Method for producing an aircraft

Abstract

The present invention relates to a method for producing an aircraft. The method includes determining structural dimensional requirements for mating components and for assembly tolerances between mating components. The method establishes a baseline specification for the mating components and establishes an index specification for mating the mating components. The method performs a variation analysis based on the baseline specification and the index specification to determine an assembly analysis result for the mating components. The method compares the assembly analysis result for the mating components with the corresponding assembly tolerances defined in the structural dimensional requirements for the mating components to verify that the baseline specification and the index specification meet the assembly tolerances between the mating components.

Description

Technical Field

[0001] The subject matter herein generally relates to a method for producing an aircraft. Background Art

[0002] Aircraft manufacturing occurs in manufacturing facilities. Many aircraft parts need to be completed and assembled. The design of aircraft parts is time-consuming and labor-intensive. The assembly process for aircraft involves assembling, aligning and joining large complex parts. For example, the wing assembly and the tail assembly can be basically assembled as a whole and then attached to the corresponding parts of the aircraft body. As another embodiment, the fuselage of the aircraft can be a combination of several body assemblies. The components of the aircraft are assembled with high precision, for example, the relative position tolerance between the components is less than 0.005 inches (about 0.1 mm). During the life of the aircraft program or in the design of a new aircraft program, various aircraft components may need to be redesigned, which may change the position of the structurally important features of the aircraft components. Other aircraft components usually need to be redesigned to adapt to such changes, thereby ensuring the assembly and construction tolerances of the aircraft. Redesigning components in the later stages of aircraft design is usually more expensive and affects the redesign of a larger number of components.

[0003] There remains a need for a method of producing aircraft in a cost-effective and reliable manner that maintains the functional integrity of the aircraft and the quality standards of the aircraft program. Summary of the invention

[0004] In one embodiment, a method for producing an aircraft is provided. The method determines structural dimensional requirements for a mating component of a first aircraft component, a mating component of a second aircraft component, and an assembly tolerance between the mating components of the first aircraft component and the second aircraft component. The method develops a baseline specification for the mating component of the first aircraft component and for the mating component of the second aircraft component, and develops an index specification for mating the mating component of the first aircraft component with the mating component of the second aircraft component. The method performs a variation analysis based on the baseline specification and the index specification to determine an assembly analysis result for the mating component of the first aircraft component and the second aircraft component. The method compares the assembly analysis result for the mating component of the first aircraft component and the second aircraft component with the assembly tolerance determined by the structural dimensional requirements for the mating component of the first aircraft component and the second aircraft component to verify that the baseline specification and the index specification meet the assembly tolerance between the mating component of the first aircraft component and the mating component of the second aircraft component.

[0005] In another embodiment, a method for producing an aircraft is provided. The method determines structural dimension requirements for a mating component of a first aircraft component, for a mating component of a second aircraft component, and for an assembly tolerance between the mating component of the first aircraft component and the mating component of the second aircraft component. The method develops a baseline specification for the mating component of the first aircraft component and for the mating component of the second aircraft component, and develops an index specification. The index specification defines an assembly sequence for mating the mating component of the first aircraft component with the mating component of the second aircraft component. The method performs a variation analysis based on the baseline specification and the index specification to determine an assembly analysis result for the mating component of the first aircraft component and the mating component of the second aircraft component. The method compares the assembly analysis result for the mating component with an assembly tolerance determined by the structural dimension requirements for the mating component of the first aircraft component and the mating component of the second aircraft component to verify that the baseline specification and the index specification meet the assembly tolerance between the mating component of the first aircraft component and the mating component of the second aircraft component. The method assembles the first aircraft component with the second aircraft component according to the assembly sequence based on the index specification and the baseline specification.

[0006] In another embodiment, a method for producing an aircraft by wing-to-body-joint assembly is provided. The method determines structural dimension requirements for wing assembly mating components of a wing assembly, the wing assembly including a center wing section, a right wing section attached to the center wing section, and a left wing section attached to the center wing section. The method determines structural dimension requirements for fuselage mating components of a fuselage including a middle body section configured to receive the wing assembly, and determines structural dimension requirements for assembly tolerances between the wing assembly mating components and the fuselage mating components. The method develops baseline specifications for the wing assembly mating components and for the fuselage mating components, and develops index specifications for mating the wing assembly mating components with the fuselage mating components. The method performs variation analysis based on the baseline specifications and the index specifications to determine assembly analysis results for the wing assembly mating components and the fuselage mating components. The method compares the assembly analysis results with the assembly tolerances determined by the structural dimension requirements to verify that the baseline specifications and the index specifications meet the assembly tolerances between the wing assembly mating components and the fuselage mating components. BRIEF DESCRIPTION OF THE DRAWINGS

[0007] Figure 1 is a perspective view of an assembly system for assembling a vehicle, such as an aircraft, according to an embodiment.

[0008] Figure 2 is a flow chart of a process for producing an aircraft according to an embodiment.

[0009] Figure 3 is a flow chart of a process for producing an aircraft according to an embodiment.

[0010] Figure 4 is a process flow diagram of producing an aircraft through a wing-to-body-joined assembly process according to an embodiment.

[0011] Figure 5 is a cross-sectional view of a portion of an aircraft showing a wing assembly positioned within a fuselage according to an embodiment.

[0012] Figure 6 is a top view of a portion of an aircraft showing a wing assembly ready for assembly with a fuselage according to an embodiment.

[0013] Figure 7 is a cross-sectional view of a 3-D solid model of a wing assembly showing a datum reference system for mating components of the wing assembly according to an embodiment.

[0014] Figure 8 is a front view of a 3-D solid model of a wing assembly showing reference specifications for mating components of the wing assembly according to an embodiment.

[0015] Fig. 9 is a rear view of a 3-D solid model of a wing assembly showing reference specifications for mating components of the wing assembly according to an embodiment.

[0016] Fig.10 is a rear perspective view of a 3-D solid model of a wing assembly showing reference planes for mating components of the wing assembly according to an embodiment.

[0017] Fig.11 is a rear view of a 3-D solid model of a wing assembly showing reference features for mating components of the wing assembly according to an embodiment.

[0018] Fig.12 is a rear perspective view of a 3-D solid model of a wing assembly showing baseline specifications for mating components of the wing assembly according to an embodiment.

[0019] Fig.13 is a cross-sectional view of a 3-D solid model of a fuselage showing a fiducial reference system for mating components of the fuselage according to an embodiment.

[0020] Fig.14 is a cross-sectional view of a 3-D solid model of a fuselage showing reference planes for mating components of the fuselage according to an embodiment.

[0021] Fig.15 is a top view of a 3-D solid model of a fuselage showing reference features for mating components of the fuselage according to an embodiment.

[0022] Fig.16 is a cross-sectional view of a 3-D solid model of a fuselage showing reference planes for mating components of the fuselage according to an embodiment.

[0023] Fig.17 is a front view of a 3-D solid model of a fuselage showing reference features for mating components of the fuselage according to an embodiment.

[0024] Fig.18 is a cross-sectional view of a 3-D solid model of a fuselage showing reference planes for mating components of the fuselage according to an embodiment.

[0025] Fig.19 is a top view of a 3-D solid model of a fuselage showing reference features for mating components of the fuselage according to an embodiment.

[0026] Fig. 20 is a cross-sectional view of a 3-D solid model of a portion of an aircraft showing wing assemblies positioned relative to a fuselage according to an embodiment.

[0027] Fig.21 is a front perspective view of a portion of an aircraft showing a wing assembly positioned relative to a fuselage according to an embodiment.

[0028] Fig. 22 is a front view of a portion of an aircraft showing a wing assembly positioned relative to a fuselage according to an embodiment.

[0029] Fig.23 is a front view of a portion of an aircraft showing a wing assembly positioned relative to a fuselage according to an embodiment.

[0030] Fig.24 is a front perspective view of a portion of an aircraft showing a wing assembly positioned relative to a fuselage according to an embodiment.

[0031] Fig.25 is a cross-sectional view of a portion of an aircraft showing a wing assembly positioned relative to a fuselage according to an embodiment.

[0032] Fig.26 is a cross-sectional view of a portion of an aircraft showing a wing assembly positioned relative to a fuselage according to an embodiment.

[0033] Fig. 27 A portion of an aircraft is shown showing a wing assembly positioned relative to a fuselage according to an embodiment.

