compressor compresses the air. The compressed air is suitably mixed with a fuel and channeled to a combustor where the fuel-air mixture is burned to generate hot combustion gases. The combustion gases are channeled to at least one turbine, where energy is extracted from the combustion gases for powering the compressor, as well as for producing useful work, such as propelling a vehicle.
Some known turbine engines are supported by an aircraft structure, for example, by a pylon extending downward beneath an aircraft wing. The engine is typically mounted and secured to the pylon in two planes normal to the engine centerline, one towards the forward end of the engine, usually just rearward of the fan section, and a second toward the aft end of the engine, typically in the turbine section. The engine is mounted by its engine frame, i.e., a static structure that supports the rotating components of the engine.
The engine frame generally has a plurality of support rings that are coupled together to the engine casing to form what may generally be called a backbone. The support rings include a forward fan frame and an aft turbine frame. Typically, the forward fan frame and the aft turbine frame support the main rotor bearings, which in turn rotatably support the rotating components of the engine. The backbone facilitates controlling engine clearance closures defined between the engine casing and the rotating components positioned radially inwardly from the backbone.
Backbone bending occurs, at least in part, due to aerodynamic loads that enter the inlet and generate pitch and yaw moments on the fan assembly. The aerodynamic loads deflect and distort the fan assembly such that the core engine between the forward and aft mounts is deflected relative to the undisturbed engine centerline. As the fan assembly size and corresponding aerodynamic loading by known turbine engines has increased, so has the magnitude of the reaction loads and bending moments. The resultant backbone bending induces a potential for increased rubbing between the rotating components and the adjacent stationary structures. The increased rubbing adversely impacts engine performance, specific fuel consumption, and the life of rotating component wear strips, and may lead to more frequent engine maintenance or reduced engine “time on wing.” Fuel costs and desirability for improved durability accentuate the need for designs and systems for reducing backbone bending.

BRIEF DESCRIPTION

In one aspect, a flexible coupling system for a gas turbine engine is provided. The flexible coupling system includes a fan hub frame having a first flange. The first flange includes a first plurality of fingers extending from the first flange. The flexible coupling system also includes a core engine having a second flange. The second flange includes a second plurality of fingers extending from the second flange. The second plurality of fingers is complementary to the first plurality of fingers extending from the first flange. The first flange and the second flange are configured to couple together in form-fitting engagement, such that the first plurality of fingers and the second plurality of fingers form an interdigitated configuration.
In another aspect, a turbine engine assembly is provided. The turbine engine assembly includes a nacelle and a fan assembly coupled together. The fan assembly includes an array of fan blades and a fan hub frame. The fan hub frame includes a first flange having a first plurality of fingers extending from the first flange. The turbine engine also includes a core engine. The core engine has a second flange that includes a second plurality of fingers extending from the second flange. The plurality of fingers is complementary to the fingers extending from the first flange. The first flange and the second flange are configured to couple together in form-fitting engagement, such that the first plurality of fingers and the second plurality of fingers form an interdigitated configuration. An applied aerodynamic load acting on the fan assembly is at least partially transmitted to the nacelle.
In yet another aspect, method of controlling backbone bending of a turbine engine is provided. The method includes coupling a first plurality of fingers to a first flange of a fan hub frame having a first longitudinal axis. The first plurality of fingers extends axially from the first flange. The method also includes coupling a second plurality of fingers to a second flange of a core engine having a second longitudinal axis. The second plurality of fingers extends axially from the second flange and is complementary to the first plurality of fingers. In addition, the method includes non-rigidly coupling the first flange and the second flange together in form-fitting engagement, such that the first plurality of fingers and the second plurality of fingers form an interdigitated configuration. The first longitudinal axis and the second longitudinal axis are substantially coaxial. Furthermore, the method includes applying an aerodynamic load to the fan hub frame such that the fan hub frame is displaced an angular value with respect to the second longitudinal axis.

