Title of the invention

Tailored Autoclave Cure cycle for better consolidation for large size wing skin of a fighter aircraft

Field of invention

Manufacturing of structural composites parts by Autoclave curing for a fighter aircraft. Composites materials are defined as two or more constituent materials which are combined together to obtain the desired engineering properties. Use of composites parts in the aeronautical applications has the following benefits:

a. Lightweight

b. High Specific strength

c. Ease of manufacture

d. Reduction in sub-assemblies

e. Corrosion resistance

Composites manufacturing involves usage of prepregs with vacuum bagging method and subsequent curing in autoclave. Getting composites parts cured in autoclave has influence on the structural performance. The cure cycle has a significant role in getting the defect free part.

Use of invention
The tailored cure cycle allows a large size component like wing skin of a aircraft having various thicknesses in different zones to be consolidated and cured without gross defects such as porosity, voids, poor consolidation and warpage.

Prior art

No prior art is available in this field which can explain how the better consolidation can be achieved by modification in the cure cycle for the large size autoclave thermoset component like wing skin of an fighter aircraft in the open literature or domain.

The reference of prior art available is United States Patent 8412380 which corresponds to cure time adjustment for a rubber article

Another prior art available is United States Patent 6696009 which explains on the usage of caul plate during layup and curing for components with variable thickness.

Draw backs of prior art

When the large size thermoset component like wing skin of an aircraft is cured in the autoclave with first well temperature at 135 Deg C , the following are the draw backs in following the recommended cure cycle of the prior art:

• Raise in the temperature (high spots) in the different zones of variable thickness of the large size wing skin component which lead to local gelling in different zones which in turn lead to non uniform consolidation and bujit up of residual stresses , porosity , warpage etc

• The defects like porosity lead to rejection of the Composite part.

• The defects like warpage lead to structural infidelity in the sub assembly of the aircraft
structure.

Comparison between prior art and present invention

i"rte prior art ^recommended cure cya'e/ does not adequately meet the requirement of good consolidation. However, it do consider the other requirements like degree of cure

The present invention does consider the uniform consolidation including degree of cure by tailoring the cure cycle for large size monolithic part like wing skin of a fighter aircraft with variable thickness zones.

Aim of the invention

The aim of the invention is to tailor the cure cyqe for component to achieve better consolidation by eliminating the raise in the temperature in the hot spots in the different zones. This in turn eliminates the non uniform consolidation, residual stresses, porosity, warpage etc.

Summary of the present invention

The present invention has been made to eliminate the non uniform consolidation in the different zones of the large size autoclave cured component like wing skin of a fighter aircraft by tailoring the cure cycle. The reduction in the dwell temperature from 135 Degree centigrade to 120 Degree centigrade helped in better consolidation of the part thereby eliminating the porosity, built up residual stresses, warpage etc.

Brief description of drawings

^Figure 1. Typical cure cycle for a monolithic part of an fighter aircraft (Recommended)

It is general practice to cure the thermoset prepregs in an autoclave with cure cycle in two dwell process for the large size autoclave component like wing skin of an aircraft. In the two dwells curing, the first dwell is an important stage which deals with trapped gasses and excess resins are made to escape .The choosing of dwell temperature is critical in the curing process which depends on the minimum viscosity of the resin system to achieve better compaction. Also, as the temperature increases, the viscosity and the gel time decreases resulting in poor consolidation of the part for a given vacuum and pressure.

When large composites components are subjected to autoclave curing with higher first dwell temperature, the different zones of the large size wing skin experiences the higher temperature (hot spots) than the rest of the zones of the part. This results in local gelling and these zones suffers non uniform consolidation resulting in porosity, warpage etc.

The. Tailored Autoclave cure cycle for large size thermoset monolithic part like wing skin of a fighter aircraft to eliminate.the above undesirable effects is given at Figure 2.

Statement of invention

This invention of tailored cure cycle for thermoset prepregs has resulted in defect free better consolidated large size wing skin for a fighter aircraft.

Detailed description of invention

In the present invention, the first dwell temperature was reduced to 120 Degree centigrade from 135 Deg centigrade. Wherein the decreasing the dwell temperature has increased the viscosity of the resin and gel time leading to better consolidation of the part for the given vacuum and temperature. This invention has eliminated the non uniform consolidation in the various zones ofthe part thereby eliminating the defects like porosity; warpage etc .The present invention has lead to a better consolidated defect free composite wing skin for a fighter aircraft.

Claims

Tailored Autoclave Cure Cycle with reduction in the dwell temperature from 135 Degree Centigrade to 120 Degree Centigrade has helped in better consolidation of the large size Thermoset autoclave cured wing skin of a fighter aircraft thereby eliminating the Porosity, built up residual stresses, warpage etc leading to a defect free part.