Turbine engine comprising an auxiliary oil supply circuit

Abstract

The invention relates to a turbine engine (1, 1', 1'') comprising: - a hydraulic control system (13) for controlling the pitch angle of the blades (2a); - a main hydraulic circuit (15) for supplying oil to the control system (13); - an auxiliary hydraulic circuit (16) for supplying oil to the control system (13); and - a system for controlling the oil supply from the auxiliary circuit (16), which system comprises: - a sensor (24) configured to acquire a signal (S1) representative of the pitch angle of the blades and transmit this signal (S1); and - a control unit (25) configured to transmit a control command (O1) to authorise the auxiliary circuit (16) to supply oil to the control system (13) if the signal (S1) transmitted by the sensor (24) to the control unit (25) corresponds to a decrease in a hydraulic pressure in the main circuit (15).

Description

[0001] DESCRIPTION

[0002] TITLE: TURBOMACHINE COMPRISING AN AUXILIARY OIL SUPPLY CIRCUIT

[0003] Technical field of the invention

[0004] The invention relates to the field of aircraft turbomachinery comprising variable pitch angle blades, a hydraulic control system for these blades and an auxiliary hydraulic circuit for supplying oil to this control system.

[0005] Technical background

[0006] An aircraft turbomachine comprises, from upstream to downstream, at least one first rotor, also called the propulsion rotor, such as a propeller when the turbomachine is a turboprop, or an unshod fan when the turbomachine is of the "open rotor" type, or a shod fan when the turbomachine is a turbojet, a compressor, a combustion chamber, and a turbine. The compressor rotor is connected to the turbine rotor and the first rotor by a drive shaft.

[0007] An airflow is compressed within the compressor, then the compressed air is mixed with fuel and burned in the combustion chamber. The gases produced by the combustion pass through the turbine, which drives the compressor rotor and the propulsion rotor.

[0008] The propeller or fan of the propulsion rotor, as well as the compressor rotor, are equipped with blades that influence the airflow. To adapt the turbomachine to flight conditions, particularly airflow, it is common practice to equip the propulsion rotor or compressor rotor with variable-pitch blades. These variable-pitch blades rotate around their axis of extension, allowing the blade pitch angle to be adjusted according to flight conditions to maximize thrust.

[0009] To modify this pitch angle, the turbomachine typically includes a variable pitch angle blade control system which includes a control unit connected to a hydraulic actuator to rotate the blades around their extension axis according to the direction of the airflow.

[0010] To supply oil to the control system, and in particular the hydraulic actuator, the turbomachine typically includes a control system oil supply system. This supply system includes, for example, a main hydraulic circuit comprising a main reservoir connected to a main pump, allowing oil to be drawn from the main reservoir and circulated to the hydraulic actuator.

[0011] Certain aircraft flight phases disrupt the oil supply to the hydraulic actuator. Specifically, the aircraft may experience flight phases during which the gravitational force is zero or negative. These flight phases are referred to in the field of this invention as the "Og condition" when the gravitational force is zero, or the "negative g condition" when the gravitational force is reversed. During such flight phases, the oil in the main reservoir, under negative g conditions, is pressed against the upper wall of the reservoir opposite the suction port. This suction port, located at the bottom of the reservoir, is connected to a main oil supply pump for the pitch control system. Under Og conditions, the oil and air form a suspension containing air bubbles, and the pump therefore draws air or oil heavily laden with air bubbles, rather than oil.Such a negative g or 0g condition degrades the oil supply to the control system and can even cause the main pump to lose prime. In all cases, the hydraulic actuator of the control system is no longer properly supplied with oil. Such degradation of the oil supply to the control system, and particularly to the hydraulic actuator, can make the pitching of the propulsion rotor blades uncontrollable, especially the propeller or unshod fan blades. This can lead to the blades being feathered by a safety system or locked by the safety system in a position different from the setpoint provided by the control unit to the hydraulic actuator. This results in a significant reduction in the turbomachine's thrust and, in some cases, can lead to a loss of control, which is unacceptable.

