• 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:
    A method of controlling an electric motor for rotating an aircraft wheel (4a, 4b) for generating a torque command for controlling the engine, the method being characterized by comprising implementing: a first servocontrol loop (23) having as its input signal a speed reference (Cons_v), for a return signal a signal representing the speed of the wheel (Vr) or the aircraft, and for an output signal an acceleration instruction (Cons_a); a second servocontrol loop (24) having as its input signal the acceleration setpoint (Cons_a), for a return signal a signal representative of the acceleration of the wheel (Ar) or of the aircraft, and for output signal the torque control
    公开号:FR3015707A1
    申请号:FR1363325
    申请日:2013-12-20
    公开日:2015-06-26
    发明作者:Clement Gorce；David Lemay
    申请人:Messier Bugatti Dowty SA；
    IPC主号:
    专利说明:

    [0001] The invention relates to a method for controlling an electric motor driving in rotation of an aircraft wheel. BACKGROUND OF THE INVENTION Currently, the control of the ground speed of an aircraft is carried out manually by a pilot of the aircraft either by a control of the engine thrust produced by the propulsion engines of the aircraft, or by using the braking system of the wheels of the aircraft, or, more rarely, by a combined control of thrust and braking. This way of controlling the ground speed of the aircraft is not very precise and obliges the pilot to adapt the piloting to the environmental conditions, to the lining of the runway, and to the structural characteristics of the aircraft (mass, etc.). . The control of the orientation of the aircraft, meanwhile, is performed manually by the pilot who uses a device for steering the front wheels of the aircraft. The pilot controls in real time a rotation angle of the front wheels so that the aircraft follows a sought-after trajectory. This way of controlling the orientation of the aircraft requires that the pilot performs many piloting operations in real time. OBJECT OF THE INVENTION The subject of the invention is a method for controlling an electric motor for rotating an aircraft wheel, making it possible to control the ground speed of the aircraft more precisely, more precisely. repeatable, and providing assistance in controlling the orientation of the aircraft. SUMMARY OF THE INVENTION In view of accomplishing this object, there is provided a method of controlling an electric motor for rotating an aircraft wheel for generating torque control for controlling the engine. According to the invention, the method comprises the implementation of: a first servo-control loop having as its input signal a speed reference, for a return signal a signal representative of the speed of the wheel or the aircraft, and for an output signal an acceleration instruction; a second servocontrol loop whose input signal is the acceleration setpoint, for a return signal a signal representative of the acceleration of the wheel or the aircraft, and for an output signal the torque command . The use of electric motors for rotating aircraft wheels and regulating the speed and acceleration of the torque control controlling these engines makes it possible to achieve precise ground speed control and very little sensitivity to the the external environment and structural characteristics of the aircraft. The implementation of such a regulation also makes it possible to apply a torque control specific to each wheel. The torque control can therefore be corrected according to the undercarriages and according to the position of the wheels within a landing gear according to the orientation angle of the desired aircraft, which makes it possible to provide assistance to the control of the aircraft. orientation. BRIEF DESCRIPTION OF THE DRAWINGS The invention will be better understood in the light of the description which follows with reference to the figures of the accompanying drawings in which: FIGS. 1a and 1b form two parts of a single figure which represents a block diagram illustrating the control method of the invention; FIG. 2 is a graph illustrating a function of limiting the acceleration of the wheel as a function of the ground speed of the aircraft that can be used in the first servo-control loop; FIG. 3 is a graph illustrating the effect of the limitation function on the torque applied by the electric motor as a function of the speed of the wheel; FIG. 4 is a graph illustrating a first maximum torque, a torque command for the electric motor and a torque margin existing between the first maximum torque and the torque command, as a function of time; - Figure 5 is a graph similar to that of Figure 4 directly representing the margin of torque; - Figure 6 is a schematic view of the wheels of the aircraft during a turn of the aircraft. DETAILED DESCRIPTION OF THE INVENTION Referring to FIGS. 1a and 1b, the invention is here implemented in an aircraft 1 of the Airbus A320 type comprising a first main undercarriage 2a provided with a first inner wheel 3a and a first wheel. 4a exterior, a second undercarriage 2b provided with a second inner wheel 3b and a second outer wheel 4b and an undercarriage with two steerable wheels (not visible in Figure 1). Each outer wheel 4a, 4b is here equipped with a respective electromechanical drive actuator 5a, 5b for driving the wheel in rotation. The aircraft 1 can thus be moved to the ground without using the propulsion engines of the aircraft. Each electromechanical drive actuator 5a, 5b comprises a respective electric motor 6a, 6b, here a synchronous three-phase permanent magnet motor. The electric motors of the electromechanical drive actuators 5 are controlled by a pilot of the aircraft 1. By acting on a ground control lever 7 of the aircraft 1 located in the cockpit of the aircraft 1, the pilot generates a speed command Ov which is transmitted to a centralized control unit 8.
