An aircraft turning system incorporating a front wheel and a main wheel

Abstract

The application provides an aircraft turning system combining front wheels and main wheels, comprising a control template, a switch input unit, an analog input unit and an analog output unit; the control template is used for executing a cooperative control strategy: a turning radius is calculated according to a front wheel turning instruction sensing signal and a real-time taxiing speed parameter; the turning radius and the structural parameters among the front wheels, the inner main wheels and the outer main wheels are used to calculate the inner main wheel turning angle and the outer main wheel turning angle, and the inner main wheel turning servo signal and the outer main wheel turning servo signal are output; the error between the inner main wheel angle / position feedback sensing signal and the inner main wheel turning angle is calculated to calculate the front wheel turning compensation angle, and the front wheel turning servo signal is output. The application proposes an inner-outer wheel asymmetric control strategy, which can realize the asymmetric cooperative control of the two-side main wheels, and simultaneously realizes the real-time steering angle feedback control of the front wheels, and realizes the synchronous control of the aircraft turning between the front wheels and the main wheels.

Description

Technical Field

[0001] This invention belongs to the field of aircraft turning technology, and relates to an aircraft turning system that combines the front wheel and the main wheel. Background Technology

[0002] Existing wide-body aircraft often employ a multi-wheel trolley-type main landing gear structure. During bottom-plane turns, wide-body passenger aircraft frequently experience skidding and tire wear. To address these issues, research has emerged on coordinated turning control strategies involving both the nose wheel and main landing gear, as seen in wide-body aircraft such as the A380 and B777. In these multi-wheel trolley-type main landing gear structures, the main landing gear is often considered as a single unit in solving the coordinated turning problem between the nose wheel and main landing gear.

[0003] However, due to the different turning radii of the two main landing gears, the steering angle of the main landing gear with the smaller turning radius is greater than that of the main landing gear with the larger turning radius. This causes additional lateral loads on the main landing gear and increases tire wear. Excessive lateral loads can even lead to the main landing gear lower locking system unlocking and retracting. Simultaneously, when the aircraft turns on the ground, there is a discrepancy between the actual coordinated turning angle of the main landing gear and the commanded turning angle of the nose wheel. This can cause the nose and main landing gear to lose synchronization in controlling the aircraft's steering, which can also easily lead to lateral loads and tire skidding.

[0004] Therefore, how to provide an aircraft turning system that can achieve asymmetric cooperative control of the two main wheels while providing real-time steering angle feedback control of the front wheel is a problem that urgently needs to be solved by those skilled in the art. Summary of the Invention

[0005] In view of this, the present invention proposes an aircraft turning system that combines the front wheel and the main wheel to solve the above-mentioned problems in the prior art.

[0006] To achieve the above objectives, the present invention adopts the following technical solution:

[0007] This invention discloses an aircraft turning system that combines a front wheel and a main wheel, comprising: a control module, a digital input unit, an analog input unit, and an analog output unit;

[0008] The digital input unit is used to acquire the front wheel wheel-mounted switch signal, the front wheel turning / sway reduction status switch signal, the inner main wheel wheel-mounted switch signal, and the outer main wheel wheel-mounted switch signal, and send them to the control template.

[0009] The analog input unit is used to acquire the front wheel turning command sensing signal, the front wheel angle / position feedback sensing signal, the inner main wheel angle / position feedback sensing signal, and the outer main wheel angle / position feedback sensing signal, and send them to the control template.

[0010] The analog output unit is used to output the front wheel turning servo signal, the inner main wheel turning servo signal, and the outer main wheel turning servo signal;

[0011] The control template is used to execute the following collaborative control strategy:

[0012] The coordinated control strategy is activated based on the front wheel on-board switch signal, the inner main wheel on-board switch signal, and the outer main wheel on-board switch signal.

[0013] The front wheel turning angle is calculated based on the front wheel turning command sensor signal;

[0014] The real-time taxiing speed parameters output by the main wheel brake control system are obtained, and the real-time turning arc length of the aircraft is calculated. The turning radius is then calculated by combining the turning angle of the front wheel.

[0015] The inner main wheel angle and outer main wheel angle are calculated using the turning radius and structural parameters between the aircraft's nose wheel, inner main wheel, and outer main wheel, and then sent to the analog output unit to output the inner main wheel turning servo signal and outer main wheel turning servo signal.

