Patent Application
20240375366
This patent application claims priority from Italian patent application no. 102023000009456 filed on May 11, 2023, the entire disclosure of which is incorporated herein by reference.
The invention relates to a method for manufacturing a box-shaped structure made of composite material for aircraft, in particular for a horizontal stabilizer of an airplane, to which explicit reference will be made in the description below without because of this lacking generality.
Structural components are known, which are used in the aviation industry, for example fuselages and parts thereof, and are made of composite materials.
In the state of the art there are aeronautical structural components that are made of lights alloys, hence of metal materials, and will then build part of the fuselage of the aircraft.
As it is known, fuselages are designed to ensure an adequate protection of the payload (crew, passengers, goods, etc.), but, at the same time, they cannot exceed set weight limits.
Furthermore, the use of metal components, despite ensuring a greater resistance, leads to an increase in overall costs.
Hence, structural components made of composite materials are needed in order to reduce the overall weight of aircraft.
As a matter of fact, the use of composite materials reduces the overall weight of aircraft and, at the same time, ensures very resistant structures.
Furthermore, the use of metal elements and the installation thereof in contact with the structures determine galvanic coupling problems, with consequent risks of corrosion of the metal and need to increase inspection levels. This leads to an increase in overall costs for the manufacturers of said components and, hence, for airlines.
Therefore, the use of composite materials stems from the need to reduce the overall weight of aircraft, to eliminate or minimize the corrosion problems affecting aeronautical structures and to increase the resistance to fatigue thereof.
As it is known, the horizontal stabilizer is a structural component with fixed surfaces that generally is part of the tail portion of the fuselage of an airplane.
In particular, the stabilizer is an aerodynamic surface that has the function of stabilizing the airplane along the pitch axis.
More in particular, a typical airplane comprises two horizontal stabilizers, each defined by a sort of wing, usually with much smaller dimensions than the main wings of the airplane, projecting from the tail portion of the fuselage.
The stabilizers extend on opposite sides of the fuselage, generally along a plane that is substantially parallel to the plane of the wings.
More in detail, the two stabilizers are coupled, at their locked ends, to a central box-shaped element known as “center box” and incorporated in the end part of the fuselage.
In other words, the center box is interposed between the two stabilizers.
Each stabilizer carries, on the inside, a box-shaped structure made of composite material.
In particular, said structure basically comprises: an upper wall; a lower wall facing the upper wall and spaced apart from the latter by a non-zero amount; and a plurality of spars extending between the upper and lower walls, defining connecting elements between them and delimiting, with the aforesaid upper and lower walls, respective elongated cavities, each extending between a front opening and a rear opening arranged in planes that are transversal to the walls and to the spars.
As mentioned above, the use of composite materials reduces the overall weight of the structural components and, hence, of aircraft and, at the same time, ensures very resistant structures.
The upper and lower walls are approximately parallel and, more precisely, slightly convergent towards the front openings or towards the rear openings.
The spars consist of partitions transversal and, more in particular, substantially orthogonal to the upper and lower walls.
In particular, each spar consists of two elongated section bars or profile beams or profiles generally having a C-shaped profile and including a back or flange and two wings orthogonally projecting from the opposite end edges of the flange.
According to an alternative embodiment, the aforesaid section bars or profile beams or profiles can have a J-shaped profile.
In order to form a spar, the two C-section profiles are joined to one another along the respective flanges so as to have opposite concavities.
The monolithic box-shaped structure described above is manufactured by means of a known method, which comprises the steps indicated below, starting from a thermosetting (resin) matrix prepreg reinforced by fibers of different nature, such as carbon fibers, aramid fibers, glass fibers, etc.
EP-A-3871869, filed by the Applicant, discloses a method for manufacturing a box-shaped monolithic structure for a horizontal stabilizer of an airplane of the type described above.
More in particular, according to the method described in EP-A-3871869, the aforesaid C-section profiles are laminated on respective elongated forming and support tools with a substantially parallelepiped-shaped profile, which are subsequently used to hold the spars in the respective predefined positions within the box-shaped structure during the manufacturing step.
Each tool consists of a substantially parallelepiped-shaped elongated body, which is delimited by flat faces and has a cross section corresponding, in terms of shape, to the one of the aforesaid elongated cavities of the structure to be manufactured.
