Method for manufacturing a profile section of a leading edge of an aircraft lifting surface; the profile section comprising a C-shaped skin, one or more auxiliary spars (33, 33', 33", 33"') and one or more longitudinal reinforcing elements (35, 37, 37') in a spanwise direction, the method comprising the following steps: a) providing: - a first laminated preform of a composite material configured with the shape foreseen for the C-shaped skin; - one or more second laminated preforms of a composite material configured with a web and two flanges (47, 49) to be joined to the first laminated preform for constituting said one or more auxiliary spars (33, 33', 33", 33"') or a part of them; - one or more pairs (51, 61, 71, 81) of third laminated preforms of a composite material comprising one or more sides or webs (53, 55, 63, 65, 73, 83) to be joined to the first laminated preform or to the second laminated preform and two flanges or feet (57, 59, 67, 69, 75, 85) for constituting the webs of said longitudinal reinforcing elements (35, 37, 37'); b) arranging said laminated preforms in a curing tooling and subjecting the assembly to an autoclave cycle to co-cure said laminated preforms; c) demoulding the curing tooling in a spanwise direction towards the root of the leading edge; characterized in that the leading edge profile section comprises a C-shaped skin, one auxiliary spar with a longitudinal stiffener and two longitudinal stringers (37, 37') reinforcing the C-shaped skin in the upper and lower zones located between the auxiliary spar and a front spar of a torsion box of the aircraft lifting surface, when the leading edge is attached to the torsion box, and wherein in step a): - a first pair of the third laminated preforms (51, 61) is configured with a first side (53, 63) to be joined to the first preform, a second side (55, 65) to be joined to the second preform, a first flange (57, 67) for constituting said longitudinal stiffener and a second flange (59, 69) for constituting a part of said two longitudinal stringers (37; 37'); - a second pair of the third laminated preforms (71, 81) is configured with webs (73, 83) to be joined to the first laminated preform and feet (75, 85) for constituting a part of said two stringers (37, 37'). a) providing: - a first laminated preform of a composite material configured with the shape foreseen for the C-shaped skin; - one or more second laminated preforms of a composite material configured with a web and two flanges (47, 49) to be joined to the first laminated preform for constituting said one or more auxiliary spars (33, 33', 33", 33"') or a part of them; - one or more pairs (51, 61, 71, 81) of third laminated preforms of a composite material comprising one or more sides or webs (53, 55, 63, 65, 73, 83) to be joined to the first laminated preform or to the second laminated preform and two flanges or feet (57, 59, 67, 69, 75, 85) for constituting the webs of said longitudinal reinforcing elements (35, 37, 37'); - a first laminated preform of a composite material configured with the shape foreseen for the C-shaped skin; - one or more second laminated preforms of a composite material configured with a web and two flanges (47, 49) to be joined to the first laminated preform for constituting said one or more auxiliary spars (33, 33', 33", 33"') or a part of them; - one or more pairs (51, 61, 71, 81) of third laminated preforms of a composite material comprising one or more sides or webs (53, 55, 63, 65, 73, 83) to be joined to the first laminated preform or to the second laminated preform and two flanges or feet (57, 59, 67, 69, 75, 85) for constituting the webs of said longitudinal reinforcing elements (35, 37, 37'); b) arranging said laminated preforms in a curing tooling and subjecting the assembly to an autoclave cycle to co-cure said laminated preforms; c) demoulding the curing tooling in a spanwise direction towards the root of the leading edge; characterized in that the leading edge profile section comprises a C-shaped skin, one auxiliary spar with a longitudinal stiffener and two longitudinal stringers (37, 37') reinforcing the C-shaped skin in the upper and lower zones located between the auxiliary spar and a front spar of a torsion box of the aircraft lifting surface, when the leading edge is attached to the torsion box, and wherein in step a): - a first pair of the third laminated preforms (51, 61) is configured with a first side (53, 63) to be joined to the first preform, a second side (55, 65) to be joined to the second preform, a first flange (57, 67) for constituting said longitudinal stiffener and a second flange (59, 69) for constituting a part of said two longitudinal stringers (37; 37'); - a second pair of the third laminated preforms (71, 81) is configured with webs (73, 83) to be joined to the first laminated preform and feet (75, 85) for constituting a part of said two stringers (37, 37'). - a first pair of the third laminated preforms (51, 61) is configured with a first side (53, 63) to be joined to the first preform, a second side (55, 65) to be joined to the second preform, a first flange (57, 67) for constituting said longitudinal stiffener and a second flange (59, 69) for constituting a part of said two longitudinal stringers (37; 37'); - a second pair of the third laminated preforms (71, 81) is configured with webs (73, 83) to be joined to the first laminated preform and feet (75, 85) for constituting a part of said two stringers (37, 37').