Patent Description:
An airframe defines the mechanical structure of an aircraft. Airframes are made of multiple components that provide desired structural properties for aircraft. For example, a portion of an airframe for a fuselage of an aircraft may include frames, skin, and stringers that are mechanically coupled together (e.g., via co-bonding, co-curing, or fasteners) in accordance with design parameters. As presently practiced, sections of fuselage may be fabricated as full-barrel sections, and these full-barrel sections may be joined via circumferential splices. However, circumferential splicing involves the addition of splice straps and numerous affixation components around the entire circumference of the fuselage sections being joined. Hence, circumferential splices add substantial amounts of weight to an aircraft, and considerable amounts of labor are involved in the installation of circumferential splices, particularly with regard to aligning the splice strap with other affixation components.

<CIT>, according to its abstract, states a shear connection for interchangeably connecting two airframe parts of an aircraft. The shear connection may include first and second interface parts. The second interface part may overlap in the longitudinal direction at an overlap section with at least one of the first interface part and the first airframe part. The shear connection may further comprise means that rigidly and non-interchangeably attach the first interface part to the first airframe part and the second interface part to the second airframe part and a shear bolt that rigidly and interchangeably connects the second interface part with the at least one of the first interface part and the first airframe part at the overlap section, thereby interchangeably connecting the first and second airframe parts.

<CIT>, according to its abstract, states a method for assembling at least one panel of large longitudinal dimensions and small thickness, to another panel. The assembly method, by overlapping two edges and riveting at a plurality of points of two panels, consists in prebonding the two edges by means of a slow-curing mastic, in superimposing the two surfaces of the two edges to be assembled and in giving them a substantially permanent position; in pre-riveting the two edges at a number of first points corresponding to a low percentage of the total number of riveting points, these first points being located at substantially equal distance from one another, and in permanently riveting at the envisaged riveting points, starting from a pre-riveted point and riveting from successive point to point.

<CIT>, according to its abstract, states a fuselage element comprising a fuselage segment comprising a skin, and junction means for connecting the skin of said segment to an adjacent segment. The segment extends along the longitudinal axis of the fuselage, and the skin includes at least one first section at least one end of said segment. The fuselage segment comprises at least one longitudinal stiffening element for the fuselage. The junction means are arranged so as to contact an outer surface of said first section of the skin, and the stiffening element partially extends at the junction means along the longitudinal axis.

The dependent claims relate to preferred advantageous features. Embodiments described herein provide for fuselage sections having skins that are dimensioned to mate with each other (e.g., via a shiplap join, rabbet, or other feature). These complementary sections of skin enhance the strength of circumferential splices between barrel sections of fuselage, and also may eliminate the need for splice straps. This may reduce the amount of added weight involved in each circumferential splice, and may additionally reduce the complexity of aligning and assembling each circumferential splice. Thus, the embodiments provided herein result in technical benefits in the form of increased (or equal) amounts of strength compared to other joins, reduced weight, and reduced labor.

One embodiment is a method for assembling an airframe of an aircraft. The method includes forming a first skin of a first circumferential section of fuselage. The first skin includes a distal portion comprising a lip and a shoulder. The method further includes aligning a second skin of a second circumferential section of fuselage with the shoulder such that the lip overlaps the second skin, and still further includes affixing the first skin and the second skin together via a circumferential splice.

A further embodiment is a system comprising a portion of an airframe of an aircraft. The system includes a first skin of a first circumferential section of fuselage, which includes a distal portion comprising a lip and a shoulder. The system also includes a second skin of a second circumferential section of fuselage that is aligned with the shoulder, such that the lip overlaps the second skin, and still further includes a circumferential splice that affixes the first skin and the second skin together.

Other illustrative embodiments can be seen with reference to the following description and drawings.

The figures and the following description provide specific illustrative embodiments of the disclosure. It will thus be appreciated that those skilled in the art will be able to devise various arrangements that, although not explicitly described or shown herein, embody the principles of the disclosure Furthermore, any examples described herein are intended to aid in understanding the principles of the disclosure, and are to be construed as being without limitation to such specifically recited examples and conditions. As a result, the invention is not limited to the specific embodiments or examples described below, but is defined by the wording of the appended claims.

