Patent Application
20100252180
This invention relates to a process for the manufacture of parts made of composite materials with two curing cycles, and more specifically parts intended for aeronautical structures.
Processes are known for the manufacture of many parts made of composite materials made with polymer resins and fiber reinforcements intended for aeronautical structures in which at least one of the subcomponents of the part is subjected to a first curing cycle in an autoclave and the assembly of the part, with all its subcomponents duly assembly, is subjected to a second curing cycle in an autoclave.
At the same time, the drawback considered due to the exothermicity of the chemical reaction taking place during the polymerization of the resin during curing in the autoclave is well known in the aeronautical industry: the heat produced by the reaction is added to the heat of the autoclave and can cause an unwanted overheating of the piece subjected to the curing cycle. Techniques such as the slow or step-wise increase of the autoclave temperature or optimization techniques such as those described in U.S. Pat. No. 5,345,397, in which the optimal autoclave temperature is periodically recalculated according to the temperature of the piece, having to that end instruments for measuring the controlled variables, are known for controlling this drawback.
However, effective techniques for reducing the effect of the exothermic reaction in the manufacture of pieces with different subcomponents, and particularly techniques reducing the manufacturing cost, are not known.
This invention is aimed at solving these drawbacks.
This invention proposes a process for manufacturing a piece of composite material made with polymer resin and fiber reinforcement from at least two subcomponents comprising the following steps:
An advantage of the process object of this invention is that it reduces the time of the first curing cycle with the subsequent savings in manufacturing costs.
Another advantage of the process object of this invention is that part of the exothermicity of the resin is released in the first curing cycle, a smaller exothermal component remaining which may be released in the second curing cycle, without excessively increasing the temperature during the process.
Other features and advantages of this invention will be understood from the following detailed description of an illustrative embodiment of its object in relation to the attached drawings.
The conventional process for manufacturing a piece of composite material such as a horizontal tailplane (HTP) skin panel of the Airbus 380 formed by a skin 11 stiffened by T stringers 13 comprises the following steps:
x) Providing the skin 11 in a cured state.
y) Providing the stringers 13 uncured.
z) Co-bonding the skin+stringers assembly.
Step x) in turn comprises the following basic steps:
Step y) in turn comprises the following basic steps:
Step z) in turn comprises the following basic steps:
In the conventional process, in the curing cycle of both step x) and step z), the temperature that is reached is 185° C. for 2 hours and the autoclave pressure is 135 psi. The skin is therefore subjected to two curing cycles.
All heat-setting resin systems develop an exothermal reaction during their polymerization with a temperature increase. The greater the mass per surface unit of resin that is heated, and therefore the greater the thickness of the material, the greater this exothermal reaction. After a certain thickness (different for each type of resin), the exothermal reaction of the resin begins to have visible effects on the resin curing cycle. When the stabilization temperature is reached and the supply of heat to the autoclave is shut off, the exothermal reaction causes a temperature increase I and therefore the resin continues heating up (see
One possible way of controlling this effect of exothermicity is to reduce the heating rate. However, reducing the heating rate leads to very long and therefore expensive manufacturing processes.
According to this invention, the process for manufacturing the same piece previously mentioned, i.e. a horizontal tailplane (HTP) skin panel of the Airbus 380 formed by a skin 11 stiffened by T stringers 13 comprises the following steps:
a) Providing the skin 11 in a partially cured state.
b) Providing the stringers 13.
c) Co-bonding the skin+stringers assembly.
As the person skilled in the art will understand, steps are used in these steps which are similar to those of the conventional process that has been omitted for the sake of simplification. The basic differences of the process according to this invention and the conventional processes are the following:
In step a) the curing cycle for the skin 11 is carried out up to a certain temperature T1 such that, at normal heating rates, on one hand a partial degree of curing is reached which allows the skin 11 to be stiff enough so that its geometry is not modified when the stringers 13 are assembled on it, and on the other hand, part of the exothermal component of the resin is released, without reaching the resin curing temperature CT, such that the exothermal component of the skin 11 during the curing cycle of step c) is so small that no visible overheating of the skin occurs.
Therefore the temperature of the piece during the first curing cycle must be on one hand greater than the resin gelling temperature GT and on the other less than the resin curing temperature CT so that a sufficient part of the exothermal component is released at a normal heating rate, for example 0.8° C./min, so that the exothermal components of the skin 11 during the curing cycle of step c) is so small that no visible overheating of the piece occurs. This step is depicted in
It is considered that the exothermal component should be released from the skin 11 in the first cycle in a degree exceeding 50%, preferably 65%.
The stringers 13, which can be in a fresh or precured state, are assembled in the skin 11 in step c). If they are fresh, tools consisting of metal angle irons 25 and plates 27 are used for that purpose. As previously mentioned, the skin 11 must be stiff enough so that said metal angle irons 25 and plates 27 doe not leave marks on it. The vacuum bag 29 is placed and then a curing cycle is carried out until the degree of curing the skin 11 and stringers 13 exceeds 90%.
As shown in
Described below are the results obtained in a specific embodiment of the invention using as a first subcomponent a panel with 300×300 mm surface area and 50 mm thick with an epoxy resin impregnated carbon fiber material designated Toray 3911-T800, and as a second subcomponent a panel with 50×50 mm surface area and 3 mm thick and a parallel example following the conventional process and using the same subcomponents.
It is understood that these subcomponents acceptably simulate the skin and a stringer of the coating of a tailplane or wing of an airplane.
In the first step of the process, the first subcomponent is partially cured following the curing cycle depicted in the lower part of
According to the measurements obtained by thermocouples located both in the center and at the ends of the panel, the exothermal reaction causes a temperature increase I2 up to 164° C., a temperature which is less than 180° C., which is the curing temperature for the material. The curing degree of the first subcomponent was also analyzed, obtaining a value of 75%.
Applying a curing cycle according to the convention process at a maximum temperature of 180° C., depicted in the upper part of
In the third step the curing cycle is carried out conventionally in both cases: the temperature is increased at a rate of 0.8° C./min until reaching 180° C., this temperature being maintained for 2 hours.
There was not overheating due to exothermicity in any of them.
The curing degree of the final piece was about 95% using both the conventional process and the process according to the invention.
The selection of the temperature, time and heating rate parameters of the first curing cycle was based on the differential scanning calorimetry (DSC) and rheometry carried out on preimpregnated material.
The process object of this invention is not only applicable to the manufacture of the HTP skin panel of an aircraft from a precured skin and a plurality of stringers in a fresh state, but to any process for the manufacture of pieces of composite materials from two subcomponents in which a partial curing of one of the subcomponents is carried out in the first step. The second subcomponent can be provided in a fresh or precured state such that co-bonding is carried out in the second step if the second subcomponent is provided in a fresh state or a secondary gluing if the second subcomponent is provided in a precured state.
The application of the process of this invention is particularly considered for the manufacture of pieces with the following subcomponents:
The stringers can have a T, Ω or U profile, or any other suitable profile.
It is considered that the proposed solution for the problem of exothermicity is shown to be particularly advantageous in the manufacture of pieces in which the first subcomponent has areas with a thickness exceeding 25 mm.
In the preferred embodiment described above any modifications comprised within the scope defined by the following claims can be introduced.
| Number | Date | Country | Kind |
|---|---|---|---|
| ES06/070123 | Jul 2006 | ES | national |