[0012] In this context, document FR-A1-3127525 proposes an auxiliary oil supply circuit for the control system to supply it when the gravitational force is zero or negative. This auxiliary circuit includes an auxiliary oil reservoir connected to an auxiliary pump that delivers oil to the control system via a hydraulic valve.

[0013] According to this document, the hydraulic valve is operated either directly by gravitational force or by the hydraulic pressure in the main supply circuit. Typically, when this pressure is below a threshold pressure, the hydraulic valve fluidly connects the control system to the auxiliary circuit.

[0014] While this solution has the advantage of ensuring a continuous power supply to the control system, independent of gravitational forces, it is not entirely satisfactory. Indeed, a valve controlled by gravitational force requires equipping the aircraft with a gravitational force sensor. However, such a solution makes the turbomachine dependent on aircraft data, particularly that provided by such a sensor. Furthermore, a hydraulic actuation system is complex to implement and can be unreliable.

[0015] Therefore, there is a need to provide an aircraft turbomachine that ensures oil supply to the variable pitch blade control system during flight phases in which the gravitational force is zero or negative in a simple and reliable manner.

[0016] Summary of the invention

[0017] To this end, the invention proposes a turbomachine for an aircraft, the turbomachine comprising:

[0018] - blades with variable pitch angle,

[0019] - a hydraulic control system for the blade pitch angle,

[0020] - an oil supply system for the control system comprising

[0021] - a main hydraulic circuit for supplying oil to the control system under conditions of positive gravitational force experienced by the aircraft, and

[0022] - an auxiliary hydraulic circuit for supplying oil to the control system under conditions of negative or zero gravitational force experienced by the aircraft, the auxiliary circuit comprising:

[0023] - an auxiliary tank, and

[0024] - an auxiliary pump connected to the auxiliary tank and the control system, and

[0025] - a system for controlling the oil supply to the auxiliary circuit.

[0026] The turbomachine is remarkable in that its control system includes:

[0027] - a sensor configured to acquire a signal representative of the blade pitch angle and transmit this signal, and

[0028] - A control unit configured to transmit a command to allow oil to be supplied to the control system via the auxiliary circuit if the signal transmitted by the sensor to the control unit corresponds to a decrease in hydraulic pressure in the main circuit. The auxiliary circuit allows the control system to be supplied from the auxiliary reservoir when the gravitational force is zero (0g condition) or negative (negative g condition).

[0029] This allows the control system to be continuously powered, regardless of the gravitational force experienced by the aircraft, and therefore limits situations of feathering or locking of the blade pitch for safety.

[0030] According to the invention, the auxiliary circuit is pressurized hydraulically by means of the control system.

[0031] The sensor acquires and transmits the signal relating to the blade pitch angle to the control unit. Depending on the control unit's interpretation of this signal, particularly when the signal indicates a pressure drop in the main circuit indicative of zero or negative gravitational force acting on the aircraft, the control unit will transmit a command to pressurize the auxiliary circuit and allow the control system to be supplied from the auxiliary tank under conditions of zero or negative gravitational force.

[0032] Such an auxiliary circuit control system therefore makes it possible to do without a hydraulic control to activate the auxiliary circuit or a gravitational force sensor, making the implementation of this auxiliary circuit simpler and more reliable.

[0033] Thanks to such a solution, the turbomachine is independent of aircraft data.

[0034] The invention may include one or more of the following features, taken individually or in combination with each other:

[0035] - The auxiliary circuit further includes a valve located between the auxiliary pump and the control system, the valve being controlled by the control unit, - the valve comprises a body having:

[0036] - an inlet connected to an outlet of the auxiliary pump,

[0037] - a first outlet connected to the auxiliary tank, and

[0038] - a second outlet connected to the control system, the valve further comprising a movable part in the body and configured to move in response to the control command from a first position in which the valve inlet is in fluidic communication with the first valve outlet to a second position in which the valve inlet is in fluidic communication with the second valve outlet,

[0039] - the auxiliary pump is mechanically driven,

[0040] - the auxiliary pump is electrically driven,

[0041] - the auxiliary pump is electrically driven in response to the control command,

[0042] - the auxiliary pump is a positive displacement pump,

[0043] - The control unit is configured to determine if the signal transmitted by the sensor represents a feathering of the blades,

[0044] - the control unit is configured to determine if the signal transmitted by the sensor reflects a difference between a measured blade pitch angle and a setpoint angle.