Many or all of the components discussed herein may be implemented as composite parts. Composite parts, such as Carbon Fiber Reinforced Polymer (CFRP) parts, are initially laid-up in multiple layers that together are referred to as a preform. Individual fibers within each layer of the preform may be aligned parallel with each other, but different layers exhibit different fiber orientations in order to increase the strength of the resulting composite part along different dimensions. Furthermore, some layers may comprise woven fabric made from fibers. The preform includes a viscous resin that solidifies in order to harden the preform into a composite part (e.g., for use in an aircraft). Carbon fiber that has been impregnated with an uncured thermoset resin or a thermoplastic resin is referred to as "prepreg. " Other types of carbon fiber include "dry fiber" which has not been impregnated with thermoset resin but may include a tackifier or binder. Dry fiber is infused with resin prior to hardening. For thermoset resins, the hardening is a one-way process referred to as curing, while for thermoplastic resins, the resin reaches a viscous form if it is re-heated, after which it can be consolidated to a desired shape and solidified. As used herein, the umbrella term for the process of transitioning a preform to a final hardened shape (i.e., transitioning a preform into a composite part) is referred to as "hardening," and this term encompasses both the curing of thermoset preforms and the forming/solidifying of thermoplastic preforms into a final desired shape.

Turning now to <FIG>, an illustration of an aircraft <NUM> is depicted for which the fabrication systems and methods described herein may be implemented. In this illustrative example, aircraft <NUM> includes wing <NUM> and wing <NUM> attached to fuselage <NUM> having a nose <NUM>. Aircraft <NUM> includes engine <NUM> attached to wing <NUM> and engine <NUM> attached to wing <NUM>. Tail section <NUM> is also attached to fuselage <NUM>. Horizontal stabilizer <NUM>, horizontal stabilizer <NUM>, and vertical stabilizer <NUM> are attached to tail section <NUM> of fuselage <NUM>. The fuselage <NUM> itself is formed from multiple barrel sections <NUM>-<NUM>, <NUM>-<NUM>, and <NUM>-<NUM> (referred to generally as "barrel sections <NUM>"). The barrel sections <NUM> have been joined together and define a circumference <NUM> of the fuselage <NUM>. In this embodiment, three different instances of barrel sections <NUM> are labeled, but any suitable number of barrel sections <NUM> may be utilized to form the fuselage <NUM> as a matter of design choice.

<FIG> depicts a cross section of a first circumferential section <NUM> (e.g., a barrel section <NUM>, a section forming a ninety-degree arc, or other curved section) of fuselage <NUM> of the aircraft <NUM> of <FIG> in an illustrative embodiment. First circumferential section <NUM> is referred to as "circumferential" because it forms a part of the circumference <NUM> of fuselage <NUM>. First circumferential section <NUM> may thus comprise one of barrel sections <NUM> of <FIG>.

As shown in <FIG>, first circumferential section <NUM> comprises a first skin <NUM>, to which stringers <NUM> and frames <NUM> are attached. In this embodiment, first circumferential section <NUM> is circular, having a center <NUM>.

<FIG> is an interior view of a circumferential splice <NUM> between circumferential sections of fuselage in an illustrative embodiment, and corresponds with view arrows <NUM> of <FIG>. The circumferential splice <NUM> extends around the entirety of the first circumferential section <NUM> (e.g., barrel section <NUM>-<NUM>), however, only a portion of the circumferential splice <NUM> is shown in this view. As shown in <FIG>, the circumferential splice <NUM> is between a first circumferential section <NUM> (including first skin <NUM>, stringers <NUM>, and frames <NUM>) and a second circumferential section <NUM> (including second skin <NUM>, stringers <NUM>, and frames <NUM>). At the circumferential splice <NUM>, the first circumferential section <NUM> and the second circumferential section <NUM> (e.g., barrel section <NUM>-<NUM>) are coupled together via a lap join <NUM>. The first circumferential section <NUM> includes a first skin <NUM> and stringers <NUM>, while the second circumferential section <NUM> includes a second skin <NUM> and stringers <NUM>. The lap join <NUM> includes a variety of components that affix the first circumferential section <NUM> to the second circumferential section <NUM>, as discussed below with regard to the following FIGS. Depending on embodiment, these various components may be affixed via fastening, co-curing, or co-bonding.

