Aeronautical propulsion system comprising an optimized fan section
A fan section of an aeronautical propulsion system includes a fan rotor including seventeen blades to twenty blades and having a solidity strictly less than 1.0. The solidity is equal to a ratio between a chord at the blade tip and an inter-blade pitch at the blade tip. The fan section has a hub-tip ratio greater than or equal to 0.22 and less than or equal to 0.32, a pressure ratio greater than or equal to 1.05 and less than or equal to 1.5, and a peripheral speed at the blade tip greater than or equal to 260 m/s and less than or equal to 400 m/s. The pressure ratio and the peripheral speed are measured at cruising speed.
1 . A propulsion system comprising:
a drive shaft movable in rotation about an axis of rotation;
a fan shaft;
a fan section comprising a fan rotor driven in rotation by the fan shaft; and
a reduction structure coupling the drive shaft and the fan shaft in order to drive the fan shaft at a rotational speed lower than the rotational speed of the drive shaft;
wherein the fan rotor comprises at least seventeen blades and at most twenty blades and has a solidity strictly less than 1.0, where the solidity is equal to a ratio between a chord at the blade tip and an inter-blade pitch at the blade tip, the fan section having a hub-tip ratio greater than or equal to 0.22 and less than or equal to 0.32, a pressure ratio greater than or equal to 1.05 and less than or equal to 1.5 and a peripheral speed at the blade tip greater than or equal to 260 m/s and less than or equal to 400 m/s, the pressure ratio and the peripheral speed being measured at cruising speed,
wherein a thrust density corresponding to each fan rotor blade of the fan section is greater than or equal to 0.5×10 4 and less than or equal to 3.0×10 4 N/m 2 , where the thrust density per blade is defined by the following formula:
Thrust
density
=
F
N
n
*
D
2
*
1
0
0
where:
FN is a thrust generated by the propulsion system and is measured when the propulsion system is stationary at cruising speed in a standard atmosphere and is expressed in Newton;
n is the number of blades in the fan rotor; and
D is the diameter of the fan rotor, measured in a plane normal to the axis of rotation at an intersection between the top and a leading edge of the fan rotor blades and is expressed in meters, the diameter being greater than or equal to 2.159 m and less than or equal to 3.048 m; and
wherein the thrust, the number of blades, and the diameter are selected such that the thrust generated by the propulsion system is comprised between 18,000 lbf and 51,000 lbf when the propulsion system is stationary, uninstalled, in take-off rating in a standard atmosphere.
2 . The propulsion system according to claim 1 , wherein the solidity of the fan rotor is also strictly greater than 0.6.
3 . The propulsion system according to claim 1 , wherein the peripheral speed at the tip of the blades is greater than or equal to 270 m/s and less than or equal to 380 m/s.
4 . The propulsion system according to claim 1 , wherein the hub-tip ratio is greater than or equal to 0.235 and less than or equal to 0.30.
5 . The propulsion system according to claim 1 housed in a fan casing, the blades of the fan rotor being fixed in rotation relative to a hub of the fan rotor so as to have a fixed pitch angle.
6 . The propulsion system according to claim 1 , further comprising a fan stator comprising at least 38 stator vanes and at most 48 stator vanes.
7 . The propulsion system according to claim 1 , wherein a bypass ratio of the propulsion system is comprised between 10 and 35 inclusive.
8 . The propulsion system according to claim 1 , wherein a bypass ratio of the propulsion system is comprised between 10 and 18 inclusive.
9 . The propulsion system according to claim 1 , wherein a drive turbine comprises at least three and at most five stages.
10 . The propulsion system according to claim 1 , wherein a compressor comprises at least two and at most four stages.
11 . The propulsion system according to claim 1 , further comprising a high-pressure turbine and a high-pressure compressor connected via a high-pressure shaft, the high-pressure shaft rotating more quickly than the drive shaft, the high-pressure turbine being a two-stage turbine.
12 . The propulsion system according to claim 11 , wherein the high-pressure compressor comprises at least eight and at most eleven stages.
13 . The propulsion system according to claim 1 , wherein the fan blades are made of a composite material comprising a fibrous reinforcement embedded in a polymer matrix.
14 . An aircraft comprising:
at least one of the propulsion system according to claim 1 ; and
a pylon,
wherein the at least one of the propulsion system is fixed to the aircraft via the pylon.
15 . A method of operating a fan section of a propulsion system, the fan section comprising a fan rotor comprising at least seventeen blades and at most twenty blades and having a solidity strictly less than 1.0 and optionally strictly greater than 0.6, where the solidity is equal to a ratio between a chord at the blade tip and an inter-blade pitch at the blade tip, the fan section having a hub-tip ratio greater than or equal to 0.22 and less than or equal to 0.32, a pressure ratio greater than or equal to 1.05 and less than or equal to 1.5, and a peripheral speed at the blade tip greater than or equal to 260 m/s and less than or equal to 400 m/s, the pressure ratio and the peripheral speed being measured at cruising speed,
wherein a thrust density corresponding to each fan rotor blade of the fan section is greater than or equal to 0.5×10 4 and less than or equal to 3.0×10 4 N/m 2 , where the thrust density per blade is defined by the following formula:
Thrust
density
=
F
N
n
*
D
2
*
1
0
0
and where:
FN is a thrust generated by the propulsion system and is measured when the propulsion system is stationary at cruising speed in a standard atmosphere and is expressed in Newton;
n is the number of blades in the fan rotor; and
D is the diameter of the fan rotor, measured in a plane normal to the axis of rotation at an intersection between the top and a leading edge of the fan rotor blades and is expressed in meters, the diameter being greater than or equal to 2.159 m and less than or equal to 3.048 m; and
wherein the method comprises:
selecting the thrust, the number of blades, and the diameter such that thrust generated by the propulsion system is comprised between 18,000 lbf and 51,000 lbf when the propulsion system is stationary, uninstalled, in take-off rating in a standard atmosphere, and
operating the fan section of the propulsion system.
16 . The method according to claim 15 , wherein the fan blades are made of a composite material comprising a fibrous reinforcement embedded in a polymer matrix.