[0045] The invention also relates to a method for supplying oil to an aircraft turbomachine according to any one of the preceding characteristics, the method comprising the following steps:

[0046] (a) acquire the signal representing the blade pitch angle,

[0047] (b) transmit this signal to the control unit,

[0048] (c) determine whether this signal corresponds to a decrease in hydraulic pressure in the main circuit, and

[0049] (d) transmit the control command to allow oil to be supplied to the control system via the auxiliary circuit if the signal corresponds to a decrease in hydraulic pressure in the main circuit. Brief description of the figures

[0050] Other features and advantages will become apparent from the following description of a non-limiting embodiment of the invention with reference to the accompanying drawings in which:

[0051] [Fig.1] Figure 1 is a schematic longitudinal sectional representation of half an aircraft turbomachine according to a first embodiment of the invention;

[0052] [Fig.2] Figure 2 is a schematic perspective representation of an aircraft turbomachine according to a second embodiment of the invention; [Fig.3] Figure 3 is a schematic longitudinal section representation of an aircraft turbomachine according to a third embodiment of the invention;

[0053] [Fig.4] Figure 4 is a schematic representation of an oil supply system according to the invention when the gravitational force is zero or negative;

[0054] [Fig.5] Figure 5 is a schematic representation of the oil supply system of Figure 4 when the gravitational force is positive;

[0055] [Fig.6] Figure 6 is a schematic representation of an oil supply system according to another embodiment, when the gravitational force is zero or negative;

[0056] [Fig.7] Figure 7 is a schematic representation of the oil supply system of Figure 6, when the gravitational force is positive;

[0057] [Fig.8] Figure 8 is a synoptic diagram of a process according to the invention.

[0058] Detailed description of the invention

[0059] A turbomachine 1, 1', 1" for an aircraft is represented for example in figures 1 to 3. The turbomachine 1, 1, 1" comprises a first rotor 2 connected to an engine M extending around a longitudinal axis X. The engine M comprises, from upstream to downstream in the direction of flow of a main airflow F along the longitudinal axis X, a compressor such as a low-pressure compressor 3 and a high-pressure compressor 4, a combustion chamber 5, a turbine such as a high-pressure turbine 6 and a low-pressure turbine 7, and a nozzle 8.

[0060] The rotor of the high-pressure turbine 6 is connected to the rotor of the high-pressure compressor 4 by a high-pressure shaft 9. The rotor of the low-pressure turbine 7 is connected to the rotor of the low-pressure compressor 3 by a low-pressure shaft 10.

[0061] The low-pressure shaft 10 and high-pressure shaft 9 are supported by bearings 12a. The bearings 12a are contained within a lubrication chamber 12 for lubrication. For example, an upstream bearing 120a is arranged radially between an upstream end of the low-pressure shaft 10 and an upstream bearing support 120b, and a downstream bearing 120a' is arranged downstream of the upstream bearing 120a and radially between the low-pressure shaft 10 and a downstream bearing support 120b'. The lubrication chamber 12 is annular. The upstream and downstream bearings 120a and 120a' are arranged within the lubrication chamber 12.

[0062] The first rotor 2 is driven in rotation by a rotor shaft 100. The rotor shaft 100 is connected to the low-pressure shaft 10. The low-pressure shaft 10 drives the rotor shaft 100 in rotation. Advantageously, the low-pressure shaft 10 is connected to the rotor shaft 100 by a speed reducer 11. This allows the first rotor 2 to be driven at a speed lower than the rotational speed of the low-pressure shaft 10. The speed reducer 11 is, for example, arranged in the lubrication chamber 12 between the upstream bearing 120a and the downstream bearing 120a'.

