Composite laminate for an airframe lifting surface and method for manufacturing thereof

Abstract

A composite laminate (2) for an airframe lifting surface comprising a set of stacked plies (10) of pre-preg material, the composite laminate (2) being surrounded by at least an edge (3) following an edge direction E along the composite laminate (2), the set of stacked plies (10) including: a first plurality (4) of plies (4.1) formed by tapes arranged parallel to the edge direction E, and a second plurality (5) of plies (5.1, 5.2, 5.3) formed by tapes arranged in a direction different to the edge direction E, wherein the thickness of at least one ply of the second plurality (5) of plies is lower than the thickness of each ply of the first plurality (4) of plies.

Claims

1. A method for manufacturing a composite laminate for an airframe lifting surface, the method comprising: laying up pre-preg material to form a set of stacked plies, wherein the set of stacked plies has an edge following an edge direction, and wherein the laying up includes: laying up a plurality of first plies one ply on another by arranging tapes of the pre-preg material parallel to the edge direction to form a first layer of a first plurality of the first plies, a second layer of a second plurality of the first plies, and a third layer of a third plurality of the first plies, and laying up a plurality of second plies by arranging tapes of the pre-preg material in one or more directions which are non-parallel to the edge direction, wherein a thickness of each of the second plurality plies is smaller than a thicknesses of each of the first plies, and wherein one ply of the second plies is sandwiched between the first and second layers, and another one ply of the second plies is sandwiched between the second and third layers; curing the set of stacked plies, and forming the composite laminate for the airframe lifting surface from the cured composite laminate.

2. The method according to claim 1, wherein the laying up of the plurality of the second plies includes arranging the tapes in a direction orthogonal to the edge direction, and the method includes: laying up a third ply with tapes of the pre-preg material oriented at an acute angle with respect to the edge direction, and laying up a fourth ply with tapes of the pre-preg material oriented at an obtuse angle with respect to the edge direction, wherein the third ply and fourth ply form outer layers of the composite material, and wherein the thickness of each of the first plies is greater than a thickness of the third ply and of the fourth ply.

3. The method according to claim 1, further comprising: joining at least one stiffening element to the composite laminate, and forming the aircraft lifting surface from the composite laminate with the joined at least one stiffening element.

4. The method according to claim 3, wherein the joining of the at least one stiffening element to the composite laminate includes co-bonding the at least one stiffening element to the cured composite laminate by an adhesive line.

5. The method according to claim 3, further comprising joining at least one stiffening element to the stacked plies, and the step of curing the stacked plies includes curing the stacked piles with the at least one stiffening element.

6. The method of claim 1, wherein the non-parallel arrangement of the second plies is orthogonal to the edge direction.

7. The method of claim 1, wherein the non-parallel arrangement of the second plies is positive or negative 45 degrees with respect to the edge direction.

8. The method of claim 1, wherein the second plies are distributed throughout the thickness of the composite laminate.

9. The method of claim 1, wherein the thicknesses of each of the second plies are in a range of 0.184 mm to 1 mm.

10. The method of claim 1, wherein the airframe lifting surface is at least one of a torsion box, horizontal tail plane, vertical tail plane, rudder, elevator, aileron, spoiler and flap.

11. The method of claim 1, wherein each of the first plies have a uniform thickness; and the thickness of each of the second plies is less than three-fourths of the uniform thickness.

12. The method of claim 1, wherein the non-parallel arrangement of each of the second plies is at an angle of at least twenty degrees with respect to the edge direction.

13. A method to form a composite laminate having an edge direction for an airframe lifting surface, the method comprising: forming a middle section of the composite laminate by laying up first plies of a pre-impregnated composite fiber material to form a first layer, a second layer and a third layer of the first plies, continuing forming the middle section by laying up a second ply of the pre-impregnated composite fiber material between the first and second layers, and a third ply of the pre-impregnated composite material between the second and third layers, wherein each of the first plies are formed of tows parallel to the edge direction, and the second and third plies are each formed of tows oriented at an angle other than parallel to the edge direction; forming a first outer section of the composite laminate by laying up a fourth ply of a pre-impregnated composite fiber material formed of tows oriented along a direction at an angle other than zero to the edge direction, forming a second outer section of the composite laminate by laying up a fifth ply of a pre-impregnated composite material and formed of tows oriented along a direction at an angle other than zero to the edge direction, stacking the first outer section, the middle section and the second outer section to form a set of stacked of plies, wherein the middle section is sandwiched between the first outer section and the second outer section, curing the set of stacked plies, and forming the composite laminate for the airframe lifting surface from the cured set of stacked plies, wherein the first outer section forms a first outer surface of the composite laminate and the second outer section forms a second outer surface of the composite laminate, wherein the first plies each have a thickness greater than a thickness of each of the second, third, fourth and fifth plies.

