• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    The invention relates to an aircraft comprising a landing gear (10) of which at least one wheel (12) is provided with an electric motor (50) configured to drive said wheel (12) in rotation, the aircraft further comprising a control system (100) of said electric motor (50), said control system (100) comprising: - a control board (102) having first control means (106) for outputting a torque or power to be applied to said engine (50) in a direction of advance (16) of the aircraft and second control means (108) provided to deliver a running speed value of the aircraft in a reversing direction (18). ) of the aircraft, - a control unit (104) connected to the control board (102) and to the electric motor (50), having first means for controlling in torque, or in power, the electric motor (50) according to the value delivered by the first control means ( 106), and second means for speed-controlling the electric motor (50) according to the running speed value supplied by the second control means (108), and - a speed sensor (110) for measuring the speed taxiing the aircraft, and transmitting this speed information to the control unit (104).
    公开号:FR3017367A1
    申请号:FR1451003
    申请日:2014-02-10
    公开日:2015-08-14
    发明作者:Raphael Renier；Xavier Guery；Christine Charbonnier；Matthieu Mayolle；Sylvain Ferro；Aurelie Treil；Rodolphe Bonet
    申请人:Airbus Operations SAS；
    IPC主号:
    专利说明:

    [0001] The present invention relates to an aircraft comprising a landing gear whose wheel is provided with an electric motor and a control system of said electric motor. It is known to equip a landing gear of an aircraft with an electric motor.
    [0002] This electric motor is used to roll the aircraft when it joins the runway or when it reaches its parking area. The implementation of such an engine makes it possible to roll the aircraft by limiting the fuel consumption since the aircraft's engines do not have to produce thrust.
    [0003] Such an arrangement is known as "etaxi". However, the control of the electric motor is currently not particularly intuitive for pilots. An object of the present invention is to propose an aircraft which does not have the drawbacks of the prior art.
    [0004] For this purpose, is proposed an aircraft comprising a landing gear of which at least one wheel is provided with an electric motor configured to drive said wheel in rotation, the aircraft further comprising a control system of said electric motor, said system control device comprising: a control panel, presenting first control means provided for delivering a torque or power value to be applied to said engine in a direction of advance of the aircraft and second control means provided for delivering a rolling speed value of the aircraft in a direction of recoil of the aircraft, - a control unit connected to the control panel and the electric motor, having first means for controlling in torque, or in power, the engine according to the value delivered by the first control means, and second means for controlling in speed the electric motor according to the the rolling speed value delivered by the second control means, and - a speed sensor for measuring the running speed of the aircraft, and transmitting this speed information to the control unit. Such a control system allows a feeling similar to the application of a thrust by the aircraft engines in forward and releases the pilot of the speed monitoring constraint in reverse.
    [0005] The characteristics of the invention mentioned above, as well as others, will emerge more clearly on reading the following description of an exemplary embodiment, said description being given in relation to the attached drawings, among which: FIG. . 1 shows an aircraft according to the invention, FIG. 2 is a schematic representation of a control system of an electric motor of a landing gear of the aircraft according to the invention, FIG. 3 is a mode of implantation of the electric motor of the control system on a landing gear when the electric motor is in a disengaged position, and FIG. 4 is a representation similar to that of FIG. 3 when the electric motor is in a meshing position. In the following description, the advancement direction corresponds to the direction in which an aircraft moves as it moves forward and the recoil direction corresponds to the direction in which the aircraft moves when it is moving backwards. Fig. 1 shows an aircraft 1 which comprises a landing gear 10 and a cockpit 2. At least one wheel 12 of the landing gear 10 is provided with an electric motor 50 configured to drive said wheel 12 in rotation. The aircraft 1 also comprises brakes for braking the wheel 12 and a brake pedal whose actuation activates the brakes. As shown in more detail in FIG. 2, the aircraft 1 comprises a control system 100 for controlling the electric motor 50. In the embodiment of the invention presented here, the electric motor 50 is equipped with a driving gear 52 and the wheel 12 of the train landing gear 10 is provided with a driven gear 14. When the driving gear 52 meshes with the driven gear 14, the wheel 12 is rotated and in the direction of rotation of the electric motor 50, the