[0063] The main airflow F passes through the turbomachine 1, 1', 1" and splits into a primary airflow F1 which passes through the engine M within a primary duct and a secondary airflow F2 which passes through the first rotor 2 in a secondary duct surrounding the primary duct.

[0064] The turbomachine 1, 1', 1" includes blades 2a which act upon the main airflow F, the primary airflow F1, or the secondary airflow F2. For example, the rotors of the low-pressure compressor 3 and high-pressure compressor 4 include blades 2a which compress the primary airflow F1 upstream of the combustion chamber 5.

[0065] In general, the 2a blades can be fixed in rotation around the longitudinal axis X or movable in rotation around the longitudinal axis X or an axis parallel to the longitudinal axis X.

[0066] According to a first embodiment shown in Figure 1, the turbomachine 1 is a turbofan engine. In this embodiment, the first rotor 2 is a shrouded fan arranged upstream of the engine M. The fan comprises blades 2a. The fan blades 2a are rotatable about the longitudinal axis X. They are, for example, supported by a disk centered on the longitudinal axis X and driven in rotation by the rotor shaft 100. The blades 2a are arranged inside a fan casing 2b. The casing 2b is surrounded by a nacelle (not shown).

[0067] According to a second embodiment shown in Figure 2, the turbomachine 1' is a turbojet engine with an unfaired fan. In this embodiment, the first rotor 2 is a fan comprising blades 2a. In this embodiment, the fan is arranged downstream of the engine M (not visible in this figure). The fan is rotatable about the longitudinal axis X. The fan blades 2a are supported by a disk that rotates about the longitudinal axis X. Furthermore, in this embodiment, a stator 2' is optionally arranged downstream of the fan to straighten the secondary airflow F2. The stator 2' forms a fixed blade about the longitudinal axis X. It comprises blades 2a that may have variable pitch. The blades 2a are mounted outside the nacelle.

[0068] According to a third embodiment shown in Figure 3, the turbomachine 1 is a turboprop. In this embodiment, the first rotor 2 is a propeller arranged upstream of the engine M. The propeller rotates about a propeller axis H parallel to the longitudinal axis X and comprises blades 2a. The blades 2a are supported by a disk centered on the propeller axis H. For example, there are at least two blades 2a, evenly distributed on the disk.

[0069] The blades 2a extend radially with respect to the longitudinal axis X. They typically comprise a blade extending along an aspect ratio Z transverse to the longitudinal axis X and a mounting element for attaching to the disk. The mounting element is, for example, a mounting foot or a platform. According to the invention, the blades 2a have a variable pitch angle. By variable pitch angle, it is understood that the blades 2a are rotatable about their aspect ratio Z. Changing the pitch of the blades 2a allows the orientation of the blades to be adapted to the airflow and thus maximizes the thrust force of the turbomachine 1, 1', 1”.

[0070] To control the pitch angle of the blades 2a, the turbomachine 1, 1', 1" according to the invention includes a control system 13 for the variable pitch-angle blades 2a. The control system 13 comprises a control unit 13a and at least one oil-supplied hydraulic actuator 13b.

[0071] The control unit 13a is, for example, fixed in rotation around the longitudinal axis X. The control unit 13a is, for example, connected to a stator of the turbomachine 1, 1', 1". The control unit 13a is known in the field of the invention by the acronym PCU for "Pitch Control Unit" in English.

[0072] The hydraulic actuator 13b is, for example, a hydraulic cylinder comprising a movable rod connected to the vanes 2a, possibly via a motion transformation mechanism. The translational movement of the rod allows the vanes 2a to rotate around their axis of extension Z. The hydraulic actuator 13b is supplied with oil. The translational movement of the rod is controlled by the control unit 13a, which supplies oil to the hydraulic actuator 13b. The hydraulic actuator 13b rotates around the longitudinal axis X or an axis parallel to the longitudinal axis X. The hydraulic actuator 13b is, for example, rotationally fixed to the vanes 2a. The hydraulic actuator 13b is, for example, arranged upstream of the control unit 13a.