<FIG> is side view of the circumferential splice <NUM> in an illustrative embodiment, and corresponds with view arrows <NUM> of <FIG>. As shown in <FIG>, the circumferential splice <NUM> includes a number of interconnected components. Starting generally from the bottom of <FIG> and moving upwards (i.e., proceeding from outboard to inboard), the circumferential splice <NUM> includes second skin <NUM>, which forms a lap join <NUM> (e.g., a ship lap join, rabbet join, etc.) with first skin <NUM>. Specifically, second skin <NUM> is disposed underneath a lip <NUM> of first skin <NUM>, and is indexed relative to shoulder <NUM> of first skin <NUM> via Determinant Assembly (DA) holes <NUM>, leaving a gap <NUM> for aerosealing. In one embodiment, DA holes <NUM> are implemented as pilot holes that are aligned at a top, left, and right of the first skin <NUM> and the second skin <NUM>. When the pilot holes are aligned with respect to the skins, the skins are known to be in a desired alignment.

The lip <NUM> and the shoulder <NUM> are disposed at a distal portion <NUM> of the first skin <NUM>. Proximate to or at the circumferential splice <NUM>, a thickness (TB) of second skin <NUM> increases at ramp <NUM>, and a thickness (TA) of first skin <NUM> increases at ramp <NUM>. This increases the strength of the lap join <NUM>. In this embodiment, TB is less than TA after the ramps have fully increased the thickness of the first skin <NUM> and the second skin <NUM>. This leaves sufficient material for lip <NUM> to remain after the first skin <NUM> is machined.

Stringer <NUM>, having flanges <NUM> and a body <NUM>, is disposed atop second skin <NUM>, while stringer <NUM>, having flanges <NUM> and a body <NUM>, is disposed atop first skin <NUM>. Flanges <NUM> extend along the lip <NUM> towards the second skin <NUM>. A filler <NUM> is disposed atop a flange <NUM> of the stringer <NUM>. A splice fitting <NUM> (e.g., an H-fitting, L-bracket, T-bracket, or other) rests partially atop the filler <NUM>, and partially atop the flange <NUM>. The splice fitting <NUM> is affixed via fasteners <NUM>, and a lowboy <NUM> and shear tie splice <NUM> protrude upward from the splice fitting <NUM>. Specifically, in location <NUM>, the fasteners <NUM> are driven from outboard to inboard through second skin <NUM>, flange <NUM>, filler <NUM>, and splice fitting <NUM>. In location <NUM>, the fasteners <NUM> are driven from outboard to inboard through second skin <NUM>, an optional shim (e.g., shim <NUM> of <FIG>), lip <NUM>, flange <NUM>, and splice fitting <NUM>. In location <NUM>, the fasteners <NUM> are driven from outboard to inboard through first skin <NUM>, flange <NUM>, and splice fitting <NUM>.

A frame <NUM> (e.g., a composite or titanium frame) has been added to this FIG. , and is affixed to the shear tie splice <NUM>. While the shear tie splice <NUM> is illustrated as being centered over the lap join <NUM>, in further embodiments the shear tie splice <NUM> is disposed to the left or the right of the lap join <NUM>. In still further embodiments, stringer <NUM> or stringer <NUM> may extend fore or aft beyond the lap join <NUM>.

Utilizing the arrangement of <FIG>, there is no need for a splice strap. Hence, a technical benefit is provided because the labor associated with aligning a splice strap is saved, and weight at the circumferential splice <NUM> may therefore be reduced because fewer components are used.

Further details of the circumferential splice <NUM> are provided with respect to <FIG>, which is a zoomed in view of the lap join <NUM> of <FIG> in an illustrative embodiment. Specifically, <FIG> corresponds with region <NUM> of <FIG>. In <FIG>, the filler <NUM> and an end <NUM> of the lip <NUM> are separated by a gap <NUM>. However, in many embodiments the filler <NUM> is butted against the end <NUM> of the lip <NUM>.

In this embodiment, an optional shim <NUM> has been included between the lip <NUM> and the second skin <NUM>. The combined thickness of the flange <NUM> and filler <NUM> is equal to a combined thickness of the shim <NUM>, lip <NUM>, and flange <NUM>. This results in a flat plane <NUM> for receiving the splice fitting <NUM>. The shim <NUM> is not a splice strap because it is not a structural component of the circumferential splice <NUM>, and also because the shim <NUM> does not form a single strap joint across two butted sections of skin.