    [0002] The centralized control unit 8 converts this speed command Ov into a first torque command Ccl for a first power unit 10a associated with the electric motor 6a of the first outer wheel 4a, and a second torque control Cc2. to a second power unit 10b associated with the electric motor 6b of the second outer wheel 4b. The power units 10 convert the torque commands into control currents supplied to the motors 6 so that they deliver the corresponding torques to the torque commands. The centralized control unit 8 comprises an acquisition module 12 and a processing module 13. The acquisition module 12 is connected to a certain number of on-board equipment communicating with the centralized control unit 8, and is intended to receive, transmit, and possibly format data exchanged with these onboard equipment. Among these embedded equipment, there is of course the joystick 7 mentioned earlier. There are also a number of sensors: wheel speed sensors 15 associated with the first outer wheel 4a and the second outer wheel 4b, current sensors 16 associated with each electric motor 6a, 6b of the electromechanical actuators 5a, 5b, an orientation sensor 17 of the steerable wheels of the undercarriage of the aircraft 1, etc. Finally, various calculators or data concentrators belonging to various systems of the aircraft 1 cooperating with the control unit are found: aircraft braking system 18, ADIRU 19 (Air Data Inertial Reference Unit) system the data control unit such as the ground speed Vs of the aircraft 1, an electric power controller 20 supplying the centralized control unit with data relating to an electrical or thermal state of the electrical power generators of the aircraft, etc. The acquisition module 12 converts the speed order Ov into a rotation speed setpoint Cons v common to the electric motor 6a of the first outer wheel 4a and to the electric motor 6b of the second outer wheel 4b. The processing module 13, for its part, implements the control method of the invention, which consists in regulating the speed and acceleration of a torque control of each electric motor 6. The processing modulus 13 receives for it the acquisition module 12 the speed setpoints Cons v and generates the first torque command Ccl and the second torque control Cc2.
    [0003] For each motor 6 of each outer wheel 4, the method of the invention comprises the implementation of a first servo-control loop 23 and a second servo-control loop 24. For each motor 6, the first control loop The servocontrol 23 has as its input signal the speed reference Cons v. The first control loop 23 comprises a first subtracter 27 which subtracts from the speed setpoint Cons v a feedback signal, in this case a signal representative of the speed of rotation of the wheel Vr, and which thus calculates an error of speed sv. The signal representative of the speed of rotation of the wheel Vr is here a measured rotation speed Vma for the first outer wheel 4a and a measured rotation speed Vmb for the second outer wheel 4b. These rotational speeds measured Vma, Vmb are measured by the speed sensors 15 of the first and second outer wheels 4a, 4b and transmitted to the interface module 12. The interface module 12 generates the signal representative of the The first servo loop 23 further includes a speed controller for converting the speed error into an acceleration command Comm a. The speed controller is here a first gain block 28 for multiplying the speed error sv by a predetermined gain. The first servocontrol loop 23 furthermore comprises a limiting function 29 for limiting the acceleration control Comm a and generating an acceleration setpoint Cons used as the input signal of the second servocontrol loop 24. The function of limitation is an acceleration slope as a function of the ground speed Vs of the aircraft 1, said ground speed Vs being transmitted by the ADIRU type system 19 to the centralized control unit 8. Referring to FIG. this acceleration slope is here defined by four controllable constants: a first acceleration constant cal, a second acceleration constant ca2, a first velocity constant cv1 and a second velocity constant cv2. With reference to FIG. 3, the first acceleration constant cal defines a stabilized pair Cs required during a constant acceleration of the outer wheel 4 of the aircraft 1. The higher the value of this first acceleration constant cal, the higher the value of this first acceleration constant cal. the stabilized torque Cs that can deliver the electric motor 6 of the outer wheel 4 is important. The adjustment of this constant cal makes it possible to adapt the stabilized torque Cs to a rolling resistance of the outer wheel 4. The second acceleration constant ca2 is used to define a maximum initial torque Cim required to drive the outer wheel 4 when the aircraft 1 is stationary. The higher the value of this second acceleration constant CA2, the greater the initial initial torque Cim that can be delivered by the electric motor 6 of the outer wheel 4.
