Method of manufacturing monolithic structure in composite material for wings or empennages of aircraft

Patent metadata

Field	Value
Publication number	US-12397928-B2
Application number	US-202118270112-A
Country	US
Kind code	B2
Filing date	Dec 23, 2021
Priority date	Jan 4, 2021
Publication date	Aug 26, 2025
Grant date	Aug 26, 2025

Abstract

The invention relates to a monolithic structure in composite material manufactured starting from a fiber-reinforced prepreg material and comprises two walls facing each other and at least one interconnection element extending transversely between the walls, connected to them and delimiting with the latter respective elongated cavities; wherein the walls extend symmetrically at opposite sides of a direction; the interconnection element is a rib extending transversely to the aforesaid direction; at least one wall has a sandwich configuration and comprises two panels facing each other and at least one spar member, which extends transversely between the panels, is connected to them, delimits with the latter respective elongated cavities and extends transversally to the rib.

First claim

The invention claimed is:

1. Method of manufacturing a monolithic structure ( 1 ) in composite material for wings ( 4 ) or empennages ( 5 ) of aircraft ( 2 ), said structure ( 1 ) being manufactured starting from a prepreg material with a fiber-reinforced polymer matrix and comprising:
a first wall ( 6 );
a second wall ( 7 ) facing, along its own surface of greatest extension, to a surface of greatest extension of said first wall ( 6 ) and arranged spaced from the first wall ( 6 ) itself by a non-zero amount; and
at least one interconnection element ( 8 ) extending transversely between said first and second wall ( 6 , 7 ), connected to them and delimiting with the first and second wall ( 6 , 7 ) themselves respective first elongated cavities ( 9 );
wherein said first and second wall ( 6 , 7 ) extend symmetrically at opposite sides of a direction (B), coinciding in use with an extension direction of said wing ( 4 ) or said empennage ( 5 ), from a root portion ( 4 a ), connected to a fuselage ( 3 ) of said aircraft ( 1 ), to a free end portion ( 4 b ) of the wing ( 4 ) or empennage ( 5 ) themselves;
wherein said interconnection element ( 8 ) is a rib ( 10 ) extending transversely to said direction (B), and
wherein at least one ( 6 , 7 ) of said first and second walls ( 6 , 7 ) has a sandwich configuration and comprises:
a first panel ( 11 ) facing another ( 7 , 6 ) of said first and second wall ( 6 , 7 );
a second panel ( 12 ) facing said first panel ( 11 ); and
at least one spar member ( 13 ), which extends transversely between said first and second panel ( 11 , 12 ), is connected to them, delimits with the first and second panel ( 11 , 12 ) themselves respective second elongated cavities ( 14 ) and extends transversely to said rib ( 10 );
said method comprising the steps of:
a) providing at least two elongated first support tools ( 21 ), each having a cross section with a polygonal external profile complementary to the profile of the first cavities ( 14 ) to be formed in said structure ( 1 );
b) placing laterally said first support tools ( 21 ) side by side so as to interpose between them one or more layers of said prepreg material destined to form said rib ( 10 );
c) providing at least two second elongated support tools ( 15 ), each having a cross section with a polygonal external profile complementary to the profile of the second cavities ( 8 ) to be formed in said structure ( 1 );
d) placing laterally said second support tools ( 15 ) side by side so as to interpose between them one or more layers of said prepreg material destined to form said spar member ( 13 );
e) forming at least three distinct skins ( 11 a , 12 a ) by laminating, for each of them, one or more layers of said prepreg material;
f) inserting between a first and a second of said skins ( 11 a , 12 a ) the assembly formed by said second support tools ( 15 ) and by said prepreg material interposed between them so as to form said one ( 6 ) of said first and second wall ( 6 , 7 );
g) inserting between said second skin ( 11 a ) and a third ( 11 a ) of said skins ( 11 a , 12 a ) the assembly formed by said first support tools ( 21 ) and said prepreg material interposed between them;
h) bringing the group formed in steps a) to g) into an autoclave at a predefined curing temperature and pressure;
wherein:
said first and second support tools ( 21 , 15 ) are internally hollow and have a composition such as to assume a rigid state at room temperature;
said first support tools ( 21 ) have prevailing directions of extension transversal to the prevailing directions of extension of said second support tools ( 15 ); and
steps b) and d) are carried out by directly laminating said prepreg material on the external surface of the walls ( 25 , 16 ) of said first and second support tools ( 21 , 15 ) in the rigid state.