wheel 12 will drive the aircraft 1 forward (arrow 16) or back (arrow 18). In the embodiment of the invention presented here, a single wheel 12 of the aircraft 1 is thus equipped, but it is possible to equip one or more wheels of each landing gear. The control system 100 comprises: - a control panel 102 arranged in the control cabin 2 of the aircraft 1 to be handled by a pilot, and - a control unit 104 connected to the control panel 102 and the electric motor 50, and configured to receive the controls from the control panel 102 and to control the electric motor 50 according to said commands. The control panel 102 comprises: first control means 106 operable by the pilot and designed to deliver a torque or power value to be applied to said motor 50 in the direction of travel 16, and second control means 108. maneuverable by the pilot and intended to deliver a rolling speed value of the aircraft 1 according to the reversing direction 18. The control unit 104 comprises: - first means for controlling in torque, or in power, the engine electrical 50 according to the value delivered by the first control means 106, that is to say when the aircraft 1 advances in the direction of advance 16, and - second means for controlling in speed the electric motor 50 according to the rolling speed value delivered by the second control means 108, that is to say when the aircraft 1 moves back in the recoil direction 18. Thus, when the pilot wishes the aircraft 1 to advance, he maneuvers the first control means 106 according to the torque, or the power, that he wants to apply to the electric motor 50, the first control means 106 then deliver to the control unit 104, the information of the value of the torque , or power, which must be applied to the electric motor 50 in the forward direction, and the control unit 104 then controls the electric motor 50 according to this setpoint torque, or power.
    [0006] Thus, when the pilot wishes the aircraft 1 to move backward, he maneuvers the second control means 108 according to the running speed at which he wishes the aircraft 1 to retreat, the second control means 108 then deliver to the aircraft unit. control 104, the information that the electric motor 50 must be controlled in speed and in reverse so that the aircraft 1 back to said running speed, and the control unit 104 then controls the electric motor 50 according to this set of driving speed. The application of a torque, or a power, to the electric motor 50 is felt by the pilot to be similar to the application of a thrust by the aircraft's engines 1 and the deceleration of the aircraft 1 is achieved using the brakes.
    [0007] The application of a speed in reverse allows the pilot to deal only with the trajectory of the aircraft 1 while the control unit 104 controls the electric motor 50 so that the speed of travel is respected and this whatever the environment, such as the inclination of the track. The application of a zero running speed using the second control means 108 slows the aircraft 1 without the need to use the brakes. The acceleration and deceleration of the aircraft 1 are controlled by the control unit 104, which avoids any sudden braking that could have an impact on the longitudinal stability of the aircraft 1 and therefore the passenger comfort.
    [0008] To know the running speed of the aircraft 1, the control system 100 comprises a speed sensor 110 for measuring the running speed of the aircraft 1. The speed sensor 110 transmits the speed information to the aircraft. control unit 104 which can then accelerate or slow down the electric motor 50 according to the value of the running speed sensed by the speed sensor 110 and the running speed to obtain.
    [0009] According to a particular embodiment of the invention, the first control means 106 and the second control means 108 consist of the same rotary knob 150 having a zero position (0 in FIG. from the zero position to a first maximum angle, in a first direction of rotation 112 is representative of the torque control, or power, of the electric motor 50, and where a rotation, from the zero position to at a second maximum angle, in a second direction of rotation 114 is representative of the speed control of the electric motor 50. According to a particular embodiment, the first maximum angle is of the order of 100 ° in the clockwise direction from the zero position, and the second maximum angle is of the order of 80 ° counterclockwise from the zero position. In the first direction of rotation 112, the rotary knob 150 takes the form of a knob of the rotary potentiometer type, which, in particular, is continuous, linear and with a constant friction force between the zero position and the first maximum angle. By the rotation of the rotary knob 150 in the first direction of rotation 112, the pilot controls the value of the torque, or of the power, from 0% in zero position to 100% of the torque, or of the power, available in the position of the first maximum angle. The return of a position different from zero (that is to say between the zero position and the first maximum angle) at the zero position causes the freewheeling of the aircraft 1 which can then be controlled only braking by the brakes.
