Abstract: The invention concerns a blade profile for aircraft rotor, comprising, between a leading edge (1A) and a trailing edge (1B), a suction face (2) and a blade face (3) whereof the geometric locus of the equidistant points of the latter defines the curvature. The invention is characterised in that the value of the ratio of the maximum curvature to the maximum thickness develops linearly with the profile (1) relative thickness and ranges between 0.13 and 0.19 for a relative thickness of 7% of the chord (C) and ranges between 0.18 and 0.24 for a relative thickness of 15% of the chord (C).

Abstract: This invention relates to a section for shrouded propeller blade, of which the relative thickness is included between 9% and 15%, wherein, the curvature of said section having a maximum value at the leading edge thereof, the upper surface line comprises three successive portions, such thatin the first portion, the curvature decreases rapidly from said maximum value at the leading edge to a value close to 20 at a first point of which the reduced abscissa is close to 1%;in the second portion, the curvature decreases more slowly from said value close to 20 at said first point up to a value close to 1 at a second point, of which the reduced abscissa is close to 25%; andin the third portion, the curvature is less than 1 between said second point and said trailing edge.

Abstract: The external part at least of a propeller blade, having a maximum relative thickness between about 2 and 6%, has a profile whose thickness variation law along the chord comprises a first zone (I) going from the leading edge to about 8% of the length of the chord in which the curvature decreases rapidly, from a maximum value at the leading edge, followed by a second zone (II) extending between about 8% of the chord and midships in which the curvature decreases linearly. In a third zone, extending from midships to 75 to 85% of the chord, the curvature is substantially constant. A fourth zone (IV), extending from the end of the third zone (typically at 80% of the chord) to the trailing edge, comprises a curvature inversion.

Abstract: The present invention relates to a blade section for rotating wings of an aircraft comprising, between the leading edge and the trailing edge, a convex lower surface and an upper surface which is convex over the greater part of its length, but concave near the trailing edge. According to the invention, the upper surface of the section comprises a region of small extent, close to the leading edge but not contiguous thereto, in which the curvature is substantially constant. The invention is applicable to the improvement of the performances in maneuvering flight, while hovering and in advance flight of the aircraft.

Abstract: The present invention relates to a blade profile for rotary wing of an aircraft, comprising, between the leading edge and the trailing edge, a convex upper surface and a non-concave lower surface, wherein, at the point of maximum curvature of the leading edge, the radius of curvature Ro is approximately defined by the expression Ro=1.7 C. e.sub.max.sup.2, in which C represents the chord and e.sub.max the relative thickness of said profile and the portion of lower surface adjacent the leading edge extends to about 20% of the chord from said leading edge and presents a very small general curvature decreasing continuously in the direction of the trailing edge up to the point of minimum curvature of the lower surface. The invention is more particularly applicable to obtaining rotary wings, for example for helicopters, capable of very high performances both when hovering, and in high speed flight, as well as during manoeuvres.

Abstract: An axial supersonic compressor comprises a casing and a hub rotating in the asing and carrying blades. The suction surface of each blade is formed with a zone in which the curvative changes and which corresponds to a supersonic-subsonic shock wave. A channel formed in each blade and opening in said zone is connected to boundary layer aspiration means.