<FIG> is a perspective view of stringers <NUM> and <NUM> that cover a portion <NUM> of a circumferential splice <NUM> in an illustrative embodiment, and corresponds with view arrows <NUM> of <FIG>. In <FIG>, flanges <NUM> of stringer <NUM> extend beyond a body <NUM> of the stringer <NUM> and across the lap join <NUM> along lip <NUM>. Meanwhile, flanges <NUM> of stringer <NUM> are covered by fillers <NUM>. The top surfaces <NUM> of the flanges <NUM>, and the top surfaces <NUM> of the fillers <NUM>, are coplanar with each other and define flat plane <NUM> of <FIG>, enabling the splice fitting <NUM> to straddle these components when installed. The orientation of flat planes <NUM> along the circumference <NUM> of the fuselage <NUM> varies as the circumference <NUM> is traversed, to accommodate the local geometry of the circumference <NUM> (i.e., the curvature of the circumference <NUM>).

<FIG> is a further view of the circumferential splice <NUM> of <FIG> in an illustrative embodiment, and corresponds with view arrows <NUM> of <FIG>. Specifically, <FIG> provides a different viewing angle of the circumferential splice <NUM>, and additionally depicts a frame <NUM> affixed to the circumferential splice <NUM>.

<FIG> are additional views that provide additional context pertaining to the components recited above. <FIG> is a zoomed in view of the circumferential splice <NUM> of <FIG> in an illustrative embodiment, and corresponds with view arrows <NUM> of <FIG>. Specifically, <FIG> provides context by depicting splice fitting <NUM>, lowboy <NUM>, and shear tie splice <NUM> from a new viewing angle. In this view, frame <NUM> has been omitted. <FIG> is a zoomed in view of a splice fitting <NUM> that forms a portion of a circumferential splice <NUM> in an illustrative embodiment, and corresponds with region <NUM> of <FIG>. In this view, frame <NUM> has been included, and frame <NUM> overlaps splice fitting <NUM>, lowboy <NUM>, and shear tie splice <NUM>.

<FIG> is a perspective view of a splice fitting <NUM> in an illustrative embodiment. <FIG> makes clear that splice fitting <NUM> is symmetrical about line <NUM>, but is asymmetrical about line <NUM> (i.e., from fore to aft). Specifically, a first portion <NUM> of a base <NUM> (e.g., an underside) of the splice fitting <NUM> rests on the filler <NUM>, and is shaped differently than a second portion <NUM> of the base <NUM> that rests on or otherwise contacts a flange <NUM> of the stringer <NUM>. This difference in shape from fore to aft accommodates placement of the splice fitting onto the filler <NUM> (e.g., on the aftward side) as well as the flanges <NUM> of stringer <NUM> (e.g., on the forward side). It will be understood that a completed circumferential splice <NUM> will include splice fittings <NUM> along the entirety of the circumference <NUM> of the fuselage <NUM>, in order to straddle the lap join <NUM> and enhance strength.

With a discussion of the overall shape of the circumferential splice <NUM> provided with regard to <FIG>, further discussion focuses on ply sequencing stackups for the first skin <NUM>, and method for fabricating the circumferential splices <NUM> discussed herein.

<FIG> depicts a sequence <NUM> of plies for a first skin <NUM> in an illustrative embodiment. The thickness of the first skin <NUM> increases from left to right as additional plies are integrated. Specifically, the sequence <NUM> dictates the number and fiber orientation of plies <NUM> of fiber reinforced material <NUM> (e.g., CFRP) that each are made from unidirectional tows arranged at a desired fiber angle. In this embodiment, the angles include ninety degrees (as shown in fiber orientation <NUM>), zero degrees (as shown in fiber orientation <NUM>), plus forty-five degrees (as shown in fiber orientation <NUM>), and minus forty-five degrees (as shown in fiber orientation <NUM>). Design constraints may apply certain requirements to a sequence <NUM>. For example, design constraints may require that the fiber orientations along the entirety of the thickness <NUM> of the sequence <NUM> be symmetrical (e.g., with respect to centerline <NUM>). This presents a difficulty because first skin <NUM> will have a lip <NUM> machined out of it, and the lip will have a different thickness than the rest of the first skin <NUM>.

To address this difficulty, the sequence <NUM> exhibits "double symmetry," wherein a region <NUM> corresponding with the lip <NUM> exhibits symmetry in fiber orientations about its centerline <NUM>, while the entirety of the thickness <NUM> of the ply sequence <NUM>, corresponding with a combined thickness of the lip <NUM> and a shoulder <NUM>, also exhibits symmetry, but with respect to centerline <NUM>. This enables a first skin <NUM> to comply with design requirements pertaining to symmetry of fiber orientations across a centerline, even when a lip <NUM> is machined from the first skin <NUM>.

