Patent Application
20170210053
This disclosure generally relates to composite structures, and more particularly to composite structures that include stiffener members and methods for making the same.
It is sometimes necessary to reinforce composite structures, such as those used in aerospace industry in order to meet needed strength and/or stiffness requirements. These structures include, for example, a skin of an aircraft such as that of a wing and/or fuselage. Skin structures are lightweight and are often thin gauged panels which need added strength and stiffness. Other structures in the aerospace industry, as well as, structures in other industries also need additional strength and/or stiffness. Adding stiffeners to a composite structure, such as to a skin structure of an aircraft, provides the needed strength and rigidity for demands placed on the skin structure of the aircraft.
Traditionally in constructing a reinforced skin, one that comprises a skin and a stiffener or a stringer structure, various fabrication processes have been employed to construct the reinforced skins. In one fabrication process, laying up composite pre-preg material for both the skin and the stiffener structures has been used. Alternatively, fabrication processes have used infusion processes wherein dry fiber was infused with resin for the stiffener elements and dry fibers were homogeneously infused with resin for the skin panel elements.
Utilizing pre-preg was advantageous for purposes of constructing the skin element since composite pre-preg material promoted tight control of optimized fiber volumes for structural efficiency and provided the opportunity to utilize automated lamination equipment to reduce labor costs. The stiffener or stringer structure, on the other hand, required non-automated and expensive hand labor lamination processes. The stiffener often required complex geometries in configuring the stiffener or stringer structure element. Stringers demanded careful placement onto the skin element to avoid fiber waviness in the stiffener structure. Fiber waviness could otherwise reduce performance of the stiffener. Additional complications arose in the fabrication of the stiffener and skin elements both being fabricated by a pre-preg layup process. Use of traditional pre-preg material in this fabrication required high temperature and high pressure curing processes which could introduce undesired results in the finished product. These high temperature and high pressure cure requirements for pre-preg material have been in the more recent past been somewhat ameliorated with utilizing pre-preg material which cures at lower temperatures and lower pressures.
Other past methods for assembling a reinforced skin structure would include making both the skin and the stiffener or stringer structures, as mentioned above, being homogenously constructed from an infused fiber fabrication process with curing the two structures at the same time. The skin structure and the stiffener structure have different fiber configurations and arrangements. The different fiber configurations and arrangements introduce different demands on the infiltrating resin during the infusion process for both of these structures. These demands provide further complications for a homogeneous co-infusion process of both the skin and stiffener structures.
Other processes for fabricating, for example a wind turbine blade, includes an outer structure constructed of dry fibers being infused with resin and an inner structure being constructed of a layup pre-preg structure positioned within the outer structure. Both of these structures are thereafter co-cured. In this process unidirectional pre-preg material is positioned within or otherwise enveloped within a fiber fabric system. The fiber fabric system and the enveloped pre-preg material are then positioned within the confinement of a vacuum bag. Infusion of resin is performed on the fiber fabric system which surrounds the pre-preg element. The infused assembly is co-cured. In this process the pre-preg material forms a connection with the infused fiber bed which surrounds the pre-preg material.
In other fabrication processes, a pre-cured stiffener is fabricated separate and apart from a pre-cured pre-preg skin which has been fabricated with a laying-up process. The pre-cured stiffener structure and the pre-cured skin structure are joined with secondary bonding. The pre-cured stiffener and pre-cured skin structures need to be independently fabricated with geometrical precision to have the surfaces of each of these pre-cured structures properly complement one another and achieve the needed geometries of the assembled structure and to promote a secure secondary bonding together of the two structures together.
An example of a method for assembling a stiffened composite structure includes a step of co-bonding a cured infused composite stiffener to a pre-preg composite laminate skin element wherein the pre-preg composite laminate skin element is dimensionally changeable.
Another example of a method for assembling a stiffened composite structure includes a step of coupling a cured infused composite stiffener which includes a preform of a plurality of braided fibers, wherein the pre-preg composite laminate skin element is dimensionally changeable.
Referring to
In referring to
It will be appreciated that automated equipment can be employed for assembling stiffened composite structure 28 which includes pre-preg composite laminate skin element or structure 20 and stiffener 22. Automation will provide labor cost savings for laying-up plies, for example, of pre-preg composite laminate skin structure 20 and for fabricating braided plurality of dry fibers into preforms, for example, for infused stiffeners 22. Automation of plurality of braided dry fiber preforms will provide the needed precision in assembling stiffener 22 such that the preform of the plurality of dry fibers will conform to complex curvatures and geometries of skin element or structure 20, as well as to other structures to be reinforced, and will avoid unwanted wrinkling configurations of the fibers within the composite material of stiffener 22 which could otherwise affect strength performance of stiffener 22.