    [0010] According to a variant, in the second direction of rotation 114, the rotary knob 150 takes the form of a switch with two stable positions, namely the zero position and a latched position corresponding to the second maximum angle, where the speed control is activated. . The return to the zero position is then performed by the pilot.
    [0011] By rotating the rotary knob 150 in the second direction of rotation 114, the pilot controls the value of the 0 knot speed (KT) in the zero position at a predetermined speed in the position of the second maximum angle. The predetermined speed is preferably less than or equal to the walking speed of a man, that is to say between 1 and 3 knots (KT) and preferably of the order of 2 knots (KT), so that a track operator can follow the pace of the aircraft 1. According to another variant, in the second direction of rotation 114, the rotary knob 150 takes the form of a switch-type button at a stable position corresponding to the zero position and an unstable position corresponding to the second maximum angle. Releasing the rotary knob 150 from a position different from the zero position causes its automatic return to the zero position. According to another variant, in the second direction of rotation 114, the rotary knob 150 takes the form of a knob of the rotary potentiometer type between the zero position and the second maximum angle. By rotating the rotary knob 150 in the second direction of rotation 114, the pilot controls the value of the speed from a zero value in the zero position to a maximum value in the position of the second maximum angle. The return to the zero position is then performed by the pilot or if the potentiometer has a single stable position corresponding to the zero position, the return to the zero position is performed automatically as soon as the driver releases the rotary knob 150.
    [0012] The return of a nonzero speed position to the zero position causes the application of a zero speed, that is to say that the control unit 104 decelerates the aircraft 1 to the stop. According to another particular embodiment of the invention not shown in the figures, the first control means and the second control means are constituted by a same lever movable in rotation about a horizontal axis and preferably perpendicular to the axis longitudinal axis of the aircraft 1, and having a zero position in which the lever is perpendicular to the plane of the control panel 102, where a rotation, from the zero position to a first maximum angle forward of the 1 is representative of the control in torque, or in power, of the electric motor 50, and where a rotation, from the zero position to a second maximum angle towards the rear of the aircraft 1 is representative of the speed control of the electric motor 50. In the direction of forward rotation, the lever takes the form of a potentiometer and the pilot controls the value of the torque, or the power, of 0% in position zero to 100% of the torque, or power, available in the position of the first maximum angle. According to a variant, in the direction of rotation towards the rear, the lever takes the form of a switch with two stable positions, namely the zero position and an engaged position corresponding to the second maximum angle where the speed control is activated. . The return to the zero position is then performed by the pilot. By rotating the lever in the direction of rotation backwards, the pilot controls the value of the speed of 0 knot (KT) in zero position at a predetermined speed in the position of the second maximum angle. According to another variant, in the direction of rotation towards the rear, the lever takes the form of a switch at a stable position corresponding to the zero position and an unstable position corresponding to the second maximum angle. Releasing the lever from a position other than the zero position causes it to automatically return to the zero position. According to another variant, in the direction of rotation towards the rear, the lever takes the form of a rotary potentiometer between the zero position and the second maximum angle. By turning the lever in the direction of rotation backwards, the pilot controls the value of the speed from a zero value in the zero position to a maximum value in the position of the second maximum angle. The return to the zero position is then performed by the pilot or if the potentiometer has a single stable position corresponding to the zero position, the return to the zero position is performed automatically as soon as the driver releases the lever. Whether in the case of the rotary knob 150 or the lever, the zero position is preferably indexed, that is to say that there is a hard point that materializes this position. To prevent the pilot from inadvertently directly from the control in torque, or power, the speed control without passing through a stop position of the electric motor 50, the transition from the zero position to the control position in speed is achieved through discontinuous kinematics. The pilot must thus perform a first manipulation of the rotary knob 150, respectively of the lever, before performing the rotation specific to the speed control. This first manipulation must not be a rotation in continuity with said own rotation. In the case of the rotary knob 150, this first manipulation can be for example: a pressing on the rotary knob 150 or an uprising of the rotary knob 150 in an axial direction. In the case of the lever, this first manipulation can be for example: a shift of the lever perpendicular to the median plane of the aircraft 1. According to a preferred embodiment, the passage of a taxi in the reversing direction 18 to a taxi according to the direction of advance 16 is carried out when the aircraft 1 is stopped, and that the pilot engages the parking brake of the aircraft 1.