In particular, each C-section profile is laminated on three flat faces of the forming and support tool, preferably after the application of a resin-based adhesive agent (known as “tackifier”).
A first face of the tool is completely coated by the prepreg, whereas the other two faces of the tool adjacent to the first face, which are orthogonal to the latter and parallel to one another, are coated only along a segment.
Hence, each tool can support, on opposite sides, two C-section profiles made of prepreg.
Depending on the composition of the tool, a so-called “costing” thereof could have to be carried out before the lamination of the section bars, through application of suitable separator films, tubular bags, ventilation fabrics.
After a vacuum consolidation at a predetermined temperature, the C-section profiles are joined two by two in the way described above so as to form the desired number of spars.
In detail, each spar is arranged so that: the flanges of the relative pair of C-section profiles making it up are interposed between two flat faces facing one another of two tools adjacent to one another; and the wings of each profile rest on respective flat faces of the relative tool parallel to one another and substantially orthogonal to the flat face supporting the flange of the section profile.
The assembly consisting of spars and tools (and possibly additional filling elements known as “noodles”) is inserted into a forming mold including a lower plate, an upper plate and two opposite side plates connecting the upper and bottom plates.
In particular, one or more layers (plies) of prepreg composite material are laminated on each of the two upper and lower plates to define the upper and lower skins, which will constitute, after the curing step, the upper and lower walls of the box-shaped structure to be manufactured.
Once the upper and lower skins are laminated, the preformed spars held in position by their respective forming and support tools are placed on the lower plate of the forming mold carrying the lower skin; subsequently, the upper plate of the forming mold carrying the upper skin is closed onto the side walls of the mold itself and onto the assembly consisting of the preformed spars and the tools.
At this point, a ventilation fabric and a tubular bag are inserted into each tool (which is longitudinally hollow).
Furthermore, known components such as separator films, ventilation fabric and a bag film are placed on the outside of the forming mold so as to create so-called front and rear sealed envelope bags.
The ends of the contiguous tubular bags are also closed and sealed to one another.
At this point, the prepared forming mold is placed inside an autoclave to carry out the curing operation at predetermined pressure and temperature values (for example, for epoxy resins, the curing temperature is circa 180° C. and the curing pressure is between 6 and 7 bar).
The curing pressure acts both outside the structure being manufactured and inside the hollow tools and, therefore, inside the cavities of the structure itself, through the tubular bags.
Once the curing is complete, the bags and the forming and support tools are removed, resulting in the box-shaped monolithic structure consisting of the upper and lower walls (skins) and the spars arranged transversely between them.
Each box-shaped structure, defining a respective stabilizer, is then joined to the aforementioned center box.
There is an increasing need to reduce the total number of components and to further reduce the overall weight of the airplane.
For example, the use of the center box poses a certain assembly difficulty at the center of the two half-wings constituting the horizontal stabilizers, especially in the case of small-sized aircraft.
In accordance with the above, construction solutions have become widespread, in which the center box, which is defined by a complex and relatively heavy box-shaped element, is replaced by a simple beam junction element, for example a relatively thin double-T-section profile beam or profile.
To this aim, each above-described box-shaped structure must be further elongated along its longitudinal direction so as to compensate for the material loss resulting from the removal of the center box (since the extension of the stabilizer must remain the same).
However, elongating the box-shaped structure entails structural challenges within the stabilizer, particularly regarding the rigidity and structural stability necessary to ensure the nominal use performance of the horizontal stabilizers.
To solve this problem, some known solutions involve the insertion of a reinforcing rib in a predetermined longitudinal position of the box-shaped structure (or at an axial coordinate, relative to a longitudinal axis of the box-shaped structure).
More in detail, the reinforcing rib consists of a series of transversal partitions, more in particular substantially orthogonal to the upper and lower walls and to the spars, aligned in series one after the other.
In other words, the reinforcing rib extends transversely to the longitudinal direction of extension of the box-shaped structure, at a given axial coordinate thereof.
In the state of the art, the insertion of the rib is rather complicated. According to a known solution, once the above-described box-shaped monolithic structure is completed, portions of material on the skins and/or at the flanges of the spars are removed to access the inside of the aforementioned cavities.