DRAWINGS

These and other features, aspects, and advantages of the present disclosure will become better understood when the following detailed description is read with reference to the accompanying drawings in which like characters represent like parts throughout the drawings, wherein:

FIG. 1 is a schematic of an aircraft gas turbine engine mounted within a nacelle and mounted to an aircraft by a mounting structure or pylon;

FIG. 2 is a sectional schematic illustrating the exemplary gas turbine engine shown in FIG. 1;

FIG. 3 is a schematic of a flexible coupling system for use with the gas turbine engine shown in FIG. 1;

FIG. 4 is a sectional schematic of the flexible coupling system shown in FIG. 3 illustrating an internal seal; and

FIG. 5 is a flow chart of an exemplary method that may be used to control backbone bending of the gas turbine engine shown in FIG. 1.

Unless otherwise indicated, the drawings provided herein are meant to illustrate features of embodiments of the disclosure. These features are believed to be applicable in a wide variety of systems comprising one or more embodiments of the disclosure. As such, the drawings are not meant to include all conventional features known by those of ordinary skill in the art to be required for the practice of the embodiments disclosed herein.

DETAILED DESCRIPTION

In the following specification and the claims, reference will be made to a number of terms, which shall be defined to have the following meanings.
The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise.
Approximating language, as used herein throughout the specification and claims, may be applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about” and “substantially”, are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value. Here and throughout the specification and claims, range limitations may be combined and/or interchanged; such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise.
As used herein, the term “core engine” will be used to refer to that portion of a gas turbine engine extending from the high pressure compressor forward flange back to the low pressure turbine. The core engine includes both the engine casing or stator and the rotating components positioned radially inwardly from the stator.
The systems and methods described herein relate to a flexible coupling system between a turbine engine fan hub frame and a core engine to facilitate reducing engine backbone bending. The flexible coupling system includes a fan hub frame having a flange with a plurality of fingers extending outward from a face of the flange. The fan hub frame is configured to couple to the core engine. The core engine has a flange including a plurality of fingers extending from a face of the core engine flange. The fingers extending from the fan hub frame flange and the core engine flange are similar and complementary to each other such that the fan hub frame and the core engine can be coupled together in form-fitting engagement, such that the fingers of the two flanges form an interdigitated configuration. The flexible coupling system facilitates enabling engine torque to be transmitted through the flexible coupling system, while accommodating off-axis pivoting, such as a pitch moment and a yaw moment, of the fan hub frame relative to the core engine. Operating a gas turbine engine with such a flexible coupling system facilitates reducing the amount of distortion occurring in the core engine which enables the gas turbine engine to maintain closer blade tip clearances, thus facilitating enhancing engine performance, specific fuel consumption, and the life of rotating component wear strips. Such a flexible coupling system may also facilitate less frequent engine maintenance or increasing engine “time on wing.”
FIG. 1 is a schematic of an aircraft gas turbine engine 10 mounted within a nacelle 12 and mounted to an aircraft (not shown) by a mounting structure or pylon 14. In the exemplary embodiment, gas turbine engine 10 is generally disposed about an engine center line 16. Gas turbine engine 10 has a forward or fan hub frame 18 that is connected to a backbone structure or core engine 20, such that fan hub frame 18 is substantially coaxial with core engine 20 about engine center line 16. Core engine 20 is further connected to an aft or turbine frame 24. Gas turbine engine 10 is attached to pylon 14, which generally extends downwardly from a wing of the aircraft (not shown), at a forward mount 26 on fan hub frame 18 and at an aft mount 28 on turbine frame 24. Alternatively, forward mount 26 may be coupled to a fan case 66.