[0016] The error between the inner main wheel angle / position feedback sensor signal and the inner main wheel rotation angle is calculated; the error between the outer main wheel angle / position feedback sensor signal and the outer main wheel rotation angle is calculated.

[0017] The front wheel turning compensation angle is calculated based on the error and sent to the analog output unit to output the front wheel turning servo signal.

[0018] Preferably, the switch input unit is further configured to acquire the front landing gear lower position lock switch signal and send it to the control template, which controls the start and stop of the system power supply according to the front landing gear lower position lock switch signal.

[0019] Preferably, the switch input unit is also used to acquire the front wheel turning gain switching switch signal and send it to the control template, which determines in real time whether the real-time coasting speed parameter output by the main wheel brake control system is lower than the set maximum coasting speed threshold.

[0020] If not, switch to the front wheel steering gain off signal to trigger the cooperative control strategy to stop;

[0021] If so, switch to the front wheel turning gain on signal, determine whether the front wheel angle / position feedback sensor signal is higher than the set minimum front wheel turning angle threshold, and if so, trigger the start of the cooperative control strategy.

[0022] Preferably, the step of acquiring the real-time taxiing speed parameters output by the main wheel brake control system, calculating the real-time turning arc length of the aircraft, and calculating the turning radius by combining the turning angle of the front wheel specifically includes:

[0023]

[0024] Where l is the turning arc length of the aircraft in ground mode, t is the real-time taxiing time of the aircraft, and V is the taxiing speed of the aircraft.

[0025]

[0026] Where R is the turning radius of the aircraft in ground mode; γ is the turning angle of the front wheels.

[0027] Preferably, the calculation of the inner main wheel angle and outer main wheel angle using the turning radius and the structural parameters between the aircraft's nose wheel, inner main wheel, and outer main wheel specifically includes:

[0028] The structural parameters between the aircraft's nose wheel, inner main wheel, and outer main wheel include: the wheelbase h between the inner and outer main wheels, and the distance s between the nose wheel and the centerline of the inner and outer main wheels;

[0029] The inner main wheel rotation angle and the outer main wheel rotation angle are calculated as follows:

[0030]

[0031]

[0032] Where: α is the inner main wheel rotation angle; β is the outer main wheel rotation angle; R is the turning radius of the aircraft in ground mode.

[0033] Preferably, the control template adopts a DSP+CPLD structure, and the digital input unit, analog input unit, and analog output unit are opto-isolated from all I / O signals of the DSP of the control template.

[0034] Preferably, the DSP outputs a PWM signal, which is then shaped by a buffer-driven optocoupler, a Smith trigger, and a filtering circuit. After passing through a V / I conversion circuit, the signal obtains the front wheel turning servo signal, the inner main wheel turning servo signal, and the outer main wheel turning servo signal, which then drive the corresponding servo valve coil.

[0035] Preferably, it also includes a switch output unit for outputting the front wheel turning status signal, the front wheel turning fault signal, the front wheel turning status switching valve status, the main wheel turning status signal, the main wheel turning fault signal, and the main wheel turning status switching valve status.

[0036] Preferably, the aircraft turning system includes a dual-channel main control device, including: a digital input unit, an analog input unit, an analog output unit, and a digital output unit, all of which are configured in dual-channel mode; the interface module of the main control module is in dual-channel hot standby mode.

[0037] The main control module includes a dual-channel logic switching circuit and a BIT fault detection circuit. The BIT fault detection circuit is used to automatically test and detect faults in various electrical components involved in the aircraft steering function.

[0038] Under normal circumstances, the dual-channel main control device relies on the communication information between the two channels to perform cross-switching and automatically find a channel without control component failure. When the detection component of this channel is fault-free, a fault-free signal is output.

[0039] When the control components of the cross passage are normal but the BIT detection components are faulty, a fault signal is output, prompting a turning operation without real-time optimization and real-time coordination.

[0040] When one or both channels of the dual-channel main controller fail simultaneously, a signal indicating failure of the front wheel and main wheel turning function is output.