<FIG> depict further sequences that are "double symmetric" arrangements of plies. In <FIG>, a uniform ply length has been provided for each ply. However, similar ramping techniques to those provided in <FIG> may be utilized in order to increase the thickness of a first skin <NUM> from left to right. In <FIG>, sequence <NUM> arranges plies <NUM> having fiber orientations <NUM>, <NUM>, <NUM>, and <NUM> to form a region <NUM> corresponding with a lip <NUM>. The region <NUM> is a portion that is left after machining off a portion (e.g., a third) of a thickness <NUM> of the sequence <NUM>. Region <NUM> is symmetrical about centerline <NUM>, and a thickness <NUM> of the sequence <NUM> is symmetrical about centerline <NUM>.

In <FIG>, additional plies <NUM> of woven fabric have been integrated into a sequence <NUM>. Sequence <NUM> arranges plies <NUM> having fiber orientations <NUM>, <NUM>, <NUM>, <NUM>, and <NUM> to form a region <NUM> corresponding with a lip <NUM>. The region <NUM> is a portion that is left after machining off a portion (e.g., one half, one third, two thirds, or any fraction) of a thickness <NUM> of the sequence <NUM>. Region <NUM> is symmetrical about centerline <NUM>, and a thickness <NUM> of the sequence <NUM> is symmetrical about centerline <NUM>.

In <FIG>, a sequence <NUM> of plies <NUM> is provided having fiber orientations of zero degrees (e.g., fiber orientation <NUM>), sixty degrees (e.g., fiber orientation <NUM>), and negative sixty degrees (e.g., fiber orientation <NUM>). The sequence <NUM> exhibits double symmetry. Specifically, region <NUM> is symmetrical about centerline <NUM>, and a thickness <NUM> of sequence <NUM> is symmetrical about centerline <NUM>.

In <FIG>, a further sequence <NUM> of plies <NUM> is provided having fiber orientations of zero degrees (e.g., fiber orientation <NUM>), sixty degrees (e.g., fiber orientation <NUM>), and negative sixty degrees (e.g., fiber orientation <NUM>). The sequence <NUM> exhibits double symmetry. Specifically, region <NUM> is symmetrical about centerline <NUM>, and a thickness <NUM> of sequence <NUM> is symmetrical about centerline <NUM>. Illustrative details of the formation of a circumferential splice <NUM> will be discussed with regard to <FIG>. Assume, for this embodiment, that first circumferential section <NUM> and second circumferential section <NUM> await joining together.

<FIG> is a flowchart illustrating a method <NUM> for forming a circumferential splice <NUM>. The steps of method <NUM> are described with reference to the components discussed in the FIGS. above, but those skilled in the art will appreciate that method <NUM> may be performed in other systems. The steps of the flowcharts described herein are not all inclusive and may include other steps not shown. The steps described herein may also be performed in an alternative order.

In step <NUM>, a first skin <NUM> of a first circumferential section <NUM> of fuselage <NUM> is formed. The first skin <NUM> includes a distal portion <NUM> comprising a lip <NUM> and a shoulder <NUM>. In optional step <NUM>, a portion of a thickness of the first skin <NUM> is machined off to form the lip <NUM> and the shoulder <NUM>. Such an operation may be performed via a mill, cutter, or other suitable machinery, such as machining tool <NUM> of <FIG>.

Step <NUM> comprises optionally inserting a shim <NUM> between the lip <NUM> and a second skin <NUM> of the second circumferential section <NUM> of fuselage <NUM>.

Step <NUM> comprises aligning the second skin <NUM> of the second circumferential section <NUM> with the shoulder <NUM> such that the lip <NUM> overlaps the second skin <NUM>. This operation is performed circumferentially along the entirety of the second circumferential section <NUM>. The result is a lap join <NUM> (e.g., a shiplap join, rabbet join, etc.), wherein the second skin <NUM> is nested against the lip <NUM> and the shoulder <NUM> of the first skin <NUM>. Hence, although the second skin <NUM> is butted against the first skin <NUM> or otherwise aligned therewith, the end result is not a butt joint.

Step <NUM> includes affixing the first skin <NUM> and the second skin <NUM> together via a circumferential splice <NUM>. In one embodiment, step <NUM> comprises aligning fillers <NUM>, splice fittings <NUM>, lowboys <NUM>, and/or shear tie splices <NUM>, and affixing these components via fasteners, co-curing, or co-bonding. Step <NUM> comprises optionally placing a filler <NUM> onto a flange <NUM> of a stringer <NUM> that is attached to the second skin <NUM>.