In the exemplary embodiment, gas turbine engine 10 is shown with engine centerline 16 being substantially horizontal. As the aircraft (not shown) takes flight and begins to climb in altitude, engine centerline 16 becomes angled to the direction of approaching airflow and aerodynamic loading M occurs on nacelle 12. Aerodynamic loading M on nacelle 12 may include vertical and lateral shear and pitch and yaw moments (not shown in FIG. 1). In the exemplary embodiment, aerodynamic loading M is shown as directed radially inward in a substantially vertical direction, but it may be from other angular directions in a full range of 0°-360° about engine centerline 16. Aerodynamic loading M on nacelle 12 facilitates pivoting gas turbine engine 10 about forward mount 26, as shown by direction arrow C. Concurrently, a thrust load N generated by gas turbine engine 10 and acting along engine centerline 16 further facilitates pivoting gas turbine engine 10 about forward mount 26, as shown by direction arrow C.
In the exemplary embodiment, pivoting of gas turbine engine 10 about forward mount 26, as shown by direction arrow C, is restrained by aft mount 28. This results in bending moments being induced and dissipated within core engine 20. Core engine 20 is deflected from its undisturbed position, which is generally concentric about engine centerline 16. In the exemplary embodiment, a flexible coupling system 100 between fan hub frame 18 and core engine 20 facilitates transferring at least a portion of aerodynamic load M into nacelle 12, and thus to pylon 14 and/or a thrust reverser assembly 72. Flexible coupling system 100 also facilitates preventing relative axial rotation or roll about engine centerline 16 between fan hub frame 18 and core engine 20 while enabling a predetermined amount of axial displacement and/or angular displacement θ (not shown in FIG. 1) between fan hub frame 18 and core engine 20. In the exemplary embodiment, angular displacement θ is in the range between about 2° and about 0°, and more particularly, in the range between about 1° and about 0°. As described above, angular displacement θ may be in any angular directions in a full range of 0°-360° about engine centerline 16.
FIG. 2 is a sectional schematic illustrating, in detail, exemplary gas turbine engine 10. In the exemplary embodiment, gas turbine engine 10 includes an engine core portion 32 including a high pressure compressor 34, a combustor 36, and a high pressure turbine 38. In the exemplary embodiment, gas turbine engine 10 is a two spool type and includes a low pressure compressor 40 coupled together with fan assembly 30, and a low pressure turbine 42. Fan assembly 30 includes an array of fan blades 44 extending radially outward from a rotor disk 46. In addition, fan assembly 30 includes fan hub frame 18 having a plurality of fan frame struts or outlet guide vanes 64 that extend radially outward to fan case 66. Gas turbine engine 10 has an intake side 48 and an exhaust side 50. Fan assembly 30 and low pressure compressor 40 are coupled by a low speed rotor shaft 52, and high pressure compressor 34 and high pressure turbine 38 are coupled by a high speed rotor shaft 54.
In the exemplary embodiment, low speed rotor shaft 52 is supported axially and radially from core engine 20 by a forward bearing assembly 56 and by an aft bearing assembly 58. High speed rotor shaft 54 is disposed concentrically about low speed rotor shaft 52 and is supported axially and radially by forward bearing assemblies 60 and aft bearing assembly 62. In the exemplary embodiment, forward bearing assembly 56 is supported within fan assembly 30, which is connected to core engine 20 through fan hub frame 18. Aft bearing assemblies 58 and 62 are supported within turbine frame 24, which is also connected to core engine 20. In the exemplary embodiment, forward bearing assemblies 60 for high speed rotor shaft 54 are detached from fan hub frame 18 and coupled to core engine 20 to isolate high speed rotor shaft 54 from