JP2019157908A - Electric hydraulic pressure actuator for aircraft - Google Patents

Abstract

To provide an electric hydraulic pressure actuator for an aircraft that can reliably lift up and down a landing gear even when being exposed to a cold environment during flight.SOLUTION: An electric hydraulic pressure actuator A for an aircraft includes: a pump 1; an electric motor 2 for driving the pump 1; an actuator body 3 telescopically operating with supply/discharge of fluid by the driven pump 1; a passage B, 20 for connecting a suction port to a discharge port of the pump 1; and an on-off valve 4 for opening and closing the passage B, 20. The electric hydraulic pressure actuator starts a warm-up operation when a flight altitude becomes equal to or lower than an altitude threshold.SELECTED DRAWING: Figure 1

Description

  The present invention relates to an electric hydraulic actuator for an aircraft.

  In aircraft, hydraulic actuators are used to raise and lower landing gears. As such a hydraulic actuator, for example, an actuator main body including a cylinder, a piston slidably inserted into the cylinder, a piston rod inserted into the cylinder and coupled to the piston, and a motor driven. And a hydraulic circuit that selects one of the contraction side oil chamber and the extension side oil chamber in the cylinder and supplies pressure oil (see, for example, Patent Document 1).

  The hydraulic circuit includes a pump that is set to a bidirectional discharge type in the middle, and a conduit that communicates the contraction-side oil chamber and the extension-side oil chamber, and a branch flow that is provided in parallel with the pump with respect to the conduit A valve and a reservoir oil chamber that communicates with the bore side oil chamber via a check valve and communicates with the drain port of the flow dividing valve. In the hydraulic actuator configured as described above, a desired thrust can be obtained by driving a pump with a motor to expand and contract an actuator body.

  Although it is not for aircraft, paying attention to the temperature of hydraulic oil in the hydraulic actuator, adjusting the number of revolutions of the motor to suppress excessive thrust of the hydraulic actuator in response to changes in the viscosity of the hydraulic oil Has been developed (see, for example, Patent Document 2).

JP 2007-239975 A JP 2013-001305 A

  Here, from the viewpoint of propulsion efficiency, aircraft often fly about 10km above the sea level from the sea, but if the altitude rises by 1km, the temperature drops by about 6.5 degrees, so the outside temperature of the aircraft in flight Reaches -50 degrees. Since the hydraulic actuator mounted on the aircraft is used in such an extremely cold environment, in the aircraft in flight, the temperature of the hydraulic oil in the hydraulic actuator becomes extremely low and the viscosity becomes high.

  The time required for the aircraft to enter the landing posture from the cruising altitude is short, and the temperature of the hydraulic oil in the hydraulic actuator does not rise sufficiently. Therefore, the temperature of the hydraulic oil in the hydraulic actuator inevitably becomes extremely low. There is.

  Therefore, the viscosity of the hydraulic oil in the hydraulic actuator is high at the time of landing, the pump may cause a suction failure, and the landing gear lowering operation in the hydraulic actuator may be slow, or the operation may not be stable. Therefore, there is an idea to stabilize the operation of the hydraulic actuator by adjusting the number of rotations of the motor, as in the hydraulic actuator disclosed in JP2013-001305A, but since the temperature of the hydraulic oil remains low, For the hydraulic actuator to be used, further improvement in reliability is demanded.

  SUMMARY OF THE INVENTION In order to solve the above problems, an object of the present invention is to provide an aircraft hydraulic actuator that can reliably lift a landing gear even if it is exposed to a cold environment during flight.

  In order to achieve the above-described object, an electric hydraulic actuator for an aircraft according to the present invention includes a pump, an electric motor that has a suction port and a discharge port, and drives a pump that sucks and discharges liquid. An actuator body that is driven by supply / discharge; a passage through which a liquid can flow by connecting a suction port and a discharge port of a pump; and an opening / closing valve that opens and closes the passage and switches between the flow and interruption of the liquid through the passage. And a controller that starts warm-up operation when the flight altitude falls below the altitude threshold. Since the electric hydraulic actuator for an aircraft configured in this way starts the warm-up operation as described above when the flight altitude falls below the altitude threshold, the liquid cooled during the flight of the aircraft enters the landing posture. Warm-up to a temperature sufficient to eliminate pump suction failure.