This procedure undoubtedly introduces complications in the process of manufacturing the stabilizer, as it is necessary to assemble the rib in quite narrow and confined spaces, without the possibility of properly viewing the assembly location. Furthermore, given the nature of the composite fiber material, there may be dimensional tolerance errors between the various components (for example, between spars and skins).
Furthermore, the assembly process of the reinforcing rib becomes increasingly complicated as the size of the aircraft and, thus, of the stabilizers to be manufactured decreases.
Therefore, the Applicant has observed that the described method and the box-shaped structure obtained thereby are open to improvements, in particular in order to strengthen the box-shaped structure, while avoiding excessive complications in the manufacturing process.
The present invention may be directed to a method for manufacturing a box-shaped structure made of composite material for an aircraft as well as a box-shaped structure, which are highly reliable and cost-effective and fulfil at least one of the needs specified above and associated with known methods for manufacturing composite material box-shaped structures.
In the following, an embodiment of the invention is described for a better understanding thereof by way of a non-limiting example and with reference to the accompanying drawings, wherein:
With reference to
In particular, the structure 1 is intended for use in (and, more precisely, for the manufacturing of) a horizontal stabilizer 2 of an aircraft fuselage 3, for example of an airplane 4, to which reference will be made in this description without because of this lacking generality.
As visible in
In the example described herein, the first axial end 1a represents a locked end of the stabilizer 2, whereas the second axial end 1b represents the free end of the stabilizer 2.
The lower 5 and upper 6 walls are, in the case shown herein, approximately parallel to one another. According to a possible alternative not shown herein, the walls 5 and 6 may also be slightly convergent towards the first end 1a or the second end 1b.
The spars 7 consist of partitions that are substantially orthogonal or, more generally, transversal to the lower 5 and upper 6 walls and are elongated in the longitudinal direction.
The structure 1 is manufactured starting from a polymer matrix prepreg, for example a thermosetting resin, reinforced with fibers of various kinds, such as carbon fibers and/or aramid fibers and/or glass fibers, etc.
Alternatively, the structure 1 could also be manufactured starting from a thermoplastic resin matrix prepreg reinforced with fibers of the type indicated above.
A method for manufacturing the structure 1 according to the invention will be disclosed below.
With reference to
With reference to
In detail, each spar 7 is obtained starting from two profiles or profile bars 12 elongated in the aforementioned longitudinal direction, having C-shaped cross sections and each consisting of a flange 13 and of two end appendages or winglets 14 transversely projecting in a substantially orthogonal direction from the opposite end edges of the flange 13 and arranged on the same side thereof. In particular, in order to form a spar 7, the two profiles 12 are joined to one another along the respective flanges 13 so as to have opposite concavities (
More in detail, the profiles 12 are obtained by: providing a plurality of support tools 15 elongated along the aforementioned longitudinal direction, each having a cross section with a polygonal external profile complementary to the profile of the cavities 8 to be formed in the structure 1 and which are intended to subsequently hold the spars 7 in the predefined positions within the structure 1 during the manufacturing step; and laminating respective second pluralities of layers of said prepreg composite material on three contiguous walls of each support tool 15.
The term “laminating” in correlation with the lamination of the second plurality of layers mentioned above includes, for the purposes of this description: both a “direct” lamination of the layers of said second plurality on the support tool 15, and the simple arrangement of the layers of said second plurality on the support tool 15 and the subsequent bending on the aforesaid three contiguous walls, and wherein the layers have already been laminated on another auxiliary tool.
In the example shown herein, each support tool 15 has a parallelepiped-like shape, which is elongated in the aforementioned longitudinal direction and is hollow in the same direction.
In an embodiment which is not shown herein, each support tool 15 may have a more complex geometry, in order to adjust to the thickness differences that may be present within the stabilizer 2 both transversely and longitudinally.
Conveniently, the support tools 15 have a composition based on reinforcing material and polymer suited to permit a shift from a rigid state to a flexible elastomer-like state and vice versa in response to a heating and a cooling, respectively, namely in response to a temperature stimulus.
