Aircraft turbomachine with a hybrid compressor

Abstract

The invention relates to an aircraft turbine engine (101) comprising an electric motor (206) downstream of a low-pressure compressor of a low-pressure body (104) of the aircraft turbine engine and upstream of a combustion chamber (102) of the aircraft turbine engine, the electric motor being configured to rotate a rotor blade ring to generate an air flow. Furthermore, the rotational speed of the electric motor rotor is independent of the rotational speed of the compressor rotor.

Description

Technical Field

[0001] The present invention relates to the general field of aircraft turbine engines, in particular to turbine engines of the turbojet type with a dual body and a dual flow. In particular, the invention relates to an aircraft turbine engine comprising, downstream of a low-pressure compressor of the low-pressure body of the aircraft turbine engine, an electric motor coaxial with the low-pressure body, the electric motor being configured to drive in rotation a movable blade ring. Background Art

[0002] The prior art includes in particular documents US-A1-2019 / 145322, US-A1-2009 / 289456 and WO-A1-2014 / 364265.

[0003] In a conventional manner, an aircraft turbine engine of the turbojet type with a dual body (i.e. with a low-pressure body and a high-pressure body) comprises, from upstream to downstream in the direction of gas flow, a fan, a low-pressure compressor, a high-pressure compressor, a combustion chamber, a high-pressure turbine and a low-pressure turbine. The rotor of the low-pressure compressor is driven by the rotor of the low-pressure turbine, and the rotor of the high-pressure compressor is driven by the rotor of the high-pressure turbine.

[0004] In turbine engines of this type, the low-pressure compressor (also known as the booster) must be able to operate over a wide range of operating conditions (typically 25% to 105% of the nominal operating conditions of the low-pressure compressor). In particular, a challenge is to avoid, over the entire range of operating conditions, the known pumping phenomenon, which causes a poor flow of gases in the turbine engine and even a reversal of the direction of these gas flows.

[0005] This objective constrains the design of the low-pressure compressor. In practice, variable geometry is often integrated into the turbine engine or the exhaust vane arrangement, in order to ensure good compressor operation, including under low-pressure conditions. In the most common case, the turbine engine includes a bleed valve, located downstream of the low-pressure compressor in the direction of gas flow, and commonly referred to as a VBV (Variable Bleed Valve). In some cases, the turbine engine also includes an input directing wheel (RDE) with variable pitch.

[0006] In any case, the addition of such systems aims to better control the flow through the low-pressure compressor under all operating conditions. However, the integration of such systems reduces the quality of the aerodynamic ducting (due to the introduction of gaps, slots, and leaks), thereby also leading to a degradation in the performance of the turbine engine at its nominal operating point. Furthermore, the addition of variable geometry (VBV, RDE, etc.) compensates for some, but not all, of the low-pressure compressor's operating losses. In fact, the deflection of the variable geometry (i.e., its ability to alter the flow) is limited by integration constraints and results in a limited correction capability.

[0007] This problem remains unchanged for turbine engines equipped with a speed reducer, whose sole function is to impose a constant speed ratio between the turbine engine's fan and the low-pressure compressor. The speed range and operating conditions of the low-pressure compressor are always affected by the low-pressure turbine.

[0008] There are also so-called three-body turbine engines, in which the intermediate compressor has a rotor driven by a turbine that is independent of the fan and the high-pressure body. However, although the operating range of this compressor is reduced to between the operating range of the high-pressure compressor and the operating range of the low-pressure compressor, the operating range of this compressor is still wide and an integrated discharge valve is still required.

[0009] Patent FR 2842565 B1 describes a turbine engine with an electric motor integrated into its forward casing. Because the motor's rotor and the compressor's rotor are fixed, the motor can provide torque to the low-pressure compressor's rotor. This power injection into the low-pressure compressor's rotor optimizes the low-pressure compressor's performance under certain operating conditions, but simultaneously degrades the high-pressure compressor's operability (i.e., the range of operating conditions under which the high-pressure compressor operates optimally). Consequently, this approach requires compromise.

[0010] Generally, currently known alternatives using variable geometry that rely on an electric motor to improve the performance of the low-pressure compressor improve the operability of the low-pressure compressor but degrade the operability of the fan and / or the high-pressure compressor. Summary of the Invention

[0011] The present invention proposes a solution based on the use of an electric motor making it possible to drive in rotation a movable ring of blades situated downstream of the low-pressure compressor in order to generate an air flow independently of the rotation speed of the compressor's rotor.