    [0013] For this purpose, the control system 100 includes a parking brake detector 116 which detects when the parking brake is engaged or not and which is connected to the control unit 104. As long as, on the one hand, the detector parking brake 116 does not indicate to the control unit 104 that the parking brake is engaged and as long as, on the other hand, the speed sensor 110 does not indicate to the control unit 104 that the running speed is zero, the control unit 104 remains in the zero speed control mode of the electric motor 50, and this even if another command is transmitted by the control panel 102. To turn on and off the system 100, it has an on-off button 118. Preferably, the on-off button 118 is a monostable push button, that is to say that a first press on the on-off button 118 puts running the control system 100 and the on-off button 118 revie In its stable position, and a second press on the start-stop button 118 stops the control system 100 and the on-off button 118 returns to its stable position.
    [0014] To inform the pilot of the on / off state of the control system 100, the on-off button 118 is equipped with a light-emitting diode which lights up when the control system 100 is on and which turns off when the control system 100 is stopped. The use of a monostable push button allows the pilot to voluntarily stop the control system 100 and also allows the control system 100 to stop on its own when certain particular conditions are met, for example, when a failure of the control system 100 is detected on one of the elements of the control system 100, or when the reactors of the aircraft 1 are in idle mode ("idle" in English), that is to say when the fan rotates and the thrust produced by the reactor is minimal and insufficient to advance the aircraft 1. The control system 100 also includes a fault indicator light 120 which lights when a failure of the control system 100 is detected. It may happen, during the progression of the aircraft 1 in reverse, that it is necessary to perform an emergency maneuver. In this case, the manipulation of the rotary knob 150, respectively of the lever, for a return to the zero position must be immediate. Or the location of the rotary knob 150, respectively of the lever, may take some time until the driver visualizes the rotary knob 150, respectively the lever, and manipulates it. To reduce this reaction time, it is advantageous that a pressure on the brake pedal, which is quickly accessible by the pilot, triggers the sending, to the electric motor 50, a deceleration control up to a speed running zero.
    [0015] For this purpose, the control system 100 comprises an actuation detector which is connected to the control unit 104 and which is provided to deliver information relating to the actuation or non-actuation of said brake pedal. Thus, when the aircraft 1 moves backwards and the brake pedal is actuated, the actuation detector informs the control unit 104 which then controls the electric motor 50 so as to decelerate it to reach a zero running speed. , which corresponds to a deactivation of the speed control and a return to zero of the speed reference. According to one variant, the control system 100 is connected to at least one proximity sensor arranged on the fuselage or the wings of the aircraft 1 and connected to the control unit 104. When, during the progression of the aircraft 1 in reverse, a proximity sensor sends a signal to the control unit 104 whose amplitude exceeds a predetermined threshold, thereby indicating that the proximity sensor has detected a nearby obstacle, the control unit 104 command then the electric motor 50 so as to decelerate to reach a running speed of zero. Fig. 3 and FIG. 4 show a particular implantation of the electric motor 50 on the landing gear 10.