Step <NUM> comprises optionally straddling the splice fitting <NUM> across the first skin <NUM> and the second skin <NUM>. In one embodiment, this comprises placing second portion <NUM> of the base <NUM> of the splice fitting <NUM> onto the flange <NUM> of the stringer <NUM> at the first skin <NUM>, and then sliding the filler <NUM> between the first portion <NUM> of the base <NUM> of the splice fitting <NUM> and the flange <NUM> of the stringer <NUM>.

Step <NUM> comprises optionally aligning the filler <NUM> with an end <NUM> of the lip <NUM>, such that a combined thickness of the filler <NUM> and the flange <NUM> of the stringer <NUM> at the second circumferential section <NUM> of fuselage <NUM> is equal to a combined thickness of the lip <NUM>, a shim <NUM> that contacts the lip <NUM>, and a flange <NUM> of a stringer <NUM> at the first circumferential section <NUM>. In one embodiment, this comprises sliding the filler <NUM> underneath the splice fitting <NUM> and atop the flange <NUM>, such that the filler <NUM> is sandwiched between the splice fitting <NUM> and the flange <NUM>.

Step <NUM> comprises optionally driving fasteners <NUM> through the first skin <NUM>, flange <NUM>, and/or splice fitting <NUM> at a location <NUM>. Step <NUM> comprise optionally driving fasteners <NUM> through the second skin <NUM>, flange <NUM>, filler <NUM>, and/or splice fitting <NUM> at a location <NUM>.

With the circumferential splice <NUM> completed, additional structure may be attached. For example, step <NUM> comprises optionally installing a frame <NUM> at the circumferential splice <NUM>. In further embodiments, the frame <NUM> is installed as a part of step <NUM>.

Step <NUM> comprises optionally forming the circumferential splice <NUM> without using a splice strap (i.e., foregoing/omitting installation of a splice strap). Although not illustrated, a splice strap is a structural component of a splice that overlaps two butted (or otherwise aligned) skins, forming a single strap joint. The circumferential splices <NUM> discussed herein forego the need for a splice strap, because lip <NUM> provides the structural purpose of strengthening that would be performed by a splice strap.

Method <NUM> provides a technical benefit by providing a strengthened join between sections of fuselage <NUM> with respect to prior systems that utilized a splice strap to form a single strap joint. Furthermore, method <NUM> reduces the amount of labor involved in aligning components of a circumferential splice <NUM> (because there is no need to align with a splice strap), and reduces weight (because the splice strap is eliminated). This results in benefits pertaining to assembly, as well as to reduced fuel consumption costs.

In the following examples, additional processes, systems, and methods are described in the context of a circumferential splice <NUM> for circumferential sections of fuselage in an illustrative embodiment.

<FIG> is a block diagram of a circumferential splice <NUM> between a first circumferential section <NUM> and a second circumferential section <NUM> of fuselage <NUM> in an illustrative embodiment. In this embodiment, first circumferential section <NUM> includes a first skin <NUM> having a lip <NUM> and a shoulder <NUM>, and further includes a stringer <NUM> having a flange <NUM>. Second circumferential section <NUM> includes second skin <NUM> that is aligned with (e.g., butted against or nested at) shoulder <NUM>, and further includes a stringer <NUM> having a flange <NUM>. A filler <NUM> rests atop the flange <NUM>, and a splice fitting <NUM> rests atop filler <NUM> and flange <NUM> of stringer <NUM>. Fasteners <NUM> affix the splice fitting <NUM> in place. Lowboy <NUM> is attached to splice fitting <NUM>, as is shear tie splice <NUM>. A frame <NUM> is attached to the shear tie splice <NUM>. Furthermore, a shim <NUM> is disposed between the lip <NUM> and the second skin <NUM>. <FIG> further depicts a machining tool <NUM> for machining the lip <NUM> from a thickness (T3) the first skin <NUM>. The machining tool <NUM> may comprise, for example, a mill, a circular or reciprocating saw, or other tool.

