Method for powering aircraft actuators having a single generator

Abstract

The invention relates to a method for powering actuators of an aircraft comprising at least one generator (500), the actuators being associated with landing gear of the aircraft and comprising at least one actuator for providing at least one of the following functions: A / extending / retracting the landing gear, B / steering the steerable wheels of the landing gear, C / normal braking of the aircraft; and one actuator for providing at least one of the following functions: D / applying the parking brake of the aircraft, E / extending the landing gear in emergency mode, F / braking in emergency mode. The method of the invention comprises the steps of: • - powering the actuator providing at least one of the functions A / , B / , C / directly with the generator; • - powering the actuator providing at least one of the functions D / , E / , F / by means of an accumulator (501) connected to the generator by a power line (502) comprising a normally closed switching member (503) for isolating the accumulator from the generator so as to have two independent and segregated power sources, each powering actuators that perform separate functions, the switching member (503) being controllable to open and allow the accumulator to be charged by the generator.

Description

[0001] METHOD FOR POWERING AIRCRAFT ACTUATORS WITH A SINGLE GENERATOR

[0002] The present invention relates to the field of aeronautical systems, and more specifically those grouped together in chapter ATA32 concerning landing systems (landing gear, braking, extension / retraction, orientation, energy distribution system and associated computers, etc.).

[0003] BACKGROUND OF THE INVENTION

[0004] On a certified aircraft, the on-board systems grouped in chapter ATA32 must ensure the following functions:

[0005] - in normal mode, extension / retraction of the landing gear, orientation of the steerable wheels for maneuvering the aircraft on the ground, braking, parking braking;

[0006] - in emergency mode, emergency extension of the landing gear, emergency braking.

[0007] These functions are performed using actuators such as locking boxes, actuators for extending / retracting the landing gear and / or associated hatches, braking actuators, wheel steering actuators, etc. Certified passenger aircraft have at least two engines, and therefore at least two hydraulic generators (pumps) to define two main hydraulic circuits, and, in general, at least one emergency hydraulic circuit powered by an accumulator, itself kept filled by one or other of the generators. The two main hydraulic circuits are used to perform the same functions alternately, or are used simultaneously to perform the same function (for example, one of the main hydraulic circuits supplies one half of the brakes, while the other of the main hydraulic circuits supplies the other half of the brakes).On single-engine aircraft, there is only one hydraulic generator. On some aircraft, such as drones, there is no hydraulic generator, and only one electric generator is available. On some single-engine aircraft, two separate hydraulic circuits are created, each powered by its own accumulator, with both accumulators being filled by the same hydraulic generator. Each circuit powers the actuators associated with the landing gear, so that all associated functions can be performed.

[0008] SUBJECT OF THE INVENTION

[0009] The object of the invention is to propose a power supply method for ensuring the functions operated by the components of the ATA32 chapter, while meeting the required segregation and safety criteria, by minimizing the number of components.

[0010] SUMMARY OF THE INVENTION

[0011] To this end, the invention proposes a method for supplying actuators of an aircraft comprising at least one generator, the actuators being associated with landing gear of the aircraft and capable of respectively ensuring at least one of the following ABC functions and one of the following DEF functions:

[0012] A / extension / retraction of the landing gear;

[0013] B / orientation of steerable wheels of the landing gears;

[0014] C / normal braking of the aircraft;

[0015] D / aircraft parking braking, or parking;

[0016] E / extension of the landing gear in emergency mode;

[0017] F / braking in emergency mode;

[0018] The method of the invention comprising the steps of: powering the actuator ensuring at least one of the functions A / , B / , C / directly with the generator; powering the actuator ensuring at least one of the functions D / , E / , F / by means of an accumulator connected to the generator by a supply line comprising a normally closed cut-off member which can be controlled to open and allow the accumulator to be charged by the generator.

[0019] Preferably, the supply line comprises a non-return member placed between the cut-off member and the accumulator to prevent any discharge of the accumulator via the supply line.

[0020] Thus, two circuits are created, normally segregated from each other by the cut-off device, the latter being opened only to ensure the charging of the accumulator, for example when starting the aircraft before any movement of the latter. If ever the generator were to fail in flight, it remains possible to ensure all or part of the functions D / , E / , F / , which allows the landing gear to be deployed and therefore allows its landing, braking and immobilization. It will be noted that the functions A / , B / , C / can be ensured as long as the generator is operating, while the functions D / , E / , F / cannot be ensured as long as the accumulator is not discharged, therefore for a limited time, which conditions the dimensioning of the latter.

