TURBOMACHINE INCLUDING AN AUXILIARY OIL SUPPLY CIRCUIT

Abstract

The invention relates to a turbomachine (1, 1', 1'') comprising: - a hydraulic control system (13) for the blade pitch angle (2a), - a main hydraulic circuit (15) for supplying oil to the control system (13), and - an auxiliary hydraulic circuit (16) for supplying oil to the control system (13), and - a control system for the oil supply to the auxiliary circuit (16) comprising: - 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 control command (O1) to allow oil supply to the control system (13) via 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). (Shorthand figure: 4)

Description

Title of the invention: TURBOMACHINE COMPRISING AN AUXILIARY OIL SUPPLY CIRCUIT Technical field of the invention

[0001] The invention relates to the field of aircraft turbomachinery comprising variable-pitch blades, a hydraulic control system for these blades, and an auxiliary hydraulic circuit for supplying oil to this control system. Technical background

[0002] A turbomachine for an aircraft comprises, from upstream to downstream, at least one first rotor, also called a 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 to the first rotor by a drive shaft.

[0003] An airflow is compressed within the compressor, then the compressed air is mixed with a fuel and burned within the combustion chamber. The gases formed by the combustion pass through the turbine, which drives the compressor rotor and the propulsion rotor.

[0004] 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 the airflow, it is known to equip the propulsion rotor or the compressor rotor with variable-pitch blades. These variable-pitch blades rotate about their axis of extension, allowing the blade pitch angle to be adjusted according to flight conditions to maximize thrust.

[0005] 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.

[0006] To supply oil to the control system, and in particular to the hydraulic actuator, the turbomachine typically includes an oil supply system for the control system. This supply system includes, for example, a main hydraulic circuit comprising a main reservoir connected to a main pump, allowing the oil to be drawn from the main reservoir and circulated to the hydraulic actuator.

[0007] Certain aircraft flight phases disrupt the oil supply to the hydraulic actuator. Indeed, 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 the invention as the "0g condition" when the gravitational force is zero, or the "negative g condition" when the gravitational force is reversed. During such flight phases, the oil contained 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 0g conditions, the oil and air form a suspension laden with air bubbles, and the pump therefore no longer draws oil but air or oil heavily laden with air bubbles.Such a negative g or 0g condition degrades the oil supply to the control system and can even cause the main pump to lose its prime. In all cases, the hydraulic actuator of the control system is no longer properly supplied with oil.

[0008] Such a degradation of the oil supply to the control system, and in particular to the hydraulic actuator, can render the pitching of the propeller blades uncontrollable, especially the blades of the propeller or the unshod fan. This can lead to the blades being feathered by a safety system or to the blades being locked by the safety system in a position different from the setpoint position 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 of the turbomachine, which is unacceptable.

[0009] In this context, document FR-A1-3127525 proposes an auxiliary oil supply circuit for the control system to supply this control system 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.

[0010] According to this document, the hydraulic valve is actuated 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.

[0011] Although this solution has the advantage of ensuring a continuous power supply to the control system, independent of gravitational force, it is not entirely satisfactory. Indeed, a valve controlled by gravitational force implies equipping the aircraft with a gravitational force sensor. However, such a solution makes the turbomachine dependent on data from the aircraft, particularly that provided by such a sensor. Also, a hydraulic actuation system is complex to implement and may lack reliability.

[0012] 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. Summary of the invention

[0013] To this end, the invention proposes a turbomachine for an aircraft, the turbomachine comprising:

[0014] - variable pitch angle blades,

[0015] - a hydraulic control system for the blade pitch angle,

[0016] - an oil supply system for the control system comprising:

[0017] - a main hydraulic circuit for supplying oil to the control system under conditions of positive gravitational force experienced by the aircraft, and

[0018] - 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 includes:

[0019] - an auxiliary tank, and

[0020] - an auxiliary pump connected to the auxiliary tank and the control system, and

[0021] - an oil supply control system for the auxiliary circuit.

[0022] The turbomachine is remarkable in that the control system comprises:

[0023] - a sensor configured to acquire a signal representative of the angle of adjustment of the paddles and transmit this signal, and

[0024] - a control unit configured to transmit a command order for allow oil supply 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.

[0025] The auxiliary circuit allows the control system to be supplied from the auxiliary tank when the gravitational force is zero (0g condition) or negative (negative g condition).

[0026] This allows the control system to be continuously powered, independently of the gravitational force experienced by the aircraft, and therefore to limit situations of feathering or locking of the blade pitch for safety.

[0027] According to the invention, the auxiliary circuit is pressurized hydraulically by means of the control system.

[0028] The sensor acquires and transmits the signal relating to the blade pitch angle to the control unit. Depending on the interpretation of this signal by this unit control, in particular when the signal reflects a pressure drop in the main circuit symptomatic of zero or negative gravitational force experienced by the aircraft, the command order will be transmitted by the control unit 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.

[0029] 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.

[0030] Thanks to such a solution, the turbomachine is independent of aircraft data.