The polymer constituting the support tools 15 advantageously is a shape memory thermosetting or thermoplastic polymer of the known kind. The polymer can be, for example, a shape memory epoxy polymer, a shape memory cyanate ester polymer, a shape memory polyurethane polymer, a shape memory vinyl polymer, a shape memory polyimide polymer, a shape memory maleimide polymer or combinations thereof, including copolymers.
Thanks to the properties of the shape memory polymer, support tools 15 can regain their original rigid shape despite extensive repeated use and after numerous heating and cooling cycles.
The reinforcing material the said support tools 15 can include one or more elastic fibres.
In particular, the reinforcing material can also contain nylon fibers, lycra fibers, polyester fibers, carbon fibers, glass fibers, aramid fibers, boron fibers, basalt fibers, polymer fibers, chopped fibers, meshes, three-dimensional fiber preforms, plain weave fabrics, twill weave fabrics or other types of fabric and combinations thereof. A suitable commercial example of nylon fibers is the nylon produced by Invista (Wichita, Kans.).
In some embodiments, the support tools 15 may contain two or more different types of reinforcing materials.
Each profile 12 is then externally laminated on three contiguous walls of the respective support tool 15 in the rigid state, preferably after the application of a resin-based adhesive agent (known as “tackifier”).
In particular, a wall of the support tool 15 in the rigid state is completely coated by the prepreg of the respective profile 12, whereas other two adjacent walls are only coated along a predefined segment (as shown in
Therefore, each support tool 15 can support, on opposite sides, two prepreg C-section profiles 12.
Hence, the lamination of the profiles 12 is performed by directly laminating the prepreg material making up said second plurality of layers on the outer surface of the walls of the tools 15 in the rigid state.
Following the lamination, the profiles 12 are compacted as described in EP-A-3871869.
Upon completion of the compaction, in order to form the spars 7, the profiles 12 are joined two by two along the respective flanges 13 by placing respective pairs of support tools 15 (
Possible recesses or gaps in the spars 7 are filled with so-called fillers or noodles as described in EP-A-3871869.
According to an alternative embodiment, the support tools 15 are defined by rigid elements, for example metal elements, namely elements not permitting the shift from a rigid state to a flexible elastomer-like state and vice versa in response to a temperature stimulus, as described in the introduction to the description of EP-A-3871869.
In particular, in this last case, each support tool consists of an elongated, rigid, solid body substantially shaped like a parallelepiped and having a cross section corresponding, in terms of shape, to that of the elongated cavities 8 of the structure 1 to be manufactured. Before placing each spar, still in the form of a preformed prepreg, between two support tools, the latter undergo a coating operation consisting of the following successive steps: applying, on each support tool, a layer of separator agent, for example in the form of film, to facilitate the following extraction of the support tool from the respective elongated cavity; fitting a tubular bag onto each support tool prepared by so doing, leaving an excess of the tubular bag at each end of the support tool for a following sealing operation; wrapping a ventilation fabric on each support tool and on the outside of the tubular bag fixing the edges with a sealant; fitting a tubular separator film onto each support tool prepared by so doing, leaving—also in this case—an excess of the separator film at each end of the support tool for the following sealing operation; sealing the ends of the tubular bag and of the tubular separator film using a sealant; applying vacuum and waiting for the tubular separator film to shrink the entire coating prepared by so doing on the relative support tool.
The assembly formed by so doing, which consists of the spars 7, the support tools 15 and any possible additional structural, support or filler element (and, if necessary, coating element, in the case of the aforementioned alternative embodiment of the support tools 15), is inserted into the forming mold 11.
In particular, according to the manufacturing method, the support tools 15 side by side and the previously obtained spars 7 are longitudinally placed on the lower wall 5 already laid on the first plate 10, as shown in
In particular, the spars 7 are laid on the lower wall 5 so that respective first winglets 14a of the spars 7 (
By so doing, a first assembly 16 of the structure 1 is defined, which includes the lower wall 5 and the spars 7 arranged on the latter.
In the case of the aforementioned alternative embodiment, wherein the support tools 15 are defined by rigid elements, for example, metal elements, a separator film, a ventilation fabric and a bag film are sequentially arranged on the entire forming mold; the bag film is sealed with a sealant at the base of the forming mold.