[0021] It will also be noted that the accumulator is preferably used to provide both normal functions (parking braking) and emergency functions (emergency braking and extension). The power supply system is thus simplified.

[0022] This ensures adequate operational safety with a reduced number of components, including a single generator and a single accumulator. This reduces the system's footprint and limits the risk of breakdown.

[0023] The invention improves the reliability and maintenance of the system. It will also be noted that the invention applies to a hydraulic power supply where the generator is a hydraulic pump driven by the aircraft engine, but also to an electrical power supply where the generator is an electrically powered motor pump assembly.

[0024] Preferably, the cut-off device is controlled to open while the aircraft is stationary on the ground, until the accumulator is fully charged.

[0025] Preferably then, the cut-off member is controlled to open in response to one of the following events: occurrence of an energy delivery signal (pressure, electrical intensity, etc.) at the output of the generator; activation of the generator; activation of an aircraft computer; occurrence of a signal indicating the increase in an aircraft electrical supply; operation by an operator of a specific button.

[0026] According to a particular embodiment of the method of the invention, the cut-off member is controlled to open in flight to keep the accumulator permanently charged.

[0027] Preferably then, the cut-off member is controlled to open: in response to the detection of a low energy level at the output of the accumulator as measured by a downstream sensor; periodically according to a predefined time interval; during the landing approach, detected as a function of at least one of the following events, for example: the extension of the landing gear, a reduction in the speed of the aircraft below a predetermined threshold, a reduction in the altitude of the aircraft below a predetermined threshold.

[0028] According to a particular embodiment of the method of the invention, the charging of the accumulator is carried out in two sub-phases separated by a brief closing phase of the cut-off device.

[0029] Other characteristics and advantages of the invention will emerge from reading the following description of a particular and non-limiting embodiment of the invention.

[0030] BRIEF DESCRIPTION OF THE DRAWINGS

[0031] Reference will be made to the figures in the attached drawings, including:

[0032] [Fig. 1] Figure 1 is a schematic view of landing gear and associated actuators fitted to an aircraft;

[0033] [Fig. 2] Figure 2 is a schematic view of a latching box provided with two ports for its actuation in normal mode and in emergency mode;

[0034] [Fig. 3] Figure 3 is a schematic hydraulic circuit showing the implementation according to the invention of a generator and an accumulator to constitute two segregated pressure sources.

[0035] DETAILED DESCRIPTION OF THE INVENTION

[0036] As illustrated in Figure 1, the invention is described here in relation to an aircraft comprising two main landing gears 100, 200 and an auxiliary landing gear 300. The main landing gears 100, 200 are maneuvered by cylinders 105, 205 to be deployed or retracted into holds closed by hatches 107, 207 actuated by cylinders 108, 208. The main landing gears 100, 200 are held in the retracted position by attachment boxes 106, 206, while the hatches are kept closed by attachment boxes 109, 209. The main landing gears 100, 200 are equipped with wheels 101, 201, themselves equipped with brakes 102, 202.

[0037] The auxiliary landing gear 300 is maneuvered by a jack 305 to be deployed or retracted into a hold closed by hatches 307 actuated by jacks

[0038] 308. The auxiliary landing gear 300 is held in the retracted position by a latching box 306, while the hatches are held closed by a latching box

[0039] 309. The auxiliary landing gear 300 further comprises a steering cylinder 310 adapted to steer the wheels to steer the aircraft on the ground. This landing gear configuration is given for illustrative purposes only and in no way limits the possibility of applying the invention to other configurations. As schematically illustrated in FIG. 2, each of the landing gear attachment boxes 106, 206, 306 and each of the hatch attachment boxes 109, 209, 309 here comprises two hydraulic ports. One of the ports is provided to power a main unlocking actuator and is powered by the aircraft extension / retraction circuit. The other of the ports is provided to power an emergency unlocking actuator allowing emergency extension of the landing gear. This emergency actuator is used in the event of unavailability of the main actuator (blockage, no power supply, etc.).Each of the actuators causes the hook to unlock so that it releases the element it was holding, hatch or landing gear. As seen in Figure 3, the aircraft is equipped with six hydraulic circuit sections, referenced from A / to F / , each closed by an isolation valve, to ensure the following functions:

[0040] A / extension / retraction of the landing gear. This function is ensured by supplying, in a determined sequence, the cylinders 105, 205, 305 to deploy or retract the cylinders 108, 208, 308 to open or close the hatches of the landing gear holds, and the main unlocking actuators of the attachment boxes 106, 206, 306, 109, 209, 309;

[0041] B / orientation of steerable wheels of the landing gears. This function is ensured by supplying a servovalve distributing the pressure in the chambers of the orientation cylinder 310 of the auxiliary landing gear 300;

[0042] C / normal braking of the aircraft. This function is carried out by supplying braking servovalves distributing pressure in the wheel brakes 102, 202

[0043] 101, 201 of the main landers 100, 200;

[0044] D / aircraft parking brake, or parking. This function is performed by applying the brakes

[0045] 102, 202 of the wheels 101, 201 of the landing gears 100, 200 a constant pressure provided by a previously inflated accumulator, so as to maintain this pressure even when the aircraft is stationary and its engine stopped;

[0046] E / extension of the landing gear in emergency mode. This function is ensured by powering the emergency release actuators of the attachment boxes 106, 206, 306, 109, 209, 309, so as to release the hatches and allow the landing gear to deploy under the effect of gravity and / or relative wind.

[0047] F / braking in emergency mode. This function is performed by supplying braking servovalves distributing pressure in the brakes 102, 202 of the wheels 101, 201 of the main landing gears 100, 200 using an alternative source of pressure, for example a second generator, or a previously inflated accumulator.

[0048] The way in which these functions are performed in practice is known per se and does not constitute the subject of the invention. The invention aims to propose a power supply circuit allowing the performance of all these functions, particularly suitable for the case of an aircraft comprising only a single propulsion engine, and therefore only a single pressure generator, in this case a hydraulic pump 500 mechanically driven by the engine. It may also be a hydraulic pump driven by an electric motor powered by an electrical energy source such as an electric generator mechanically driven by the engine, an electromagnetic circuit associated with the engine, a battery, etc.

[0049] As can be seen in Figure 3, the hydraulic circuit of the invention consists of connecting the pump 500 to a hydraulic accumulator 501 by a supply line 502 comprising a cut-off valve 503 and a non-return valve 504. The cut-off valve 503 is normally in a closed state, but can be controlled to be placed in an open state in which the pump 500 can recharge the accumulator 501. The non-return valve 504 prevents the discharge of the accumulator 501 via the supply line 502 in the event of a leak in the upstream hydraulic circuit, or failure or leakage of the cut-off valve 503. Upstream 505 and downstream 506 pressure sensors make it possible to monitor the pressure in the circuit on either side of the cut-off valve 503. Specifically, the upstream pressure sensor 505 measures the pressure delivered by the pump 500, while the downstream pressure sensor 506 measures the pressure delivered by the accumulator 501.As illustrated, and according to a preferred arrangement, the pump 500 serves to power the actuators performing the functions A / , B / , C / , while the accumulator 501 serves to power the actuators performing the functions D / , E / , F / . Two independent and segregated power sources have thus been created, powering each of the actuators performing distinct functions.

[0050] The operation of the circuit of the invention is as follows. The cut-off valve 503 is controlled open when the aircraft is powered up while it is stationary on the ground, for a predetermined duration sufficient to allow the accumulator 501 to be fully filled. This duration depends on the characteristics of the pump 500 (pressure / flow rate) and the volume of the accumulator 501. It will be noted that the opening of the cut-off valve 503 occurs in a non-critical phase where the aircraft is on the ground, stationary on the tarmac, before departure. The filling phase of the accumulator 501 can thus be carried out automatically on a signal such as the presence of pressure at the outlet of the pump 500, or on activation of the pump 500 or the braking computer or any other computer (flight control, etc.), or on a signal indicating the rise of a specific electrical power supply of the aircraft.

[0051] In particular, the filling of the accumulator 501 can also be controlled by a signal generated by an operator pressing a specific button placed in the cockpit of the aircraft or behind a hatch accessible by the maintenance team. Of course, in the latter case, it is first necessary to ensure that a hydraulic generation is operational, either the pump 500, or a remote generator brought by the maintenance team.