[0012] To this end, according to a first embodiment, the invention relates to an aircraft turbine engine comprising at least one combustion chamber and a body, in particular a low-pressure body, said body comprising, upstream of said chamber, an axial compressor intended to deliver air to said chamber, and, downstream of said chamber, a turbine, in particular an axial low-pressure compressor, which receives hot gases from said chamber and is intended to drive the rotor of said compressor, said compressor comprising at least one compression stage, at least one compression stage having a first stationary blade ring and a first movable blade ring, the first stationary blade ring being fixed to the casing, the first movable blade ring extending radially at the periphery of said rotor of the compressor,

[0013] The turbine engine is characterized in that it further includes at least one motor coaxial with or parallel to the main body downstream of the compressor and upstream of the chamber, the stator of the motor being fixed to the casing, the rotor of the motor driving the second movable blade ring to rotate, the second movable blade ring being configured to generate an air flow when the second movable blade ring is set to rotate, the rotor of the motor being further guided in rotation by at least one bearing fixed to the casing, so that the rotational speed of the rotor of the motor is independent of the rotational speed of the rotor of the compressor.

[0014] The device according to the invention may include one or more of the following features, taken alone or in combination with one another:

[0015] - The turbine engine is of the twin-body type, in which the high-pressure compressor of the high-pressure body is located downstream of the low-pressure compressor and the electric motor and upstream of the combustion chamber, and is also of the dual-flow type, in which the low-pressure body and the high-pressure body are located in the main flow duct.

[0016] The electric motor is of annular type and comprises a hollow shaft configured to drive the second movable blade ring in rotation.

[0017] The turbine engine comprises a plurality of electric motors configured to rotate the second movable blade ring via gears.

[0018] -The compressor comprises a plurality of compression stages, the electric motor is located downstream of the last stage of the compressor and upstream of the combustion chamber, the rotor of the electric motor rotates in the same direction of rotation as the rotor of the compressor, and the second fixed blade ring is located downstream of and adjacent to the second movable blade ring.

[0019] The compressor comprises a single compression stage constituted by a radial sector of a fan of a turbine engine, the electric motor being located downstream of said stage and the second stationary blade ring being located downstream of and adjacent to the second movable blade ring.

[0020] - the compressor comprises a plurality of compression stages, each comprising a movable blade ring, the electric motor being located downstream of and adjacent to the last movable blade ring of the compressor, the rotor of the electric motor rotating in a direction opposite to that of the rotor of the compressor.

[0021] The second fixed blade ring is located downstream of and adjacent to the second movable blade ring.

[0022] According to a second embodiment, the invention also relates to a method for controlling an electric machine of an aircraft turbine engine according to the first embodiment, said method comprising the following steps executed by a control unit:

[0023] a) receiving data characteristic of an operating point of a turbine engine at a given time;

[0024] b) determining target operating conditions of the motor from the received data;

[0025] c) determining the target power of the motor based on the target operating conditions;

[0026] d) comparing the instantaneous power of the motor with the determined target power (704), and if the difference between the instantaneous power and the target power is less than a determined threshold, returning to step b),

[0027] otherwise,

[0028] e) changing (705) the operating conditions of the motor to achieve the target power; and,

[0029] f) Determine (706) the instantaneous power of the motor and return to step d).

[0030] Furthermore, the characteristic data of the operating point of the turbine engine may include at least one of the following data:

[0031] - operating conditions of the low-pressure compressor of the turbine engine;

[0032] - operating conditions of the high-pressure compressor of the turbine engine;

[0033] - the pressure measured at the inlet of the high-pressure compressor of the turbine engine;

[0034] - the pressure measured at the inlet of the low-pressure compressor of the turbine engine;

[0035] - the temperature measured at the inlet of the low-pressure compressor of the turbine engine; and

[0036] - Temperature measured at the inlet of the high-pressure compressor of the turbine engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0037] The invention will be better understood and other details, features and advantages of the invention will become clearer through the following description made by way of non-limiting example and with reference to the accompanying drawings, in which:

[0038] [ Figure 1 ] Figure 1 is a simplified schematic depiction of an aircraft turbine engine according to an embodiment of the invention;

