Monolithic Composite Aircraft Leading Edge Co-Curing
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Solution Overview
Problem
The existing manufacturing methods for aircraft leading edges are costly due to the need for multiple separate elements and complex tooling, which increases production costs and assembly complexity.
Innovation Solution
A monolithic leading edge profile section is manufactured using composite prepreg technology, incorporating a leading edge skin, auxiliary spar, and stringers, which are co-cured in an autoclave to reduce the number of components and simplify assembly, with the auxiliary spar's position and configuration optimized for structural and aerodynamic performance.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Manufacturing precision
If multiple separate elements (ribs, spars, stiffeners) are manufactured and assembled using traditional prepreg technology, then manufacturing precision and structural integrity are maintained, but manufacturing cost and assembly complexity increase significantly
Solution Approach 1:
The patent merges multiple separate leading edge elements (ribs, spars, stiffeners, and skin) into a single monolithic composite structure manufactured using RTM technology. This integration eliminates the need for separate manufacturing and assembly operations, reducing assembly complexity while maintaining structural integrity through the unified monolithic design.
Solution Approach 2:
The monolithic leading edge structure performs multiple functions simultaneously: it provides structural support, aerodynamic shaping, and stiffening all in one integrated component. This multi-functionality replaces what previously required multiple specialized elements, reducing both manufacturing steps and assembly operations.
2Productivity
If Resin Transfer Moulding (RTM) technology is used to integrate all reinforcements in a closed mould, then the number of assembly operations is reduced, but tooling complexity and manufacturing cost remain high
Solution Approach 1:
The patent segments the tooling system into reusable components: a permanent male tooling for the spar and removable female tooling for the skin and ribs. This segmentation allows the complex RTM process to be broken down into manageable stages with standardized, reusable tooling elements, reducing overall tooling complexity while maintaining manufacturing efficiency.
Solution Approach 2:
The spar is pre-formed using permanent male tooling before the final RTM operation. This preliminary action allows the most complex structural element to be prepared in advance with standardized tooling, simplifying the subsequent integration steps and reducing the complexity of the main moulding operation.
3Manufacturing precision
If classical hot-forming processes are used for each element separately, then manufacturing precision is maintained, but the number of independent manufacturing steps and costs increase
Solution Approach 1:
The patent combines multiple separate hot-forming operations into a single RTM process where all elements (spar, ribs, stiffeners, and skin) are formed and integrated in one continuous manufacturing step. This merging maintains precision through the controlled RTM process while eliminating the need for multiple independent manufacturing cycles, thereby reducing costs.
4Strength
If multiple elements are manufactured and assembled with rivets, then structural loads are supported, but manufacturing cost and production time increase
Solution Approach 1:
The patent merges multiple load-bearing elements into a single monolithic structure where loads are distributed throughout the integrated composite design. This eliminates the need for time-consuming rivet assembly operations while maintaining or enhancing load-bearing capacity through the continuous fiber reinforcement and unified structural design.
Solution Approach 2:
The patent utilizes composite materials in a monolithic RTM-manufactured structure to achieve high strength-to-weight ratio and load-bearing capacity without requiring mechanical fasteners. The composite construction provides inherent structural integrity and load distribution, replacing traditional metal assembly methods with rivets.
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 reduces manufacturing and assembly costs while enabling an optimized design for the aircraft lifting surface, improving structural integrity and reducing bird impact risks.
Implementation Method 1
arranged together in an autoclave and subjected to a co-curing process
Implementation Method 2
Using this RTM technology, all dry laminates are formed to the final shape by means of classical hot-forming processes. Then, all formed laminates are co-injected together in a closed mould.
Data Source
Figure 1a~2b
Figure 3a~4b
Figure 5a~6b
AI summary
A method for manufacturing a leading edge (11) profile section (30) of an aircraft lifting surface is provided. The method comprises the following steps: a) providing a set of laminated preforms (31,33) of a composite material configured with a suitable shape for constituting the leading edge profile section; b) arranging said laminated preforms in a curing tooling and subjecting the assembly to an autoclave cycle to co-cure said laminated preforms; c) demoulding the curing tooling in a spanwise direction towards the aircraft symmetry plane. The invention also comprises a leading edge profile section (30) manufactured by said method comprising in addition to the skin of the leading edge profile section, one or more of the following structural elements: an auxiliary spar (33), a longitudinal stiffener (35) reinforcing an auxiliary spar, a longitudinal stringer (37,37') reinforcing the skin of the leading edge profile.