  Further, the electric hydraulic actuator for aircraft may start the warm-up operation on the condition that the controller has the flight altitude of the aircraft at or below the altitude threshold and the temperature of the liquid is at or below the temperature threshold. In the electric hydraulic actuator for aircraft configured in this way, if the temperature of the internal liquid is a temperature that does not require warming up, it will not be warmed up even if the flight altitude falls below the altitude threshold. Since no warm-up operation is avoided, energy consumption can be reduced.

  Furthermore, the electric hydraulic actuator for aircraft may end the warm-up operation when the temperature of the liquid becomes equal to or higher than a preset temperature threshold after the controller starts the warm-up operation. In the electric hydraulic actuator for an aircraft configured as described above, when the liquid is warmed and there is no need for warm-up, the warm-up operation can be terminated, so that energy consumption can be reduced.

  The electric hydraulic actuator for aircraft may be determined by the duration of the warm-up operation when the controller determines the end of the warm-up operation. In the electric hydraulic actuator for an aircraft configured as described above, when the liquid is warmed and there is no need for warm-up, the warm-up operation can be terminated, so that energy consumption can be reduced.

  Furthermore, the electric hydraulic actuator for aircraft may be provided with an opening / closing valve in a passage connecting the suction port and the discharge port of the pump via the actuator body. In the electric hydraulic actuator for an aircraft configured in this way, the liquid circulates through the pump and the bypass passage through the actuator body having a large heat capacity during the warm-up operation. Proper telescopic operation is guaranteed.

  The electric hydraulic actuator for aircraft may be provided with an orifice in the passage, or may function as an orifice at a communication position where the on-off valve opens the passage. In the electric hydraulic actuator for an aircraft configured as described above, heat is generated when the liquid passes through the orifice, and the temperature of the liquid can be efficiently increased to shorten the warm-up operation.

  According to the electric hydraulic actuator for aircraft of the present invention, the landing gear can be lifted and lowered reliably even if it is exposed to a cold environment during flight.

It is the figure which showed the electric hydraulic actuator for aircrafts in one embodiment of this invention. It is the figure which showed the electric hydraulic actuator for aircrafts in the 1st modification of one embodiment of this invention. It is the figure which showed the electric hydraulic actuator for aircrafts in the 2nd modification of one embodiment of this invention.

  Hereinafter, as shown in FIG. 1, an electric hydraulic actuator for aircraft A according to an embodiment of the present invention includes a pump 1, an electric motor 2 that drives the pump 1, and an actuator that is driven by supplying and discharging liquid by the pump 1. A bypass passage B as a passage through which a liquid can flow by connecting a main body actuator body 3, a suction port and a discharge port of the pump 1, an opening / closing valve 4 for opening / closing the bypass passage B, an electric motor 2 and an opening / closing valve 4 And a controller C for controlling the above.

  The electric hydraulic actuator A for aircraft is mounted on an aircraft. In this example, the pump 1 is driven by the electric motor 2 and the actuator main body 3 is expanded and contracted to lift the landing gear (not shown) of the aircraft. Used for.

  Hereinafter, each part of the electric hydraulic actuator A for aircraft will be described in detail. In the present embodiment, the actuator body 3 is movably inserted into the cylinder 10, the piston 11 that is movably inserted into the cylinder 10 and partitions the cylinder 10 into the expansion side chamber R1 and the contraction side chamber R2, and the cylinder 10. It is a direct acting single rod type cylinder provided with a piston rod 12 which is inserted and connected at one end to the piston 11. The actuator body 3 is set to a telescopic type in which the piston 11 moves in the axial direction with respect to the cylinder 10 and the piston rod 12 enters and exits the cylinder 10 to expand and contract. The actuator body 3 may be a double rod cylinder in which the piston rod 12 protrudes on both sides of the piston 11 and is inserted into the extension side chamber R1 and the contraction side chamber R2. The actuator main body 3 is an actuator that is set so that the pressure receiving areas for receiving the pressures in the expansion side chamber R1 and the contraction side chamber R2 are equal to each other as in the actuator main body disclosed in Japanese Patent Application Laid-Open No. 2007-71363 filed by the applicant. The body may be used. Therefore, the structure of the actuator body 3 is not limited as long as it can be extended and contracted by supplying hydraulic pressure.