[0039] [ Figure 2 ] Figure 2 is a schematic half-section view of the front portion of an aircraft turbine engine according to a first embodiment of the invention;

[0040] [ Figure 3 ] Figure 3 is a schematic diagram of an embodiment of a motor according to the present invention;

[0041] [ Figure 4 ] Figure 4 is a schematic half-section view of the front portion of an aircraft turbine engine according to a second embodiment of the invention;

[0042] [ Figure 5 ] Figure 5 is a schematic half-section view of the front portion of an aircraft turbine engine according to a third embodiment of the invention;

[0043] [ Figure 6 ] Figure 6 is a schematic half-section view of the front portion of an aircraft turbine engine according to a fourth embodiment of the invention;

[0044] [ Figure 7 ] Figure 7 is a block diagram of the steps of a method for controlling an aircraft turbine engine according to an embodiment of the invention.

[0045] Elements having the same function in different exemplary embodiments are provided with the same reference numerals in the figures. DETAILED DESCRIPTION

[0046] Reference Figure 1 and Figure 2 , an aircraft turbine engine according to one embodiment of the present invention will now be described. Figure 1 and Figure 2In the figure, only one low-pressure body is shown. However, those skilled in the art will appreciate that the present invention is also applicable to a twin-body turbine engine, i.e., a turbine engine comprising a low-pressure body including a fan, a low-pressure compressor, and a low-pressure turbine, and a high-pressure body including a high-pressure compressor and a high-pressure turbine. Furthermore, such a turbine engine may be, for example, a twin-body, dual-flow turbojet engine (also known as a turbofan engine), in which the low-pressure body and the high-pressure body are located in the main flow duct.

[0047] In the following, the terms “upstream” and “downstream” are defined with reference to the direction of gas flow in the turbine engine.

[0048] An aircraft turbine engine 101 comprises a combustion chamber 102 and a low-pressure body 103. The low-pressure body 103 comprises, upstream of the chamber 102, an axial low-pressure compressor 104 intended to deliver air to the chamber, and, downstream of the chamber, a turbine 105 receiving the hot gases from the chamber and intended to drive a rotor 106 of the compressor.

[0049] exist Figure 2 In the illustrated example, the compressor includes four compression stages 201, 202, 203, and 204. Each of the four compression stages has a first stationary blade ring 201b, 202b, 203b, and 204b, and a first movable blade ring 202a, 203a, and 204a. The first stationary blade ring is fixed to a casing 205 and extends radially around the outer circumference of the compressor rotor. A fan 201a located upstream of the compressor drives the airflow straightened by the first stationary blade ring 201b. However, those skilled in the art will appreciate that the present invention is not limited to this number of stages and is generally applicable to aircraft turbine engines whose low-pressure compressors include at least one compression stage.

[0050] In any case, the turbine engine further comprises a motor 206 coaxial with the low-pressure body 103, downstream of the compressor and upstream of the chamber, the stator 207 of the motor being fixed to the housing 205, and the rotor 208 of the motor driving the movable blade ring 212 to rotate. The motor is configured to generate an air flow through the movable blade ring when the motor rotates. In addition, the rotor of the motor is guided in rotation by a bearing 209 fixed to the housing (via a shroud 211), so that the rotation speed of the rotor of the motor is independent of the rotation speed of the rotor of the compressor. In other words, the motor can be started, stopped, accelerated or decelerated regardless of the operating conditions of the low-pressure compressor. In addition, according to different embodiments of the turbine engine, the direction of rotation of the rotor of the motor can be the same as or opposite to the direction of rotation of the rotor of the low-pressure compressor.

[0051] exist Figure 2In the example shown in , the fixed blade ring 210 is located downstream of and adjacent to the second movable blade ring 212. These fixed blades straighten the flow generated by the movable blades. Therefore, the motor with the movable blade ring and the adjacent fixed blade ring can be regarded as an independent last stage of the low-pressure compressor. As will become clear later, the speed of the rotor of the motor can be controlled and adjusted according to selected parameters to optimize the aerodynamic operation of the assembly including the low-pressure compressor and the motor. Therefore, the control of the speed of the rotor of the motor enables greater flexibility in the use of the entire low-pressure compressor.

[0052] Furthermore, in the case of a twin-body turbine engine, the electric machine can be powered by a generator driven by the high-voltage body via power electronics.