[0041] As can be seen from the above technical solution, compared with the prior art, the present invention has the following beneficial effects:

[0042] 1. This invention fully considers the asymmetric difference in steering angle between the inner and outer main wheels when a wide-body passenger aircraft with a multi-wheel trolley structure turns, and proposes an asymmetric control strategy for the inner and outer wheels;

[0043] 2. This invention utilizes the error between the actual feedback sensing signals of the inner and outer main wheel rotation angles and the calculated inner and outer main wheel rotation angles to determine whether they match the commanded rotation angle of the front wheel in real time, thereby effectively solving the problem of the inability to synchronously control the aircraft's steering between the front and main wheels;

[0044] 3. This invention uses the nose landing gear lower position lock switch signal as the trigger signal for starting and stopping the power supply of the aircraft steering control system, effectively avoiding excessive lateral loads that could even cause the main landing gear lower position lock to unlock and retract.

[0045] 4. The present invention also has fault detection and channel switching functions to ensure normal system operation. If the control strategy cannot be implemented, the system can be automatically switched to a yaw reduction state to ensure aircraft safety. Attached Figure Description

[0046] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on the provided drawings without creative effort.

[0047] Figure 1 This is a schematic diagram of the overall architecture of an aircraft turning system combining the front wheel and the main wheel in an embodiment of the present invention;

[0048] Figure 2 This is a flowchart of the collaborative control strategy for controlling template execution in an embodiment of the present invention;

[0049] Figure 3 This is a structural diagram of the control template in an embodiment of the present invention. Detailed Implementation

[0050] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0051] See Figure 1 This invention discloses an aircraft turning system that combines a front wheel and a main wheel, comprising: a control module, a digital input unit, an analog input unit, and an analog output unit;

[0052] The digital input unit is used to acquire the front wheel on-board switch signal, the front wheel turning / sway reduction status switch signal, the inner main wheel on-board switch signal, and the outer main wheel on-board switch signal, and send them to the control module;

[0053] The analog input unit is used to acquire the front wheel turning command sensing signal, the front wheel angle / position feedback sensing signal, the inner main wheel angle / position feedback sensing signal, and the outer main wheel angle / position feedback sensing signal, and send them to the control module; all the sensing signals are LVDT dual signals.

[0054] The analog output unit is used to output the front wheel steering servo signal, the inner main wheel steering servo signal, and the outer main wheel steering servo signal.

[0055] See Figure 2 The control template is used to execute the following collaborative control strategies:

[0056] The coordinated control strategy is activated based on the front wheel on-board switch signal, the inner main wheel on-board switch signal, and the outer main wheel on-board switch signal to ensure that turning is only possible in ground mode.

[0057] The front wheel turning angle is calculated based on the front wheel turning command sensor signal;

[0058] The real-time taxiing speed parameters output by the main wheel brake control system are obtained, and the real-time turning arc length of the aircraft is calculated. The turning radius is then calculated by combining the turning angle of the front wheel.

[0059] The turning radius and structural parameters between the aircraft's nose wheel, inner main wheel, and outer main wheel are used to calculate the inner main wheel angle and the outer main wheel angle, and then sent to the analog output unit to output the inner main wheel turning servo signal and the outer main wheel turning servo signal.

[0060] The error between the inner main wheel angle / position feedback sensor signal and the inner main wheel rotation angle is calculated; the error between the outer main wheel angle / position feedback sensor signal and the outer main wheel rotation angle is calculated.

[0061] The front wheel turning compensation angle is calculated based on the error and sent to the analog output unit to output the front wheel turning servo signal.

[0062] In one embodiment, the digital input unit is also used to acquire the nose landing gear lower position lock switch signal and send it to the control module, which controls the start and stop of the system power supply (DC28V) according to the nose landing gear lower position lock switch signal.

[0063] In one embodiment, the digital input unit is also used to acquire the front wheel steering gain switching switch signal and send it to the control module. The control module determines in real time whether the real-time coasting speed parameter output by the main wheel brake control system is lower than the set maximum coasting speed threshold, such as <45km / h, and switches to high gain.

[0064] If not, switch to the front wheel steering gain off signal to trigger the cooperative control strategy to stop;

[0065] If so, switch to the front wheel turning gain on signal, determine whether the front wheel angle / position feedback sensor signal is higher than the set minimum front wheel turning angle threshold, and if so, trigger the start of the cooperative control strategy.

[0066] If the front wheels make a large turn (e.g., > ±20°), the main wheel coordinated turning function is activated. The software calculates the turning angles of the inner and outer main wheels and enters a three-linkage control of the turning angles of the front wheel, inner main wheel, and outer main wheel. At the same time, three turning control signals are output, and the actual turning angles of the three are detected in real time and dynamically adjusted (the adjustment cycle is tentatively set at 50ms).