[0034] Fig.28 is a perspective view of a portion of an aircraft showing a wing assembly positioned relative to a fuselage according to an embodiment. DETAILED DESCRIPTION

[0035] Embodiments described herein provide systems and methods for performing producibility analysis for engineering design (product) and construction (process), such as for aircraft. The systems and methods are used to manufacture parts of an aircraft, such as for joining parts of an aircraft together. In various embodiments, the systems and methods are used for wing-to-body-joined assembly of an aircraft. In other various embodiments, the systems and methods can be used for other parts of an aircraft, such as for joining the tail to the fuselage, for joining the main body of the fuselage together, for joining the nose to the fuselage, for joining the engine to the wing, for joining the wing to the center wing section, for joining the wing leading edge panels together, for joining the pylon to the wing, for joining the internal aircraft parts together, or for joining other parts of the aircraft.

[0036] Embodiments of a producibility analysis system utilize a variation analysis process to analyze the manufacturability of an engineering design and build to verify that the engineering design and build are within the manufacturing capabilities of a manufacturing facility used to produce an aircraft. Embodiments described herein utilize baseline specifications and target specifications to perform a variation analysis to verify that the design and build sequence are within the structural dimensional requirements of mating components for the components being joined and within the assembly tolerances for the components being joined. Embodiments of a producibility analysis system define structural dimensional requirements for mating components being joined and compare the structural dimensional requirements to the manufacturing capabilities to negotiate assembly tolerances with the engineering design.

[0037] Embodiments of the systems and methods described herein may be employed in the early design stages to focus design resources on problem areas and avoid redesign or redoing in later stages. A producibility analysis method is performed to determine whether one or more components contribute to undesirable geometric changes that may result in negative margins (e.g., interference of components), require design modifications, require tool modifications, or require modifications to the construction index specifications for joining components. The producibility analysis system identifies and avoids expensive design flaws or poorly planned manufacturing methods in the early design stages. The producibility analysis system verifies the construction sequence, prioritizes the placement of key features of the components, and analyzes the design and construction constraints to verify that the components can be efficiently assembled or disassembled. The producibility analysis method is performed virtually, such as using a 3-D solid model, so that graphical visualization and direct manipulation can be performed in a simulated manufacturing environment. The producibility analysis method has the ability to be easily updated as the engineering design and construction mature over time.

[0038] Figure 11 is a perspective view of an assembly system 100 for assembling a vehicle, such as an aircraft 10, according to an embodiment. Although the assembly system 100 is shown and described herein with reference to the design and manufacture of the aircraft 10, it is recognized that the assembly system 100 may be used for other types of vehicles, such as motor vehicles, water vehicles, military vehicles, other types of aerospace vehicles, etc. The assembly system 100 may be disposed in a manufacturing facility 102. The aircraft 10 may be manufactured in various stages at various stations 104 within the manufacturing facility 102. For example, various components of the aircraft 10 may be pre-assembled at one or more stations (or at other remote manufacturing facilities) and joined at other stations within the manufacturing facility 102 to complete the aircraft 10. The assembly system 100 utilizes a manufacturability analysis system 110 to analyze the manufacturability of the various components to verify that the engineering design of the components and the build sequence for joining the various components are within the manufacturing capabilities of the manufacturing facility 102 for producing the aircraft 10. Manufacturability analysis system 110 may be disposed within manufacturing facility 102 , or may be located remotely from manufacturing facility 102 .

[0039] The aircraft 10 may be a commercial aircraft, a military aircraft, a helicopter, or another type of aircraft 10. In the illustrated embodiment, the aircraft 10 includes a propulsion system 12 having two engines 14 for propelling the aircraft 10. The engines 14 may be gas turbine engines. Alternatively, the propulsion system 12 may include more engines 14 than shown. The engines 14 may be carried by the wings 16 of the aircraft 10. In other embodiments, the engines 14 may be carried by the fuselage 18 and / or the tail 20. The tail 20 includes a horizontal stabilizer 22 and a vertical stabilizer 24. The fuselage 18 of the aircraft 10 may define an interior compartment or area, such as a passenger cabin, a cockpit, a cargo area, etc. The aircraft 10 includes a nose 26 and a tail 28 located at opposite ends of the fuselage 18.

[0040] In an embodiment, the aircraft 10 includes a wing assembly 30 that is configured to be joined to a mid-body section 32 of the fuselage 18. The wing assembly 30 may be preassembled and joined to the mid-body section 32 of the fuselage 18 during a wing-to-body-joining process. In various other embodiments, the right and left wings may be directly joined to the mid-body section 32 of the fuselage 18 rather than being preassembled like the wing assembly 30. The wing assembly 30 includes a center wing section 34, a right wing section 36, and a left wing section 38. The center wing section 34, the right wing section 36, and the left wing section 38 are individually assembled and joined at one or more wing building stations.

[0041] In various embodiments, the tail 28 is joined to the main body of the fuselage 18 at the tail joining station. The wing assembly 30 is transported as a complete unit to be joined to the fuselage 18 at the wing joining station. The engine 14 is attached to the wing 16 at the engine joining station. The aircraft 10 can be transported to one or more aircraft processing stations downstream of the engine joining station. During the manufacture of the aircraft 10, other components can be joined or assembled at these or other stations, and the manufacturability analysis system 110 analyzes the manufacturability of various components (e.g., the main fuselage section, tail, nose, wing assembly, engine, etc.) to verify that the engineering design (e.g., the size, shape, location of the structural features of the mating components of the components) and the construction sequence (e.g., the construction sequence of putting the mating components together includes the moving steps to position the mating components relative to each other) meet the assembly and design tolerances of the components for joining the assembly.

[0042] In an embodiment, the manufacturability analysis system 110 determines the structural dimension requirements for various mating components and for the assembly tolerances between the mating components. The manufacturability analysis system 110 formulates the benchmark specifications for the mating components and formulates the index specifications for making the mating components fit. The benchmark specifications include the benchmark features for the mating components and the benchmark reference planes for the mating components, and the benchmark reference planes define the benchmark reference system. The benchmark reference system defines the coordinate system for the mating components to define the orientation of the mating components in space. The index specifications include the construction sequence for orienting the benchmark reference systems relative to each other. The manufacturability analysis system 110 performs a variation analysis based on the benchmark specifications and the index specifications to determine the assembly analysis results for the mating components, and compares the assembly analysis results for the mating components with the corresponding assembly tolerances defined in the structural dimension requirements for the mating components. The manufacturability analysis system 110 verifies that the benchmark specifications and the index specifications meet the assembly tolerances between the mating components.

[0043] The manufacturability analysis system 110 may include or represent hardware and related instructions (e.g., software stored on a tangible and non-transitory computer-readable storage medium such as a computer hard drive, ROM, RAM, etc.) that perform the operations described herein. The hardware may include electronic circuitry that includes and / or is connected to one or more logic-based devices such as microprocessors, processors, controllers, etc. These devices may be off-the-shelf devices that perform operations from the instructions described herein. Additionally or alternatively, one or more of these devices may be hardwired with logic circuitry to perform these operations.

[0044] In an embodiment, assembly system 100 includes positioning system 134 having a controller 136 for controlling the position of components at various stations within manufacturing facility 102. Positioning system 134 is communicatively coupled to manufacturability analysis system 110. Manufacturability analysis system 110 can determine a build sequence for the components, for example, based on a virtual joint simulation using a 3-D solid model, which enables graphical visualization and direct manipulation in a simulated manufacturing environment. The validated build sequence can be shared with positioning system 134 (e.g., uploaded to positioning system 134). Positioning system 134 uses the build sequence to control the positioning of components during manufacture of aircraft 10.

[0045] The assembly system 100 includes a motion system 140 having a component support tool 142 for supporting components and moving components within a station, for example, during a joining process. The component support tool 142 can be moved within a work area of ​​the manufacturing facility 102, such as between and within each station, for positioning components relative to each other. In various embodiments, the component support tool 142 can be computer-controlled and programmable. For example, the component support tool 142 can be operably coupled to a controller 136 of the positioning system 134. The controller 136 can control the movement and positioning of the component support tool 142 according to a defined build sequence. The component support tool 142 can move along a predetermined path. In various embodiments, in addition to or instead of the controller 136, the component support tool 142 can be driven and manipulated by an operator.

[0046] In an embodiment, the component support tool 142 may include a crane or other type of overhead support for supporting the component from an overhead position. In various other embodiments, the component support tool 142 may include a jack tower, a pogo stick support, or other type of support for supporting the component from below. The component support tool 142 may be supported by a crawler traction device that allows the component support tool 142 to move between various stations. In various other embodiments, the component support tool 142 may be supported by a bracket on a track to facilitate movement between various stations. In alternative embodiments, other types of component support tools 142 may be used to support the component and allow movement between various stations.