[0053] Advantageously, the better control of the gas flow achieved by the integration of the electric motor thus makes it possible to better avoid pumping phenomena. Furthermore, the use of an electric motor also makes it possible, if necessary, to maintain an air flow during the so-called engine shutdown (hot) phase, for example in the main duct of a turbojet engine with dual flow. This flow can be used, for example, for maintenance functions related to cleaning the engine, to facilitate startup, to generate energy in the event of fan autorotation, or to cool the engine between two consecutive flights to prevent deformation of the rotor elements related to overheating (overheating imposes a waiting time before restarting).

[0054] Reference Figure 3 , embodiments of a motor according to the present invention will now be described.

[0055] Figure 3 The left side portion of FIG. 3 shows a ring-type motor 301 having a hollow shaft with a large inner diameter configured to drive a movable blade ring 302 to rotate.

[0056] Figure 3 The right part of the diagram depicts a plurality of motors 303 configured to rotate a movable blade ring 305 via gears 304. In this embodiment, the motors 303 are parallel to the low pressure body, ie they extend along axes parallel to the longitudinal extension axis of the body.

[0057] According to an embodiment of one or more electric machines, one or more bearings guide the rotor of each electric machine and may be integrated into the electric machine or positioned between the rotor of the electric machine and an element of the casing of the turbine engine.

[0058] Figure 4 、 Figure 5 as well as Figure 6 Different embodiments of an aircraft turbine engine according to the invention are shown.

[0059] In reference Figure 4 In the illustrated embodiment, the low-pressure compressor comprises only one compression stage 401, which comprises a radial segment of a turbine engine's fan 401a and a stationary blade ring 401b. An electric motor 404 is located downstream of this compression stage, and another stationary blade ring 403 is located downstream of and adjacent to the movable blade ring 402. Advantageously, this embodiment enables high compression ratios to be maintained in turbine engines whose low-pressure compressor comprises only one compression stage and in which the integration of systems such as variable bleed valves is complex or even impossible.

[0060] In reference Figure 5 In the illustrated embodiment, the low-pressure compressor includes a plurality of compression stages 501, 502, 503, and 504, each of which includes a movable blade ring. A motor 505 is located downstream of and adjacent to the last movable blade ring of the compressor (which constitutes the last compression stage of the low-pressure compressor), and the direction of rotation of the motor's rotor is opposite to that of the compressor's rotor. In this case, the motor is controlled so that the movable blade ring 506 minimizes the swirl of the flow at the outlet of the movable blade ring. This makes it possible to avoid the use of a straightener (stator) located downstream of the movable blade ring 506.

[0061] Reference Figure 6 The last embodiment shown uses the same Figure 5 The embodiment shown has a similar architecture, but in addition, the last embodiment comprises a fixed blade ring 607 located downstream and adjacent to the movable blade ring 606. This movable blade ring then straightens the flow.

[0062] Reference Figure 7 , we will now describe an embodiment of a method for controlling an electric machine of an aircraft turbine engine according to the invention. The steps of the method are performed by a control unit.

[0063] Step 701 includes receiving characteristic data of the operating point of the turbine engine at a given time. The characteristic data may, for example, include at least one of the following data:

[0064] - operating conditions of the low-pressure compressor of the turbine engine;

[0065] - operating conditions of the high-pressure compressor of the turbine engine;

[0066] - the pressure measured at the inlet of the high-pressure compressor of the turbine engine;

[0067] - the pressure measured at the inlet of the low-pressure compressor of the turbine engine;

[0068] - the temperature measured at the inlet of the low-pressure compressor of the turbine engine; and

[0069] - Temperature measured at the inlet of the high-pressure compressor of the turbine engine.

[0070] Furthermore, a person skilled in the art will know how to select data, for example measured by sensors or determined from the electric machine, suitable for enabling the aerodynamic performance of the low-pressure compressor to be optimized by controlling the electric machine.

[0071] Step 702 comprises determining target operating conditions for the electric machine based on the received data.The target operating conditions are operating conditions that enable optimization of the flow in the aerodynamic duct and avoidance of pumping.

[0072] Step 703 includes determining a target power of the motor from the target operating conditions. In other words, the target power is the power corresponding to the target operating conditions.

[0073] During a step 704 , the control unit compares the instantaneous power of the electric machine with the target power determined in step 703 .