[0031] The invention may comprise one or more of the following features, taken individually or in combination with each other:

[0032] - the auxiliary circuit further includes a valve located between the auxiliary pump and the control system, the valve being controlled by the control unit,

[0033] - the valve comprises a body having:

[0034] - an input connected to an output of the auxiliary pump,

[0035] - a first outlet connected to the auxiliary tank, and

[0036] - a second output connected to the control system,

[0037] the valve further comprising a movable member 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,

[0038] - the auxiliary pump is mechanically driven,

[0039] - the auxiliary pump is electrically driven,

[0040] - the auxiliary pump is electrically driven in response to the control command,

[0041] - the auxiliary pump is a positive displacement pump,

[0042] - the control unit is configured to determine if the signal transmitted by the The sensor represents a feathering of the blades.

[0043] - the control unit is configured to determine if the signal transmitted by the The sensor translates a difference between a measured blade pitch angle and a setpoint angle.

[0044] 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:

[0045] (a) acquire the signal representing the blade pitch angle,

[0046] (b) transmit this signal to the control unit,

[0047] (c) determine whether this signal corresponds to a decrease in hydraulic pressure in the main circuit, and

[0048] (d) transmit the command to authorize an oil supply 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

[0049] 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:

[0050] [Fig-1] [Fig.1] is a schematic longitudinal sectional representation of a half-turbomachine of aircraft according to a first embodiment of the invention;

[0051] [Fig.2] [Fig.2] is a schematic perspective representation of an aircraft turbomachine according to a second embodiment of the invention;

[0052] [Fig.3] [Fig.3] is a schematic longitudinal cross-sectional representation of an aircraft turbomachine according to a third embodiment of the invention;

[0053] [Fig.4] [Fig.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] [Fig.5] is a schematic representation of the oil supply system of [Fig.4] when the gravitational force is positive;

[0055] [Fig.6] [Fig.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] [Fig.7] is a schematic representation of the oil supply system of [Fig.6], when the gravitational force is positive;

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

[0058] 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.

[0059] 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.

[0060] 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, 120a' are arranged within the lubrication chamber 12.

[0061] 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'.

[0062] The main airflow F passes through the turbomachine 1, 1', 1" and splits into a primary airflow Fl which passes through the engine M within a primary channel and a secondary airflow F2 which passes through the first rotor 2 in a secondary channel surrounding the primary channel.

[0063] The turbomachine 1, 1', 1" includes blades 2a which enable action to be exerted on the main airflow F, the primary airflow Fl or the secondary airflow F2. For example, the rotors of the low-pressure compressor 3 and high-pressure compressor 4 include blades 2a which enable compression of the primary airflow Fl upstream of the combustion chamber 5.

[0064] In general, the blades 2a can be fixed in rotation around the longitudinal axis X or movable in rotation around the longitudinal axis X or of an axis parallel to the longitudinal axis X.

[0065] According to a first embodiment shown in [Fig. 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 housing 2b. The housing 2b is surrounded by a nacelle (not shown).

[0066] According to a second embodiment shown in [Fig. 2], the turbomachine 1' is a turbojet engine with an unfaired fan. According to this embodiment, the first rotor 2 is a fan comprising blades 2a. According to this embodiment, the 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 also rotates about the longitudinal axis X. Furthermore, according to this embodiment, a straightener 2' is optionally arranged downstream of the fan to straighten the secondary airflow F2. The straightener 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.

[0067] According to a third embodiment shown in [Fig. 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.

[0068] The blades 2a extend radially with respect to the longitudinal axis X. They typically comprise a blade extending along an elongation axis Z transverse to the longitudinal axis X and a mounting element to the disk. The mounting element is, for example, a mounting foot or a platform.

[0069] According to the invention, the blades 2a have a variable pitch angle. By variable pitch angle, it is understood that the blades 2a are movable in rotation about their elongation axis Z. Changing the pitch of the blades 2a makes it possible to adapt the orientation of the blades according to the airflow and thus to maximize the thrust force of the turbomachine 1, 1', 1”.

[0070] In order to control the pitch angle of the blades 2a, the turbomachine 1, 1', 1" according to the invention comprises 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, optionally via a motion transformation mechanism. The translational movement of the rod allows the vanes 2a to rotate about their extension axis Z. The hydraulic actuator 13b is supplied with oil. The translational movement of the movable rod is controlled by the control unit 13a, which supplies oil to the hydraulic actuator 13b. The hydraulic actuator 13b rotates about the longitudinal axis X or about an axis parallel to the longitudinal axis X. The hydraulic actuator 13b is for example rotationally fixed to the blades 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 fixed 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 in the lubrication chamber 12.

[0074] Furthermore, according to the invention, the turbomachine 1, 1', 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 the invention. The main pump 19 is, for example, a positive displacement pump. The positive displacement pump is, for example, of fixed or 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, in order 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.

[0078] During a first phase of operation of the turbomachine 1, 1', 1”, in particular when the gravitational force experienced by the aircraft is positive, the main pump 19 draws oil from the main reservoir 18 and allows the circulation of the oil in the main circuit 15 to the control system 13.