    [0016] Fig. 3 shows a disengaging position, when the driving gear 52 does not mesh with the driven gear 14, and FIG. 4 shows a meshing position, when the driving gear 52 meshes with the driven gear 14. The transition from the meshing position to the disengaging position is effected by means of a tilting system 200 of the control system 100 and which is provided to allow the passage of the meshing position to the disengaged position and vice versa on command of the control unit 104. Here the tilting system 200 comprises a fixed base 202 and fixed on the landing gear 10 , a first link 204, a second link 206, elastic means comprising for example two compression springs and a jack 208 mounted in parallel with said elastic means. For reasons of readability, a single spring is shown, and this spring and the cylinder are each represented by two parallel lines bearing the reference 208, but the two springs and the cylinder are arranged one behind the other in a direction perpendicular to the plane of the leaf. The electric motor 50 is rotatably mounted on the base 202 about an axis parallel to the axis of the wheel 12. An end of the first link 204 is rotatably mounted on the electric motor 50.
    [0017] One end of the second link 206 is rotatably mounted on the base 202. The other end of the first link 204 and the other end of the second link 206 are rotatably mounted to each other. One end of each spring 208 is rotatably mounted at said other ends, and one end of the jack 208 is also rotatably mounted at said other ends. The other end of each spring 208 and the other end of the jack 208 are rotatably mounted on the base 202. The springs and the jack 208 are arranged in the angle formed between the two links 204 and 206. In position d meshing, the jack 208 is activated by the control unit 104 and pushes said other ends, which tends to bring the two rods 204 and 206 closer and thus to rotate the electric motor 50 to bring it closer to the driven gear 14, and the compression springs are then stretched.
    [0018] In the disengaged position, the jack 208 is deactivated by the control unit 104 and the compression springs contract, which reduces their lengths and brings said other ends closer together, which tends to separate the two rods 204 and 206 and thus to make pivoting the electric motor 50 away from the driven gear 14. For safety reasons, only the compression springs are used in the disengaged position and a possible failure of the cylinder will therefore not influence the position of the electric motor 50. The jack 208 may be an electric jack controlled directly by the control unit 104, or a hydraulic cylinder controlled by the control unit 141 through the introduction of a hydraulic supply derived from an existing hydraulic circuit on the aircraft 1. Apart from the out of service mode when a failure of the control system 100 is detected, the use of the tilt system t to the control system 100 to present three modes of operation: A non-activated mode, in which the control system 100 is not running and wherein the tilting system 200 maintains the disengaging position. A standby mode, wherein the control system 100 is running and wherein the tilt system 200 maintains the disengaging position. An activated mode in which the control system 100 is in operation and wherein the tilt system 200 maintains the meshing position. The transition from standby mode to activated mode is carried out for example according to the following diagram: - the two gears are unsheathed and the tilting system 200 maintains the disengaged position, - the control unit 104 checks the speed of the wheel 14 thanks to the speed sensor 110 and the control unit 104 accelerates the electric motor 50 until it reaches the speed of the wheel 14, - when the difference in speed between the wheel 14 and the electric motor 50 is below a threshold, the control unit 104 activates the jack 208 to move into the meshing position, and - the driving gear 52 and the driven gear 14 mesh with each other to roll the gear. aircraft 1.
    [0019] The transition from the activated mode to the standby mode is effected for example by deactivation of the jack 208 by the control unit 104, which causes the tilting system 200 to move to the disengaged position under the action of the springs 208, then by stopping the electric motor 50.5
    权利要求:
    Claims (10)
    [0001]
    CLAIMS1) Aircraft (1) comprising a landing gear (10) of which at least one wheel (12) is provided with an electric motor (50) configured to drive said wheel (12) in rotation, the aircraft (1) further comprising a control system (100) of said electric motor (50), said control system (100) comprising: - a control board (102) having first control means (106) for outputting a value of torque or power to be applied to said engine (50) in a direction of advance (16) of the aircraft (1) and second control means (108) provided to deliver a running speed value of the aircraft ( 1) in a reversing direction (18) of the aircraft (1), - a control unit (104) connected to the control panel (102) and to the electric motor (50), having first means for controlling the torque, or power, the electric motor (50) according to the value delivered by the first means control device (106), and second means for speed-controlling the electric motor (50) according to the running speed value supplied by the second control means (108), and - a speed sensor (110) for measuring the taxiing speed of the aircraft (1), and transmitting this speed information to the control unit (104).