[0067] In one embodiment, the real-time taxiing speed parameters output by the main wheel brake control system are obtained, and the real-time turning arc length of the aircraft is calculated. The turning radius is then calculated by combining this with the nose wheel turning angle. Specifically, this includes:

[0068]

[0069] Where l is the turning arc length of the aircraft in ground mode, t is the real-time taxiing time of the aircraft, and V is the taxiing speed of the aircraft.

[0070]

[0071] Where R is the turning radius of the aircraft in ground mode; γ is the turning angle of the front wheels.

[0072] In one embodiment, the turning angles of the inner and outer main wheels are calculated using the turning radius and structural parameters between the aircraft's nose wheel, inner main wheel, and outer main wheel. Specifically, this includes:

[0073] The structural parameters between the aircraft's nose wheel, inner main wheel, and outer main wheel include: the wheelbase h between the inner and outer main wheels, and the distance s between the nose wheel and the centerline of the inner and outer main wheels;

[0074] The inner main wheel rotation angle and the outer main wheel rotation angle are calculated as follows:

[0075]

[0076]

[0077] Where: α is the inner main wheel rotation angle; β is the outer main wheel rotation angle; R is the turning radius of the aircraft in ground mode.

[0078] In one embodiment, see Figure 3 The control module adopts a DSP+CPLD structure, and the digital input unit, analog input unit, and analog output unit are opto-isolated from all I / O signals of the DSP in the control module.

[0079] The DSP chip is designed for embedded control, integrating all necessary hardware and peripheral functional components internally, requiring no additional functionality, thus implementing SoC (System-on-a-Chip) technology. The DSP chip has a built-in watchdog circuit that enables instant restart if the program crashes. Simultaneously, it utilizes an external CPLD to monitor the CPU in real time; if the CPU crashes three times within a specified time, it is considered to have an internal CPU fault, thus achieving a dual watchdog monitoring function.

[0080] The analog input signal uses linear opto-isolation, the frequency pulse sequence signal and the analog PWM output signal use high-speed opto-isolation, the ordinary switch input uses low-speed opto-isolation, and the hydraulic lock output control uses Darlington high-power opto-isolation.

[0081] In one embodiment, the DSP outputs a PWM signal (with a duty cycle that smoothly varies from 0 to 100%), which is then driven by a buffer to an optocoupler, a Smith trigger for shaping, and a filtering circuit. After passing through a V / I conversion circuit, the signal generates the front wheel steering servo signal, the inner main wheel steering servo signal, and the outer main wheel steering servo signal, which in turn drive the corresponding servo valve coils. The control circuit for the servo valve coils operates at 0–40 mA.

[0082] To reduce the instability of the control current caused by changes in the resistance and inductance of the servo valve control coil, constant current control is implemented in the circuit.

[0083] In one embodiment, the aircraft turning system includes a dual-channel main control device, comprising: a digital input unit, an analog input unit, an analog output unit, and a digital output unit, all of which are configured in dual-channel mode; and the interface module of the main control module is in dual-channel hot standby mode.

[0084] The main control module includes a dual-channel logic switching circuit and a BIT fault detection circuit. The BIT fault detection circuit is used to automatically test and detect faults in various electrical components involved in the aircraft steering function.

[0085] Under normal circumstances, the dual-channel main control device relies on the communication information between the two channels to perform cross-switching and automatically find a channel without control component failure. When the detection component of this channel is fault-free, a fault-free signal is output.

[0086] When the control components of the cross passage are normal but the BIT detection components are faulty, a fault signal is output, prompting a turning operation without real-time optimization and real-time coordination.

[0087] When one or both channels of the dual-channel main controller fail simultaneously, a signal indicating failure of the front wheel and main wheel turning function is output.

[0088] In one embodiment, it further includes a digital output unit connected to the main control module, used to output the front wheel turning status signal, the front wheel turning fault signal, the front wheel turning status switching valve status, the main wheel turning status signal, the main wheel turning fault signal, and the main wheel turning status switching valve status according to the main control module control strategy, the dual-channel logic switching result, and the BIT fault detection result.

[0089] The specific output switch signals are as follows:

[0090] Front wheel turning indicator light (green);

[0091] Front wheel turning warning light (red);

[0092] Front wheel turning state switching valve (DC28V):

[0093] Main wheel turning status light (green);

[0094] Main wheel turning malfunction indicator (red);

[0095] Main wheel turning state switching valve (DC28V).