  The cylinder 10 is closed at the right end in FIG. 1, and the other end of the piston rod 12 protrudes outside the cylinder 10 from the left end. Further, the cylinder 10 includes a port 10 a that communicates with the expansion side chamber R <b> 1 and a port 10 b that communicates with the contraction side chamber R <b> 2, and these ports 10 a and 10 b are connected by a supply passage 14 outside the cylinder 10. The extension side chamber R1 and the contraction side chamber R2 provided in the cylinder 10 are filled with liquid such as hydraulic oil, but the liquid used for the operation of the electric hydraulic actuator A for aircraft is used as hydraulic oil. Without being limited, depending on the use environment, it is possible to use liquids other than hydraulic oil such as aqueous solution, electrorheological fluid, and magnetorheological fluid.

  A bidirectional discharge pump 1 is provided in the supply passage 14. The pump 1 is driven by the electric motor 2. When the pump 1 is rotated forward, the pump 1 sucks liquid from the contraction side chamber R2 through the supply passage 14 and sends the liquid to the extension side chamber R1. Then, the liquid is discharged from the extension side chamber R1 so that the liquid is sucked and sent out to the contraction side chamber R2. Therefore, the suction port and the discharge port in the pump 1 are interchanged depending on the rotation direction of the pump 1.

  Further, in this example, a damping passage 15 is provided in parallel with the supply passage 14 to communicate the extension side chamber R1 and the contraction side chamber R2 in the cylinder 10, and an electromagnetic valve 16 is provided in the middle of the damping passage 15. It has been. The electromagnetic valve 16 takes a communication position that opens the damping passage 15 when energized and takes a blocking position that shuts off the damping passage 15 when not energized. When the solenoid valve 16 takes the communication position, it functions as an orifice so as to provide resistance to the flow of hydraulic oil going back and forth between the expansion side chamber R1 and the contraction side chamber R2. The damping passage 15 may be installed so that the extension side chamber R1 and the contraction side chamber R2 communicate with each other outside the cylinder 10 independently of the supply passage 14, but if connected to the supply passage 14, the damping passage 15 is provided in the cylinder 10. The number of ports 10a and 10b to be provided can be reduced.

  Further, the bypass passage B is in parallel with the supply passage 14 and communicates the suction port and the discharge port of the pump 1. In the middle of the bypass path B, an on-off valve 4 is provided. This on-off valve 4 is an electromagnetic on-off valve, which takes a communication position that opens the bypass path B to allow the flow of the liquid when not energized, shuts off the bypass path B when energized, and allows the liquid to flow through the bypass path B. It is a normally closed on-off valve that takes a blocking position to block flow. Further, when the on-off valve 4 is in the communication position, it allows liquid to flow through the bypass passage B and functions as an orifice to provide resistance to the flow of hydraulic oil passing through the bypass passage B. Further, the bypass passage B may be installed so as to communicate the extension side chamber R1 and the contraction side chamber R2 outside the cylinder 10 independently of the supply passage 14, but if connected to the supply passage 14, the bypass passage B is provided in the cylinder 10. The number of ports 10a and 10b to be provided can be reduced.

  Further, in the electric hydraulic actuator for aircraft A of the present embodiment, compensation passages 17 connected in a loop shape are provided on both sides of the supply passage 14 across the pump 1. The compensation passage 17 is connected to the reservoir 19 via a low-pressure priority shuttle valve 18. The compensation passage 17 may be installed so as to communicate the extension side chamber R1 and the contraction side chamber R2 outside the cylinder 10 independently of the supply passage 14, but the compensation passage 17 is connected to the supply passage 14 as described above. This configuration is advantageous in that the number of ports 10a and 10b provided in the cylinder 10 can be reduced.

