Detection and disposal method for single rotor failure of tail stand coaxial dual-rotor aircraft and flight control system

Abstract

The application discloses a tail seat type coaxial dual-rotor aircraft single-rotor failure detection and disposal method and a flight control system, and belongs to the technical field of electric vertical take-off and landing aircraft flight control. The method sets a failure detection threshold according to the wind suspension state, determines suspected failure when the coaxial upper and lower rotors are asymmetric in tension command, the body yaw angular velocity meets the direction locking condition and lasts for a confirmation window; the real thrust of the high side pumped by the thrust back reading device falls below the specified proportion and lasts for a confirmation window to lock the failure side; the real thrust tracking command cancels the suspected determination, and the gust condition is structurally rejected; then the failure side is cleared, the healthy side bears the total tension demand, the yaw is handed over to the wing tip propulsion paddle, the roll is borne by the roll control surface, the failure lock is set as the highest priority, the remaining disposal chain is immediately reset when competing for the same shot, and the hierarchical emergency exit is executed according to the wind condition and height. The application is fast in detection and safe in full-line safety disposal.

Description

Technical Field

[0001] This invention relates to the field of flight control technology for electric vertical takeoff and landing aircraft, specifically to a method for failure detection and handling of a tail-seat coaxial dual-rotor aircraft under single-rotor failure conditions, and a flight control system for implementing this method. Background Technology

[0002] In a tail-seat coaxial twin-rotor electric vertical takeoff and landing aircraft (coaxial main rotor assembly combined with wingtip propeller and slipstream inner aileron configuration), the yaw channel relies on the speed difference (anti-torque difference) between the coaxial upper and lower rotors in hovering and low-speed modes. When either the upper or lower rotor fails (motor or electronic speed controller malfunction), traditional closed-loop control cannot identify the failed side: the thrust command of the failed rotor is continuously pumped up to saturation by the closed loop, while the surviving rotor is suppressed, causing the entire aircraft to lose control and crash to the ground within seconds, leaving the flight control system with an extremely short response window.

[0003] Furthermore, the kinematic characteristics of single rotor failure (asymmetric coaxial thrust command, yaw spin and its growth trend) highly overlap with the kinematic characteristics of gust impact and gust correction: the yaw moment applied by gust impact causes the yaw control loop to output the same asymmetric command and spin growth. Detection methods based solely on kinematic statistics cannot distinguish between the two, and the false trigger rate is unacceptable—false locking will misjudge normal gust correction as rotor failure, thus causing an accident.

[0004] Furthermore, after a single rotor fails, the coaxial differential speed capability of the yaw channel is halved or even reduced to zero. If the conventional emergency response chain (such as the gust roll recovery response chain and the strong wind upwind hovering maintenance response chain) adjusted according to the healthy configuration is directly applied to the failed configuration, it will interfere with the failure response.

[0005] The problem is more prominent when hovering in the wind (wind-hovering) during strong winds: failure handling and fighting against the wind compete for propeller authority. When the propeller rotates at full speed to maintain its position against the wind, its yaw differential authority is physically reduced to zero, forming a "dead spiral" and accelerating the entire aircraft's fall to the ground. Summary of the Invention

[0006] The technical problem to be solved by this invention is to provide a method and flight control system for detecting and handling single rotor failure in a tail-seat coaxial twin-rotor aircraft, so that single rotor failure can be detected and locked on the failed side quickly and accurately—especially in the case where kinematic characteristics and gust impacts are highly overlapping, it can distinguish between real failure and gust disturbance, and avoid false locking; and after locking, it can complete control redistribution and graded emergency handling to achieve safe handling of the entire envelope.