[0096] The working principle of the main control module's electrical dual redundancy is as follows:

[0097] (1) When the aircraft is in air mode and the nose landing gear lower position lock is open, the DC28V power supply (main and backup dual-path automatic switching) of the main control module is automatically cut off, and the system is in a power-off non-working state.

[0098] (2) When the aircraft is in ground mode and the nose landing gear lower position lock is closed, the DC28V power supply of the main control module is automatically turned on, the DSP runs the power-on active BIT program, the dual channels perform self-tests simultaneously, and apply detection signals to detect related accessories.

[0099] (3) Fault detection and channel switching functions

[0100] If a single fault is detected in the normal channel (fault alarm light flashes), switch to the fault-free hot standby channel.

[0101] If the hot standby channel also fails, the following situations will be handled:

[0102] a) At this point, assuming the dual-channel main control devices are functioning normally, the two channels will be cross-switched based on their communication information, automatically finding a channel without a faulty control component. When the detection component of this channel is fault-free, the fault alarm light will turn off.

[0103] b) When the control components of the cross passage are normal but the detection components are faulty, the fault alarm light will illuminate, prompting a turning operation without real-time optimization and real-time coordination.

[0104] c) When the main control device of the backup channel fails or both main control devices fail simultaneously, the external monitoring circuit and CPLD arbitration logic will issue a fault alarm light and an alarm sound to indicate to the driver that the turning function of the front wheels and main wheels has failed and the turning function can only be achieved by using differential braking.

[0105] Before executing the coordinated control strategy, the main control module obtains the current front wheel turning / sway reduction status switch signal. When it is determined that the coordinated control strategy cannot be implemented, the front wheel turning / sway reduction status switch signal automatically causes the system to turn into the sway reduction state.

[0106] The various embodiments in this specification are described in a progressive manner, with each embodiment focusing on the differences from other embodiments. The same or similar parts between the various embodiments can be referred to each other.

[0107] The foregoing has provided a detailed description of the aircraft turning system combining the front wheel and the main wheel provided by the present invention. Specific examples have been used in this embodiment to illustrate the principle and implementation of the present invention. The description of the above embodiments is only for the purpose of helping to understand the method and core idea of ​​the present invention. At the same time, for those skilled in the art, there will be changes in the specific implementation and application scope based on the idea of ​​the present invention. Therefore, the content of this specification should not be construed as a limitation of the present invention.

[0108] The above description of the disclosed embodiments enables those skilled in the art to make or use the invention. Various modifications to these embodiments will be readily apparent to those skilled in the art, and the general principles defined in these embodiments may be implemented in other embodiments without departing from the spirit or scope of the invention. Therefore, the invention is not to be limited to the embodiments shown herein, but is to be accorded the widest scope consistent with the principles and novel features disclosed herein.

Claims

1. An aircraft turning system that combines a front wheel and a main wheel, characterized by, The utility model relates to a kind of aircraft ground steering control system, including: Control template, switch quantity input unit, analog quantity input unit and analog quantity output unit; The switch quantity input unit is used to obtain front wheel load switch signal, front wheel turning / anti-sway state switch signal, inner main wheel load switch signal and outer main wheel load switch signal, and is sent to the control template; The analog quantity input unit is used to obtain front wheel turning instruction sensing signal, front wheel angle / position feedback sensing signal, inner main wheel angle / position feedback sensing signal and outer main wheel angle / position feedback sensing signal, and is sent to the control template; The analog quantity output unit is used to output front wheel turning servo signal, inner main wheel turning servo signal and outer main wheel turning servo signal; The control template is used to execute the following cooperative control strategy: According to front wheel load switch signal, inner main wheel load switch signal and outer main wheel load switch signal, cooperative control strategy is triggered to start; According to front wheel turning instruction sensing signal, the front wheel turning angle is calculated; Real-time sliding speed parameter output by main wheel brake control system is obtained, and the real-time turning arc length of aircraft is calculated, and the turning radius is calculated in combination with the front wheel turning angle; Inner main wheel turning angle and outer main wheel turning angle are calculated using the turning radius and structural parameters between aircraft front wheel, inner main wheel and outer main wheel, and the inner main wheel turning servo signal and outer main wheel turning servo signal are output to the analog quantity output unit; According to inner main wheel angle / position feedback sensing signal, the error between it and the inner main wheel turning angle is calculated;According to outer main wheel angle / position feedback sensing signal, the error between it and the outer main wheel turning angle is calculated; According to the error, the front wheel turning compensation angle is calculated, and the front wheel turning servo signal is output to the analog quantity output unit.

