Aircraft turbine engine with a hybrid compressor

1 . An aircraft turbine engine comprising at least one combustion chamber, a high-pressure compressor and a low-pressure body, said high-pressure compressor upstream of said at least one combustion chamber, said low-pressure body comprising an axial low-pressure compressor; upstream of said high-pressure compressor, said high-pressure compressor configured to deliver air to said at least one combustion chamber and, downstream of said at least one combustion chamber, a turbine receiving hot gases from said at least one combustion chamber and configured to drive a rotor of said axial low-pressure compressor, said axial low-pressure compressor comprising a plurality of compression stages each having a ring of rotor blades extending radially at a periphery of said rotor of the axial low-pressure compressor and a respective ring of stator vanes secured to a casing,

wherein a first compression stage of the plurality of compression stages is a fan stage comprising a ring of fan rotor blades and a respective first ring of stator vanes, and downstream of the fan stage are booster stages of the plurality of compression stages, each of the booster stages comprising a respective ring of booster rotor blades and a respective ring of booster stator vanes,

wherein said aircraft turbine engine further, comprises at least one electric machine located downstream of the last of the booster stages of said axial low-pressure compressor and upstream of said high-pressure compressor, the at least one electric machine is coaxial or parallel to said low-pressure body, a stator of the at least one electric machine is secured to said casing and a rotor of the at least one electric machine drives in rotation an additional ring of rotor blades downstream of the last of the booster stages and upstream of the high-pressure compressor, the additional ring of rotor blades configured to generate a flow of air when the additional ring of rotor blades is rotated, a respective additional ring of stator vanes secured to the casing being located downstream of and adjacent to the additional ring of rotor blades, and said rotor of the at least one electric machine being furthermore guided in rotation by at least one bearing secured to said casing so that a speed of rotation of said rotor of the at least one electric machine is independent of a speed of rotation of the rotor of the axial low-pressure compressor and independent of a speed of rotation of the high-pressure compressor, and

wherein the rotor of the at least one electric machine rotates in the same direction of rotation as the rotor of said axial low-pressure compressor.

2 . The aircraft turbine engine according to claim 1 , said aircraft turbine engine being of a dual-body type, and wherein a high-pressure body comprises the high-pressure compressor, the high-pressure compressor is located downstream of the axial low-pressure compressor, of the additional ring of rotor blades and the respective additional ring of stator vanes, and of the at least one electric machine and upstream of the at least one combustion chamber, said aircraft turbine engine further being double-flow with the low-pressure body and the high-pressure body located in a flow duct of a primary flow.

3 . The aircraft turbine engine according to claim 1 , wherein the at least one electric machine is of an annular type and comprises a hollow shaft configured to drive the additional ring of rotor blades in rotation.

4 . The aircraft turbine engine according to claim 1 , comprising a plurality of electric machines configured to rotate, by means of gears, the additional ring of rotor blades.

5 . The aircraft turbine engine of claim 1 , wherein the at least one electric machine is configured to rotate the rotor of the at least one electric machine in two directions of rotation.

6 . The aircraft turbine engine of claim 1 , wherein the at least one bearing is integrated to the at least one electric machine.

7 . A method for controlling an electric machine of an aircraft turbine engine according to claim 1 , said method comprising the following steps, executed by a control unit:

a) receiving characteristic data of an operating point of the aircraft turbine engine at a given time;

b) determining, from the received data, a target operating regime of the at least one electric machine;

c) determining, from the target operating regime, a target power of the at least one electric machine;

d) comparing instantaneous power of the at least one electric machine and the target power and, if a difference between the instantaneous power and the target power is less than a determined threshold, returning to the step b),

otherwise,

e) changing the target operating regime of the at least one electric machine intended to achieve the target power; and,

f) determining the instantaneous power of the at least one electric machine and returning to the step d).

8 . The method according to claim 7 , said aircraft turbine engine being of a dual-body type, and wherein a high-pressure body comprises the high-pressure compressor, the high-pressure compressor is located downstream of the axial low-pressure compressor, of the additional ring of rotor blades and the respective additional ring of stator vanes, and of the at least one electric machine and upstream of the at least one combustion chamber, said aircraft turbine engine further being double-flow with the low-pressure body and the high-pressure body located in a flow duct of a primary flow wherein the characteristic data of an operating point of the aircraft turbine engine at a given time comprises, at least one of the following data:

an operating regime of the axial low-pressure compressor of the aircraft turbine engine;

an operating regime of the high-pressure compressor of the aircraft turbine engine;

a pressure measured at an inlet of the high-pressure compressor of the aircraft turbine engine;

a pressure measured at an inlet of the axial low-pressure compressor of the aircraft turbine engine;

a temperature measured at the inlet of the axial low-pressure compressor of the aircraft turbine engine; and,

a temperature measured at the inlet of the high-pressure compressor of the aircraft turbine engine.