The aforementioned tubular separator film arranged on each support tool is opened at the ends; the excess length with respect to the respective support tool is cut off.
The aforementioned tubular bags present on the respective support tools are folded at the ends and sealed.
Hence, a so-called “envelope bag” is obtained by sealing the edge of the external bag film arranged on the forming mold to the ends of the tubular bags of the support tools and by further sealing the ends of the contiguous tubular bags to one another.
At this point, vacuum is applied inside the envelope bag until the material of the external bag shirks towards the external surface of the forming mold.
The ends of the tubular bags are then opened and, by continuing to apply vacuum, the tubular bags detach themselves from the respective support tools, tending to minimize the volume enclosed inside the envelope bag.
With reference to
According to the invention, the upper wall or second wall 6 comprises a main portion 17 and a closing portion 18, which are coplanar to one another and define two distinct pieces with discontinuity.
In other words, the wall 6 is defined by two distinct and separate parts, which, at the same time, can be coupled to one another, as better explained below.
The method according to the invention entails laminating a third plurality 17a of layers of prepreg composite material on a second forming plate 19 of the forming mold 11, thus obtaining a main portion 17 of a second composite material skin 6a, said second skin 6a defining the upper wall 6 (
According to this first preferred embodiment, the method according to the invention further entails laminating a fourth plurality 18a of layers of prepreg composite material onto a third forming plate 20 of the forming mold 11, which is separate and distinct from the second plate 19 (
In accordance with what is specified in the introduction to the description and with reference to
In detail, the reinforcing rib 21 consists of a series of transversal partitions 22, more in particular substantially orthogonal to the walls 5 and 6 and transversal to the spars 7, aligned in series one after the other.
More in detail, the rib 21 is located in the area of a transversal sector S of the structure 1, which defines a given longitudinal position (or “axial coordinate”) along the structure 1, interposed between the first end 1a and the second end 1b.
The reinforcing elements 22 are aligned with one another transversely to the axis A and in the area of (namely, along) the transversal sector S.
The reinforcing elements 22 can be made both of composite material and of metal material.
In an embodiment that is not shown herein, the rib 21 could be oriented orthogonally to the axis A.
Furthermore, according to an embodiment that is not shown herein, the reinforcing elements 22 could have different geometries, which are not shown herein, for example they could have a C-shaped section, a double T-shaped section, etc.
In other words, the reinforcing rib 21 extends transversely to the longitudinal direction of extension of the box-shaped structure 1.
The presence of rib 21 strengthens the structure 1 and increases the stability and rigidity thereof, in view of the removal of the aforementioned “center box”.
Furthermore, the presence of rib 21 allows manufacturers to provide, in the area of the transversal sector S, control systems (for example, pneumatic, hydraulic or electric pistons) for the inclination of stabilizer 2 (namely, the incidence of the half-wing), which are known per se and are not shown or described in detail. This configuration allows some of the stresses arising from the aforementioned control systems to be offloaded onto the structure 1 and, in particular, onto the reinforcing rib 21, thereby improving the structural properties and the resistance to stresses of the structure 1.
With reference to
It should be pointed out that
In particular, the method according to the aforementioned first embodiment of the invention involves (
The group comprising the first assembly 16 and the main portion 17 defines a second assembly 23 of the structure 1.
In this condition, the closing portion 18 remains separated from the second assembly 23, for example supported by the third plate 20.
The second assembly 23 obtained by so doing then has an opening 24 (
According to this first embodiment, following the placement of main portion 17 on the first assembly 16, a curing of the second assembly 23 is performed in a known manner, which is not described in detail, by applying a predetermined temperature and pressure to the first plurality, second plurality and third plurality 17a of composite material layers respectively laminated in the ways described above.
In other words, the layers constituting the second assembly 23 are cured together.
As schematically shown in
More precisely, during the step of curing the second assembly 23, the curing pressure is applied both inside the forming mold 11 and inside the support tools 15, whose walls are made flexible by the shift from the rigid state to the flexible elastomer-like state and are therefore forced, by the curing pressure itself, to adhere to said profiles 12 and to said first wall 5 and main portion 17.
Conveniently, the support tools 15 are configured to assume the flexible elastomer-like state at a temperature lower than the curing temperature and higher than 50° C.