[0047] In an embodiment, the assembly system 100 includes a metrology system 150 having at least one tracking device 152 for locating a component in a work area of ​​the manufacturing facility 102. The controller 136 is communicatively coupled to the tracking device 152 and receives position data from the tracking device 152. In various embodiments, the tracking device 152 is a laser tracking device configured to determine the position of a component using one or more laser beams. The component may include a reflector, such as a retroreflector, for locating by the tracking device 152. In other various embodiments, the tracking device 152 may be an image tracking device, such as a camera configured to detect the position of a component based on an image acquired by the camera. Other types of tracking devices 152 may be used in alternative embodiments. In an embodiment, the controller 136 controls the relative positioning of the component support tool 142 in the work area of ​​the manufacturing facility 102 based on the position data acquired by the tracking device 152.

[0048] Figure 2 is a flow chart of a process for producing aircraft 10 according to an embodiment. Figure 2 Processes such as those performed by the manufacturability analysis system 110 are shown. The manufacturability analysis system 110 may perform various processes sequentially, and may repeat the processes multiple times, such as with one or more changes to input parameters to verify an operable build sequence and / or prioritize the placement of critical features of a component and / or analyze design and build constraints to verify that the component can be efficiently assembled or disassembled. The manufacturability analysis system 110 performs process steps to determine whether one or more components contribute to undesirable geometric changes that may result in negative slack (e.g., interference of components), require design modifications, require tool modifications, or require modifications to build metric specifications for joined components.

[0049] During the design process for the aircraft 10, the manufacturability analysis system 110 is used to perform a manufacturability analysis to demonstrate or determine the feasibility of one or more build sequences. The manufacturability analysis system 110 uses variation analysis to predict the effects of variations in the assembly process to verify that the build sequence can be used for production of the aircraft 10. In a first stage 200, a structural concept is created that defines the structural features of the components and the interrelationships between the components. In a second stage 202, a manufacturability analysis is performed to verify that the engineering design and build sequence are feasible. In a third stage 204, the structural concept is refined and / or the manufacturability analysis is repeated to provide continuous improvements to the engineering design and build sequence, thereby providing one or more recommendations.

[0050] In the first phase 200, a product concept is developed at step 210, and a model, such as a CAD model, of the initial concept is created at step 212. In an embodiment, developing the product concept includes defining the size of mating components of a component, defining the shape of mating components of a component, defining the location of structural features of mating components of a component, defining the relative location of mating components of a component, etc. After developing and modeling the product concept, the manufacturability analysis system 110 is used to verify that the mating components can be joined and manufactured.

[0051] In the second phase 202 , structural dimensioning requirements are developed at step 220 , baseline specifications are developed at step 222 , target specifications are developed at step 224 , variation analysis is performed at step 226 , and manufacturability recommendations are determined at step 228 .

[0052] In an embodiment, developing the structural dimension requirements (step 220) includes establishing working assumptions to be used in the variation analysis, such as the component is a rigid body, no deflection in the component, no deflection in the component support tool, no temperature variation, the material of the component, etc. Developing the structural dimension requirements (step 220) includes defining structural features of the mating components, such as size, surface shape, location of structural features (e.g., surface edges, holes, brackets, etc.), etc. Developing the structural dimension requirements (step 220) includes defining tolerances, such as manufacturing tolerances for components, assembly tolerances between components, etc. Developing the structural dimension requirements (step 220) includes defining failure conditions and consequences due to failure.

[0053] In an embodiment, developing a reference specification (step 222) includes identifying reference features for mating components. A reference feature is a physical feature of a mating component. A reference feature may be a surface, an edge, an opening, a protrusion, or other reference feature associated with a mating component. Each mating component includes at least one reference feature. The reference features are used to define a reference plane that passes through or along the reference specification of the mating component. The mating components have corresponding reference features that are aligned and positioned relative to each other during engagement of the mating components. The reference specification may be developed by defining a primary reference feature of a first mating component of a first aircraft component and a secondary reference feature of a first mating component of a second aircraft component. The reference specification may be developed by defining a reference reference system for each mating component. The reference reference system is defined by three mutually perpendicular planes that establish a coordinate system for the mating components. Tolerances of the positions of the mating components may be defined relative to the reference reference system. The reference specification may include three mutually perpendicular first aircraft component reference planes for the mating component of the first aircraft component, and three mutually perpendicular second aircraft component reference planes for the mating component of the second aircraft component.

[0054] In an embodiment, formulating the index specification (step 224) includes establishing a construction sequence for positioning the mating components relative to each other. The construction sequence is a step-by-step process of moving the mating components relative to each other. Optionally, one mating component can be in a fixed position, and the other mating component can be moved relative to the fixed mating component. In other embodiments, the two mating components can be moved sequentially or simultaneously. The construction sequence includes movement in three-dimensional space, for example, with six degrees of freedom. For example, the movable mating component can be moved by translating along any one of the mutually perpendicular X, Y and Z axes, and can be moved by changing the orientation between these axes by rotating with a roll angle, a pitch angle and a yaw angle; wherein the movement along the X axis defines the forward and backward movement; the movement along the Y axis defines the left and right movement; the movement along the Z axis defines the up and down movement; the rolling movement defines the rotation around the X axis; the pitch movement defines the rotation around the Y axis; and the yaw movement defines the rotation around the Z axis. The index specification may be established by establishing an assembly sequence to position primary datum features (eg, a datum reference system of a first mating component) and secondary datum features (eg, a datum reference system of a second mating component) relative to each other for engagement.

[0055] In an embodiment, the variation analysis (step 226) is based on the baseline specification and the index specification. The variation analysis is performed using a control system, such as a 3-D control system. The variation analysis virtually simulates the assembly of the mating components according to the build sequence to determine the assembly analysis results between the mating components. The variation analysis determines whether the build sequence is feasible (e.g., within the scope of manufacturing capabilities). The variation analysis determines whether one or more mating components contribute to undesirable geometric changes that may result in negative margins (e.g., interference from mating components). The variation analysis determines whether all baseline features can be located at appropriate locations within the assembly tolerances defined by the structural dimensional requirements.

[0056] In an embodiment, the producibility suggestion (step 228) includes providing a report of findings regarding the feasibility of manufacturing a component according to a particular build sequence. If the build sequence is approved, the producibility suggestion (step 228) may include providing an approval suggestion that the build sequence is feasible. If the build sequence is rejected, the producibility suggestion (step 228) may include providing a report indicating areas of concern or areas of failure. The producibility suggestion (step 228) may include providing a result report as a suggestion indicating the assembly tolerance for each datum feature of the component. The producibility suggestion (step 228) may include providing a remedy as a suggestion. The design team may use the producibility suggestion to evaluate the build sequence. For example, the design team may approve the build sequence and send the build sequence to the assembly system 100 to build the aircraft according to the approved build sequence. The design team may modify the structural dimension requirements of one or more mating components.

[0057] In the third stage 204, a feedback loop is provided for use throughout the life cycle of the design of the aircraft 10 and / or throughout the life cycle of the manufacture of the aircraft 10. The design may be changed or updated to continuously improve the design of the aircraft 10. In the third stage 204, the manufacturability analysis system 110 receives manufacturability data in step 230. The manufacturability analysis system 110 iterates the product design in step 232, for example, to modify or change the shape of the mating components, the location of one or more structural features, etc. In step 234, the manufacturability analysis system 110 re-evaluates the structural dimension requirements to correspond to the modification in the product design. If it is necessary to correspond to the modification in the product design, the manufacturability analysis system 110 updates the baseline specification in step 236. If it is necessary to correspond to the modification in the product design, the manufacturability analysis system 110 updates the target specification in step 238. The manufacturability analysis system 110 repeats the variation analysis in step 240 based on the modified baseline specification and the modified target specification. In step 242, the manufacturability analysis system 110 determines a modified manufacturability recommendation. Recommendations are sent to the design team to obtain additional feedback for continued improvement of the engineering design and construction sequence.

[0058] Figure 3 is a flow chart of a process for producing aircraft 10 according to an embodiment. Figure 3 Processes such as those performed by the manufacturability analysis system 110 are shown. The manufacturability analysis system 110 may perform various processes sequentially, and may repeat the processes multiple times, such as with one or more changes to input parameters to verify an operable build sequence and / or prioritize the placement of critical features of a component and / or analyze design and build constraints to verify that the component can be efficiently assembled or disassembled. The manufacturability analysis system 110 performs process steps to determine whether one or more components contribute to undesirable geometric changes that may result in negative slack (e.g., interference of components), require design modifications, require tool modifications, or require modifications to build metric specifications for joining components.