  In this example, the reservoir 19 is pressurized by being filled with gas in addition to the hydraulic oil as a liquid, and exerts internal pressure on the expansion side chamber R1 and the contraction side chamber R2 in the cylinder 10 so that the aircraft flies. Even in this case, the inside of the cylinder 10 does not become lower than the external pressure. In the reservoir 19, gas-liquid separation is performed by partitioning the gas and the liquid with a free piston, a diaphragm, a bladder, a bellows, or the like so that the gas and the liquid do not mix even when the yaw, roll, or pitch of the aircraft changes. Structure is adopted.

  When the controller C receives a command to deploy a landing gear (not shown) from an aircraft (not shown), the controller C controls the electric motor 2 to drive the actuator main body 3 to lower the landing gear and deploy the landing gear. Is received, the electric motor 2 and the on-off valve 4 are controlled to drive the actuator body 3 to lift and store the landing gear. When the landing gear is lifted or lowered, the controller C drives the actuator body 3 by driving the pump 1 with the electric motor 2 with the solenoid valve 16 being in the communication position while the on-off valve 4 is in the shut-off position. As an emergency operation, when the landing gear is lowered from the retracted position and moved to the unfolded position, the opening / closing valve 4 is used as the communication position without driving the electric motor 2, and the actuator body 3 is functioned as a damper. You may make it move to a deployment position with dead weight.

  In addition to the above-described operation, the controller C performs pumping with the electric motor 2 when the flight altitude of the aircraft falls below a preset altitude threshold regardless of the landing gear deployment command or storage command input from the aircraft. 1 is driven and the on-off valve 4 is opened to perform the warm-up operation to the communication position.

  The flight altitude of the aircraft is input to the controller C from an altimeter installed in the aircraft, but the electric hydraulic actuator A for aircraft may have its own altimeter. Then, the controller C compares the received flight altitude with the altitude threshold, and when the flight altitude falls below the altitude threshold, the controller C controls the electric motor 2 and the on-off valve 4 as described above to perform warm-up operation.

  In the warm-up operation, the pump 1 is rotationally driven by the electric motor 2 and the bypass passage B is opened by the on-off valve 4, so that the hydraulic oil discharged from the pump 1 is sucked by the pump 1 through the bypass passage B. . That is, the hydraulic oil circulates through the pump 1 and the bypass passage B. The pump 1 may be driven in the forward rotation direction or in the reverse rotation direction. When the pump 1 is driven in this way, the hydraulic oil is warmed by the heat generated by the pump 1. In the present embodiment, since the on-off valve 4 functions as an orifice at the communication position, heat is generated when the hydraulic oil passes through the orifice, so that the hydraulic oil is efficiently warmed. The altitude threshold is determined when the aircraft hydraulic hydraulic actuator A is landing the temperature of the hydraulic oil by warm-up until the aircraft descends to an altitude at which the landing gear starts to be deployed when entering the landing posture at the normal descent speed of the aircraft. The altitude is set to such an extent that a time for raising the gear to a temperature that does not hinder the lifting and lowering of the gear can be secured. In other words, if the altitude threshold is set to X feet and the altitude at which the landing gear starts to be deployed is set to Y feet, it will operate if the aircraft warms up during the time required to descend from X feet to Y feet in normal flight conditions. The altitude threshold may be set so that the temperature of the oil can be sufficiently warmed.

  Then, when the flight altitude falls below the altitude threshold, the warm-up operation as described above is started, so that the hydraulic oil that has cooled during the flight of the aircraft eliminates the suction failure of the pump 1 before entering the landing posture. Since the aircraft is warmed up to a sufficient temperature, the aircraft hydraulic actuator A can reliably lift and lower the landing gear. That is, the aircraft hydraulic hydraulic actuator A starts warm-up operation when the flight altitude falls below the altitude threshold. Therefore, when the aircraft lands, the hydraulic oil (liquid) in the actuator body 3 is sufficiently warmed up. Thus, the aircraft hydraulic actuator A can lift and lower the landing gear reliably.