In the case of the alternative embodiment discussed above, in which the support tools 15 are rigid and are not made of a polymer that is deformable in response to a temperature stimulus, the so-called envelope bag is manufactured, as mentioned above. Then, vacuum is applied inside the envelope bag until the material of the external bag shirks towards the external surface of the forming mold. In this case, therefore, the curing pressure is applied through the tubular bags and not through deformation of the support tools 15.
Once the second assembly 23 is cured, the support tools 15 can be removed, thus defining the aforementioned cavities 8.
In accordance with the above, according to this first embodiment of the invention, the layers constituting the closing portion 18 are cured separately from the main portion 17 and the second assembly 23.
In particular, a curing of the closing portion 18 is performed by applying a predetermined temperature and pressure to the fourth plurality of layers 18a placed (and previously laminated) on the third plate 20, according to a known procedure that is not described in detail.
In order to implement the rib 21, following the placement of the main portion 17 on the first assembly 16 and, in particular, following the curing of the second assembly 23, the method according to the aforementioned first embodiment of the invention involves inserting the aforementioned reinforcing elements 22 into the cavities 8 and between the spars 7 in the area of the transversal sector S (
More precisely, each reinforcing element 22 is inserted into the respective cavity 8 in the area of the opening 24 or in the area of the first end 1a and is caused to longitudinally slide (along the axis A) until it reaches the transversal sector S, as schematically shown in
Each reinforcing element 22 preferably comprises a back 22a and two winglets 22b (an upper one and a lower one) projecting transversely and, more in particular, orthogonally from the back 22a.
Once the reinforcing elements 22 are inserted, the winglets 22b are coplanar to and in contact with the first wall 5 and the main portion 17, respectively.
Each reinforcing element 22 further includes a pair of side flanges 22c, which are opposite one another and orthogonal to the back 22a and to the winglets 22b.
The flanges 22c are configured to be placed in contact with the flanges 13 of the spars 7, once the rib 21 is in position in the area of the transversal sector S, as shown in
Up to this point, the closing portion 18 remains separated from the second assembly 23, as schematically shown in
This allows the reinforcing elements 22 and, hence, the rib 21 to be inserted in a simple and smooth manner, without having to use probes to view the inside of the finished structure and without the need to drill holes and/or make openings in the finished structure. This results in a significant time saving and in a substantial reduction of assembly errors, thereby improving the quality of the assemblies and avoiding damages to the structure 1.
Following the insertion of the rib 21, the method according to the aforementioned first embodiment of the invention then involves (
In particular, the closing portion 18 is also placed so as to be coplanar to and in contact with the winglets 22b of each reinforcing element 22 constituting the rib 21.
Thus, the rib 21 is simultaneously in contact with both the closing portion 18 and the main portion 17, by means of the (upper) winglets 22b.
In light of the above, the closing portion 18 defines a sort of “cap” portion designed to close the aforementioned opening 24, as schematically shown in
The aforementioned placement of the closing portion 18 on the assembly 23 is appropriately carried out subsequent to the aforementioned curing of the fourth plurality 18a of composite material layers on the third plate 20.
In light of the above, the second wall or upper wall 6 is manufactured in two parts separated from one another in the area of the transversal sector S, with the insertion of the rib 21 in the area of said transversal sector S between the placement of the first one of said two parts (the main portion 17) and the placement of the second one of said two parts (the closing portion 18).
Thanks to the method described above, placing the rib 21 inside the structure 1 is easier compared to the case in which the rib is inserted after the construction and the firm and integral coupling of the walls 5 and 6 and of the spars 7.
As shown in
Therefore, the method according to the invention entails, following the placement of the closing portion 18, joining the main portion 17 and the closing portion 18 together.
In particular, as schematically shown in
In an embodiment, the splicing body 25 is made of metal material.
Once positioned, the splicing body 25 is nailed or riveted to the wall 6, namely to the main portion 17 and to the closing portion 18 (
Appropriately, the splicing body 25 is also nailed or riveted to the winglets 22b lying underneath, so that the rib 21 is fixed in its position in the area of the transversal sector S.
In an alternative embodiment, the splicing body 25 is made of prepreg composite material, for example the same material making up the layers of the aforementioned pluralities of composite material layers.