aerodynamic load M experienced by fan assembly 30. Aerodynamic load M is transmitted throughout fan assembly 30 by fan hub frame 18, through outlet guides vanes 64, and to fan case 66, and thus could adversely affect the operation of high speed rotor shaft 54 if it were not at least partially isolated from aerodynamic load M. For example, if bearing assemblies 60 were attached to fan hub frame 18, deflections of fan hub frame 18 due to aerodynamic load M would be transmitted to high speed rotor shaft 54. Deflections of high speed rotor shaft 54 may result in increased rubbing of high pressure compressor 34 and high pressure turbine 38, thus adversely impacting engine performance, specific fuel consumption, and the life of rotating component wear strips, and may lead to more frequent engine maintenance or reduced engine “time on wing.”
In the exemplary embodiment, fan assembly 30 of gas turbine engine 10 is coupled to nacelle 12 and thrust reverser assembly 72. Thrust reverser assembly 72 is coupled to an outer V-groove joint 68, which is formed substantially around a periphery of fan case 66, and to an inner V-groove joint 70, which is formed substantially around a periphery of fan hub frame 18. Thrust reverser assembly 72 may be attached to pylon 14 by hinged lugs (not shown) that allow fore and aft movement of thrust reverser assembly 72. In the exemplary embodiment, outer V-groove joint 68 and inner V-groove joint 70 form part of a connection between fan assembly 30 and thrust reverser assembly 72 to facilitate transferring at least a portion of aerodynamic load M (shown in FIG. 1) between fan assembly 30 and thrust reverser assembly 72.
During operation, air (not shown) flows axially through nacelle 12 and fan assembly 30 in a direction that is substantially parallel to engine centerline 16, which extends through gas turbine engine 10. The air is supplied to low pressure compressor 40 and high pressure compressor 34 where it is compressed. The compressed air is delivered to combustor 36 where it is burned with a fuel (not shown) to generate a combustion gas flow (not shown). The combustion gas flow from combustor 36 drives high pressure turbine 38 and low pressure turbine 42 and generates thrust load N (shown in FIG. 1). High pressure turbine 38 drives high pressure compressor 34 by way of high speed rotor shaft 54 and low pressure turbine 42 drives fan assembly 30 by way of low speed rotor shaft 52.
FIG. 3 is a schematic of flexible coupling system 100 for use with gas turbine engine 10 (shown in FIG. 1). In the exemplary embodiment, flexible coupling system 100 facilitates joining fan hub frame 18 to core engine 20. The exemplary flexible coupling system 100 accommodates off-axis pivoting, such as a pitch moment Mz and a yaw moment My, of fan hub frame 18 relative to core engine 20, which is generally collinear to engine centerline 16, while transmitting engine torque or a roll moment Mx and thrust load N between fan hub frame 18 and core engine 20.
In the exemplary embodiment, core engine 20 includes a high pressure compressor (HPC) flange 102 positioned on a forward end of core engine 20 proximate fan hub frame 18. A plurality of HPC fingers 104 extend axially from a face 106 of HPC flange 102. The circumferential spaces between HPC fingers 104 are HPC grooves 108. In addition, fan hub frame 18 includes a fan hub frame (FHF) flange 110 positioned opposite HPC flange 102 and coupled to an aft end of fan hub frame 18. A plurality of FHF fingers 112, complementary to HPC fingers 104, extends axially from a face 114 of FHF flange 110. The circumferential spaces between FHF fingers 112 are FHF grooves 116. In the exemplary embodiment, when core engine 20 and fan hub frame 18 are coupled together, each of HPC fingers 104 form-fittingly engage a respective FHF groove 116, and each of FHF fingers 112 form-fittingly engage a respective HPC groove 108, thereby forming an interdigitated configuration. As a result, roll moment Mx is transmitted between fan hub frame 18 and core engine 20, while fan hub frame 18 may pitch or yaw through angular displacement θ. In the exemplary embodiment, HPC fingers 104 and FHF fingers 112 are formed with curved outer surfaces to facilitate enabling fan hub frame 18 to pitch or yaw through angular displacement θ. Alternatively, HPC fingers 104 and FHF fingers 112 may be formed in any shape that enables HPC flange 102 and FHF flange 110 to operate as described herein. As described above, angular displacement θ may be any angular direction in a full range of 0°-360° about engine centerline 16. In the exemplary embodiment, HPC fingers 104 and FHF fingers 112 are slidingly coupled together in the circumferential direction of HPC flange 102 and FHF flange 110 respectively. Alternatively, HPC fingers 104 and FHF fingers 112 may be fabricated such that when coupled together there is a gap between the fingers in the circumferential direction.