[0059] In an embodiment, at step 300, structural dimensional requirements for a mating component of a first aircraft component are determined. At step 302, structural dimensional requirements for a mating component of a second aircraft component are determined. At step 304, structural dimensional requirements for assembly tolerances between the mating component of the first aircraft component and the corresponding mating component of the second aircraft component are determined. In an embodiment, the structural dimensional requirements may be determined by the manufacturability analysis system 110. The structural dimensional requirements may define working assumptions to be used in the variation analysis, such as the component being a rigid body, no deflection in the component, no deflection in the component support tool, no temperature variation, the material of the component, etc. The structural dimensional requirements may relate to structural features of the mating component, such as size, surface shape, location of structural features (e.g., surface edge, hole, bracket, etc.), etc. The structural dimensional requirements may relate to tolerances associated with the mating component, such as manufacturing tolerances for the component, assembly tolerances between the components, etc. The structural dimensional requirements may define failure conditions of the mating component and the consequences resulting from the failure of the mating component.

[0060] In an embodiment, at step 306, a reference specification is developed for the mating component of the first aircraft component and the mating component of the second aircraft component. The reference specification may be based on structural dimension requirements. The reference specification may be developed by identifying reference features for the mating components. The reference features may be surfaces, edges, openings, protrusions, or other reference features associated with the mating components. Each mating component includes at least one reference feature. The reference features may be reference points or reference planes passing through or along the mating components. The mating component of the first aircraft component and the mating component of the second aircraft component have corresponding reference features that are aligned and positioned relative to each other during engagement of the mating components.

[0061] In an embodiment, at step 308, a specification is developed for mating a mating component of a first aircraft component with a mating component of a second aircraft component. The specification establishes a build sequence for positioning the mating components relative to each other. The build sequence is a step-by-step process of moving the mating components into position relative to each other. The build sequence includes movement in three-dimensional space, for example, with six degrees of freedom.

[0062] In an embodiment, at step 310, a variation analysis is performed based on the baseline specification and the target specification to determine assembly analysis results for the mating components. The variation analysis is performed using a control system, such as a 3D control system. The variation analysis virtually simulates the assembly of the mating components according to the build sequence. The variation analysis determines whether the build sequence is feasible.

[0063] In an embodiment, at step 312, the assembly analysis results for the mating components are compared with the corresponding assembly tolerances defined in the structural dimensional requirements for the mating components. The assembly analysis results are compared with the assembly tolerances to verify that the reference specifications and the index specifications meet the assembly tolerances between the mating components of the first aircraft component and the corresponding mating components of the second aircraft component. The assembly analysis results are compared with the assembly tolerances to verify that the assembly analysis results are within the manufacturing capabilities of the manufacturing facility used to produce the aircraft. The assembly analysis results are compared with the assembly tolerances to determine whether one or more mating components contribute to undesirable geometric changes that may result in negative margins (e.g., interference from mating components). The assembly analysis results are compared with the assembly tolerances to determine whether all reference features can be located at appropriate locations within the assembly tolerances defined by the structural dimensional requirements.

[0064] If the assembly analysis results are favorable (e.g., within assembly tolerances), the build sequence is approved at step 314, and the mating components may be assembled according to the build sequence at step 316. If the assembly analysis results are unfavorable (e.g., one or more mating surfaces are outside of assembly tolerances, or the specification cannot proceed because assembly requirements are not met, such as due to interference of mating components), the build sequence is rejected at step 318, and the analysis or design is modified.

[0065] In an embodiment, at step 320, the structural dimension requirement for the mating component of the first aircraft component is modified. At step 322, the structural dimension requirement for the mating component of the second aircraft component is modified. At step 324, the structural dimension requirement for the assembly tolerance between the mating components is modified. In an embodiment, the structural dimension requirement for the mating component of the first aircraft component, the structural dimension requirement for the mating component of the second aircraft component, and the structural dimension requirement for the assembly tolerance are all modified. In another embodiment, only one or two modifications are made to the structural dimension requirements, and no additional modifications are made to the structural dimension requirements.

[0066] In an embodiment, the baseline specification for the mating component is modified based on the modified structural dimension requirements for the mating component of the first aircraft component and / or the mating component of the second aircraft component at step 326. At step 328, the index specification for the mating component is modified based on the modified structural dimension requirements and / or the modified baseline specification.

[0067] In an embodiment, at step 330, a variation analysis is performed based on the modified baseline specification and / or the modified index specification to determine a modified assembly analysis result for the mating component. At step 332, the modified assembly analysis result for the mating component is compared with a corresponding modified assembly tolerance defined in the modified structural dimension requirement for the mating component. The modified assembly analysis result is compared with the modified assembly tolerance to verify that the modified baseline specification and the modified index specification meet the modified assembly tolerance between the mating component of the first aircraft component and the corresponding mating component of the second aircraft component. If the modified assembly analysis result is favorable (e.g., within the modified assembly tolerance), the modified build sequence is approved at step 334, and the mating components can be assembled according to the modified build sequence at step 316. If the modified assembly analysis result is unfavorable (e.g., one or more modified mating features are outside the modified assembly tolerance, or the modified index specification cannot proceed due to interference of the mating components), the modified build sequence is rejected at step 336, and the analysis is modified again.

[0068] Figure 4 is a process flow diagram of producing the aircraft 10 through a wing-to-body-joining assembly process according to an embodiment. Figure 4 A process such as that performed by the manufacturability analysis system 110 is shown. The manufacturability analysis system 110 performs process steps to determine if one or more mating components or the body 32 of the wing assembly 30 or fuselage 18 contributes to undesirable geometric variations that may result in negative margins (e.g., interference of components), require design modifications, require tooling modifications, or require modifications to build performance specifications for the joined components.

[0069] In an embodiment, at step 400, structural dimensional requirements for mating components of the wing assembly 30 are determined. At step 402, structural dimensional requirements for mating components of the fuselage 18 are determined. At step 404, structural dimensional requirements for assembly tolerances between mating components of the wing assembly 30 and corresponding mating components of the fuselage 18 are determined. In an embodiment, the structural dimensional requirements of the wing assembly 30 relate to a lower rear spar split epsilon fitting, a lower front spar split epsilon fitting, an over wing chord, a wing assembly keel beam, or other mating components of the wing assembly 30. The structural dimensional requirements of the fuselage 18 relate to an upper rear spar split epsilon fitting, an upper front spar split epsilon fitting, a fuselage longeron, a fuselage keel beam, or other mating components of the fuselage 18.

[0070] In an embodiment, at step 406, baseline specifications are developed for the mating components of the wing assembly 30 and the mating components of the fuselage 18. The baseline specifications may be based on structural dimension requirements. The baseline specifications may be developed by identifying baseline features for the mating components. The baseline features may be surfaces, edges, openings, protrusions, or other baseline features associated with the mating components.

[0071] In an embodiment, at step 408, a specification is developed for mating mating components of the wing assembly 30 with mating components of the fuselage 18. The specification establishes a build sequence for positioning the mating components relative to a reference frame for each assembly. The build sequence includes movement in three-dimensional space, for example, with six degrees of freedom. In an embodiment, the main body 32 of the fuselage 18 is fixed during the build sequence, and the wing assembly 30 moves relative to the main body 32 of the fuselage 18 according to the build sequence.

[0072] In an embodiment, at step 410, a variation analysis is performed based on the baseline specification and the target specification to determine assembly analysis results for the mating components. The variation analysis is performed using a control system, such as a 3D control system. The variation analysis virtually simulates the assembly of the mating components according to the build sequence. The variation analysis determines whether the build sequence is feasible.

[0073] In an embodiment, the assembly analysis results for the mating components are compared to corresponding assembly tolerances defined in the structural dimension requirements for the mating components at step 412. The assembly analysis results are compared to the assembly tolerances to verify that the baseline specification and the target specification meet the assembly tolerances between the mating component of the first aircraft component and the corresponding mating component of the second aircraft component. The assembly analysis results are compared to the assembly tolerances to verify that the assembly analysis results are within the manufacturing capabilities of the manufacturing facility used to produce the aircraft.

[0074] If the assembly analysis result is unfavorable (e.g., one or more baseline specifications are outside of assembly tolerances, or the target specifications cannot proceed due to interference with mating components), the build sequence is modified, such as by modifying one or more structural dimension requirements and / or the baseline specifications. If the assembly analysis result is favorable (e.g., within assembly tolerances), then at step 414, the build sequence is approved and the mating components can be assembled according to the build sequence. In an embodiment, at step 414, the wing assembly 30 is assembled with the fuselage 18 according to a wing-to-body-joined assembly sequence based on the target specifications and the baseline specifications.

[0075] Figure 5 is a cross-sectional view of a portion of aircraft 10 showing wing assembly 30 positioned within fuselage 18. Wing assembly 30 is coupled to fuselage 18 according to a build sequence verified by manufacturability analysis system 110. Figure 5A center wing section 34 and a left wing section 38 of the wing assembly 30 are shown positioned within the fuselage 18. Portions of the fuselage 18 are removed to show the wing assembly 30 positioned within the fuselage 18.