    [0004] The first rate constant cv1 is used to define a transition torque Ct corresponding to the difference between the initial torque and the stabilized torque. The value of the first speed constant cv1 must be close to a speed value of the outer wheel 4 when the constant acceleration is reached, which makes it possible to avoid a significant torque drop between the initial torque and the stabilized torque. Finally, the second speed constant cv2 is used to define a time T necessary for the motor 6 to, from the moment when the initial torque is applied, reach the stabilized torque: the higher the value of this second speed constant is high. the more this time T is important. The second control loop 24 has for entry the acceleration setpoint Cons generated by the first control loop 23. The second control loop 24 comprises a second subtractor 31 which subtracts the acceleration setpoint Cons has a return signal, in this case a signal representative of the acceleration of the wheel Ar, and which thus calculates an acceleration error sa. The signal representative of the acceleration of the wheel Ar is here obtained by deriving the signal representative of the speed of the wheel Vr, that is to say by deriving the speed Vma for the first outer wheel 4a and the speed Vmb for the second outer wheel 4b. The derivation is carried out by a shunting block 32. The second servocontrol loop further comprises a first branch 33 comprising a low-pass filter 34 and a second branch 35 parallel to the first branch 33 having an integrator 36. Each of the branches for input the acceleration error sa. The low-pass filter 34 of the first branch 33 is a first-order filter intended to make the control sufficiently responsive to disturbances such as those associated with driving the wheel at a slope of a track on which roll the aircraft, wind, etc. The low-pass filter 34 has for input the acceleration error sa and for output a first filtered pair C fill. The low-pass filter 34 has the following transfer function: K. 11-T.z-1, where K1, Ti are adjustable constants. The integrator 36, meanwhile, generates no static error and allows the regulation to have an acceptable response time. The integrator 36 has for input the acceleration error sa and for output a second filtered pair C fil2. The integrator 36 here has the transfer function:, where K2 is an adjustable constant. The first branch 33 furthermore comprises a first torque saturation function 39 having, as input, the first filtered torque C fill and for output a first torque output C satl. The second branch 34 has a second torque saturation function 40 having the second filtered torque C wire 2 for input and outputting a second torque output C sat2. The first saturation function 39 defines a first torque threshold Scl which limits the first filtered pair. The second saturation function defines a second torque threshold Sc2 having for value the value of the first torque threshold Scl at which the value of the output of the low-pass filter, that is to say the first filtered pair C, is subtracted. fill.
    [0005] Thus, if the first filtered pair C fill has a maximum value equal to 120 Nm (newtons-meter) and if the value of the first torque threshold Scl is 200 Nm, then the second output of torque C sat2 is limited by a couple 80 Nm
    [0006] The second control loop 24 finally comprises a first summer 42 which adds the first torque output C satl and the second torque output C sat2 and thus generates the torque command for the motor of the electromechanical actuator of the wheel. The method of the invention is thus implemented to generate a first torque command Ccl to the power unit 10a of the electric motor 6a associated with the first outer wheel 4a and a second torque control Cc2 to the power unit 10b of the electric motor 6b of the second outer wheel 4b. The method of the invention further comprises the implementation of an assistance function 41 to the orientation of the aircraft 1. This assistance function 41 is intended to correct the first torque command Ccl and the second torque control Cc2 according to an orientation angle has steerable wheels of the undercarriage of the aircraft 1, so as to assist the steerable wheels of the undercarriage to orient the aircraft 1. The assistance function 41 is implemented in the processing module 13, and has as inputs the first torque command Ccl, the second torque command Cc2 and the orientation angle cx. The assistance function 41 comprises a third and a fourth subtracter 43, 44, a unit delay block 45, a second, a third and a fourth gain block 46, 47, 48, a second summator 49 and two multipliers 50, 51 The third subtractor 43 subtracts from the orientation angle at a delayed orientation angle cxr resulting from the action of the unit delay block 45 on the angle of rotation cx. The result of this subtraction is an angle variation Ac, which is zero when the orientation angle a is constant in time and not zero otherwise. This angle variation Acx is then multiplied by the second gain block 46 having a configurable gain and transforming the angle variation Acx into a correction torque C corr. This correction torque C corr is applied at the input of the second summer 49 and at the input of the fourth subtractor 44. The third and fourth gain blocks 47, 48, which have a unit gain, are respectively connected to a second input of second summator 49 and a second input of the fourth subtractor 44. The output of the second summator 49 is connected to a first input of the first multiplier 50. The output of the fourth subtractor 44 is connected to a first input of the second multiplier 51.