A method for assembling stiffened composite structure 28, as seen in
Step 38 of providing pre-preg composite laminate skin element or structure 20 further includes a step of laying up a plurality of pre-preg composite plies. As mentioned above, this step of laying up a plurality of pre-preg composite plies can be a fully automated process. Plies of composite material include fibers that are constructed of a material selected from a wide variety of materials such as glass, aramid or carbon. Similarly, the plies are constructed of a resin selected from a wide variety of resins such as epoxy or bismaleimide resins which may also include toughening additives or components such as thermoplastics or silicon or other particles. The laminate can be assembled with a number of plies that are needed for the construction of a particular composite element or structure and the fiber orientation for the each ply can be positioned as needed for the construction of a particular composite element or structure as well. In this embodiment, pre-preg composite laminate skin element or structure 20, as mentioned above, is at a Stage B of curing, permitting structure 20 to be in a state or condition wherein the dimensions of structure 20 are changeable.
The method for assembling stiffened composite structure 28, as seen in the flow chart of
In step 30, in this example, the composition of the plurality of dry fibers are selected from fibers constructed of one of carbon, aramid and glass. In this example, carbon fibers are employed. The dry carbon fibers are positioned in one of a configuration of braided, tape and fabric utilizing automated equipment and in this example the carbon fibers are braided. The automated equipment, with use of mandrels, if needed, provides accurate construction of the preform so as to provide manufacturing efficiency and precision. Precision is needed to construct a preform that will accommodate the complex geometries and curvatures stiffener 22 in order to address the proper strength support to skin element or structure 20 in fabricating the stiffened composite structure 28. The reliability of implementing use of automated braiding equipment and mandrels promotes dimensional accuracy of stiffener 22 and reduces the occurrence of unwanted fiber waviness.
In this example, the braided preform of plurality of dry fibers is positioned on a tool that accurately replicates the size and shape needed for stiffener 22 to operate with skin element 20. Step 32 of positioning the preform of plurality of dry fibers within a resin barrier and infusing the preform is conducted for fabricating the infused composite stiffener 22. The resin barrier, for example, includes a consumable such as a vacuum bagging film. A vacuum is applied, in this example, to the interior of the bagging film and an epoxy based resin, such as Hexcel RTM6, manufactured by Hexcel Corp. of Dallas, Tex., is drawn into the resin barrier or bagging film infusing the preformed braided, in this example, plurality of dry fibers forming an infused composite stiffener 22. The epoxy based resin is infused, in this example, at a temperature of approximately two hundred and fifty degrees Fahrenheit (250° F.).
The method for assembling composite stiffened structure 28 further includes step 34 of curing infused composite stiffener 22 separate from pre-preg composite laminate skin element 20 and prior to coupling infused composite stiffener 22 to pre-preg composite laminate skin element 20. The curing of stiffener 22 will include, in this example, heating infused stiffener 22 to a temperature up to approximately three hundred and fifty degrees Fahrenheit (350° F.). The pressure applied stiffener 22 in this curing process includes approximately one half an atmosphere of pressure for lower pressure infusion processes up to multiple atmospheres of pressure for higher pressure injection molding processes. The cured infused composite stiffener 22, in this example, results in mass fraction of resin of approximately thirty to thirty five percent (30% to 35%) and a fiber volume fraction of approximately fifty five to sixty percent (55% to 60%).
Once infused stiffener 22 has been cured, in this example, step 36 of preparing surface of cured stiffener 22 takes place. In preparation of co-bonding stiffener 22 and skin 20, in this example, a surface of stiffener 22 is either abraded to form a roughened surface on stiffener 22 or a layer of material of stiffener 22 is removed. The prepared surface of stiffener 22 facilitates coupling stiffener 22 and skin 20, in this example, with co-bonding of stiffener 22 to skin element 20.
As described above, pre-preg composite laminate skin element 20 and infused cured composite stiffener 22 are constructed separately from one another. Once pre-preg composite laminate skin element 20 has been assembled, step 38 of providing the pre-preg composite laminate skin element 20 is taken so as to be able to proceed with the step of coupling skin element 20 and infused composite stiffener 22 together.
An example of a composite stiffener structure 28 is shown in
In a second example of this method, an exploded schematic view of a composite stiffener structure 28 is shown in
The method for assembling stiffened composite structure 28 further includes step 42, as seen in
In this example, low temperature and low pressure curing pre-preg is used for pre-preg composite laminate skin element 20 in the method for assembling stiffened composite structure 28. This low temperature and low pressure pre-preg material is often referred to as out of autoclave pre-preg. Step 42 of co-bonding the cured infused composite stiffener 22 to the pre-preg composite laminate skin element 20 includes a step of applying heat to the cured infused composite stiffener 22 and the pre-preg composite laminate skin element 20 to a temperature within a temperature range which includes two hundred and ten degrees Fahrenheit (210° F.) up to and including two hundred and eighty degrees Fahrenheit (280° F.). This relatively lower temperature range for co-bonding of stiffener 22 to pre-preg skin element 20 reduces the introduction of additional defects in the fabrication of composite stiffened structure 28 which otherwise can be introduced with use of higher temperatures. The step of co-bonding cured infused composite stiffener 22 to pre-preg composite laminate skin element 20 includes a step of applying pressure to the cured infused composite stiffener 22 and pre-preg composite laminate skin element 20 which includes a pressure within a pressure range which includes atmospheric pressure up to and including forty five pounds per square inch (45 psi) of pressure.
While various embodiments have been described above, this disclosure is not intended to be limited thereto. Variations can be made to the disclosed embodiments that are still within the scope of the appended claims.