[0076] Figure 6 1 is a top view of a portion of the aircraft 10 showing the wing assembly 30 ready for assembly with the fuselage 18. The wing assembly 30 is coupled to the fuselage 18 according to a build sequence verified by the manufacturability analysis system 110. During manufacturing, many mating components of the wing assembly 30 and the fuselage 18 need to be aligned and positioned relative to each other for the wing-to-body-joining process. In an embodiment, the mating components include split epsilon fittings, over-wing chords, stringers, short frame tension straps, flap panel top surfaces, pressure deck edges, keel extensions, and other various surfaces and edges of the various mating components.

[0077] Figures 7 to 19 Various embodiments of 3-D solid models of an aircraft component and a mating component of the aircraft component are shown. Manufacturability analysis system 110 uses the 3-D solid model to virtually simulate the assembly of a first aircraft component and a second aircraft component when performing variation analysis to determine assembly analysis results between the corresponding mating components. Manufacturability analysis system 110 establishes datum features and datum planes as part of datum specifications. The datum specifications are used to perform variation analysis.

[0078] Figure 7 is a cross-sectional view of a 3-D solid model of the wing assembly 30 , illustrating a datum reference system for datum specifications of mating components of the wing assembly 30 . Figure 7 A primary datum plane 500, a secondary datum plane 502, and a tertiary datum plane 504 are shown. Datum planes 500, 502, 504 are mutually perpendicular wing assembly datum planes. Primary datum plane 500 is a horizontal datum plane extending from front to back and from side to side. Second datum plane 502 is a vertical datum plane extending from top to bottom and from side to side. Third datum plane 504 is a vertical datum plane extending from top to bottom and from front to back. Datum planes 500, 502, 504 are defined by datum features on corresponding mating components of wing assembly 30.

[0079] Figure 8 is a rear view of a 3-D solid model of the wing assembly 30 , showing reference specifications for the mating components of the wing assembly 30 . Fig. 9 is a front view of a 3-D solid model of the wing assembly 30 , showing reference specifications for the mating components of the wing assembly 30 . Figure 8 and Fig. 9A primary datum plane 500 and datum features 510, 512, 514 used to define the primary datum plane 500 are shown. The primary datum plane 500 extends from front to rear and from side to side within the wing assembly 30. In an embodiment, the datum features 510, 512, 514 may be defined by a front spar 550 and a rear spar 552. The datum features 510, 512, 514 may be defined by surfaces, edges, holes, or other features of the front spar 550 and the rear spar 552. The datum features 510, 512 may be defined by a hole 554 in the front spar 550. The datum feature 514 may be defined by a center hole 556 in the rear spar 552. In other embodiments, other datum features may be used to define the primary datum plane 500.

[0080] Fig.10 is a rear perspective view of a 3-D solid model of the wing assembly 30 , showing reference specifications for the mating components of the wing assembly 30 . Fig.11 is a rear view of a 3-D solid model of the wing assembly 30 , showing reference specifications for the mating components of the wing assembly 30 . Fig.10 and Fig.11 A second reference plane 502 and reference features 520, 522, 524 used to define the reference plane 502 are shown. The second reference plane 502 extends from top to bottom and from side to side within the wing assembly 30. In an embodiment, the reference features 520, 522, 524 may be defined by the rear spar 552. The reference features 520, 522, 524 may be defined by a center hole 558 in the rear spar 552 or other features of the rear spar 552. Fig.11 Datum features 520, 522, 524 are shown on the rear surface of the rear spar 552. In other embodiments, other datum features may be used to define the second datum plane 502.

[0081] Fig.12 is a rear perspective view of a 3-D solid model of the wing assembly 30 , showing reference specifications for the mating components of the wing assembly 30 . Fig.12 A third reference plane 504 and a reference feature 530 for defining the third reference plane 504 are shown. The third reference plane 504 extends from top to bottom and from front to back within the wing assembly 30. In an embodiment, the reference feature 530 may be defined by a hole in the rear spar 552. The reference feature 530 may be defined by a surface, an edge, a hole, or other feature of the rear spar 552. Fig.12 A datum feature 530 is shown at the center hole 558 of the rear spar 552, and the third datum plane 504 is perpendicular to the primary datum plane 500 and the second datum plane 502 at the datum feature 530. In other embodiments, other datum features may be used to define the third datum plane 504.

[0082] Fig.13 is a cross-sectional view of a 3-D solid model of the fuselage 18 showing a datum reference system for datum specifications of mating components of the fuselage 18 . Fig.13 A primary datum plane 600, a secondary datum plane 602, and a tertiary datum plane 604 are shown. Datum planes 600, 602, 604 are mutually perpendicular fuselage datum planes. Primary datum plane 600 is a horizontal datum plane extending from front to back and from side to side. Second datum plane 602 is a vertical datum plane extending from top to bottom and from side to side. Third datum plane 604 is a vertical datum plane extending from top to bottom and from front to back. Datum planes 600, 602, 604 are defined by datum features on corresponding mating components of fuselage 18.

[0083] Fig.14 is a cross-sectional view of the 3-D solid model of the fuselage 18 showing reference planes for mating components of the fuselage 18 . Fig.15 is a top view of a 3-D solid model of the fuselage 18 showing baseline specifications for the mating components of the fuselage 18 . Fig.14 and Fig.15 A primary datum plane 600 and datum features 610, 612, 614, 616 used to define the primary datum plane 600 are shown. The primary datum plane 600 extends from front to rear and from side to side within the fuselage 18. In an embodiment, the datum features 610, 612, 614, 616 may be defined by a respective upper split ε fitting 650. The datum features 610, 612, 614, 616 may be surfaces, edges, holes, or other features of the upper split ε fitting 650. The datum features 610, 612 may be defined by the upper front spar split ε fitting 650, and the datum features 614, 616 may be defined by the upper rear spar split ε fitting 650. In other embodiments, other datum features may be used to define the primary datum plane 600.

[0084] Fig.16 is a cross-sectional view of the 3-D solid model of the fuselage 18 showing reference planes for mating components of the fuselage 18 . Fig.17 is a front view of a 3-D solid model of the fuselage 18 showing datum features for mating components of the fuselage 18 . Fig.16 and Fig.17 A second reference plane 602 and reference features 620, 622, 624 for defining the second reference plane 602 are shown. The second reference plane 602 extends from top to bottom and from side to side within the fuselage 18. In an embodiment, the reference features 620, 622, 624 may be defined by a corresponding upper split ε fitting 650, such as an upper rear spar split ε fitting. The reference features 620, 622, 624 may be defined by a surface, edge, hole, or other feature of the upper rear spar split ε fitting. Fig.17Datum features 620, 622, 624 are shown on the front surface 652 of the two upper rear spar split epsilon fittings. In other embodiments, other datum features may be used to define the second datum plane 602.

[0085] Fig.18 is a cross-sectional view of the 3-D solid model of the fuselage 18 showing reference planes for mating components of the fuselage 18 . Fig.19 is a top view of a 3-D solid model of the fuselage 18 showing baseline specifications for the mating components of the fuselage 18 . Fig.18 and Fig.19 A third reference plane 604 and reference features 630, 632 for defining the third reference plane 604 are shown. The third reference plane 604 extends from top to bottom and from front to rear within the fuselage 18. In an embodiment, the reference features 630, 632 may be defined by a corresponding upper split ε fitting 650, such as an upper rear spar split ε fitting. The reference features 630, 632 may be defined by a surface, edge, hole, or other feature of the upper rear spar split ε fitting. Fig.19 Datum features 630, 632 are shown on the inboard surface 654 of the upper rear spar split epsilon fitting 650 on the right hand and left hand sides. In other embodiments, other datum features may be used to define the third datum plane 604.

[0086] Figures 20 to 28 Various embodiments of 3-D solid models of aircraft parts and mating parts of the aircraft parts are shown. Manufacturability analysis system 110 uses the 3-D solid models to virtually simulate the assembly sequence of the first aircraft part and the second aircraft part when performing variation analysis to determine assembly analysis results between the corresponding mating parts. Figures 20 to 28 Aircraft components are shown as wing assemblies 30 and fuselage 18; however, as described above, variation analysis may be performed on other aircraft components during manufacture of the aircraft 10. The manufacturability analysis system 110 develops an index specification for mating a mating component of a first aircraft component and a mating component of a second aircraft component. The index specification is defined by a plurality of index moves. In an embodiment, the index moves are controlled by constraining the first aircraft component and the second aircraft component relative to each other within six degrees of freedom (X, Y, Z, pitch, roll, yaw) so as to meet functional requirements. The index specification defines a build sequence used during manufacture and assembly of the aircraft 10. The index specification is used to perform variation analysis.