  As described above, the electric hydraulic actuator A for aircraft according to the present invention includes the pump 1, the electric motor 2 that drives the pump 1, the actuator main body 3 that expands and contracts by supplying and discharging liquid by the driving of the pump 1, and the pump. 1 includes a bypass passage (passage) B that connects the suction port and the discharge port 1 and an on-off valve 4 that opens and closes the bypass passage (passage) B, and starts a warm-up operation when the flight altitude falls below an altitude threshold. Therefore, according to the electric hydraulic actuator for aircraft A of the present invention, the landing gear can be lifted and lowered reliably even if it is exposed to a cold environment during flight.

  In the aircraft hydraulic hydraulic actuator A of the present embodiment, the bypass passage B connecting the suction port and the discharge port of the pump 1 without using the actuator body 3 is used as a passage, and the on-off valve 4 is connected to the bypass passage B. Is provided. In this way, when the warm-up operation is executed, the hydraulic oil (liquid) circulates through the pump 1 and the bypass passage B without passing through the actuator body 3 having a large heat capacity. Necessary minimum hydraulic fluid (liquid) can be warmed up. In other words, if the temperature of the hydraulic oil in the actuator body 3 having a large volume and a large heat capacity is to be raised by the warm-up operation, it takes time to raise the temperature, so that the time required for the warm-up operation becomes longer. Then, since it is necessary to set the altitude threshold value to a high value inevitably, the drive time of the electric motor 2 becomes longer during the warm-up operation. However, if the aforementioned bypass path B is provided, the warm-up operation time can be shortened. This makes it possible to reduce the energy consumption during the warm-up operation, which is economical, and allows the altitude threshold value to be set low.

  In addition, as shown in FIG. 2, the bypass path B of FIG. 1 is abolished, and a passage 20 for communicating the extension side chamber R1 and the contraction side chamber R2 is provided separately from the supply passage 14 as a passage used for the warm-up operation. The on-off valve 4 may be provided in the passage 20 and the passage 20 may be used for warm-up operation. In this way, the passage 20 communicates the suction port and the discharge port of the pump 1 via the actuator body 3. In this case, since the hydraulic oil (liquid) discharged from the pump 1 is returned to the suction port of the pump 1 through the extension side chamber R1 and the contraction side chamber R2 of the actuator body 3, the warm-up operation time becomes long. Since the liquid in the actuator body 3 is also warmed, the smooth expansion / contraction operation of the actuator body 3 is guaranteed. Note that the effect that the landing gear can be lifted and lowered reliably is not lost even in the actuator A including the passage 20 that communicates the suction port and the discharge port of the pump 1 through the actuator body 3 as described above.

  Further, in the aircraft hydraulic actuator A according to the present embodiment, the on-off valve 4 functions as an orifice at a communication position that opens the bypass passage (passage) B, so that hydraulic fluid (liquid) passes through the on-off valve 4. When heat is generated, warm-up operation can be shortened. Instead of the on-off valve 4 functioning as an orifice at the communication position, an orifice 21 may be provided in the bypass path (passage) B separately from the on-off valve 4 as shown in FIG. In the aircraft hydraulic actuator A configured as described above, when hydraulic oil (liquid) passes through the on-off valve 4 or the orifice 21, heat is generated to efficiently raise the temperature of the liquid and warm up operation. Can be shortened.

  In addition, the electric hydraulic actuator A for aircraft may continue the warm-up operation until a command to lift the landing gear is input after the warm-up operation is started. When the temperature of the hydraulic oil (liquid) becomes equal to or higher than a preset temperature threshold, the warm-up operation may be terminated. The temperature of the hydraulic oil (liquid) may be detected by providing a temperature sensor in the vicinity of the pump 1 or in the bypass passage B and passage 20 used for warm-up operation. The temperature threshold may be set to a temperature at which the suction failure of the pump 1 is eliminated. In this way, when the hydraulic oil (liquid) is sufficiently warmed and there is no need for warm-up, the warm-up operation can be terminated, so that the energy consumption of the aircraft electric hydraulic actuator A can be reduced.