[0052] Other filling logics are possible, more sophisticated, but providing more operational flexibility. For example, it may be planned to control the cut-off valve 503 to open during the flight in order to keep the accumulator completely filled at all times, in other words in order to compensate for leaks from the equipment connected to the accumulator. These reopenings must however be managed with care; they will preferably only take place after ensuring the absence of leaks from the emergency systems downstream of the cut-off valve 503, and will preferably be brief (only the time to compensate for the small volume of fluid corresponding to the anticipated leaks from the various connected equipment).The absence of leakage downstream of the cut-off valve 503 is verified by comparing the pressure drop read at the downstream pressure sensor 506 with a theoretical maximum pressure drop corresponding to the theoretical maximum leakage level of the various connected equipment during the period of time considered.

[0053] The reopening of the cut-off valve 503 in flight can be based on a low pressure threshold at the downstream sensor 506, or can still be carried out periodically according to a predefined time interval, or even during the extension of the landing gear before landing, or based on a temperature threshold of the fluid in the reservoir of the pump 500 while restarting the pump 500 at the same time to heat the fluid and thus not excessively increase the extension time of the landing gear. The temperature of the fluid is here measured by an RTD probe based on a platinum winding and is theoretically included in a temperature range extending from -60 °C to +135 °C.

[0054] Once the filling of the accumulator 501 is complete, the control of the cut-off valve 503 is stopped and it is returned to the closed position, thus isolating the pump 500 from the accumulator 501. The accumulator 501 is sized so as to autonomously perform all of the functions assigned to it.

[0055] The presence of the downstream pressure sensor 506 makes it possible to verify that the cut-off valve 503 is not stuck open, which prevents it from performing its segregation function. In practice, while the cut-off valve 503 is commanded to open, the pressure measured by the downstream pressure sensor 506 must increase, until it reaches the nominal pressure value normally delivered by the pump 500, or more precisely, until it reaches the value of the pressure supplied by the upstream pressure sensor 505 (to within a few units due to the flow rate). It may then be advisable to divide the accumulator filling phase into two sub-phases interspersed with an intermediate phase during which the cut-off valve 503 is briefly closed. The purpose of this re-closing is to verify that the pressure measured by the downstream pressure sensor 506 stops increasing due to its good isolation from the pump 500.

[0056] Of course, other logics can also be imagined in order to detect a blockage of the cut-off valve 503 in the open position. For example, in the case of a gas accumulator 501, the gas tends to heat up in contact with the fluid heated by the pump 500. Therefore, when the cut-off valve 503 closes, a slight reduction in the downstream pressure must be observed due to the cooling of the gas, indicating the effective closing of the cut-off valve 503.

[0057] It should also be noted that each time the parking brake is applied (function D / ), the pressure measured by the downstream pressure sensor 506 must decrease due to the consumption of fluid from the accumulator 501. If it does not decrease, this means that the cut-off valve 503 is stuck open and that the pump 500 continues to charge the accumulator 501.

[0058] The presence of the non-return valve 504 makes it possible to overcome the cases, which are rare in practice, of rupture or tearing of upstream piping, or even of a crack in the cut-off valve 503. In such a situation, the non-return valve 504 makes it possible to maintain the pressure level in the accumulator 501, preventing it from emptying.

[0059] The invention is not limited to what has just been described, but on the contrary encompasses any variant falling within the scope defined by the claims.

[0060] In particular, although the invention has been illustrated in application to an aircraft equipped with a generator in the form of a hydraulic pump for powering hydraulic actuators, the invention may also be applied to an aircraft equipped with an electric generator for powering electric actuators, the hydraulic accumulator being replaced by an electric accumulator (battery, capacitors, etc.), the cut-off member then being a controlled switch, while the non-return member is a unidirectional electrical component, for example of the diode type. The pressure sensors will be replaced by electrical intensity and / or voltage sensors, capable of also measuring an energy level at the output of the generator or accumulator (voltage, electrical intensity, etc.).

[0061] The aircraft may comprise at least one thermal propulsion engine (with a hydraulic or electric generator) or at least one electric propulsion engine (with a hydraulic generator). The invention also applies to an aircraft comprising an electric propulsion engine and an electric generator, the actuators being electric actuators.

[0062] Any type of valve can be used, including three-way / two-position or two-way / two-position valves, single- or double-stage hydraulic valves, valves with one control channel or two control channels (redundancy).