[0074] If the difference between the instantaneous power and the target power is smaller than the given threshold, the method returns to step 702 .

[0075] Otherwise, execute step 705. This step includes modifying the operating conditions of the motor to achieve the target power.

[0076] Finally, step 706 includes determining the instantaneous power of the motor and returning to step 704 to again compare this power to the target power.

Claims

1. An aircraft turbine engine (101) comprising at least one combustion chamber (102) and a body (103), said body comprising, upstream of said combustion chamber, an axial compressor (104) intended to deliver air to said combustion chamber, and, downstream of said combustion chamber, a turbine (105) receiving hot gases from said combustion chamber and intended to drive the rotor of said axial compressor, said axial compressor comprising at least one compression stage, said at least one compression stage having a first fixed blade ring and a first movable blade ring, said first fixed blade ring being fixed to a casing (205), said first movable blade ring extending radially at the outer periphery of the rotor of said axial compressor, The aircraft turbine engine is characterized in that it further comprises at least one electric motor coaxial with or parallel to the main body, downstream of the axial compressor and upstream of the combustion chamber, the stator (207) of the electric motor being fixed to the housing, the rotor of the electric motor driving the second movable blade ring to rotate, the second movable blade ring being configured to generate an air flow when the second movable blade ring rotates, the rotor of the electric motor being further guided in rotation by at least one bearing (209) fixed to the housing, so that the rotation speed of the rotor of the electric motor is independent of the rotation speed of the rotor of the axial compressor, The axial compressor includes a plurality of compression stages, the motor is located downstream of the last compression stage of the axial compressor and upstream of the combustion chamber, the rotor of the motor rotates in the same rotation direction as the rotor of the axial compressor, and the second fixed blade ring (210) is located downstream of the second movable blade ring and adjacent to the second movable blade ring. 2 . The aircraft turbine engine according to claim 1 , wherein the main body is a low-pressure main body and the axial compressor is an axial low-pressure compressor.

3. An aircraft turbine engine according to claim 2, being of the twin-body type and wherein: The high-pressure compressor of the high-pressure body is located downstream of the axial low-pressure compressor and the electric machine and upstream of the combustion chamber, the aircraft turbine engine also being of the dual-flow type, wherein the low-pressure body and the high-pressure body are located in the main flow duct.

4. An aircraft turbine engine according to any one of claims 1 to 3, wherein The motor is of annular type and comprises a hollow shaft configured to drive the second movable blade ring in rotation.

5. The aircraft turbine engine according to any one of claims 1 to 3, comprising a plurality of electric motors configured to rotate the second movable blade ring via a gear (304).

6. A method for controlling an electric machine of an aircraft turbine engine according to any one of claims 1 to 5, comprising the following steps performed by a control unit: a) receiving ( 701 ) data characteristic of the operating point of the aircraft turbine engine at a given time; b) determining (702) target operating conditions of the motor from the received characteristic data; c) determining (703) a target power of the motor from the target operating condition; d) comparing the instantaneous power of the motor with the determined target power (704), and if the difference between the instantaneous power and the target power is less than a determined threshold, returning to step b), otherwise, e) changing (705) the operating conditions of the motor to achieve the target power; as well as, f) Determine (706) the instantaneous power of the motor and return to step d).

7. The method according to claim 6, wherein: The main body is a low-pressure main body, the axial compressor is an axial low-pressure compressor, the aircraft turbine engine is of a twin-body type, wherein the high-pressure compressor of the high-pressure main body is located downstream of the axial low-pressure compressor and the electric machine and upstream of the combustion chamber, the aircraft turbine engine is also of a dual-flow type, wherein the low-pressure main body and the high-pressure main body are located in a main flow duct, and the characteristic data of the operating point of the aircraft turbine engine include at least one of the following data: - operating conditions of the axial low-pressure compressor of the aircraft turbine engine; - operating conditions of the high-pressure compressor of the aircraft turbine engine; - the pressure measured at the inlet of the high-pressure compressor of the aircraft turbine engine; - the pressure measured at the inlet of the axial low-pressure compressor of the aircraft turbine engine; - the temperature measured at the inlet of the axial low-pressure compressor of the aircraft turbine engine; and - the temperature measured at the inlet of the high-pressure compressor of the aircraft turbine engine.

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