[0087] Fig. 20 A cross-sectional view of a 3-D solid model of a portion of the aircraft 10 showing the wing assembly 30 positioned relative to the fuselage 18 . Fig.211 is a front perspective view of a portion of an aircraft 10 showing the wing assembly 30 positioned relative to the fuselage 18. A first index movement of an index specification is used to initially position the wing assembly 30 relative to the fuselage 18. In an embodiment, the first index movement of the index specification aligns the front surfaces 700, 702 of the upper rear spar split ε fitting 704 and the lower rear spar split ε fitting 706 on the right hand side of the aircraft 10. Fig.21 The front surfaces 700, 702 of the upper rear spar split epsilon fitting 704 and the lower rear spar split epsilon fitting 706 are shown aligned relative to each other. In an embodiment, the first index movement of the index specification sets the position of the wing assembly 30 relative to the fuselage 18 on the right hand side of the aircraft 10 in the X direction (e.g., from front to back). In an embodiment, the wing assembly 30 undergoes one or more translational movements of the corresponding mating components in the X direction and / or one or more translational movements of the corresponding mating components in the Y direction and / or one or more translational movements of the corresponding mating components in the Z direction to position the wing assembly 30 relative to the fuselage 18. For example, the wing assembly 30 may be moved to simulate movement from one station within a manufacturing facility to another station within the manufacturing facility by the component support tool 142.

[0088] Fig. 22 1 is a front view of a portion of the aircraft 10 showing the wing assembly 30 positioned relative to the fuselage 18. The second index movement of the index specification is used to further align the wing assembly 30 relative to the fuselage 18. In an embodiment, the second index movement of the index specification aligns the front surfaces 700, 702 of the upper rear spar split ε fitting 704 and the lower rear spar split ε fitting 706 on the left hand side of the aircraft 10. The second index movement includes a rotational movement of the wing assembly 30. In an embodiment, the second index movement of the index specification is used to set the yaw angle of the wing assembly 30. For example, the wing assembly 30 can be rotated about the Z axis to align the front surfaces 700, 702 of the upper rear spar split ε fitting 704 and the lower rear spar split ε fitting 706 on the left and right hand sides.

[0089] The third index movement of the index specification is used to further align the wing assembly 30 relative to the fuselage 18. In an embodiment, the third index movement of the index specification aligns the upper rear spar split ε fitting and the lower rear spar split ε fitting. The third index movement involves a rotational movement of the wing assembly 30. In an embodiment, the third index movement of the index specification is used to set the roll angle of the wing assembly 30. For example, the wing assembly 30 can be rotated about the X-axis to align the front surfaces 700, 702 of the upper rear spar split ε fitting 704 and the lower rear spar split ε fitting 706 on the left hand side.

[0090] Fig.23 is a front view of a portion of the aircraft 10 showing the wing assembly 30 positioned relative to the fuselage 18 . Fig.24 is a front perspective view of a portion of the aircraft 10 showing the wing assembly 30 positioned relative to the fuselage 18. A fourth index movement of the index specification is used to further align the wing assembly 30 relative to the fuselage 18. The fourth index movement involves a translational movement of the wing assembly 30 in the Y direction. In an embodiment, the fourth index movement centers the skin of the fuselage 18 between the left and right hand sides of the wing assembly 30 past the wing chord 710 so that the gap 712 between the fuselage skin 714 and the past the wing chord 710 on both sides is equal. In an embodiment, the fourth index movement of the index specification sets the position of the wing assembly 30 relative to the fuselage 18 in the Y direction (e.g., from side to side).

[0091] Fig.25 is a cross-sectional view of a portion of aircraft 10 showing wing assembly 30 positioned relative to fuselage 18 . Fig.26 1 is a cross-sectional view of a portion of an aircraft 10 showing the wing assembly 30 positioned relative to the fuselage 18. The fifth index movement of the index specification is used to further align the wing assembly 30 relative to the fuselage 18. In an embodiment, the fifth index movement of the index specification mates the upper rear spar split ε fitting 704 and the lower rear spar split ε fitting 706. The top surface 716 of the lower rear spar split ε fitting 706 mates with the bottom surface 718 of the upper rear spar split ε fitting 704. The fifth index movement involves a translational movement of the wing assembly 30 in the Z direction. In an embodiment, the fifth index movement of the index specification is used to set the position of the wing assembly 30 in the Z direction (e.g., from top to bottom) relative to the fuselage 18.

[0092] Fig. 27 A portion of the aircraft 10 is shown, illustrating the wing assembly 30 positioned relative to the fuselage 18 . Fig.28 1 is a perspective view of a portion of the aircraft 10 showing the wing assembly 30 positioned relative to the fuselage 18. The sixth index movement of the index specification is used to further align the wing assembly 30 relative to the fuselage 18. In an embodiment, the sixth index movement of the index specification aligns the top surface 720 of the front end 722 of the seat track 724 on the wing assembly 30 with the top surface 730 of the rear end 732 of the seat track 734 on the fuselage 18. The sixth index movement involves a rotational movement of the wing assembly 30. In an embodiment, the sixth index movement of the index specification is used to set the pitch angle of the wing assembly 30. For example, the wing assembly 30 can be rotated about the Y axis to align the seat tracks 724, 734.

[0093] In an embodiment, the sixth index move may affect the Z positioning of the upper rear spar split ε fitting and the lower rear spar split ε fitting. After the sixth index move of the index specification, the fifth index move of the index specification may be repeated to readjust the Z positioning of the upper rear spar split ε fitting 704 and the lower rear spar split ε fitting 706. The sixth index move may be repeated. The fifth and sixth index moves may be repeated multiple times to achieve an optimal fit between a set pitch angle and a set Z positioning of the components. In an embodiment, the index specification ends with a Z positioning move, such as the fifth index move, because the assembly tolerance between the rear spar split ε fittings 704, 706 is tighter than the assembly tolerance of the seat rails 724, 734.

[0094] In other embodiments, the index moves may be performed in a different order. Additional index moves may be utilized when aligning different datum features of the wing assembly 30 and fuselage 18. Different index moves may be utilized when performing variation analysis on different aircraft components.

[0095] The target movement defines a build sequence for assembly of the aircraft components. A variation analysis is performed to virtually simulate assembly of the aircraft components to determine that the assembly analysis results meet assembly tolerances. The variation analysis is used to verify that the build sequence is within the manufacturing capabilities of a manufacturing facility used to produce the aircraft 10. If one or more of the assembly analysis results do not meet specifications, the structural dimensional requirements may be modified, and a different build sequence may be developed by performing a variation analysis based on the modified structural dimensional requirements using the modified baseline specifications and the target specifications.

[0096] Furthermore, the present disclosure includes embodiments according to the following clauses:

[0097] 1. A method for producing an aircraft, the method comprising:

[0098] determining structural dimensional requirements for a mating component of a first aircraft component, for a mating component of a second aircraft component, and for an assembly tolerance between the mating component of the first aircraft component and the mating component of the second aircraft component;

[0099] developing a baseline specification for the mating component for the first aircraft component and for the mating component for the second aircraft component;

[0100] developing specifications for mating the mating component of the first aircraft component with the mating component of the second aircraft component;

[0101] performing a variation analysis based on the baseline specification and the target specification to determine assembly analysis results for the mating component of the first aircraft component and the mating component of the second aircraft component; and

[0102] The assembly analysis results for the mating component of the first aircraft component and the mating component of the second aircraft component are compared with the assembly tolerances determined by the structural dimensional requirements for the mating component of the first aircraft component and the mating component of the second aircraft component to verify that the baseline specification and the target specification meet the assembly tolerances between the mating component of the first aircraft component and the mating component of the second aircraft component.

[0103] 2. The method of clause 1, wherein the performing variation analysis comprises using a 3-D control system to virtually simulate assembly of the first aircraft component and the second aircraft component to determine the assembly analysis results for the mating component of the first aircraft component and the mating component of the second aircraft component.

[0104] 3. The method of clause 1, further comprising verifying that the assembly analysis results are within the manufacturing capabilities of a manufacturing facility used to produce the aircraft.

[0105] 4. The method according to clause 1, further comprising modifying at least one of: a) the structural dimensional requirement of the mating component of the first aircraft component; b) the structural dimensional requirement of the mating component of the second aircraft component; and c) the structural dimensional requirement for the assembly tolerance.

[0106] 5. The method according to clause 4, further comprising:

[0107] modifying the baseline specification based on the modified structural dimensional requirements; and

[0108] The index specification is modified based on the modified structural size requirement and the modified baseline specification.