    [0007] The first torque command Ccl and the second torque command Cc2 are respectively connected to a second input of the first multiplier 50 and to a second input of the second multiplier 51. The output of the first multiplier is a first corrected torque command C corrl to the engine 6a of the first outer wheel 4a. The output of the second multiplier 51 is a corrected second corrected torque command C to the motor 6b of the second outer wheel 4b.
    [0008] The difference between the first corrected torque control C corrl and the second corrected torque control C corr2 is a differential torque C diff which tends to orient the aircraft 1 in the direction of the steering angle cx.
    [0009] Advantageously, the assistance function 41 to the orientation of the aircraft 1 is activated and deactivated by pin-programming. The centralized control unit is equipped for this purpose with an integrated configuration connector. By connecting to this integrated configuration connector an external configuration connector provided with pins having a first electrical configuration, the assistance function 41 is activated to the orientation of the aircraft 1. By connecting to this connector of FIG. In the integrated configuration, the external configuration connector whose pins have a second electrical configuration, the assistance function 41 is activated to the orientation of the aircraft 1. With reference to FIGS. 4 and 5, the method of the invention further comprises the implementation, for each electric motor, an anti-saturation function. The anti-saturation function is performed in the power unit 10 associated with the electric motor 6. The anti-saturation function calculates a margin of torque AC existing between a first maximum torque C max for the electric motor 6 in question. , and corrected torque control C corrl, C corr2 at the output of the second servo loop 24. If the torque margin AC is positive, the anti-saturation function allows a larger torque command. If the AC torque margin is negative, the anti-saturation function decreases the torque control to the maximum allowable torque. The first maximum torque C max1 depends on the characteristics of the motor and the speed of rotation of the motor. The torque margin AC visible in FIG. 4 corresponds to a ground displacement of the aircraft 1 towards the front. To define the AC torque margin in the case of a displacement of the aircraft 1 towards the rear, an absolute value of the corrected torque command Ccorr is used which is subtracted at the first maximum torque C max to obtain a margin of torque AC similar to the margin corresponding to a ground movement of the aircraft forward. The method further comprises the implementation of a limitation of the corrected torque control C corrl, C corr2 as a function of the electric current consumed by the motors to avoid overloading the electrical power generators of the aircraft that provide the units of power the electric power necessary for the operation of the electric motors.
    [0010] The torque is thus limited by a second maximum torque C max2. The overload data of the generators are supplied to the centralized control unit 8 by the electric power controller 20.
    [0011] The method further comprises the implementation of a limitation of the corrected torque control C corrl, C corr2 as a function of the temperature of the electric generators, to prevent this temperature from becoming too important. The torque is thus limited by a third maximum torque C max3. The data relating to the temperature of the generators are also provided to the centralized control unit 8 by the electric power controller 20.
    [0012] The anti-saturation function and the limitations of the corrected torque control are carried out in the power unit 10 associated with the electric motor 6. In another embodiment, the anti-saturation function and the limitations of the corrected torque control are realized. simultaneously by implementing a combined limitation function. This combined limitation function calculates a maximum permissible maximum torque C maxg equal to the minimum of the first, second and third maximum couples C max, C max2, C max3, and limits the torque control corrected by this maximum permissible torque C maxg . It will be noted that the process of the invention is hereby inhibited in certain particular situations. A first particular situation occurs when the aircraft 1 is moving forward on the ground and a negative speed is controlled. A second particular situation occurs when the aircraft 1 is moving backwards and a positive speed is controlled. A third particular situation occurs when a speed lower than a present speed of the aircraft 1 is controlled. A fourth particular situation occurs when braking of an aircraft wheel 1 is controlled via the braking system. A fifth particular situation occurs when too much wheel speed is controlled. The process is inhibited in the first three situations to prevent a regenerative braking phenomenon from occurring: the power supply of the engine 6 would not lead to driving the wheels but would tend to slow them down and produce energy tending to increase the temperature of the electric motor. The method is inhibited in the fourth situation to avoid any loss of braking efficiency and any mechanical damage impacting the motor 6 or the drive actuator 5.
