Composite Fuselage Panel Joining via Injection Bonded Tabs
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Solution Overview
Problem
The labor-intensive and time-consuming process of joining composite fuselage sections in large aircraft using traditional splice plate methods requires extensive tooling and a high number of fasteners, leading to increased assembly time and complexity, especially in widebody aircraft where the inner diameter is large and the number of fasteners exceeds 5,000, demanding over 2,000 man-hours for assembly.
Innovation Solution
The method involves using composite structural panels with longitudinally extending tabs on one panel that interfit with corresponding recessed areas on another, allowing for bonding with a bonding material injected through injection holes, reducing the need for extensive fastening and simplifying the assembly process by aligning mating edges in edgewise alignment.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Reliability
If traditional splice plate methods with extensive fasteners are used to join composite fuselage sections, then structural strength and reliability are improved, but assembly time and labor requirements increase significantly
Solution Approach 1:
The patent combines multiple fastening functions into a single integrated splice plate structure. The splice plate integrates longitudinal stiffener functions, lateral frame functions, and fastening functions into one component, eliminating the need for separate fasteners at each location. This merging reduces the number of fastening operations from thousands to a manageable number while maintaining structural integrity through the integrated design.
Solution Approach 2:
The splice plate is divided into multiple modular components including longitudinal extensions, lateral extensions, and orthogonal flanges that can be separately manufactured and then assembled. This segmentation allows for easier manufacturing, quality control, and assembly while the complete splice plate structure provides the necessary structural strength. The modular design enables the complex joining function to be achieved through coordinated simpler components.
2Manufacturing precision
If traditional splice plate methods with extensive tooling fixtures are used, then manufacturing precision is improved, but device complexity increases
Solution Approach 1:
The splice plate design incorporates self-aligning features and self-positioning capabilities that eliminate the need for complex external tooling fixtures. The orthogonal flanges, longitudinal extensions, and lateral extensions are designed to automatically align with corresponding features on the fuselage sections during assembly. This self-service approach maintains manufacturing precision through inherent geometric constraints rather than requiring elaborate external positioning equipment.
3Strength
If compound splice plate systems with longitudinal extensions and fittings are used, then structural strength is improved, but ease of manufacture deteriorates
Solution Approach 1:
The compound splice plate system is segmented into distinct manufacturable components: a central splice plate, longitudinal extensions, lateral extensions, and orthogonal flanges. Each component can be manufactured using standard composite fabrication processes without requiring complex tooling. The components are designed with standardized attachment interfaces that simplify assembly, allowing the high-strength joint to be achieved through coordinated simpler parts rather than a single complex monolithic structure.
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 significantly reduces the labor and part count required for joining composite fuselage sections, enhancing efficiency and reducing assembly time by eliminating the need for extensive fastening, thus lowering the overall man-hours needed for assembly.
Implementation Method 1
a bonding material bonding the first and second panels together
Data Source
Figure 1~2
Figure 3~5
Figure 6~7
AI summary
An aircraft fuselage assembly (100) comprising: first (112) and second (114) skin sections each further comprising a plurality of plies of composite materials and having a mating edge, with the mating edges (116, 118) disposed in edgewise alignment with each other; and a plurality of tabs (120) each extending longitudinally from the mating edge (116) of the first skin (112) section along an interior surface of the second skin (114) section, each tab (120) including an exterior surface (124) in facing registration with a corresponding portion of the interior surface (132) of the second skin (114) section, each of said tabs (120) further including at least one injection hole (142) formed therethrough for injection of bonding material (140) between said exterior surface (124) and said corresponding portion for attaching the first (112) and second (114) skin sections.