2. An aircraft turning system incorporating a front wheel and a main wheel as claimed in claim 1 wherein, The switch quantity input unit is also used to obtain front landing gear lower lock switch signal, and is sent to the control template, and the control template controls the start-stop of system power supply according to front landing gear lower lock switch signal.

3. An aircraft turning system incorporating a front wheel and a main wheel as claimed in claim 1 wherein, The switch quantity input unit is also used to obtain front wheel turning gain switching switch signal, and is sent to the control template, and the control template judges whether real-time sliding speed parameter output by main wheel brake control system is lower than set maximum sliding speed threshold value in real time; If not, switch to front wheel turning gain off signal, and trigger cooperative control strategy to stop; If yes, switch to front wheel turning gain on signal, and judge whether front wheel angle / position feedback sensing signal is higher than set minimum front wheel turning angle threshold value, if yes, trigger cooperative control strategy to start.

4. The aircraft turning system incorporating a front wheel and a main wheel as defined in claim 1, wherein, The real-time sliding speed parameter output by main wheel brake control system is obtained, and the real-time turning arc length of aircraft is calculated, and the turning radius is calculated in combination with the front wheel turning angle, specifically including: Wherein, l is the turning arc length of aircraft in ground mode, t is the real-time sliding time length of aircraft;V is the sliding speed of aircraft; Wherein, R is the turning radius of aircraft in ground mode;γ is the front wheel turning angle.

5. The aircraft turning system incorporating a front wheel and a main wheel as defined in claim 1, wherein, The inner main wheel turning angle and outer main wheel turning angle are calculated using the turning radius and structural parameters between aircraft front wheel, inner main wheel and outer main wheel, specifically including: The structural parameters between the front wheel, the inner main wheel and the outer main wheel of the aircraft include a wheel track h between the inner main wheel and the outer main wheel, and a distance s between the front wheel and the middle axis of the inner main wheel and the outer main wheel; The inner main wheel rotation angle and the outer main wheel rotation angle are calculated as follows: Wherein, α is the inner main wheel rotation angle, β is the outer main wheel rotation angle, and R is the turning radius of the aircraft in the ground mode.

6. An aircraft turning system incorporating a front wheel and a main wheel as claimed in claim 1 wherein, The control template adopts a DSP+CPLD structural mode, and the switching value input unit, the analog value input unit and the analog value output unit are optoelectronically isolated from all I / O signals of the DSP of the control template.

7. An aircraft turning system incorporating a front wheel and a main wheel as claimed in claim 6 wherein, The DSP outputs PWM signals, which are driven by a buffer, a photo-coupler, a Schmitt trigger shaper and a filter circuit, and then are converted by a V / I conversion circuit to obtain front wheel turning servo signals, inner main wheel turning servo signals and outer main wheel turning servo signals, and then drive corresponding servo valve coils.

8. An aircraft turning system incorporating a front wheel and a main wheel as claimed in claim 1 wherein, Further, the switching value output unit is used to output front wheel turning state signals, front wheel turning fault signals, front wheel turning state switching valve states, main wheel turning state signals, main wheel turning fault signals and main wheel turning state switching valve states.

9. An aircraft turning system incorporating a front wheel and a main wheel as claimed in claim 8 wherein, The aircraft turning system includes a double-channel main control device, which includes a switching value input unit, an analog value input unit, an analog value output unit and a switching value output unit, all of which are double-channel configurations, and an interface module of a main control template is a double-channel hot standby mode; The main control template includes a double-channel logic switching circuit and a BIT fault detection circuit, which is used for automatically testing and fault detecting each electrical component participating in the aircraft turning function; Under normal conditions of the double-channel main control device, cross switching of two channels is performed by relying on communication information between each other, and a channel without control component faults is automatically found, and when the detection component of the channel has no faults, a fault-free signal is outputted; When the control component of the cross channel is normal and the BIT detection component has faults, a fault signal is outputted, prompting to perform turning operation without real-time optimization and real-time cooperation; When one channel or both channels of the double-channel main control device fail, a front wheel and main wheel turning function failure signal is outputted.

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