    [0013] The method is inhibited in the fifth situation to avoid any mechanical or electrical damage affecting the motor 6 or the drive actuator 5. In a second embodiment, and with reference to FIG. 6, the speed reference Cons_v and the acceleration setpoint Cons_a of the two outer wheels 4a, 4b are specific to each of the wheels: the regulation in speed and acceleration of the two outer wheels 4a, 4b is achieved via two parallel and independent regulations.
    [0014] In the second embodiment, a speed reference Cons v1 of the first outer wheel 4a and a speed setpoint Cons_v2 of the second outer wheel 4b are obtained from the ground speed of the aircraft V, the angle d Orientation to Orientable Wheels and Structural Characteristics of the Aircraft 1. Thus, the speed setpoint Cons_v1 of the first outer wheel 4a is equal to: L Cons (1- tan cr) .Vs, the setpoint velocity v2 Cons of the second outer wheel 4b is equal to: L 1 + w Cons = (1+ L = tan a), Its with L = 2 where Ly is the distance between a longitudinal central axis of the aircraft 1 and a center point Pc of an outer wheel 4a or 4b, ly is the distance between the centers Cel and Cet of the axles of the two main landing gear, and w is the distance between two central points of the two wheels of the same main landing gear. The invention is not limited to the particular embodiments that have just been described, but, on the contrary, covers any variant within the scope of the invention as defined by the claims. Although it has been chosen to equip drive actuators with the outer wheels of the main landing gears of the aircraft, the method of the invention can of course be implemented on one or more wheels. Similarly, the invention is of course applicable to aircraft having a different number of landing gear, a different number of wheels, or a different wheel arrangement on the undercarriages.
    [0015] To illustrate the invention, a control architecture comprising a centralized unit comprising an acquisition module and a processing module, power units, etc. has been used. The invention can of course be implemented in a different architecture having a different distribution of functions inside different equipment. For example, the centralized unit could have a centralized power module replacing the power units, and so on.
    [0016] It is furthermore observed that the control method of the invention can be implemented very simply in any type of aircraft of which at least one wheel is equipped with an electric motor for rotating said wheel. To implement the control method of the invention in an existing control architecture, it suffices to program the control loops in a centralized unit of this architecture. It is therefore not necessary to add electrical equipment to the existing architecture, to replace cables, etc. Similarly, although it has been indicated that the speed command is transmitted directly to the centralized control unit following a pilot action on the joystick, the speed command can also be generated by a calculator , especially in the context of an automatic rolling of the aircraft. Although it has been indicated that the signal representative of the acceleration of the wheel Ar is obtained by derivation of the signal representative of the speed of the wheel Vr, this signal can also be obtained by a measurement of the acceleration of the wheel, made for example by an accelerometer positioned on the wheel. Similarly, although it has been described that the feedback signal of the first servo loop is a signal representative of the speed of the wheel, and that the feedback signal of the second servo loop is a representative signal of the acceleration of the wheel, it is possible to use as return signals a signal representative of the speed of the aircraft and a signal representative of the acceleration of the aircraft, the speed of the aircraft aircraft and acceleration of the aircraft being preferably a longitudinal speed of the aircraft and a longitudinal acceleration of the aircraft. These representative signals are obtained by measurement, or estimation, or by calculation, etc. A first torque saturation function as well as an anti-saturation function and limitations of the torque control corresponding to limitations by maximum couples C max, C max2, C max3 have been mentioned. It is furthermore possible to provide additional limiting and saturation functions to implement additional protection measures, by defining at least one new maximum limiting torque, without modifying either structurally or operationally the regulation. These additional protective measures are intended, for example, to protect mechanical elements of the actuator, or to protect the electrical power generators of the aircraft against overvoltage, etc. When the anti-saturation function and the limitations of the corrected torque control are realized simultaneously by the implementation of a combined limitation function, it is then sufficient to recalculate the maximum permissible torque C maxg taking into account of this new maximum torque limitation.