[0073] The control system 13 advantageously includes an oil transfer device 13c from the control unit 13a to the hydraulic actuator 13b. The oil transfer device 13c ensures the transfer of oil from the stationary control unit 13a to the rotating hydraulic actuator 13b. The oil transfer device 13c is known by the acronym OTB for "Oil Transfer Bearing." The oil transfer device 13c is, for example, arranged within the lubrication chamber 12.

[0074] Furthermore, according to the invention, the turbomachine 1, T, 1” includes an oil supply system 14 for the control system 13.

[0075] The supply system 14 includes a main circuit 15 and an auxiliary circuit 16 for supplying oil to the control system 13. The supply system 14 may further include an oil recovery circuit 17 from the control system 13.

[0076] The main circuit 15 is a hydraulic circuit. It comprises a main reservoir 18 and a main pump 19 connected to the main reservoir 18 and to the control system 13. The main pump 19 is a hydraulic pump. The main pump 19 is dedicated to supplying oil to the pitch control system 13 and is commonly referred to as a "pitch pump" in English within the field of this invention. The main pump 19 is, for example, a positive displacement pump. Positive displacement pumps can, for example, have a fixed displacement or a variable displacement.

[0077] Advantageously, the main circuit 15 may include an additional pump 20 mounted between the main reservoir 18 and the main pump 19, to ensure sufficient oil pressure at the inlet of the main pump 19. The main circuit 15 may also include a heat exchanger 21 located between the additional pump 20 and the main pump 19. The heat exchanger 21 is, for example, of the oil / fuel type. During an initial phase of operation of the turbomachine 1, particularly when the gravitational force experienced by the aircraft is positive, the main pump 19 draws oil from the main reservoir 18 and allows the oil to circulate through the main circuit 15 to the control system 13.

[0078] During a second phase of operation of the turbomachine 1, 1', 1", particularly when the gravitational force acting on the aircraft is zero or negative, typically during negative gravity (or inverted gravity) flight, the oil is trapped in the upper part of the main tank 18, while the lower part, connected to the main pump 19, is filled with air. In zero or negative gravity conditions, an air-oil mixture may also be suspended in the main tank 18. The main pump 19 may therefore draw in air or oil heavily laden with air bubbles. This is unacceptable because the control system 13 must be supplied with oil that is relatively free of air bubbles in order to avoid compromising the operation of the control unit 13a and, consequently, the hydraulic actuator 13b, which controls the blade pitch 2a. Furthermore, the presence of air can lead to the main pump 19 losing its prime.

[0079] Such conditions can therefore lead to a drop in hydraulic pressure in the main circuit 15 and prevent the blades 2a from being set according to the control unit 13a setpoint.

[0080] In this context, the turbomachine 1, 1', 1" advantageously includes a blade safety system 2a. The safety system includes, for example, a unit for feathering the blades 2a in response to a decrease in hydraulic pressure in the main circuit 15, or a unit for locking the blade pitch angle 2a in response to a decrease in hydraulic pressure in the main circuit 15. The feathering unit includes, for example, weights that allow the blades 2a to be positioned in the feathered position when a decrease in hydraulic pressure occurs in the main circuit 15. The blade safety system 2a may result in a loss of thrust, and it is necessary to supply oil to the control system 13 during the second phase of operation of the turbomachine 1, 1.

[0081] In this context, the auxiliary circuit 16 provides the supply to the control system 13 in this second phase of operation of the turbomachine 1, 1', 1”. The auxiliary circuit 16 includes an auxiliary tank 22 and an auxiliary pump 23.

[0082] The auxiliary pump 23 includes an inlet 23a connected to the auxiliary tank

[0083] 22 and an output 23b connected to the control system 13. The auxiliary pump

[0084] 23 is preferably a volumetric pump.

[0085] The auxiliary pump 23 is, according to a first embodiment, mechanically driven. According to this example, it can be driven by the low-pressure shaft 10 or the high-pressure shaft 9.

[0086] In another example, the auxiliary pump 23 is electrically driven. According to this example, the turbomachine 1, 1', 1'" includes an electric motor for driving the auxiliary pump 23.