[0007] To solve the above-mentioned technical problems, the present invention adopts the following technical solution: a method for detecting and handling single rotor failure in a tail-seat coaxial twin-rotor aircraft, wherein the tail-seat coaxial twin-rotor aircraft includes an upper rotor and a lower rotor arranged coaxially, a left propeller and a right propeller respectively located at the left and right wingtips, and a roll control surface. The upper rotor and the lower rotor are each driven by a drive motor. The aircraft is equipped with a thrust feedback device, which is used to read back the actual thrust of the upper rotor and the lower rotor. The thrust feedback device is at least one of the following: telemetry feedback of the electronic speed controller of the drive motor, a current sensor, a voltage sensor, a speed sensor, and a thrust sensor. The roll control surface is located between the upper rotor and the lower rotor. The method involves at least one of the aileron, flap aileron, and left / right differential thrust unit within the slipstream of the lower rotor, comprising the following steps: Step 1: Real-time acquisition of the thrust commands of the upper and lower rotors, the actual thrust read back by the thrust feedback device, the airframe kinematic characteristics, and the windshield status. The airframe kinematic characteristics include the airframe yaw rate, and at least one of the roll rate and attitude angle change rate; Step 2: Setting failure detection thresholds according to the windshield status. When the windshield status is inactive, a static windshield threshold group is used; when the windshield status is active, a windshield threshold group is used. The yaw spin threshold of the windshield threshold group is higher than the yaw spin threshold of the static windshield threshold group, and the windshield threshold group is indeed... The confirmation window duration is not shorter than the confirmation window duration of the static windshield threshold group; when the difference between the thrust commands of the upper rotor and the lower rotor exceeds the asymmetry threshold, and at least one of the body kinematic characteristics meets the failure characteristic conditions and continues to reach the confirmation window duration, it is determined that there is a suspected single rotor failure, and the side with the larger thrust command between the upper rotor and the lower rotor is the side that was pumped up; the failure characteristic conditions include the direction locking condition, which requires that the direction of the body yaw rate is consistent with the spin direction caused by the residual counter-torque of the whole machine when the rotor on the side that was pumped up fails; Step 3: Use the actual thrust read back by the thrust readback device as confirmation evidence, when the actual thrust of the rotor on the side that was pumped up drops When the drop falls below the specified proportion of the thrust command and continues for the duration of the confirmation window, the suspected single rotor failure is confirmed, the side rotor is determined to be faulty, and the faulty side is locked; when the actual thrust tracks the thrust command within the confirmation window duration, the determination of the suspected single rotor failure is revoked and control is maintained according to the healthy configuration, and the gust flapping and gust correction conditions are thus structurally rejected; Step 4: Perform health perception control redistribution according to the locked faulty side: clear the thrust command of the faulty side rotor to zero, and the healthy side rotor undertakes the total thrust demand of the upper rotor and the lower rotor, switch the yaw channel from coaxial differential speed to differential thrust of the left propeller and the right propeller, and the roll channel is undertaken by the roll control surface;Step 5: Set the latch for the single rotor failure as the highest priority disabling event. During the latch's active period, suspend other emergency response chains and prohibit the suspended emergency response chains from rearming. When the latch and other emergency response chains compete within the same control frame, the latch takes priority, and the other emergency response chains are immediately reset, with the redistributed control configuration taking full control. Step 6: Execute tiered emergency exits: freeze the heading target and switch to emergency control gain, handling according to wind conditions—in calm wind conditions, land in place after attitude stabilizes; in wind-suspended conditions, maintain attitude first, and after attitude stabilizes, divert traffic according to altitude, either accelerating to fixed-wing cruise or landing in place under controlled conditions.

[0008] Furthermore, the tail-seat coaxial dual-rotor aircraft also includes a wind speed estimation module. The wind suspension state is determined by the output of the wind speed estimation module. The wind suspension state is activated when the incoming wind speed estimated by the wind speed estimation module is not less than the wind suspension boundary wind speed. The calm wind barrier threshold group and the wind suspension barrier threshold group are selected by the activation and deactivation of the wind suspension state, respectively.

[0009] Furthermore, when the altitude above the ground is lower than the prescribed armed altitude, the failure detection in steps 2 and 3 will not be initiated.

[0010] Furthermore, in step 2, the kinematic characteristic of the aircraft is the yaw rate of the aircraft, and the failure characteristic condition is the direction lock condition.

[0011] Furthermore, in step 3, the specified ratio is 40% to 80%.

[0012] Furthermore, in step 4, when the yaw channel is switched to the differential thrust, the counter-torque of the surviving rotor is included as a feedforward term in the distribution of the differential thrust; the roll control surface is an aileron located in the slipstream of the upper rotor and the lower rotor; when the single rotor fails in the fixed-wing cruise mode, the thrust command of the failed rotor is cleared to zero, and the surviving rotor takes on the pitch trim function of the dynamic horizontal stabilizer.

[0013] Furthermore, in step 5, the other emergency response chains that are suspended include the roll-out response chain and the wind suspension response chain.

[0014] Furthermore, in step 6, freezing the heading target means freezing the heading target to the current heading at the time of latching; the criterion for attitude stabilization is: the aircraft's attitude angular velocity is lower than the specified angular velocity value, the attitude deviation is lower than the specified angle value and continues for a specified duration; under calm wind conditions, after attitude stabilization is confirmed, a vertical landing is performed in place.

[0015] Furthermore, in step 6, under wind suspension conditions, the wind suspension holding chain is first engaged to maintain the attitude and allow it to drift with the wind. After the attitude stabilizes and is confirmed: when the ground clearance is not lower than the specified altitude value, it enters the forward transition acceleration and switches to fixed-wing cruise supported by the wing; when the ground clearance is lower than the specified altitude value, it performs a controlled landing in place.

