Abstract: The centrifugal turbo-compressor (2) includes a hermetic casing (3); a drive shaft (6) having a longitudinal axis and rotatably arranged within the hermetic casing (3); a compression stage including an impeller (17) connected to the drive shaft (6); a gas suction inlet (42); and a gas flow path (P) fluidly connected to the gas suction inlet (42) and configured to supply the compression stage with a gas flow. The gas flow path (P) includes a relaxation chamber (46) at least partially surrounding the drive shaft (6), the gas suction inlet (42) emerging substantially radially into the relaxation chamber (46); and a plurality of inlet flow guide channels (51) fluidly connected to the relaxation chamber (46) and angularly distributed around the longitudinal axis of the drive shaft (6), the inlet flow guide channels (51) extending radially towards the drive shaft (6) and being axially offset from the gas suction inlet (42) and the relaxation chamber (46).

Abstract: The centrifugal turbo-compressor (2) includes a hermetic casing (3); a drive shaft (6) having a longitudinal axis and rotatably arranged within the hermetic casing (3); a compression stage including an impeller (17) connected to the drive shaft (6); a gas suction inlet (42); and a gas flow path (P) fluidly connected to the gas suction inlet (42) and configured to supply the compression stage with a gas flow. The gas flow path (P) includes a relaxation chamber (46) at least partially surrounding the drive shaft (6), the gas suction inlet (42) emerging substantially radially into the relaxation chamber (46); and a plurality of inlet flow guide channels (51) fluidly connected to the relaxation chamber (46) and angularly distributed around the longitudinal axis of the drive shaft (6), the inlet flow guide channels (51) extending radially towards the drive shaft (6) and being axially offset from the gas suction inlet (42) and the relaxation chamber (46).

Abstract: A variable-pitch vane including a plurality of propeller blades (2). Each blade has a variable pitch according to a blade rotational axis (A1) and a root (201). A plurality of rotor connecting shafts (6), each shaft having a foot (602) and a head (601). The root (201) of each blade is mounted on the head (601) of a rotor connecting shaft via a pivot (8) in such a way as to allow each blade (2) to rotate according to the blade rotational axis (A1), in which each blade (2) has a blade pitch, such that the blade rotational axis (A1) is inclined relative to a radial axis (A2) passing through the foot (602) of the corresponding shaft.

Abstract: A method for modeling at least a part of a blade of a non-streamlined propeller, part of the blade having an offset. The method includes (a) Parameterization of at least one Bezier curve representing a deformation of the blade characterizing the offset, defined by: a. First and second end control points (PCU1, PCUK); b. At least one intermediate control point (PCUi, i?[[2,K?1]]) disposed between the end points (PCU1, PCUK). The parameterization being performed according to at least one deformation parameter and the cutting height in the blade, on the basis of which the abscissa of the intermediate control point (PCUi) and the ordinate of the second end point (PCUK) are expressed. Optimized values of the deformation parameter or parameters are determined and then output.

Abstract: An aircraft turbopropeller includes two coaxial contra-rotating unducted propellers, upstream and downstream respectively, each propeller including an annular row of blades. The blades of the downstream propeller have a reverse sweep in relation to that of the blades of the upstream propeller. The leading edges of the blades of the downstream propeller extend radially outwards from downstream to upstream, at least over a radially outer portion of the blades.

Abstract: The present invention concerns a method for modelling at least a part of a blade (2) of a non-streamlined propeller (1), the part of the blade (2) having an offset (3), the method being characterised in that it comprises implementing, using data processing means (11) of a piece of equipment (10), steps of: (a) Parameterisation of at least one Bézier curve representing a deformation of said blade (2) characterising the offset (3), defined by: a. First and second end control points (PCU1, PCUK); b.

Abstract: An aircraft turbopropeller includes two coaxial contra-rotating unducted propellers, upstream and downstream respectively, each propeller including an annular row of blades. The blades of the downstream propeller have a reverse sweep in relation to that of the blades of the upstream propeller. The leading edges of the blades of the downstream propeller extend radially outwards from downstream to upstream, at least over a radially outer portion of the blades.