[0087] The auxiliary pump 23 can be driven continuously, i.e. in both operating phases of the turbomachine 1, 1', 1" or according to the operating phases of the turbomachine 1, 1', 1". When the auxiliary pump 23 is driven continuously, it can be driven according to a first regime in the first operating phase of the turbomachine 1, 1', 1" and according to a second regime higher than the first regime in the second operating phase of the turbomachine 1, 1', 1", a pump regime corresponding to a pump flow rate.

[0088] The auxiliary tank 20 is configured to deliver oil during the second phase of operation of the turbomachine 1, 1', 1". The auxiliary tank 20 is therefore configured to deliver oil under 0g and / or negative g conditions. It includes a suction port preferably located in the upper part of the auxiliary tank 23 and connected to the auxiliary pump 23. This ensures oil suction when the latter is pressed against the upper part of the auxiliary tank 23 when the aircraft is subjected to a negative gravitational force.

[0089] Advantageously, the auxiliary tank 22 is connected to the main tank 18 by a pipe 22a. This ensures the evacuation of oil from the auxiliary tank 22 in the first phase of operation in particular.

[0090] In addition, the recovery circuit 17 fluidly connects the control system 13 to the auxiliary tank 22. This minimizes oil losses while filling the auxiliary tank 22.

[0091] The turbomachine 1, T, 1” further includes a control system for the supply of the auxiliary circuit 16. The control system includes a sensor 24 and a control unit 25.

[0092] The sensor 24 is configured to acquire a signal S1 relating to the angle of pitch of the blades 2a and transmit this signal S1 to the control unit 25.

[0093] Sensor 24, for example, is a position sensor configured to measure the position of the foot or platform of blades 2a. In this example, sensor 24 is of the electromagnetic type.

[0094] In another example, sensor 24 is a linear LVDT (Linear Variable Differential Transformer). Sensor 24 measures the position of the foot or platform of the blades 2a according to the position of the hydraulic actuator 13b.

[0095] The control unit 25 includes a computer. The computer is typically a FADEC (for "Full Authority Digital Engine Control").

[0096] The control unit 25 is configured to transmit an oil supply command 01 from the control system 13 via the auxiliary circuit 16 according to the signal S1 from the sensor 24.

[0097] As illustrated in Figure 4, when the transmitted signal S1 is a signal S11 corresponding to a decrease in hydraulic pressure in the main circuit 15, then the control order 01 to supply oil to the control system 13 by the auxiliary circuit 16 is transmitted to the auxiliary circuit 16. The auxiliary circuit 16 then supplies the control system 13.

[0098] As illustrated in Figure 5, when the transmitted signal S1 is a signal S12 corresponding to a standard hydraulic pressure in the main circuit 15, then the control order 01 to supply oil to the control system 13 by the auxiliary circuit 16 is not transmitted to the auxiliary circuit 16. The main circuit 15 then supplies the control system 13.

[0099] The command order 01 is addressed according to a first embodiment to the auxiliary pump 23 which is electrically activated in response to the command order 01.

[0100] Advantageously, the auxiliary circuit 16 may include a check valve mounted between the auxiliary pump 23 and the control system 13. According to a first embodiment, the control unit 25 is configured to determine whether the signal S1 represents a feathering of the blades 2a. According to this embodiment, the sensor 24 measures the position of the blades 2a, and the control unit determines whether this position is representative of a feathered position of the blades 2a under the aircraft's flight conditions. A feathered position of the blades 2a under conditions where the blades 2a should be in a different position indicates a decrease in pressure in the main circuit 15.

[0101] According to a second embodiment, the control unit 25 is configured to determine whether the signal S1 indicates a difference between the measured blade pitch angle 2a and a setpoint angle. In this embodiment, the sensor 24 measures the position of the blades 2a, and the control unit 25 compares this position to a setpoint value imposed by the control unit 13a of the control system 13. If a difference is detected by the control unit 25, then the hydraulic actuator 13b has locked in response to a decrease in pressure in the main circuit 15.

