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

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 (4a), connected to a fuselage (3) of said aircraft (1), to a free end portion (4b) 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 (11a, 12a) 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 (11a, 12a) 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 (11a) and a third (11a) of said skins (11a, 12a) 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.