Integrated inlet design

This invention relates generally to gas turbine engine nacelle assemblies and, more particularly, to methods and apparatus for an integrated inlet design for a nacelle assembly.

Some known nacelle assemblies used with turbine engines include a plurality of components that disrupt aerodynamic flow of the nacelle assembly with a plurality of circumferential and axial gaps and steps defined between the components. For example, at least one known nacelle assembly includes an inlet lip, an inlet outer barrel, and a fan cowl that define an outer flow surface of the nacelle assembly. Moreover, within some nacelle assemblies, some known components include a plurality of segments, which define additional gaps and steps on the outer flow surface. For example, at least one known nacelle assembly used in a large high-bypass engine includes an inlet lip defined by a plurality of radial segments.

Within some turbine engines, at least some of the components and at least some segments of the components used in the nacelle assembly are coupled in position within the nacelle assembly with a plurality of mechanical fasteners. For example, mechanical fasteners may secure a bulkhead that is positioned internal to the inlet lip. The mechanical fasteners, because of their orientation, also disrupt aerodynamic flow of the nacelle assembly. More specifically, in the bulkhead example, heads of the mechanical fasteners are exposed on the outer surface of the nacelle assembly and are positioned directly in the outer flow path. In addition to the exposed heads of the mechanical fasteners, other gaps and/or steps in the nacelle assembly may inhibit laminar flow over the outer flow surface of the nacelle assembly, and may increase aerodynamic drag. Operating a turbine engine with increased aerodynamic drag may reduce fuel burn efficiency.

In one aspect, a method for assembling a nacelle assembly is provided. The method includes providing an outer structure that extends at least between a highlight defined by a forward end of the outer structure and a point defined by a maximum diameter of the nacelle assembly and coupling the outer structure to an inner barrel.

In another aspect, an outer structure of a nacelle assembly is provided. The outer structure includes a highlight defined by a forward end of the outer structure and an outer aft section that includes a point defined by a maximum diameter of the nacelle assembly, wherein the nacelle assembly extends at least between the highlight and the point.

In yet another aspect, a nacelle assembly is provided. The nacelle assembly includes an inner barrel and an outer structure comprising a highlight and an outer aft section, wherein the highlight is defined by a forward end of the outer structure, wherein the outer aft section includes a point defined by a maximum diameter of the nacelle assembly, wherein the nacelle assembly extends at least between the highlight and the point.

FIG. 1 is a perspective view of an exemplary aircraft;

FIG. 2 is a cross-sectional view of a portion of an exemplary turbine engine that may be used with the aircraft shown in FIG. 1 ;

FIG. 3 is a schematic illustration of a nacelle assembly used with the turbine engine shown in FIG. 2 ;

FIG. 4 is another schematic illustration of the nacelle assembly shown in FIG. 3 ; and

FIG. 5 is a schematic illustration of an outer structure used with the nacelle assembly shown in FIG. 3 .

The systems and methods described herein provide an exemplary integrated inlet design that defines an aerodynamic nacelle assembly design. The description should enable one of ordinary skill in the art to make and use the disclosure, and the description describes several embodiments, adaptations, variations, alternatives, and uses of the disclosure, including what is presently believed to be the best mode of carrying out the disclosure. The disclosure is described herein as being applied to a preferred embodiment, namely, an integrated inlet design.

FIG. 1 is a schematic illustration of an exemplary aircraft 8 that includes at least one turbine engine 10 used to provide thrust for aircraft 8 . In the exemplary embodiment, turbine engine 10 is coupled to a wing 11 . More specifically, turbine engine 10 is coupled to a wing pylon (not shown) at wing 11 . In an alternate embodiment, turbine engine 10 is coupled to a tail 13 . More specifically, turbine engine 10 is coupled to a fuselage (not shown) at tail 13 .

FIG. 2 is a cross-sectional view of a portion of turbine engine 10 . In the exemplary embodiment, turbine engine 10 includes a fan assembly 16 that is rotatably coupled about a longitudinal centerline axis 32 . In the exemplary embodiment, fan assembly 16 is positioned at a forward end 33 of turbine engine 10 . In an alternative embodiment, fan assembly 16 is positioned at an aft end 35 of turbine engine 10 . Fan assembly 16 includes a plurality of rows of fan blades 19 positioned within a nacelle assembly 12 . In one embodiment, nacelle assembly 12 houses various operating components (not shown) of turbine engine 10 .

In the exemplary embodiment, turbine engine 10 also includes a core engine 17 that is positioned downstream from fan assembly 16 . Core engine 17 includes a compressor 18 , a combustor 20 , and a turbine 22 that is coupled to compressor 18 via a core rotor shaft 26 .

During operation, core engine 17 generates combustion gases that are channeled downstream to a turbine 24 which extracts energy from the gases for powering fan assembly 16 through a shaft 28 .