    [0002]
    2) Aircraft (1) according to claim 1, characterized in that it comprises a cockpit (2), and in that the control panel (102) is disposed in the cockpit (2).
    [0003]
    3) Aircraft (1) according to claim 1 or claim 2, characterized in that the first control means (106) and the second control means (108) are constituted by the same rotary knob (150) having a zero position (0), wherein a rotation from the zero position to a first maximum angle in a first direction of rotation (112) is representative of the torque or power control of the electric motor (50), and wherein a rotation from the zero position to a second maximum angle in a second direction of rotation (114) is representative of the speed control of the electric motor (50).
    [0004]
    4) Aircraft (1) according to claim 1 or claim 2, characterized in that the first control means and the second control means are constituted by the same movable lever having a zero position, where a rotation of the lever, from the zero position towards the front of the aircraft (1) is representative of the control in torque, or in power, of the electric motor (50), and where a rotation of the lever, from the zero position to the rear of the aircraft (1) is representative of the speed control of the electric motor (50).
    [0005]
    5) Aircraft (1) according to one of claims 1 to 4, characterized in that the aircraft (1) comprises brakes for braking said at least one wheel (12) and a brake pedal whose active actuation said the control system (100) includes a brake pedal actuation sensor connected to the control unit (104), and that when the aircraft (1) moves in the reverse direction ( 18) and that the actuation sensor detects an actuation of the brake pedal, the control unit (104) controls the electric motor (50) to decelerate to a zero running speed.
    [0006]
    6) Aircraft (1) according to one of claims 1 to 5, characterized in that the control system (100) comprises at least one proximity sensor connected to the control unit (104), and that when the aircraft (1) moves in the reversing direction (18) and that the at least one proximity sensor detects a near obstacle, the control unit (104) controls the electric motor (50) so as to decelerate to zero speed.
    [0007]
    7) Aircraft (1) according to claim 6, characterized in that the sensor is arranged on the wings or the fuselage of the aircraft
    [0008]
    8) Aircraft (1) according to one of claims 1 to 7, characterized in that the aircraft (1) comprises a parking brake, said control system (100) comprises a parking brake sensor (116) connected to the control unit (104), and in that the control unit (104) is provided to remain in a zero speed control mode of the electric motor (50) as long as the parking brake sensor (116) ) does not indicate that the parking brake is engaged and as long as the speed sensor (110) does not indicate that the running speed is zero.
    [0009]
    9) Aircraft (1) according to one of claims 1 to 8, characterized in that said control system (100) comprises an on-off button (118), the type of stable mono push button.
    [0010]
    10) Aircraft (1) according to one of claims 1 to 9, characterized in that the electric motor (50) is equipped with a driving gear (52), in that the wheel (12) is equipped with a driven gear (14), and in that the control system (100) comprises a tilt system (200) controlled by the control unit (104) and which is provided to allow the passage of a meshing position wherein the drive gear (52) meshes with the driven gear (14) at a disengaging position in which the drive gear (52) does not engage the driven gear (14) and vice versa.
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    同族专利:
    公开号 | 公开日
    FR3017367B1|2017-02-24|
    US20150225075A1|2015-08-13|
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    法律状态:
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    FR1451003A|FR3017367B1|2014-02-10|2014-02-10|AIRCRAFT COMPRISING A LANDING TRAIN WITH A WHEEL PROVIDED WITH AN ELECTRIC MOTOR AND A CONTROL SYSTEM OF SAID ELECTRIC MOTOR|FR1451003A| FR3017367B1|2014-02-10|2014-02-10|AIRCRAFT COMPRISING A LANDING TRAIN WITH A WHEEL PROVIDED WITH AN ELECTRIC MOTOR AND A CONTROL SYSTEM OF SAID ELECTRIC MOTOR|
    US14/618,356| US20150225075A1|2014-02-10|2015-02-10|Aircraft comprising a landing gear having one wheel provided with an electric motor and control system for said electric motor|
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