  The electric hydraulic actuator for aircraft A may determine the end of the warm-up operation based on the duration of the warm-up operation regardless of the temperature of the hydraulic oil (liquid). That is, the warm-up operation may be terminated when a preset warm-up operation time has elapsed after the start of the warm-up operation. The warm-up operation time may be set to such a time that the temperature of the hydraulic oil (liquid) reaches the temperature at which the suction failure of the pump 1 is eliminated by the warm-up operation. Thus, even if the warm-up operation is finished by determining whether or not the warm-up operation has been sufficiently performed, the energy consumption of the electric hydraulic actuator for aircraft A can be reduced.

  Furthermore, regarding the start condition of the warm-up operation, not only the condition for the flight altitude of the aircraft but also the temperature of the hydraulic oil (liquid) may be set as the condition. Specifically, the electric hydraulic actuator for aircraft A starts warm-up operation on condition that the flight altitude of the aircraft is lower than the altitude threshold and the temperature of the hydraulic oil (liquid) is lower than the temperature threshold. Also good. The temperature threshold is the same as the above-described temperature threshold, and may be set to a temperature at which the suction failure of the pump 1 is eliminated. Also, the altitude threshold is set in the same manner as described above. In the aircraft hydraulic hydraulic actuator A configured as described above, when the temperature of the internal hydraulic fluid (liquid) is a temperature that does not require warming up in the first place, the warming-up operation is performed even if the flight altitude is below the altitude threshold. Therefore, unnecessary warm-up operation is avoided and energy consumption can be reduced.

  Note that the circuits (the supply passage 14 and the compensation passage 17) for connecting the actuator main body 3, the pump 1, and the reservoir 19 shown in the present embodiment are merely examples, and the design can be changed as appropriate. In the present embodiment, the actuator body that is the drive target of the pump 1 is a direct acting hydraulic cylinder, but may be a rotary hydraulic motor. Further, although the pump 1 is a bidirectional discharge type pump, liquid is selectively supplied to the expansion side chamber R1 and the contraction side chamber R2 in the cylinder 10 on the circuit side by using a one-way discharge type pump. And may be able to discharge. Even when a one-way discharge type pump is used in this way, if a passage is provided to connect the discharge port and suction port of the pump and the pump is driven, the liquid circulates between the pump and the passage, so that a warm-up operation can be performed.

  Although the preferred embodiments of the present invention have been described in detail above, modifications, variations, and changes can be made without departing from the scope of the claims.

DESCRIPTION OF SYMBOLS 1 ... Pump, 2 ... Electric motor, 3 ... Actuator main body, 4 ... On-off valve, 20 ... Passage, 21 ... Orifice, A ... Electric hydraulic actuator for aircraft, B ... Bypass (passage), C ... Controller

Claims (6)

An aircraft hydraulic actuator mounted on an aircraft,
A pump having a suction port and a discharge port for sucking and discharging liquid;
An electric motor for driving the pump;
An actuator body driven by supplying and discharging liquid by the pump;
A passage through which the liquid can flow by connecting the suction port and the discharge port of the pump;
An on-off valve that opens and closes the passage to switch between liquid flow and shut-off through the passage; and
And a controller that opens the on-off valve and starts the warm-up operation to drive the pump when the flight altitude of the aircraft is equal to or lower than an altitude threshold. 2. The aircraft hydraulic fluid according to claim 1, wherein the controller starts the warm-up operation when a flight altitude of the aircraft becomes an altitude threshold value or less and a temperature of the liquid becomes a temperature threshold value or less. Pressure actuator. The electric hydraulic actuator for an aircraft according to claim 1, wherein the controller ends the warm-up operation when a temperature of the liquid becomes a temperature threshold value or more during the warm-up operation. The aircraft according to any one of claims 1 to 3, wherein the controller ends the warm-up operation when a preset warm-up operation time has elapsed since the start of the warm-up operation. Electric hydraulic actuator for use. The aircraft hydraulic pressure actuator according to any one of claims 1 to 4, wherein the passage connects the suction port and the discharge port of the pump via the actuator body. 6. The aircraft electric machine according to claim 1, wherein the passage includes an orifice in the middle, or the opening / closing valve functions as an orifice at a communication position where the passage opens the passage. 6. Hydraulic actuator.

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