<FIG> further illustrates that a combined thickness (T2) of the filler <NUM> and a flange <NUM> of the stringer <NUM> at the second circumferential section <NUM> of fuselage <NUM> is equal to a combined thickness (T1) of the lip <NUM>, a shim <NUM> contacting the lip <NUM>, and a flange <NUM> of a stringer <NUM> at the first circumferential sectional <NUM>. Furthermore, a thickness (T4) of the second skin <NUM> and a shim <NUM> equals a thickness (T5) of the shoulder <NUM>.

Referring more particularly to the drawings, embodiments of the disclosure may be described in the context of aircraft manufacturing and service in method <NUM> as shown in <FIG> and an aircraft <NUM> as shown in <FIG>. During pre-production, method <NUM> may include specification and design <NUM> of the aircraft <NUM> and material procurement <NUM>. During production, component and subassembly manufacturing <NUM> and system integration <NUM> of the aircraft <NUM> takes place. Thereafter, the aircraft <NUM> may go through certification and delivery <NUM> in order to be placed in service <NUM>. While in service by a customer, the aircraft <NUM> is scheduled for routine work in maintenance and service <NUM> (which may also include modification, reconfiguration, refurbishment, and so on). Apparatus and methods embodied herein may be employed during any one or more suitable stages of the production and service described in method <NUM> (e.g., specification and design <NUM>, material procurement <NUM>, component and subassembly manufacturing <NUM>, system integration <NUM>, certification and delivery <NUM>, service <NUM>, maintenance and service <NUM>) and/or any suitable component of aircraft <NUM> (e.g., airframe <NUM>, systems <NUM>, interior <NUM>, propulsion system <NUM>, electrical system <NUM>, hydraulic system <NUM>, environmental system <NUM>).

As shown in <FIG>, the aircraft <NUM> produced by method <NUM> may include an airframe <NUM> with a plurality of systems <NUM> and an interior <NUM>. Examples of systems <NUM> include one or more of a propulsion system <NUM>, an electrical system <NUM>, a hydraulic system <NUM>, and an environmental system <NUM>. Any number of other systems may be included. Although an aerospace example is shown, the principles of the invention may be applied to other industries, such as the automotive industry.

As already mentioned above, apparatus and methods embodied herein may be employed during any one or more of the stages of the production and service described in method <NUM>. For example, components or subassemblies corresponding to component and subassembly manufacturing <NUM> may be fabricated or manufactured in a manner similar to components or subassemblies produced while the aircraft <NUM> is in service. Also, one or more apparatus embodiments, method embodiments, or a combination thereof may be utilized during the subassembly manufacturing <NUM> and system integration <NUM>, for example, by substantially expediting assembly of or reducing the cost of an aircraft <NUM>. Similarly, one or more of apparatus embodiments, method embodiments, or a combination thereof may be utilized while the aircraft <NUM> is in service, for example and without limitation during the maintenance and service <NUM>. Thus, the invention may be used in any stages discussed herein, or any combination thereof, such as specification and design <NUM>, material procurement <NUM>, component and subassembly manufacturing <NUM>, system integration <NUM>, certification and delivery <NUM>, service <NUM>, maintenance and service <NUM> and/or any suitable component of aircraft <NUM> (e.g., airframe <NUM>, systems <NUM>, interior <NUM>, propulsion system <NUM>, electrical system <NUM>, hydraulic system <NUM>, and/or environmental system <NUM>).

Claim 1:
A method for assembling an airframe of an aircraft, the method comprising:
forming a first skin (<NUM>) of a first circumferential section (<NUM>) of fuselage, the first skin including a distal portion comprising a lip (<NUM>) and a shoulder (<NUM>);
aligning a second skin (<NUM>) of a second circumferential section (<NUM>) of fuselage with the shoulder (<NUM>) such that the lip (<NUM>) overlaps the second skin (<NUM>); and
affixing the first skin (<NUM>) and the second skin (<NUM>) together via a circumferential splice (<NUM>),
wherein affixing the first skin (<NUM>) and the second skin (<NUM>) together comprises
placing a filler (<NUM>) onto a flange (<NUM>) of a stringer (<NUM>) at the second circumferential section (<NUM>) of fuselage, wherein a combined thickness (T2) of the filler (<NUM>) and the flange (<NUM>) of the stringer (<NUM>) at the second circumferential section (<NUM>) of fuselage is equal to a combined thickness (T1) of the lip (<NUM>), an optional shim (<NUM>) contacting the lip (<NUM>), and a flange (<NUM>) of a stringer (<NUM>) at the first circumferential section (<NUM>), thereby forming a flat plane for receiving a splice fitting (<NUM>).