2. Method according to claim 1 , wherein said first and second support tools ( 21 , 15 ) have a composition based on reinforcing material and polymer suitable to allow the passage from said rigid state to a flexible elastomeric state and vice versa in response to heating and cooling respectively; wherein said first and second support tools ( 21 , 15 ) are configured to assume said flexible elastomeric state at a temperature below the curing temperature and above 50° C.; and wherein, during step h), the curing pressure is applied both inside the autoclave and inside said first and second support tools ( 21 , 15 ), whose walls ( 25 , 16 ) are made flexible by the passage from the rigid state to the flexible elastomeric state and are therefore pushed by the curing pressure itself to adhere to said adjacent prepreg material.

3. Method according to claim 2 , wherein the polymer of said first and second support tools ( 21 , 15 ) is a shape memory thermosetting or thermoplastic polymer, and wherein the reinforcing material of said first and second support tools ( 21 , 15 ) includes one or more elastic fibers.

4. Method according to claim 1 , wherein said first and/or second support tools ( 21 , 15 ) are made of a low-melting thermoplastic material, in particular polystyrene, and are externally coated with tubular bags ( 27 ) in such a way that, during step h), said first and/or second support tools ( 21 , 15 ) dissolve and the curing pressure is applied through said tubular bags ( 27 ).

5. Method according to claim 1 , wherein said first and/or second support tools ( 21 , 15 ) are made of a material soluble in a given solvent, for example water.

6. Method according to claim 1 , wherein also said other ( 7 , 6 ) of said first and second walls ( 6 , 7 ) of said structure ( 1 ) comprises the same sandwich configuration as said one of said first and second walls ( 6 , 7 ), said method comprising the further steps of:
i) forming a fourth skin ( 12 a ) by laminating one or more layers of said prepreg material; and
1) inserting between the said third and fourth skin ( 11 a , 12 a ) a further assembly formed by said second support tools ( 15 ) and by said prepreg material interposed between them so as to form in a sandwich configuration also said another ( 7 , 6 ) of said first and second wall ( 6 , 7 ).

7. Method according to claim 1 , wherein the distance between said first and second panels ( 11 , 12 ) of said one ( 6 , 7 ) of said first and second wall ( 6 , 7 ) is smaller than the distance of said first panel ( 11 ) from said another ( 7 , 6 ) of said first and second wall ( 6 , 7 ).

8. Method according to claim 1 , wherein the distance between said first and second panel ( 11 , 12 ) of each of said first and second wall ( 6 , 7 ) is smaller than the distance between said first panels ( 11 ) facing each other.

9. Method according to claim 1 , wherein said rib ( 10 ) is a profiled element with an open or closed cross section.

10. Method according to claim 1 , wherein said spar member ( 13 ) has an open or closed cross section.

Assignees

• Leonardo Spa

Inventors

• Corvaglia Stefano Giuseppe
• Gallo Nicola
• Baroni Antonio
• Ricciardi Alfredo

Classifications

• B64C3/20

  Integral or sandwich constructions · CPC title

• B29L2031/3085

  Wings · CPC title

• B29K2105/089

  fabric · CPC title

• B29K2025/06

  PS, i.e. polystyrene · CPC title

• B29C70/682

  Preformed parts characterised by their structure, e.g. form · CPC title