In the exemplary embodiment, flexible coupling system 100 includes a flange coupling member 118. Flange coupling member 118 extends circumferentially around the entire periphery of HPC flange 102 and FHF flange 110 when they are coupled together. In the exemplary embodiment, flange coupling member 118 includes a rigid C-shaped channel 120 configured to receive HPC flange 102 and FHF flange 110 within C-shaped channel 120 and to couple them together in form-fitting engagement to facilitate preventing relative axial rotation between fan hub frame 18 and core engine 20. Flange coupling member 118 also includes a compliant element 124 coupled to an inner portion 122 of C-shaped channel 120. In the exemplary embodiment, compliant element 124 is a rigid elastomer. Alternatively, compliant element 124 may include any compliant member or biasing element that enables flexible coupling system 100 to operate as described herein, e.g., without limitation, spring-loaded balls or ball-tipped plungers.
In the exemplary embodiment, flange coupling member 118 flexibly couples HPC flange 102 and FHF flange 110 together in form-fitting engagement, such that HPC fingers 104 and FHF fingers 112 form an interdigitated configuration, to prevent relative rotation between HPC flange 102 and FHF flange 110 while enabling a predetermined amount of axial displacement and/or angular displacement θ between fan hub frame 18 and core engine 20. In addition, compliant element 124 operates to seal leakage of a process fluid (not shown) between HPC flange 102 and FHF flange 110. In operation, as fan hub frame 18 moves in response to pitch moment Mz and/or yaw moment My, compliant element 124 compresses and/or expands to maintain a seal around HPC flange 102 and FHF flange 110. When fan hub frame 18 reaches a predetermined amount of movement, compliant element 124 ceases to comply and transfers pitch moment Mz and/or yaw moment My to rigid C-shaped channel 120. C-shaped channel 120 facilitates preventing fan hub frame 18 from moving beyond a predetermined angular displacement θ, thereby facilitating enhancing engine performance, specific fuel consumption, and the life of rotating component wear strips. Such a flexible coupling system may also facilitate less frequent engine maintenance or increasing engine “time on wing.”
FIG. 4 is a sectional schematic of flexible coupling system 100 illustrating a sealing member or internal seal 126. In the exemplary embodiment, flexible coupling system 100 includes internal seal 126 to seal against leakage of a process fluid (not shown) between HPC flange 102 and FHF flange 110. In the exemplary embodiment, core engine 20 includes a recessed groove 128 formed in face 106 of HPC flange 102. Recessed groove 128 extends circumferentially about engine centerline 16. In addition, fan hub frame 18 includes a recessed groove 132 formed in face 114 of FHF flange 110. Recessed groove 132 extends circumferentially about engine centerline 16. In the exemplary embodiment, internal seal 126 is fitted within recessed grooves 128 and 132 to regulate the flow of air (not shown) between the flexible joint formed by HPC flange 102 and FHF flange 110. In the exemplary embodiment, internal seal 126 is a rigid elastomer. Alternatively, internal seal 126 may be any flexible material that permits internal seal 126 to operate as described herein.
