Composite Splice Joints for Fuselage Thermal Expansion
Find Innovative SolutionsGenerate Solutions
Solution Overview
Problem
Existing methods for joining composite fuselage sections in large aircraft are limited, particularly in using metal splice plates which can be cumbersome and prone to thermal expansion issues, and lack efficient methods for maintaining structural integrity and aerodynamics.
Innovation Solution
A method and structure for joining composite fuselage sections using a strap and fittings that extend across the splice joint, with the strap and fittings made from composite materials to match thermal expansion coefficients, providing a redundant load path and ensuring aerodynamic continuity.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Reliability
If aluminum splice plates are used to join composite fuselage sections, then the structural strength and reliability of the joint is improved, but the device complexity and manufacturing difficulty increase due to the need for special automated splice machines and laser alignment systems
Solution Approach 1:
The patent applies homogeneity by making the fitting and stiffener components identical in material composition (composite materials matching the fuselage sections). This eliminates the need for special alignment machinery because the identical materials expand and contract uniformly with temperature changes, naturally maintaining alignment without complex automated positioning systems.
Solution Approach 2:
The patent changes the material parameter from dissimilar (aluminum splice plates) to similar (composite materials matching the fuselage). This parameter change resolves the thermal expansion mismatch problem and eliminates the need for complex laser alignment systems and automated splice machines, while maintaining joint reliability through thermally compatible materials.
2Weight of moving object
If composite materials are used for fuselage sections, then the strength-to-weight ratio is improved, but the ease of manufacture decreases due to difficulties in joining sections together
Solution Approach 1:
The patent uses composite materials for both the fuselage sections and the joining components (fittings and stiffeners). This allows all components to be manufactured using the same composite manufacturing processes, eliminating the difficulty of joining dissimilar materials and maintaining the weight advantages of composite construction throughout the entire assembly.
Solution Approach 2:
By making the fittings and stiffeners from the same composite materials as the fuselage sections, the patent creates a homogeneous structure that can be manufactured and joined using consistent processes. This homogeneity dramatically improves ease of manufacture compared to using aluminum splice plates with composite sections.
3Reliability
If aluminum splice plates are used for joining, then the structural integrity is improved, but the aerodynamic surface quality worsens due to the complexity of maintaining smooth surfaces at splice joints
Solution Approach 1:
The patent applies homogeneity by using identical composite materials for the fittings, stiffeners, and fuselage sections. This ensures uniform thermal expansion characteristics, preventing differential movement that would disrupt the aerodynamic surface. The homogeneous structure maintains smooth aerodynamic surfaces while preserving structural integrity.
Solution Approach 2:
The patent changes the material parameter from aluminum to composite materials, matching the fuselage sections. This parameter change eliminates thermal expansion mismatch, allowing the aerodynamic surface to remain smooth and continuous across splice joints while maintaining structural integrity through thermally compatible materials.
4Ease of manufacture
If different materials are used for fuselage sections and splice plates, then the ease of manufacture is improved through standardized aluminum components, but the reliability decreases due to thermal expansion mismatches
Solution Approach 1:
The patent applies homogeneity by using the same composite materials for both fuselage sections and joining components. This eliminates thermal expansion mismatches between dissimilar materials, ensuring reliable joints under thermal conditions while maintaining ease of manufacture through standardized composite manufacturing processes.
Solution Approach 2:
The patent changes the material parameter from aluminum splice plates to composite materials matching the fuselage sections. This parameter change resolves the thermal expansion compatibility issue, improving joint reliability under thermal conditions while maintaining manufacturing efficiency through consistent material usage.
Applied Scientific Principles
This section explains which scientific principles are used to turn an abstract innovation direction into a practical engineering solution.
Function Achieved in This Case
This approach enhances the structural integrity and aerodynamic smoothness of the fuselage by using composite materials that match thermal expansion, offering a fail-safe load path and simplifying installation while maintaining design flexibility.
Implementation Method 1
strap and fittings made from composite materials to match thermal expansion coefficients
Data Source
AI summary
Structures and methods for joining composite fuselage sections and other panel assemblies together are disclosed herein. In one embodiment, a shell structure configured in accordance with the present invention includes a first panel portion positioned adjacent to a second panel portion. The first panel portion can include a first stiffener attached to a first composite skin, and the second panel portion can include a second stiffener attached to a second composite skin. The shell structure can further include a fitting extending across a first edge region of the first panel portion and a second edge region of the second panel portion. A first end portion of the fitting can be attached to the first stiffener and the first composite skin, and a second end portion of the fitting can be attached to a second stiffener and a second composite skin, to join the first panel portion to the second panel portion.