14. The method of claim 13, wherein the direction of the tows of the second and third plies is orthogonal to the edge direction.

15. The method of claim 13, wherein the direction of the tows of the fourth and fifth plies is at 45 degrees to the edge direction.

Description

DESCRIPTION OF THE DRAWINGS

(1) These and other characteristics and advantages of the invention will become clearly understood in view of the detailed description of the invention which becomes apparent from the disclosed embodiments of the invention, given just as examples and not being limited thereto, with reference to the drawings.

(2) FIG. 1 shows a perspective view of an airframe lifting surface according to a particular embodiment of the present invention.

(3) FIG. 2 shows a schematic representation of a composite laminate for an airframe lifting surface according to the state of the art.

(4) FIG. 3 shows a schematic representation of a composite laminate for an airframe lifting surface according to an embodiment of the present invention.

(5) FIG. 4 shows an aircraft according to an embodiment of the present invention.

DETAILED DESCRIPTION

(6) The present invention provides an improved composite laminate design for an aircraft airframe lifting surface manufactured with high deposition pre-preg material formats.

(7) Airframe Lifting Surface

(8) FIG. 1 shows a perspective view of an interior side of an airframe lifting surface that comprises a composite laminate (2) comprising a set of stacked plies (10) made of high deposition pre-preg material formats. A plurality of edges (3) bounds the composite laminate (2) as it can be observed in this figure. Each edge (3) defines an edge direction E along the end of the composite laminate, that is, each edge (3) follows an edge direction E along the composite laminate (2).

(9) As shown in FIG. 3, the lifting surface (1) includes a composite laminate (2) that comprises a first plurality (4) of plies formed by tapes that are laid up parallel to the depicted edge direction E, and a second plurality (5) of plies formed by tapes that are laid up in a direction different to the depicted edge direction E. Particularly, the plies (5.1, 5.2, 5.3) that are laid up in a direction different to the edge direction E have a thickness that is lower than the thickness of the plies (4.1) that are laid up in the edge direction E. The new design of composite laminate (2) provides a decrease in the thickness of some of the plies as compared to state of the art if high deposition pre-preg material formats.

(10) The airframe lifting surface shown in FIG. 1 further comprises a plurality of stiffening elements (6), e.g. stringers, located along an upper surface of the composite laminate (2) and bonded to such composite laminate surface. The stiffening elements (6) may be co-cured with a fresh set of stacked plies (10) of the composite laminate (2). In another example, the stiffening elements (6) are co-bonded with a cured composite laminate (2). In another example, the stiffening elements (6) are bonded to the composite laminate (2) by an adhesive line.

(11) Composite Laminate (2)

(12) FIGS. 2 and 3 show a schematic cross sectional view of a composite laminate (2, 22) composed by a plurality of stacked plies forming a set of stacked plies (10, 20) of high deposition pre-preg material formats. In particular, a prior art design according to FIG. 2 is compared to the new design shown in FIG. 3.

(13) FIG. 2 shows a prior art design wherein the composite laminate (22) is composed by a set of stacked plies (20) each with the same thickness. These plies (20) are formed by tapes that are laid up, for each ply, in either a direction parallel to the edge direction E or in a direction different to the edge direction E. In this particular example, the composite laminate (22) comprises in order of laying up the following plies formed by tapes arranged according to a particular direction with respect to the edge direction E: (i) a fourth ply (25.3) formed by tapes that are laid up in a −45° direction, (ii) a third ply (25.2) formed by tapes that are laid up in a +45° direction, (iii) two first plies (24.1) formed by tapes that are laid up parallel to the edge direction E (0° plies), (iv) a second ply (25.1) formed by tapes that are laid up in a direction orthogonal to the edge direction E (90° plies), (v) two first plies (24.1) formed by tapes that are laid up parallel to the edge direction E (0° plies), (vi) a third ply (25.1) formed by tapes that are laid up in a direction orthogonal to the edge direction E (90° plies), (vii) (two first plies (24.1) formed by tapes that are laid up parallel to the edge direction E (0° plies), (viii) a third ply (25.2) formed by tapes that are laid up in a +45° direction, and (ix) a fourth ply (25.3) formed by tapes that are laid up in a −45° direction.