FIG. 5 is a flow chart of an exemplary method 500 that may be used to control backbone bending of gas turbine engine 10 (shown in FIG. 1). Referring to FIGS. 1-3 and 5, to facilitate controlling backbone bending of gas turbine engine 10, in the exemplary embodiment, FHF fingers 112 are coupled 502 to FHF flange 110 of fan hub frame 18. FHF fingers 112 extend axially outward from flange face 114. In addition, HPC fingers 104 are coupled 504 to HPC flange 102 of core engine 20. HPC fingers 104 extend axially outward from flange face 106. HPC fingers 104 are formed substantially the same and complementary to FHF fingers 112. HPC flange 102 and FHF flange 110 are non-rigidly coupled 506 together in form-fitting engagement, such that HPC fingers 104 and FHF fingers 112 form an interdigitated configuration. In operation, aerodynamic load M is applied 508 to fan hub frame 18, which facilitates flexing fan hub frame 18 a predetermined amount of axial displacement and/or angular displacement θ (not shown in FIG. 1) relative to engine centerline 16. As described above, angular displacement θ may be in any angular directions in a full range of 0°-360° about engine centerline 16.
In the exemplary embodiment, internal seal 126 is coupled between fan hub frame 18 and core engine 20 to regulate a predetermined flow rate of a process fluid (not shown) between FHF flange 110 and HPC flange 102. Furthermore, FHF flange 110 may be non-rigidly coupled to HPC flange 102 in form-fitting engagement with flange coupling member 118, which extends at least partially about the periphery of FHF flange 110 and HPC flange 102. Flange coupling member 118 includes compliant element 124 positioned between flange coupling member 118 and at least one of FHF flange 110 and HPC flange 102. Compliant element 124 facilitates sealing leakage of a process fluid (not shown) between HPC flange 102 and FHF flange 110 by compressing and/or expanding to maintain a seal around HPC flange 102 and FHF flange 110 as fan hub frame 18 moves in response to aerodynamic moment M.
The systems and methods as described herein facilitate controlling backbone bending of a gas turbine engine. Specifically, the systems and methods described facilitate controlling backbone bending by providing a flexible coupling system between the fan hub frame and the stator of the gas turbine engine. The flexible coupling system includes a plurality of fingers extending axially from flange faces of the fan hub frame and the core engine, such that the fingers of the two flanges form an interdigitated configuration, to enable off-axis flexing of the fan hub frame in relation to the core engine. Therefore, in contrast to known gas turbine engines without such a flexible coupling, the system and methods described herein facilitate reducing the amount of distortion occurring in the core engine which enables the gas turbine engine to maintain closer blade tip clearances, thus facilitating enhancing engine performance and specific fuel consumption.
An exemplary technical effect of the systems and methods described herein includes at least one of: (a) reducing the amount of distortion occurring in the core engine of a gas turbine engine assembly by at least partially isolating the core engine from inlet aerodynamic loads and transmitting those aerodynamic loads to the engine mounting structure of the aircraft, such as to a pylon; (b) controlling engine clearance closures defined between the engine casing and the rotating components positioned radially inwardly from the core engine; and (c) enhancing engine performance and specific fuel consumption.
Exemplary embodiments of a flexible coupling system for a gas turbine engine are described above in detail. The methods and systems are not limited to the specific embodiments described herein, but rather, components of systems and/or steps of the methods may be utilized independently and separately from other components and/or steps described herein. For example, the systems and apparatus described herein may have other industrial or consumer applications and are not limited to practice with submersible pumps as described herein. Rather, one or more embodiments may be implemented and utilized in connection with other industries.
Although specific features of various embodiments of the invention may be shown in some drawings and not in others, this is for convenience only. In accordance with the principles of the invention, any feature of a drawing may be referenced and/or claimed in combination with any feature of any other drawing.
This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.