    权利要求:
    Claims (8)
    [0001]
    REVENDICATIONS1. A method of controlling an electric motor for rotating an aircraft wheel (4a, 4b) for generating a torque command for controlling the engine, the method being characterized by comprising a first servo loop (23) having as its input signal a speed reference (Cons_v), for a return signal a signal representative of the speed of the wheel (Vr) or of the aircraft, and for an output signal an acceleration instruction (Cons_a); a second servocontrol loop (24) having as its input signal the acceleration setpoint (Cons_a), for a return signal a signal representative of the acceleration of the wheel (Ar) or of the aircraft, and for output signal the torque command.
    [0002]
    2. Control method according to claim 1, wherein the first control loop comprises a first subtractor (27) for calculating a speed error (ev) by subtracting at the speed reference the signal representative of the wheel speed, a gain block for multiplying the speed error (ev) by a predetermined gain so as to generate an acceleration command, and a limiting function for limiting the acceleration control and generating the instruction acceleration.
    [0003]
    3. Control method according to claim 2, wherein the limiting function is an acceleration slope as a function of a ground speed (Vs) of the aircraft (1).
    [0004]
    4. Control method according to one of the preceding claims, wherein the second control loop (24) comprises a second subtractor (31) for calculating an acceleration error (Ea) byretrancenant to the acceleration setpoint signal representative of the acceleration of the wheel, a first branch (33) comprising a low-pass filter (34), a second branch (35) parallel to the first branch (33) comprising an integrator (36), and a summator (42) for adding signals from the first and second branches.
    [0005]
    5. Control method according to one of the preceding claims for controlling the electric motors of a first and a second aircraft wheels, wherein, for the motor (6a) of the first wheel (4a). and for the motor (6b) of the second wheel (4b), the speed reference is a setpoint common to both engines.
    [0006]
    6. Control method according to claim 5 further comprising the implementation of an assistance function (41) to the orientation of the aircraft (1), said assistance function having as inputs the output of the second control loop (24) of the motor of the first wheel (4a), the output of the second servo control loop (24) of the motor of the second wheel (4b), and an orientation angle (a) d a front wheel of the aircraft, and for output a corrected first torque command of the motor of the first wheel and a second corrected torque control of the motor of the second wheel.
    [0007]
    7. Control method according to claim 6, wherein the assistance function (41) to the orientation of the aircraft (1) is activated and deactivated by pin-programming.
    [0008]
    8. Control method according to one of claims 1 to 4 for controlling the electric motors of a first and second aircraft wheels, wherein for the motor of the first wheel (4a) and for the motor of the second wheel (4b), the velocity setpoint is specific to each wheel and the signal representative of the speed of each wheel is a measure of speed of each wheel.
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    同族专利:
    公开号 | 公开日
    US9452826B2|2016-09-27|
    EP2886454A1|2015-06-24|
    US20150175257A1|2015-06-25|
    CN104724293B|2017-11-21|
    FR3015707B1|2017-04-21|
    EP2886454B1|2018-04-04|
    CN104724293A|2015-06-24|
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    法律状态:
    2015-06-26| PLFP| Fee payment|Year of fee payment: 3 |
    2016-12-22| PLFP| Fee payment|Year of fee payment: 4 |
    2017-06-23| CD| Change of name or company name|Owner name: MESSIER-BUGATTI-DOWTY, FR Effective date: 20170518 |
    2017-12-21| PLFP| Fee payment|Year of fee payment: 5 |
    2019-11-20| PLFP| Fee payment|Year of fee payment: 7 |
    2020-11-20| PLFP| Fee payment|Year of fee payment: 8 |
    2021-11-17| PLFP| Fee payment|Year of fee payment: 9 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1363325A|FR3015707B1|2013-12-20|2013-12-20|METHOD FOR CONTROLLING AN ELECTRIC MOTOR DRIVING ROTATION OF AN AIRCRAFT WHEEL|FR1363325A| FR3015707B1|2013-12-20|2013-12-20|METHOD FOR CONTROLLING AN ELECTRIC MOTOR DRIVING ROTATION OF AN AIRCRAFT WHEEL|
    EP14197769.4A| EP2886454B1|2013-12-20|2014-12-12|A method of controlling an electric motor for driving rotation of an aircraft wheel|
    US14/568,185| US9452826B2|2013-12-20|2014-12-12|Method of controlling an electric motor for driving rotation of an aircraft wheel|
    CN201410811680.5A| CN104724293B|2013-12-20|2014-12-19|The method for controlling the electric notor of the rotation for driving aircraft wheel|
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