[0109] 6. The method according to clause 5, further comprising:

[0110] repeating the variation analysis based on the modified baseline specification and the modified target specification to determine a modified assembly analysis result; and

[0111] The modified assembly analysis results are verified to be within manufacturing capabilities of a manufacturing facility used to produce the aircraft.

[0112] 7. The method of any of clauses 1 to 6, further comprising assembling the first aircraft component and the second aircraft component according to an assembly sequence defined by the target specification and the reference specification.

[0113] 8. A method according to any one of clauses 1 to 6, wherein the establishing the datum specification includes defining primary datum features of a first mating component of the first aircraft component and defining secondary datum features of a first mating component of the second aircraft component, and wherein the establishing the index specification includes establishing an assembly sequence to position the primary datum features and the secondary datum features relative to each other for engagement.

[0114] 9. A method according to any one of clauses 1 to 6, wherein the establishing of the reference specification comprises defining three mutually perpendicular first aircraft component reference planes for the mating component of the first aircraft component, and defining three mutually perpendicular second aircraft component reference planes for the mating component of the second aircraft component.

[0115] 10. A method according to any one of clauses 1 to 6, wherein the establishing of the datum specification includes defining a primary datum feature at a surface of a first mating component of the first aircraft component and defining a secondary datum feature at a surface of a first mating component of the second aircraft component, and wherein the establishing of the datum specification includes establishing an assembly sequence to position the primary datum feature and the secondary datum feature relative to each other for engagement.

[0116] 11. A method according to any one of clauses 1 to 6, wherein the establishing of the baseline specification comprises establishing the baseline specification for a lower rear wing spar split ε fitting, a lower front wing spar split ε fitting, an upper rear wing spar split ε fitting and an upper front wing spar split ε fitting, wherein the lower rear wing spar split ε fitting and the lower front wing spar split ε fitting are mating parts of the first aircraft component, and wherein the upper rear wing spar split ε fitting and the upper front wing spar split ε fitting are mating parts of the second aircraft component.

[0117] 12. A method according to any one of clauses 1 to 6, wherein the establishing of the index specification includes defining an assembly sequence for the first aircraft component and the second aircraft component, and the assembly sequence includes: a) one or more translational movements of the mating component of at least one of the first aircraft component or the second aircraft component in the X direction; b) one or more translational movements of the mating component of at least one of the first aircraft component or the second aircraft component in the Y direction; c) one or more translational movements of the mating component of at least one of the first aircraft component or the second aircraft component in the Z direction; d) one or more rotational movements of the mating component of at least one of the first aircraft component or the second aircraft component at a pitch angle; e) one or more rotational movements of the mating component of at least one of the first aircraft component or the second aircraft component at a roll angle; and f) one or more rotational movements of the mating component of at least one of the first aircraft component or the second aircraft component at a yaw angle.

[0118] 13. A method according to any one of clauses 1 to 6, wherein the first aircraft component is a wing assembly including a center wing portion, a right wing portion and a left wing portion, the mating components of the wing assembly include a lower rear wing spar split ε fitting, a lower front wing spar split ε fitting, a wing chord and a wing assembly keel beam, and wherein the second aircraft component is a fuselage including a middle body portion configured to receive the wing assembly, the mating components of the fuselage include an upper rear wing spar split ε fitting, an upper front wing spar split ε fitting, a fuselage longitudinal beam and a fuselage keel beam, and the performing variation analysis includes: a) determining the assembly analysis result between the lower rear wing spar split ε fitting and the upper rear wing spar split ε fitting; b) determining the assembly analysis result between the lower front wing spar split ε fitting and the upper front wing spar split ε fitting; c) determining the assembly analysis result between the wing chord and the fuselage longitudinal beam; and d) determining the assembly analysis result between the wing assembly keel beam and the fuselage keel beam.

[0119] 14. A method for producing an aircraft, the method comprising:

[0120] determining structural dimensional requirements for a mating component of a first aircraft component, for a mating component of a second aircraft component, and for an assembly tolerance between the mating component of the first aircraft component and the mating component of the second aircraft component;

[0121] developing a baseline specification for the mating component for the first aircraft component and for the mating component for the second aircraft component;

[0122] developing a specification defining an assembly sequence for mating the mating component of the first aircraft component with the mating component of the second aircraft component;

[0123] performing a variation analysis based on the baseline specification and the target specification to determine assembly analysis results for the mating component of the first aircraft component and the mating component of the second aircraft component;

[0124] comparing the assembly analysis results for the mating component of the first aircraft component and the mating component of the second aircraft component with the assembly tolerances determined by the structural dimensional requirements for the mating component of the first aircraft component and the mating component of the second aircraft component to verify that the baseline specification and the target specification satisfy the assembly tolerances between the mating component of the first aircraft component and the mating component of the second aircraft component; and

[0125] The first aircraft component and the second aircraft component are assembled according to the assembly sequence based on the index specification and the reference specification.

[0126] 15. A method for producing an aircraft by wing-to-body-joint assembly, the method comprising:

[0127] determining structural dimensional requirements of wing assembly mating components for a wing assembly, the wing assembly comprising a center wing section, a right wing section attached to the center wing section, and a left wing section attached to the center wing section;

[0128] determining structural dimensional requirements for fuselage mating components of a fuselage including a mid-body section configured to receive the wing assembly;

[0129] Determining structural dimension requirements for assembly tolerances between the wing assembly mating components and the fuselage mating components;

[0130] developing baseline specifications for the wing assembly mating components and for the fuselage mating components;

[0131] Establishing specifications for mating the wing assembly mating components with the fuselage mating components;

[0132] performing a variation analysis based on the baseline specification and the target specification to determine assembly analysis results for the wing assembly mating component and the fuselage mating component; and

[0133] The assembly analysis results for the wing assembly mating component and the fuselage mating component are compared with the assembly tolerances determined by the structural dimensional requirements for the wing assembly mating component and the fuselage mating component to verify that the baseline specification and the index specification meet the assembly tolerances between the wing assembly mating component and the fuselage mating component.

[0134] 16. The method of clause 15, wherein performing the variation analysis comprises using a 3-D control system to virtually simulate assembly of the wing assembly and the fuselage to determine assembly analysis results between mating components of the wing assembly and mating components of the fuselage.

[0135] 17. The method of clause 15, further comprising verifying that the assembly analysis results are within the manufacturing capabilities of a manufacturing facility used to produce the aircraft.

[0136] 18. The method of clause 15, further comprising modifying at least one of: a) the structural dimensional requirement for the wing assembly mating component; b) the structural dimensional requirement for the fuselage mating component; and c) the structural dimensional requirement for the assembly tolerance.

[0137] 19. The method according to clause 18, further comprising:

[0138] modifying the baseline specification based on the modified structural dimensional requirements; and

[0139] The index specification is modified based on the modified structural size requirement and the modified baseline specification.

[0140] 20. The method according to clause 19, further comprising:

[0141] repeating the variation analysis based on the modified baseline specification and the modified target specification to determine a modified assembly analysis result; and

[0142] The modified assembly analysis results are verified to be within manufacturing capabilities of a manufacturing facility used to produce the aircraft.

[0143] 21. A method according to any of clauses 15 to 20, further comprising assembling the wing assembly with the fuselage according to an assembly sequence defined by the target specification and the reference specification.

[0144] 22. A method according to any one of clauses 15 to 20, wherein the establishing of the benchmark specifications comprises defining major benchmark features of a first mating component of the wing assembly and defining major benchmark features of a first mating component of the fuselage based on the assembly tolerances, and wherein the establishing of the benchmark specifications comprises establishing an assembly sequence to position the major benchmark features of the first mating components relative to each other for engagement.

[0145] 23. A method according to any one of clauses 15 to 20, wherein the establishing of the reference specifications comprises defining three mutually perpendicular wing assembly reference planes for the wing assembly mating components and defining three mutually perpendicular fuselage reference planes corresponding to the wing assembly reference planes for the fuselage mating components.

[0146] 24. A method according to any one of clauses 15 to 20, wherein the establishing the benchmark specification includes defining a first benchmark feature at an upper surface of a first wing assembly mating component and at a rear edge of the upper surface, and defining a second benchmark feature at a lower surface of a first fuselage mating component and at a rear edge of the lower surface, and wherein the establishing the benchmark specification includes establishing an assembly sequence to position the first benchmark feature and the second benchmark feature relative to each other for engagement.

[0147] 25. A method according to any one of clauses 15 to 20, wherein the establishing of the baseline specification comprises establishing the baseline specification for a lower rear spar split ε fitting, a lower front spar split ε fitting, an upper rear spar split ε fitting and an upper front spar split ε fitting, wherein the lower rear spar split ε fitting and the lower front spar split ε fitting are wing assembly mating parts, and wherein the upper rear spar split ε fitting and the upper front spar split ε fitting are fuselage mating parts.