In this case, the splicing body 25 made of composite material is preferably manufactured and cured separately and then joined as described above.
Alternatively, after the placement of the splicing body 25 in the area of the transversal sector S and above at least part of the joining zone G, the assembly defined by the splicing body 25, the main portion 17 and the closing portion 18 is cured by applying a predetermined temperature and pressure (namely, the splicing body 25 made of composite material is “co-bonded” with the main portion 17 and the closing portion 18 already cured).
Whatever the nature of the splicing body 25, this leads to the creation of the second wall or upper wall 6 in its entirety, by permanently joining the main portion 17 to the closing portion 18.
The box-shaped structure 1 is thus obtained.
Thanks to the configuration in accordance with this first preferred embodiment of the invention, it is possible to obtain a box-shaped structure 1 for a stabilizer 2 of an airplane 4 including a reinforcing rib 21 and devoid of the aforementioned “center box” in a simple and quick manner, thereby strengthening the structure 1 while avoiding introducing excessive complications into the manufacturing process. This results in a significant time saving and in a substantial reduction of assembly errors, thereby improving the quality of the assemblies and avoiding damages to the structure 1. Furthermore, the overall ergonomics of the operators involved in the manufacturing process is improved, thus enhancing their safety.
With reference to
In particular, only the differences between the method according to the first embodiment and the method according to the second embodiment will be explicitly explained below. Unless expressly indicated, the two methods are to be considered identical.
More in particular, an alternative process for manufacturing the upper wall 6 will be described, in accordance with the aforementioned second embodiment of the invention.
In accordance with said second preferred embodiment, the method according to the invention involves laminating the fourth plurality 18a of layers of prepreg composite material on the aforementioned second forming plate 19, namely on the same plate on which the third plurality 17a of layers constituting the main portion 17 are laminated, so that the main portion 17 and the closing portion 18 are coplanar and adjacent on the second plate 19 (
Preferably, the second plate 19 has a longitudinal axis B.
More precisely, according to the second embodiment of the invention, the following operational steps are carried out one after the other: the layers of the third plurality 17a are laminated on the second plate 19 (
Preferably, part of the separator film 26 is also folded so that at least a part thereof is interposed between the second plate 19 and the laminated closing portion 18, so as to prevent possible resin movements from the main portion 17 to the closing portion 18 from somehow causing an undesired “sticking” of the layers of the closing portion 18 during the lamination.
Preferably, the aforementioned curing of the main portion 17 and of the closing portion 18 is carried out after having placed the main portion 17 and the closing portion 18 on the aforementioned first assembly 16 (as schematically shown in
More in detail, the main portion 17 and the closing portion 18 have dimensions that are such that, once placed on the first assembly 16, the main portion 17 extends from the second end 1b to the transversal sector S and the closing portion 18 extends from the transversal sector S to the first end 1a.
During this curing step, the tools 15 remain inside the forming mold 11, as already explained in accordance with the first embodiment.
The tools 15 are preferably removed once the curing step has ended.
Once the aforementioned curing is carried out, the method according to the second preferred embodiment involves decoupling the closing portion 18 from the main portion 17 and from the spars 7, using the separator film 26, thus defining the aforementioned opening 24.
Indeed, thanks to the presence of the separator film 26, the closing portion 18 can easily be separated from the main portion 17 and from the spars 7, even though it has been cured together with these components (which, together with the first wall 5, define the aforementioned second assembly 23).
Following the temporary decoupling of the closing portion 18, the aforementioned procedure for inserting the rib 21 is carried out (
After that, the closing portion 18 is again positioned so as to cover the opening 24, obtaining the box-shaped structure 1 with the rib 21 on the inside (
Subsequently, the aforementioned final permanent joining of the main portion 17 to the closing portion 18 is carried out using the splicing body 25 (
Thanks to the configuration in accordance with the second embodiment of the invention, it is possible to obtain a structure 1 with minimal tolerance errors, despite a slightly more complex lamination compared to the previously described case.