(14) The above design, corresponding to the prior art, would experience the edge effect or delamination on the edges (30) of the composite laminate (20) since the application of high deposition material formats for manufacturing an airframe lifting surface.

(15) FIG. 3 shows a new composite laminate design (2) wherein the composite laminate (2) is composed by a set of stacked plies (10) with different thickness. The plies (10) are formed by tapes that are laid up in either a direction parallel to the edge direction E or in a direction different to the edge direction E. In this particular example (shown in FIG. 3), the composite laminate (2) comprises in order of laying up the following plies formed by tapes arranged according to a particular direction with respect to the edge direction E: (i) a forth ply (5.3) formed by tapes that are laid up in a −45° direction, (ii) a third ply (5.2) formed by tapes that are laid up in a +45° direction, (iii) a second ply (5.1) formed by tapes that are laid up in a direction orthogonal to the edge direction E, (iv) two first plies (4.1) formed by tapes that are laid up parallel to the edge direction E, (v) a second ply (5.1) formed by tapes that are laid up in a direction orthogonal to the edge direction E, (vi) two first plies (4.1) formed by tapes that are laid up parallel to the edge direction E, (vii) a second ply (5.1) formed by tapes that are laid up in a direction orthogonal to the edge direction E, (viii) two first plies (4.1) formed by tapes that are laid up parallel to the edge direction E, (ix) a second ply (5.1) formed by tapes that are laid up in a direction orthogonal to the edge direction E, (x) a third ply (5.2) formed by tapes that are laid up in a +45° direction, and (xi) a forth ply (5.3) formed by tapes that are laid up in a −45° direction.

(16) To reduce or even avoid the edge effect or delamination of the composite laminate (2), the thickness of the plies that are laid up in a direction different to the edge direction E is reduced compared to the prior art laminate shown in FIG. 2. Particularly, the thicknesses of the second plies (5.1) may be 0.184 mm to 1 mm and is reduced as compared to the thicknesses of the plies (4.1) and as compared to the second plies (25.1) of the prior art design shown in FIG. 2. In addition, the new design also includes two more second plies (5.1) formed by tapes that are laid up in a direction orthogonal to the edge direction E. Furthermore, the thickness of the third plies (5.2) and the fourth plies (5.3) is reduced as compared to the thicknesses of the first plies (4.1) and as compared to the plies (25.1) in the prior art design shown in FIG. 2. The thicknesses of each of the second, third and fourth plies (5.1, 5.2 and 5.3) may be three fourths, two-thirds, one-half, one-third or one-quarter, the thickness of any one of the thicknesses of the first plies (4.1). Moreover, the thicknesses of the first plies (4.1) may be uniform for each of the plies. Similarly, the thickness of each of the second plies (5.1) may be uniform for each of the plies, and the thickness of each of the fourth plies (5.3) may be uniform. Moreover, the thicknesses of the second plies (5.1) may be uniform with the thicknesses of each of the third plies (5.2) and the fourth plies (5.3). In an alternate approach, the thicknesses of the second plies (5.1) may be smallest at the second plies (5.1) at or near the center of the composite laminate (2) and may be thicker towards the upper or lower surface of the composite laminate. Similarly, the third and fourth plies (5.2 and 5.3) which are positioned towards the upper or lower surfaces of the composite laminate may be thicker than the second plies (5.1) which are positioned towards the center of the composite laminate.

(17) Therefore, in the composite laminate (2) design shown in this FIG. 3, the thickness of each cured ply where the tapes are arranged in a direction different to the edge direction E is lower than the thickness of each cured ply where the tapes are arranged parallel to the edge direction E.