[0148] 26. A method according to any one of clauses 15 to 20, wherein the establishing of the specification comprises defining an assembly sequence for the wing assembly and the fuselage, the assembly sequence comprising: a) one or more translational movements of the wing assembly mating components in the X direction; b) one or more translational movements of the wing assembly mating components in the Y direction; c) one or more translational movements of the wing assembly mating components in the Z direction; d) one or more rotational movements of the wing assembly mating components at a pitch angle; e) one or more rotational movements of the wing assembly mating components at a roll angle; and f) one or more rotational movements of the wing assembly mating components at a yaw angle.

[0149] 27. A method according to any one of clauses 15 to 20, wherein the wing assembly mating parts include a lower rear wing spar split ε fitting, a lower front wing spar split ε fitting, a wing chord-crossing and a wing assembly keel beam, and wherein the fuselage mating parts include an upper rear wing spar split ε fitting, an upper front wing spar split ε fitting, a fuselage longitudinal beam and a fuselage keel beam, and wherein performing variation analysis includes: a) determining the assembly analysis result between the lower rear wing spar split ε fitting and the upper rear wing spar split ε fitting; b) determining the assembly analysis result between the lower front wing spar split ε fitting and the upper front wing spar split ε fitting; c) determining the assembly analysis result between the wing chord-crossing and the fuselage longitudinal beam; and d) determining the assembly analysis result between the wing assembly keel beam and the fuselage keel beam.

[0150] It should be understood that the above description is illustrative, not restrictive. For example, the above embodiments (and / or aspects thereof) can be used in combination with each other. In addition, without departing from the scope of the invention, many modifications can be made to adapt specific situations or materials to the teachings of the present disclosure. The sizes, material types, orientations, and the number and position of the various components described herein are intended to define the parameters of certain embodiments, and are by no means restrictive, but are merely embodiments. After reading the above description, many other embodiments and modifications within the spirit and scope of the claims will be apparent to those skilled in the art. Therefore, the scope of the present disclosure should be determined with reference to the attached claims together with the full range of equivalents to which these claims are entitled. In the attached claims, the terms "including" and "wherein" are used as the simple English equivalents of the corresponding terms "including" and "wherein". In addition, in the attached claims, the terms "first", "second", and "third", etc. are used only as marks, and are not intended to impose numerical requirements on their objects. Furthermore, the appended claim limitations are not written in means-plus-function format and are not intended to be interpreted under 35 USC §112(f) unless and until such claim limitations expressly use the phrase "means to" followed by a recitation of a function without further structure.

Claims

1. A method for producing an aircraft (10), the method comprising: Determining (300, 302, 304) structural dimension requirements (220) for a mating component of a first aircraft component (30), for a mating component of a second aircraft component (18), and for an assembly tolerance between the mating component of the first aircraft component and the mating component of the second aircraft component; developing (306) baseline specifications for the mating component for the first aircraft component and for the mating component for the second aircraft component; developing (308) a virtual index specification, the virtual index specification including a step-by-step process for moving the mating components into position relative to each other, the virtual index specification defining a virtual assembly sequence for simulating mating of the mating components of the first aircraft component with the mating components of the second aircraft component; performing ( 310 ) a virtual variation analysis based on the baseline specification and the virtual index specification to virtually simulate and determine assembly analysis results for the mating component of the first aircraft component and the mating component of the second aircraft component; as well as The assembly analysis results for the mating component of the first aircraft component and the mating component of the second aircraft component are compared (312) to the assembly tolerance determined by the structural dimensional requirements for the mating component of the first aircraft component and the mating component of the second aircraft component to verify that the baseline specification and the target specification meet the assembly tolerance between the mating component of the first aircraft component and the mating component of the second aircraft component.

2. The method according to claim 1, wherein: The performing (310) variation analysis includes using a 3-D control system to virtually simulate assembly of the first aircraft component (30) and the second aircraft component (18) to determine assembly analysis results for the mating components of the first aircraft component and the mating components of the second aircraft component.

3. The method according to claim 1, further comprising: The assembly analysis results are verified (318) to be within manufacturing capabilities of a manufacturing facility (102) for producing the aircraft (10).

4. The method according to claim 1, further comprising: Modifying (320, 322, 324) at least one of: a) the structural dimensional requirement for the mating component of the first aircraft component (30); b) the structural dimensional requirement for the mating component of the second aircraft component (18); and c) the structural dimensional requirement for the assembly tolerance.

5. The method according to claim 4, further comprising: modifying (326) the baseline specification based on the modified structural dimension requirements; as well as The index specification is modified (328) based on the modified structural dimension requirement and the modified baseline specification.

6. The method according to claim 5, further comprising: repeating ( 330 ) the variation analysis based on the modified baseline specification and the modified index specification to determine a modified assembly analysis result; as well as The modified assembly analysis results are verified (332) to be within manufacturing capabilities of a manufacturing facility (102) for producing the aircraft (10).

7. The method according to any one of claims 1 to 6, further comprising: The first aircraft component and the second aircraft component are assembled ( 316 ) according to an assembly sequence defined by the target specification and the baseline specification.

8. The method according to any one of claims 1 to 6, wherein: The developing (306) datum specifications include defining primary datum features of a first mating component of the first aircraft component (30) and defining secondary datum features of a first mating component of the second aircraft component (18), and wherein the developing (308) datum specifications includes developing an assembly sequence to position the primary datum features and the secondary datum features relative to each other for engagement.

9. The method according to any one of claims 1 to 6, wherein: The developing (306) datum specifications includes defining three mutually perpendicular first aircraft component datum planes (500, 502, 504) for the mating component of the first aircraft component (30) and defining three mutually perpendicular second aircraft component datum planes (600, 602, 604) for the mating component of the second aircraft component (18).

10. The method according to any one of claims 1 to 6, wherein: The developing (306) datum specifications includes defining a primary datum feature at a surface of a first mating component of the first aircraft component (30) and defining a secondary datum feature at a surface of a first mating component of the second aircraft component (18), and wherein the developing (308) datum specifications includes developing an assembly sequence to position the primary datum feature and the secondary datum feature relative to each other for engagement.

11. The method according to any one of claims 1 to 6, wherein: The developing (306) of the benchmark specification includes developing benchmark specifications for a lower rear wing spar split ε fitting, a lower front wing spar split ε fitting, an upper rear wing spar split ε fitting, and an upper front wing spar split ε fitting, wherein the lower rear wing spar split ε fitting and the lower front wing spar split ε fitting are mating parts of the first aircraft component (30), and wherein the upper rear wing spar split ε fitting and the upper front wing spar split ε fitting are mating parts of the second aircraft component (18).

12. The method according to any one of claims 1 to 6, wherein: The developing (308) specification includes defining an assembly sequence for the first aircraft component (30) and the second aircraft component (18), the assembly sequence including: a) one or more translational movements of the mating component of at least one of the first aircraft component or the second aircraft component in an X direction; b) one or more translational movements of the mating component of at least one of the first aircraft component or the second aircraft component in a Y direction; c) one or more translational movements of the mating component of at least one of the first aircraft component or the second aircraft component in a Z direction; d) one or more rotational movements of the mating component of at least one of the first aircraft component or the second aircraft component at a pitch angle; e) one or more rotational movements of the mating component of at least one of the first aircraft component or the second aircraft component at a roll angle; and f) one or more rotational movements of the mating component of at least one of the first aircraft component or the second aircraft component at a yaw angle.

13. The method according to any one of claims 1 to 6, wherein: The first aircraft component (30) is a wing assembly (30) including a center wing section (34), a right wing section (36) and a left wing section (38), the mating components of the wing assembly including a lower rear spar split ε fitting, a lower front spar split ε fitting, a wing chord and a wing assembly keel beam, and wherein the second aircraft component (18) is a fuselage (18) including a middle body section configured to receive the wing assembly, the mating components of the fuselage including an upper rear spar split ε fitting, a lower front spar split ε fitting, a wing chord and a wing assembly keel beam. The embodiment of the present invention relates to a split ε accessory, an upper front wing spar split ε accessory, a fuselage longitudinal beam and a fuselage keel beam, wherein the performing variation analysis comprises: a) determining the assembly analysis result between the lower rear wing spar split ε accessory and the upper rear wing spar split ε accessory; b) determining the assembly analysis result between the lower front wing spar split ε accessory and the upper front wing spar split ε accessory; c) determining the assembly analysis result between the wing chord and the fuselage longitudinal beam; and d) determining the assembly analysis result between the wing assembly keel beam and the fuselage keel beam.

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