Indeed, since the main portion 17, the closing portion 18 and the spars 7 are cured together, the shape coupling between the spars 7 and said portions 17 and 18 and between the portions 17 and 18 themselves is more precise than in the previously described case
In fact, the main portion 17 and the closing portion 18, for they are laminated on the same tool (i.e. the second plate 19), fit optimally compared to the case in which they are laminated on two distinct plates (as in the previous case), despite a slightly more complex lamination compared to the previously described case.
Based on the above description, the method according to the invention (both in accordance with the first embodiment and in accordance with the second embodiment) results in a box-shaped structure for fuselages and aircraft wings and, in particular, for a horizontal stabilizer, which is made of a fiber-reinforced polymer matrix prepreg composite material, said structure having a longitudinal axis, extending longitudinally between a first axial end and a second axial end, and comprising: first longitudinal wall; a second longitudinal wall facing the first wall and spaced apart from the first wall by a non-zero amount; a plurality of spars extending longitudinally and transversely between said first wall and second wall, connected to them and delimiting, with the first wall and the second wall, respective longitudinal cavities; and a reinforcing rib including a series of transversal reinforcing elements, each inserted in one of said cavities and between two adjacent spars, the transversal elements being aligned with one another transversely to said longitudinal axis and in the area of a transversal sector of said box-shaped structure interposed between said first axial end and second axial end; wherein the second wall comprises a main portion and a closing portion coplanar to one another and defining two distinct pieces with discontinuity and separated from one another in the area of said transversal sector.
Preferably, the closing portion is coupled to the main portion by means of a splicing body; said splicing body is simultaneously nailed or riveted to the main portion and to the closing portion.
Alternatively, the splicing body is made of a prepreg composite material and is manufactured and cured separately or cured in co-bonding with the main portion 17 and the closing portion 18 already cured, as described above.
The features of method for manufacturing the structure 1 and of the structure 1 according to the invention reveal evident advantages that can be obtained with them.
In particular, it is possible to obtain a box-shaped structure 1 for a stabilizer 2 of an airplane 4 including a reinforcing rib 21 and devoid of the aforementioned “center box” in a simple and quick manner, thereby strengthening the structure 1 while avoiding introducing excessive complications into the manufacturing process.
Indeed, thanks to the manufacturing method according to the invention, it is not necessary to carry out complicated, expensive and lengthy processes of detection and/or measurement of the inside of the finished box-shaped structure.
Moreover, it is not necessary to make holes, openings or the like in the finished structure to carry out assembly operations in general or to insert detection probes and reinforcing elements, as these are inserted before the placement and the final joining of the closing portion 18.
Therefore, the method according to the invention results in a significant time saving and in a substantial reduction of assembly errors, thereby improving the quality of the assemblies and avoiding damages to the structure 1. Furthermore, the overall ergonomics of the operators involved in the manufacturing process is improved, thus enhancing their safety (as no additional holes and/or openings have to be made in the finished structure 1, tools designed for this purpose are not required).
Therefore, thanks to the method and the structure 1 according to the invention, the elimination of the “center box” and the replacement thereof with a simple beam-like joining element, for example a relatively thin double T-section profile, are easier and simpler to be carried out, besides not introducing structural critical elements within the stabilizer 2.
To this regard, it should be pointed out that the assembly of said beam-like joining element (schematically shown in
In light of the above, the structure 1 can be elongated along the longitudinal axis A to compensate for the material loss resulting from the removal of the “center box”, however without introducing excessive complications into the manufacturing process of the stabilizer 2.
Furthermore, the presence of rib 21 allows manufacturers to provide, in the area of the transversal sector S, control systems (for example, pneumatic, hydraulic or electric pistons) for the inclination of stabilizer 2, which are known per se and are not shown or described in detail. This configuration allows some of the stresses arising from the aforementioned control systems to be offloaded onto the structure 1 and, in particular, onto the reinforcing rib 21, thereby improving the structural properties and the resistance to stresses of the structure 1.
Therefore, the closing portion 18 defines a sort of structural “cap”, through which the loads and stresses exerted, in use, upon the stabilizer 2 pass and are distributed.
The method and the structure 2 disclosed and shown herein can be subjected to changes and variants, without for this reason going beyond the scope of protection set forth in the appended claims.
| Number | Date | Country | Kind |
|---|---|---|---|
| 102023000009456 | May 2023 | IT | national |