(18) In a particular example, the 50% of the plies of the set of stacked plies (10) corresponds to first plies (4.1) formed by tapes parallel to the edge direction E; the 33% of the plies corresponds to third and fourth plies (5.2, 5.3) formed by tapes in the laying up direction of ±45° with respect to the edge direction E; and the 17% of plies corresponds to the second plies (5.1) formed by tapes in a direction orthogonal to the edge direction E.

(19) Method for Manufacturing a Composite Laminate (2)

(20) The present invention may be embodied as a method for manufacturing a composite laminate (2) for an airframe lifting surface, the method comprising the steps of: (a) laying up pre-preg material forming a set of stacked plies (10), (b) curing the set of stacked plies (10) obtained in the step a), and (c) obtaining a composite laminate (2) for an airframe lifting surface in the form of a cured composite laminate (2); wherein the cured composite laminate (2) obtained in the step c) is surrounded by at least an edge (3), the edge (3) following an edge direction E along the composite laminate (2);

(21) The step (a) of laying up pre-preg material may comprise laying up at least: (i) a first plurality (4) of plies (4.1) with tapes laid up in the edge direction E, and (ii) a second plurality (5) of plies (5.1, 5.2, 5.3), each ply (5) with tapes laid up in a laying up direction different to the edge direction E; and the thickness of each ply of the second plurality (5) of plies is lower than the thickness of each ply of the first plurality (4) of plies.

(22) The step (a) of laying up the pre-preg material may comprise laying up high deposition pre-preg material formats in the form of tapes (4) that makes up each ply of a plurality of stacked plies (10). According to at least an edge (3) of the composite laminate (2), in the step of laying up, some plies are laid up in the edge direction E and other plies are laid up in a direction different to the edge direction E.

(23) In a particular example, some plies are laid up in a laying up direction of ±45° direction with respect to the edge direction E; other plies are laid up in another laying up direction that is orthogonal to the edge direction E, and some other plies are laid up in a direction that is parallel to the edge direction E.

(24) The thickness of the laid up plies is a particular thickness for each ply so that the thickness of the plies formed by tapes that are laid up in a direction different to the edge direction E is lower than the thickness of the plies formed by tapes that are laid up in the edge direction E. That is, according to the edge (3) of the composite laminate (2) and both its edge direction E and the laying up direction of each ply, the thickness of the laid up tapes varies for each ply respectively in view of the prior art solutions.

(25) In a particular example, after a cured composite laminate (2) is obtained, the method further comprises trimming the composite laminate (2) at its edges (3) in order to obtain the desired form.

(26) Method for Manufacturing an Airframe Lifting Surface

(27) The present invention provides a method for manufacturing an airframe lifting surface comprising the following steps: (a) providing at least a composite laminate (2), (b) providing stiffening elements (6), (c) joining the stiffening elements (6) to the composite laminate (2), and (d) obtaining the airframe lifting surface.

(28) The cured composite laminate (2), may be joined to at least a stiffening element (6) by a co-bonding process wherein the stiffening element (6) is bonded to the composite laminate (2) by an adhesive line.

(29) In another particular example, the set of stacked plies (10), in the form of a fresh composite laminate, is joined to at least a stiffening element (6) and both components are cured together in a same curing cycle by means of a co-curing process.

(30) Aircraft

(31) FIG. 4 shows an aircraft (7) comprising a plurality of airframe lifting surfaces corresponding to a wing torsion box (8), a horizontal tail plain (9), a vertical tail plain (11) and a rudder (12). These airframe lifting surfaces (1) are manufactured from a composite laminate (2).

(32) While at least one exemplary embodiment of the present invention(s) is disclosed herein, it should be understood that modifications, substitutions and alternatives may be apparent to one of ordinary skill in the art and can be made without departing from the scope of this disclosure. This disclosure is intended to cover any adaptations or variations of the exemplary embodiment(s). In addition, in this disclosure, the terms “comprise” or “comprising” do not exclude other elements or steps, the terms “a” or “one” do not exclude a plural number, and the term “or” means either or both. Furthermore, characteristics or steps which have been described may also be used in combination with other characteristics or steps and in any order unless the disclosure or context suggests otherwise. This disclosure hereby incorporates by reference the complete disclosure of any patent or application from which it claims benefit or priority.