Claims

What is claimed is:

1. A flexible coupling system for a turbine engine comprising:

a fan hub frame comprising a first flange comprising a first plurality of fingers extending therefrom; and

a core engine comprising a second flange comprising a second plurality of fingers extending therefrom and complementary to said first plurality of fingers, wherein said first flange and said second flange are configured to couple together in form-fitting engagement, such that said first plurality of fingers and said second plurality of fingers form an interdigitated configuration.

2. The flexible coupling system of claim 1 further comprising a sealing member positioned between said fan hub frame and said core engine, said sealing member configured to regulate a predetermined flow rate of a process fluid between said first flange and said second flange.

3. The flexible coupling system of claim 1 further comprising a flange coupling member extending at least partially about a periphery of said first flange and a periphery of said second flange, said flange coupling member comprising a compliant element between said flange coupling member and at least one of said first flange and said second flange.

4. The flexible coupling system of claim 3, wherein said flange coupling member is configured to couple said first flange and said second flange together in form-fitting engagement, said first flange and said second flange configured to facilitate limiting relative axial rotation between said fan hub frame and said core engine.

5. The flexible coupling system of claim 3, wherein said compliant element is configured to facilitate a predetermined value of an angular displacement between said fan hub frame and said core engine.

6. The flexible coupling system of claim 3, wherein said compliant element is configured to regulate a predetermined flow rate of a process fluid between said first flange and said second flange.

7. The flexible coupling system of claim 1, wherein said core engine further comprises a bearing structure coupled to said core engine.

8. A turbine engine assembly comprising:

a nacelle;

a fan assembly coupled to said nacelle and comprising:

an array of fan blades; and

a core engine comprising:

a second flange comprising a second plurality of fingers extending therefrom and complementary to said first plurality of fingers, wherein said first flange and said second flange are configured to couple together in form-fitting engagement, such that said first plurality of fingers and said second plurality of fingers form an interdigitated configuration, wherein an applied aerodynamic load acting on said fan assembly is at least partially transmitted to said nacelle.

9. The engine assembly of claim 8, wherein said fan assembly further comprises a fan case and a plurality of outlet guide vanes extending between said fan hub frame and said fan case.

10. The engine assembly of claim 9, wherein said fan hub frame is configured to transmit the applied aerodynamic load through said outlet guide vanes to said fan case.

11. The engine assembly of claim 8 further comprising a sealing member positioned between said fan assembly and said core engine, said sealing member configured to regulate a predetermined flow rate of a process fluid between said first flange and said second flange.

12. The engine assembly of claim 8 further comprising a flange coupling member extending at least partially about a periphery of said first flange and a periphery of said second flange, said flange coupling member comprising a compliant element between said flange coupling member and at least one of said first flange and said second flange.

13. The engine assembly of claim 12, wherein said flange coupling member is configured to couple said first flange and said second flange together in form-fitting engagement, said first flange and said second flange configured to facilitate limiting relative axial rotation between said fan hub frame and said core engine.

14. The engine assembly of claim 12, wherein said compliant element is configured to facilitate a predetermined value of an angular displacement between said fan assembly and said core engine.

15. The engine assembly of claim 12, wherein said compliant element is configured to regulate a predetermined flow rate of a process fluid between said first flange and said second flange.

16. The engine assembly of claim 8, wherein said core engine further comprises a bearing structure coupled to said core engine.

17. A method of controlling backbone bending of a gas turbine engine, said method comprising:

coupling a first plurality of fingers to a first flange of a fan hub frame having a first longitudinal axis, wherein the first plurality of fingers extend axially therefrom;

coupling a second plurality of fingers to a second flange of a core engine having a second longitudinal axis, wherein the second plurality of fingers extend axially therefrom complementary to the first plurality of fingers;

non-rigidly coupling the first flange and the second flange together in form-fitting engagement, such that the first plurality of fingers and the second plurality of fingers form an interdigitated configuration, wherein the first longitudinal axis and the second longitudinal axis are substantially coaxial; and

applying an aerodynamic load to the fan hub frame such that the fan hub frame is displaced an angular value with respect to the second longitudinal axis.

18. The method of claim 17 further comprising coupling a sealing member between the fan hub frame and the core engine to regulate a predetermined flow rate of a process fluid between the first flange and the second flange.

19. The method of claim 17, wherein non-rigidly coupling the first flange to the second flange comprises non-rigidly coupling the first flange to the second flange with a flange coupling member, wherein the first flange has a first periphery and the second flange has a second periphery, the flange coupling member extending at least partially about the first periphery and the second periphery of the first flange and the second flange respectively, wherein the flange coupling member includes a compliant element between the flange coupling member and at least one of the first flange and the second flange.

20. The method of claim 17, wherein coupling a second plurality of fingers to a second flange of a core engine includes coupling a second plurality of fingers to a second flange of a core engine including a bearing structure coupled to the core engine.