FIGS. 3 and 4 are schematic illustrations of nacelle assembly 12 . In the exemplary embodiment as shown in FIG. 3 , nacelle assembly 12 is generally annular and defines an opening 32 sized and configured to channel air through nacelle assembly 12 . Nacelle assembly 12 has a diameter 34 at forward end 33 measured with respect to an inner surface 43 that is smaller than a maximum diameter 36 of an outer surface of nacelle assembly 12 .

Nacelle assembly 12 includes an inner barrel 38 and an outer structure 40 . In the exemplary embodiment, outer structure 40 is fabricated from a continuous fiber reinforced thermoplastic material. Alternatively, outer structure 40 may be fabricated from a continuous fiber reinforced thermoset material. In one embodiment, outer structure 40 incorporates an integral surfacing film to facilitate protecting nacelle assembly 12 from lighting strikes.

Outer structure 40 includes a highlight 42 that is defined by forward end 33 of outer structure 40 , an outer aft section 44 that is defined by highlight 42 , and an inner aft section 46 that is defined by highlight 42 . Outer aft section 44 includes a point 45 on an outer surface 47 defined by the maximum diameter 36 of nacelle assembly 12 . In another embodiment, outer structure 40 is coupled to a hoop frame (not shown) at an aft end 35 of outer aft section 44 proximate to a fan cowl 54 , such that the hoop frame facilitates coupling outer structure 40 to fan cowl 54 .

In the exemplary embodiment as shown in FIG. 4 , inner aft section 46 includes a first flange 48 that is used to couple outer structure 40 to inner barrel 38 . In another embodiment, outer structure 40 includes a second flange 50 that is used to couple outer structure 40 to a fan case 53 .

In the exemplary embodiment, nacelle assembly 12 includes a plurality of bulkheads 57 , 58 , and 59 that facilitate structurally supporting nacelle assembly 12 . Bulkheads 57 , 58 , and 59 also facilitate absorbing any aftward forces induced to outer surface 47 of outer structure 40 , such as may be created during a bird-strike for example. In the exemplary embodiment, bulkhead 57 is coupled to outer aft section 44 and inner aft section 46 , bulkhead 58 is coupled to outer aft section 44 and inner barrel 38 , and bulkhead 59 is coupled to outer aft section 44 and fan case 53 . In the exemplary embodiment, bulkhead 59 is coupled to an aft end 35 of outer aft section 44 such that bulkhead 59 facilitates coupling outer structure 40 to fan cowl 54 . Bulkheads 57 , 58 , and 59 may be coupled through a plurality of coupling processes such as through a mechanical fastening process, a thermoplastic welding process, and an adhesive bonding process.

FIG. 5 is a schematic illustration of outer structure 40 . In the exemplary embodiment, outer structure 40 includes at least one stiffener 56 that is oriented either circumferentially and/or longitudinally to facilitate structurally supporting outer aft section 44 . In the exemplary embodiment, stiffener 56 is an independent component that is coupled to inner surface 43 of outer structure 40 via, for example, thermoplastic welding and/or an adhesive bonding. In an alternate embodiment, stiffener 56 may be formed integrally with outer structure 40 .

In the exemplary embodiment, nacelle assembly 12 includes a structural foam 60 that is oriented and positioned to facilitate absorbing any aftward forces induced to outer surface 47 of outer structure 40 , such as may be created during a bird-strike for example. In the exemplary embodiment, structural foam 60 is coupled downstream 35 from highlight 42 , against inner surface 43 of outer structure 40 using at least one of a thermoplastic welding process and an adhesive bonding process.

During operation, air flows from forward end 33 to aft end 35 of nacelle assembly 12 . Outer structure 40 defines an aerodynamic surface over outer surface 47 and facilitates maintaining laminar flow over outer surface 47 from highlight 42 to point 45 defined by maximum diameter 36 . Outer surface 47 facilitates reducing drag and improving aerodynamics and fuel efficiency. Moreover, outer structure 40 facilitates reducing assembly time and a number of components of nacelle assembly 12 as compared to known nacelle assemblies, thereby facilitating reducing manufacturing costs and improving manufacturing cycle time. The part count reduction and the use of advanced materials, such as thermoplastic or graphite, facilitate a weight reduction for nacelle assembly 12 as compared to known nacelle assemblies.

The methods, apparatus, and systems described herein for an integrated inlet design are not limited to the specific embodiments described herein. The integrated inlet design described herein provides a more aerodynamic design that facilitates reducing drag and increasing fuel efficiency. Practice of the methods, apparatus, or systems described or illustrated herein is not limited to nacelle assemblies. Rather, the methods, apparatus, and systems described or illustrated herein may be utilized independently and separately from other components and/or steps described herein.

The 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 languages of the claims.

While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.