[0079] During a second phase of operation of the turbomachine 1, 1', 1”, particularly when the gravitational force experienced by the aircraft is zero or negative, typically 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 is therefore at risk of drawing 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 not to compromise the operation of the control unit 13a and thus the hydraulic actuator 13b, which controls the pitch of the vanes 2a. Furthermore, the presence of air can lead to the main pump 19 losing its prime.

[0080] 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.

[0081] 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 pitch angle of the blades 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.

[0082] The safety shutdown of the blades 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', 1”.

[0083] 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.

[0084] The auxiliary pump 23 includes an inlet 23a connected to the auxiliary reservoir 22 and an outlet 23b connected to the control system 13. The auxiliary pump 23 is preferably a positive displacement 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] According to 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, that is, 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 operating mode in the first operating phase of the turbomachine 1, 1', 1" and according to a second higher regime than the first regime in the second phase of operation 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, 1', 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 SI relating to the angle of pitch of the blades 2a and transmit this signal SI to the control unit 25.

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

[0094] According to another example, the sensor 24 is a linear LVDT type sensor, which stands for "Linear Variable Differential Transformer". The 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 a control order 01 for oil supply from the control system 13 via the auxiliary circuit 16 according to the signal S1 from the sensor 24.

[0097] As illustrated in [Fig.4], when the transmitted signal SI is a signal SI 1 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 [Fig.5], when the transmitted signal SI is an S12 signal corresponding to a standard hydraulic pressure in the main circuit 15, then the control order 01 for supplying 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 non-return valve mounted between the auxiliary pump 23 and the control system 13.

[0101] According to a first embodiment, the control unit 25 is configured to determine whether the signal SI 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 according to 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.

[0102] According to a second embodiment, the control unit 25 is configured to determine whether the signal SI 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.

[0103] 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.

[0104] 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.

[0105] The valve 26 further comprises a movable element 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 comprises, for example, a spring of reminder allowing the moving part to be returned from the second position to the first position.

[0106] It is thus understood that in the first position as illustrated in [Fig.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 circulation of the oil in the main circuit 15. In the second position as illustrated in [Fig.6], the auxiliary pump 23 draws oil from the auxiliary reservoir 22 and the oil is routed to the control system 13.

[0107] Regardless of the position of the moving part of the valve 26, the auxiliary pump 23 can therefore be active.

[0108] 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”.

[0109] Also, the difference in operating conditions makes it possible to minimize the consumption of electrical energy when the power requirements of the control system 13 by the auxiliary circuit 16 are low.

[0110] 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 supply of oil 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.

[0111] According to this embodiment, the valve 26 is controlled by the control unit 25 according to the SI signal transmitted by sensor 24.

[0112] In particular, as illustrated in [Fig. 6], when the transmitted signal SI is the signal SU corresponding to a decrease in hydraulic pressure in the main circuit 15, then the control command 01 for oil circulation in the circuit auxiliary 16 is transmitted to valve 26 and the moving body is moved into the second position.

[0113] As illustrated in [Fig. 7], when the transmitted signal SI 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.

[0114] A method for supplying oil to the turbomachine 1, 1', 1" will now be described. The method is illustrated for example in [Fig. 8].

[0115] The process comprises the following steps:

[0116] (a) acquire an SI signal representative of the blade pitch angle 2a,

[0117] (b) transmit to the control unit 25 the SI signal measured in step (a),

[0118] (c) determine whether the SI signal corresponds to a decrease in hydraulic pressure in the main circuit 15, and

[0119] (d) transmit command order 01 to authorize the oil supply to control system 13 by the auxiliary circuit 16 if the signal SI corresponds to a decrease in hydraulic pressure in the main circuit 15.

[0120] Thanks to the invention, it is possible to ensure an 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 0g or negative g conditions. The auxiliary circuit 16 ensures an oil supply to the control system 13 in response to a drop in hydraulic pressure in the main circuit 15.

[0121] Thanks to the control system, the operation of the 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 blades (2a), - a hydraulic control system (13) for the pitch 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 tank (22) and to the control system (13), and - a control system for the oil supply to the auxiliary circuit (16),characterized in that the control system comprises: - a sensor (24) configured to acquire a signal (SI) representative of the blade pitch angle and transmit this signal (SI), and - a control unit (25) configured to transmit a control command (01) to allow oil to be supplied to the control system (13) via the auxiliary circuit (16) if the signal (SI) 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 whether the signal (SI) 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 (SI) 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, 1', 1") according to any one of the preceding claims, the method comprising the following steps: (a) acquiring the signal (SI) representing the blade pitch angle (2a), (b) transmitting this signal (SI) to the control unit (25), (c) determining whether this signal (SI) corresponds to a decrease in hydraulic pressure in the main circuit (15), and (d) transmit the command order (01) to allow oil supply to the control system (13) through the auxiliary circuit (16) if the signal (SI) corresponds to a decrease in hydraulic pressure in the main circuit (15).

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