[0102] According to an embodiment of the invention illustrated in Figures 6 and 7, the auxiliary circuit 16 may include a hydraulic valve 26. The valve 26 is a 3 / 2-way distributor, i.e., having three ports and two positions. The valve 26 is, for example, spring-returned.

[0103] The valve 26 has a body having an inlet 26a connected to the outlet 23b of the auxiliary pump 23 and a first outlet 26b connected to the auxiliary tank 22 and a second outlet 26c connected to the control system 13. The second outlet 26c is connected for example to the outlet of the main pump 19.

[0104] The valve 26 further includes a movable member in the body configured to move between a first position in which the inlet 26a of the valve 26 is in fluidic communication with the first outlet 26b of the valve 26 and a second position in which the inlet 26a of the valve 26 is in fluidic communication with the second outlet 26c of the valve 26. The valve 26 includes, for example, a return spring enabling the movable member to be returned from the second position to the first position.

[0105] It is thus understood that in the first position as illustrated in Figure 7, the auxiliary pump 23 draws oil from the auxiliary reservoir 22 and the oil is redirected to the auxiliary reservoir 22. The oil supply to the control circuit 13 is ensured by the main pump 19, which allows the oil to circulate in the main circuit 15. In the second position as illustrated in Figure 6, the auxiliary pump 23 draws oil from the auxiliary reservoir 22 and the oil is conveyed to the control system 13.

[0106] Regardless of the position of the moving part of the valve 26, the auxiliary pump 23 can therefore be active.

[0107] In the first position, the auxiliary pump 23 can be in the first operating mode, and in the second position, the auxiliary pump 23 can be in the second operating mode. Preferably, the auxiliary pump 23 therefore remains active regardless of the position of the moving part of the valve 26. This eliminates the need for a priming time for the auxiliary pump 23 and ensures a rapid oil supply to the control system 13 during the second operating phase of the turbomachine 1, 1', 1”.

[0108] Also, the difference in operating conditions helps to minimize electrical energy consumption when the power requirements of the control system 13 by the auxiliary circuit 16 are low.

[0109] Thus, when the turbomachine 1, 1', 1" is in the first operating phase, particularly when the aircraft is in a so-called "normal" flight phase (i.e., under positive g conditions), the valve 26 is in the first position and the auxiliary pump 23 can be in the first operating regime. When the turbomachine 1, 1', 1" is in the second operating phase, particularly when the aircraft is in a flight phase in which the gravitational force is zero (referred to as "0g") or negative (referred to as "negative g"), the valve 26 is in the second position and the auxiliary pump 23 can be in the second operating regime, which is higher than the first regime. This ensures the oil supply to the control system 13 from the auxiliary reservoir 22 and prevents any interruption in the oil supply to the control system 13 under these conditions.

[0110] According to this embodiment, the valve 26 is controlled by the control unit 25 according to the signal S1 transmitted by the sensor 24.

[0111] In particular, as illustrated in Figure 6, when the transmitted signal S1 is the signal S11 corresponding to a decrease in hydraulic pressure in the main circuit 15, then the control order 01 for oil circulation in the auxiliary circuit 16 is transmitted to the valve 26 and the moving body is moved into the second position.

[0112] As illustrated in Figure 7, when the transmitted signal S1 is the signal S12 corresponding to a standard hydraulic pressure in the main circuit 15, then the control command 01 for oil circulation in the auxiliary circuit 16 is not transmitted to the valve 26, and the moving body is held in the first position. The auxiliary pump 23 remains active, preferably according to the first operating mode, and the oil is returned to the auxiliary reservoir 22. A method for supplying oil to the turbomachine 1, 1', 1" will now be described. The method is illustrated, for example, in Figure 8.

[0113] The process includes the following steps:

[0114] (a) acquire a signal S1 representative of the blade pitch angle 2a,

[0115] (b) transmit to control unit 25 the signal S1 measured in step (a),

[0116] (c) determine whether signal S1 corresponds to a decrease in hydraulic pressure in the main circuit 15, and

[0117] (d) transmit command order 01 to allow the supply of oil to the control system 13 by the auxiliary circuit 16 if the signal S1 corresponds to a decrease in hydraulic pressure in the main circuit 15.