[0063] The invention can be used with single-port housings, equipped upstream with a shuttle valve selecting between the two hydraulic circuits.

[0064] The upstream sensor 505 can also be a simpler discrete sensor providing OK / KO type operating information. The upstream pressure sensor 505 is optional. In its absence, to detect a lack of pressure from the hydraulic pump, it is necessary to use information relating to one of the actuators performing functions A / , B / , C / .

[0065] The invention covers the case where the actuators controlled by the method of the invention only provide one or more of the ABC functions and the DEF functions respectively: it is thus possible to have an actuator for one of the ABC functions and an actuator for one of the DEF functions.

[0066] The downstream pressure sensor 506 can be replaced by another sensor giving an indication of the state of the accumulator and its capacity, for example, a level sensor.

[0067] The reopening of the cut-off valve 503 in flight can also be based on any event occurring on approach to landing and in particular the reduction of altitude below a predetermined threshold, the reduction of the aircraft speed below a predetermined threshold, etc.

[0068] The non-return device 504 placed between the cut-off device and the accumulator to prevent any discharge of the accumulator via the supply line is optional if the risk of such a discharge is accepted.

[0069] Although the invention allows the implementation of the six functions using a single generator and a single accumulator, it will be possible to be satisfied with only having the generator perform some of the associated functions. Similarly, it will be possible to be satisfied with only having the accumulator perform some of the associated functions.

Claims

CLAIMS 1. Method for powering actuators (102, 105, 106, 109, 202, 205, 206, 209, 305, 306, 309, 310) of an aircraft comprising at least one generator (500) driven in rotation, the actuators being associated with landing gears (100, 200, 300) of the aircraft and comprising at least one actuator for ensuring at least one of the ABC functions and an actuator for ensuring at least one of the following DEF functions: A / extension / retraction of the landing gear; B / orientation of steerable wheels of the landing gear; C / normal braking of the aircraft; D / aircraft parking braking, or parking; E / extension of the landing gear in emergency mode; F / braking in emergency mode; The method of the invention comprising the steps of: powering the actuator providing at least one of the functions A / , B / , C / directly with the generator; powering the actuator providing at least one of the functions D / , E / , F / by means of an accumulator (501) connected to the generator by a power supply line (502) comprising a normally closed cut-off member (503) to isolate the accumulator from the generator so as to have two independent and segregated power sources, each powering actuators performing distinct functions, the cut-off member (503) being controllable to open and allow the accumulator to be charged by the generator.

2. Method according to claim 1, in which the supply line (502) comprises a non-return member (504) placed between the cut-off member (503) and the accumulator (501) to prevent any discharge of the accumulator via the power supply line.

3. Method according to claim 1 or 2, in which the cut-off member (503) is controlled to open while the aircraft is stationary on the ground, until the accumulator (501) is completely filled.

4. Method according to claim 3, in which the cut-off member (503) is controlled to open in response to one of the following events: occurrence of an energy delivery signal at the output of the generator (500); activation of the generator (500); activation of a computer of the aircraft; occurrence of a signal indicating the increase of an electrical power supply of the aircraft; operation by an operator of a specific button.

5. Method according to claim 1, in which the cut-off member (503) is controlled to open in flight to keep the accumulator (501) permanently charged.

6. Method according to claim 5, in which the cut-off member (503) is controlled to open: in response to the detection of a low energy level at the output of the accumulator as measured by a downstream sensor (506); periodically according to a predefined time interval; during a landing approach.

7. The method of claim 6, wherein the landing approach is detected based on one of the less than the following events: extension of the landing gear, a reduction in aircraft speed below a predetermined threshold, a reduction in aircraft altitude below a predetermined threshold.

8. Method according to claim 1, in which the charging of the accumulator (501) is carried out in two sub-phases separated by a brief closing phase of the cut-off member (503).

9. Method according to claim 1, in which the aircraft comprises a thermal propulsion engine and the generator is a hydraulic generator, the actuators being hydraulic actuators.

10. The method of claim 1, wherein the aircraft comprises an electric propulsion motor and the generator is a hydraulic generator, the actuators being hydraulic actuators.

11. Method according to claim 1, in which the aircraft comprises a thermal propulsion engine and the generator is an electric generator, the actuators being electric actuators.

12. The method of claim 1, wherein the aircraft comprises an electric propulsion motor and the generator is an electric generator, the actuators being electric actuators.

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