[0118] Thanks to the invention, it is possible to ensure a continuous oil supply to the control system 13 during all operating phases of the turbomachine 1, 1', 1", and in particular during the aircraft's flight phases in Og or negative g conditions. The auxiliary circuit 16 ensures a continuous oil supply to the control system 13 in response to a drop in hydraulic pressure in the main circuit 15.

[0119] Thanks to the control system, the operation of auxiliary circuit 16 is simple and reliable. It eliminates the need for gravitational force sensors and hydraulic control.

Claims

DEMANDS 1. Turbomachine (1, 1', 1") for an aircraft, the turbomachine (1, 1', 1") comprising: - variable pitch angle blades (2a), - a hydraulic control system (13) for the angle of the blades (2a), - an oil supply system (14) for the control system (13) comprising: - a main hydraulic circuit (15) for supplying oil to the control system (13) under conditions of positive gravitational force experienced by the aircraft, and - an auxiliary hydraulic circuit (16) for supplying oil to the control system (13) under conditions of negative or zero gravitational force experienced by the aircraft, the auxiliary circuit (16) comprising: - an auxiliary tank (22), and - an auxiliary pump (23) connected to the auxiliary reservoir (22) and to the control system (13), and - an oil supply control system for the auxiliary circuit (16), characterized in that the control system comprises: - a sensor (24) configured to acquire a signal (S1) representative of the blade pitch angle and transmit this signal (S1), and - a control unit (25) configured to transmit a command order (01) to allow oil to be supplied to the control system (13) through the auxiliary circuit (16) if the signal (S1) transmitted by the sensor (24) to the control unit (25) corresponds to a decrease in hydraulic pressure in the main circuit (15).

2. Turbomachine according to the preceding claim, characterized in that the auxiliary circuit (16) further comprises a valve (26) located between the auxiliary pump (23) and the control system (13), the valve (26) being controlled by the control unit (25).

3. Turbomachine according to the preceding claim, characterized in that the valve (26) comprises a body having: - an inlet (26a) connected to an outlet (23b) of the auxiliary pump (23), - a first outlet (26b) connected to the auxiliary tank (22), and - a second outlet (26c) connected to the control system (13), the valve (26) further comprising a movable member in the body and configured to move in response to the control order (01) from a first position in which the inlet (26a) of the valve (26) is in fluidic communication with the first outlet (26b) of the valve (26) to a second position in which the inlet (26a) of the valve (26) is in fluidic communication with the second outlet (26c) of the valve (26).

4. Turbomachine according to any one of the preceding claims, characterized in that the auxiliary pump (23) is mechanically driven.

5. Turbomachine according to any one of claims 1 to 3, characterized in that the auxiliary pump (23) is electrically driven.

6. Turbomachine according to claim 1, characterized in that the auxiliary pump (23) is electrically driven in response to the control command (01).

7. Turbomachine according to any one of the preceding claims, characterized in that the auxiliary pump (23) is a positive displacement pump.

8. Turbomachine according to any one of the preceding claims, characterized in that the control unit (25) is configured to determine if the signal (S1) transmitted by the sensor (24) represents a feathering of the blades (2a).

9. Turbomachine according to any one of claims 1 to 7, characterized in that the control unit (25) is configured to determine whether the signal (S1) transmitted by the sensor (24) reflects a deviation between a measured blade pitch angle (2a) and a setpoint angle.

10. A method for supplying oil to an aircraft turbomachine (1, T, 1”) according to any one of the preceding claims, the method comprising the following steps: (a) acquire the signal (S1) representative of the blade pitch angle (2a), (b) transmit this signal (S1) to the control unit (25), (c) determine whether this signal (S1) corresponds to a decrease in hydraulic pressure in the main circuit (15), and (d) transmit the control order (01) to allow oil supply to the control system (13) through the auxiliary circuit (16) if the signal (S1) corresponds to a decrease in hydraulic pressure in the main circuit (15).

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