[0016] Another technical solution of the present invention is: a method for detecting and handling single rotor failure in a tail-seat coaxial dual-rotor aircraft. The tail-seat coaxial dual-rotor aircraft includes an upper rotor and a lower rotor arranged coaxially, a left propeller and a right propeller respectively located at the left and right wingtips, and roll control surfaces. The upper rotor and the lower rotor are each driven by a drive motor. The aircraft is equipped with a thrust feedback device, which is used to feedback the actual thrust of the upper rotor and the lower rotor. The method includes the following steps: Step 1: Real-time acquisition of the thrust index of the upper rotor and the lower rotor. The thrust readback device reads the actual thrust and airframe kinematic characteristics, including at least one of yaw rate, roll rate, and attitude angle change rate; Step 2: When the difference between the thrust commands of the upper rotor and the lower rotor exceeds the asymmetry threshold, and at least one of the airframe kinematic characteristics meets the failure characteristic condition and continues to meet the confirmation window duration, a suspected single rotor failure is determined, and the side with the larger thrust command between the upper rotor and the lower rotor is the side that was pumped up; Step 3: The thrust readback device reads the actual thrust and airframe kinematic characteristics, including at least one of the airframe kinematic characteristics, yaw rate, roll rate, and attitude angle change rate; Actual thrust is used as confirmatory evidence. When the actual thrust of the pumped-up rotor falls below a specified proportion of its thrust command, the suspected single-rotor failure is confirmed, the rotor on that side is determined to be faulty, and the faulty side is locked. When the actual thrust tracks the thrust command, the determination of the suspected single-rotor failure is revoked, and control is maintained according to the healthy configuration. Step 4: Perform health-aware control redistribution according to the locked faulty side: clear the thrust command of the faulty rotor to zero, and let the healthy rotor bear the combined thrust demand of the upper and lower rotors, and change the yaw channel from coaxial differential... The system quickly switches to the differential thrust of the left and right propellers, with the roll channel being handled by the roll control surfaces; Step 5: Set the latch of the single rotor failure as the highest priority disabling event, suspend other emergency response chains, and have the redistributed control configuration take over; Step 6: Execute graded emergency exits: freeze the heading target and switch to emergency control gain, and handle according to wind conditions—in calm wind conditions, wait for the attitude to stabilize before landing in place; in wind-suspended conditions, maintain the attitude first, and after the attitude stabilizes, divert according to the altitude, execute acceleration to switch to fixed-wing cruise or in-place controlled landing.

[0017] This invention also provides a flight control system for handling single-rotor failure in a tail-seat coaxial twin-rotor aircraft, comprising: a flight control computer; a coaxial upper rotor power unit and a coaxial lower rotor power unit, each including a rotor, a drive motor, and an electronic speed controller; a thrust readback device for reading back the actual thrust of the coaxial upper rotor power unit and the coaxial lower rotor power unit, wherein the thrust readback device is at least one of the following: telemetry readback of the electronic speed controller of the drive motor, a current sensor, a voltage sensor, a speed sensor, and a thrust sensor, and the signal output terminal of the thrust readback device is connected to the signal input terminal of the flight control computer; a left propeller power unit and a right propeller power unit respectively located at the left and right wingtips; a roll control surface, wherein the roll control surface is at least one of the following: an aileron, a flap aileron, and a left and right differential thrust unit located within the slipstream of the coaxial upper rotor power unit and the coaxial lower rotor power unit; and an attitude inertial measurement unit connected to the flight control computer; the flight control computer is configured to execute the above-described detection and handling method.

[0018] The beneficial effects of this invention are as follows: First, the collapse of actual thrust read back by the thrust readback device is used as evidence of failure confirmation. The thrust readback device can be implemented by at least one of the following: telemetry readback of the electronic speed controller, current sensor, voltage sensor, speed sensor, and tension sensor. It utilizes the structural difference between failure and gust in the causal relationship of "command-actual thrust"—when a single rotor fails, the actual thrust on the side pumped up collapses and no longer tracks the command, while during gust impact and gust correction, both upper and lower rotors are intact and the actual thrust tracks the command. Therefore, the gust condition is structurally negated, and the false locking problem that cannot be overcome by pure kinematic detection is eliminated. Second, the trigger condition constrains the direction of the yaw rate of the aircraft with the direction locking condition, making it consistent with the spin direction caused by the residual anti-torque of the whole aircraft when the rotor on the side pumped up fails, and can take at least one of the roll rate and attitude angle change rate. The system employs a comprehensive approach to judgment, eliminating misjudgments caused by reverse disturbances and avoiding reliance on a single kinematic quantity. Thirdly, failure detection thresholds are set in stages according to the wind suspension state, with the calm wind threshold group ensuring detection speed and the wind suspension threshold group ensuring interference immunity, balancing both operating conditions. Fourthly, the health perception control redistribution after latching involves zeroing the failed side, the healthy side bearing the combined thrust demand of the upper and lower rotors, yaw transfer to the propeller differential, and roll return to the roll control surface to reconstruct the control configuration, ensuring complete three-channel attitude control capability and even fixed-wing cruise capability even after failure. Fifthly, failure latching, as the highest priority disability event, suspends other handling chains adjusted according to the health configuration and prohibits their rearmament, immediately resetting when competing for the same control signal to avoid multiple handling chains interfering with each other. Sixthly, graded emergency exits are diverted according to wind conditions and altitude, allowing for in-situ landing in calm winds and accelerated escape or controlled landing after wind suspension ceasefire, achieving safe handling across the entire envelope.

[0019] It should be noted that the claims of this invention and the applicant's related applications are as follows: the switching between hovering mode and cruise mode (including the transition corridor) is claimed by the related applications, and the disposal chain of this invention reuses its transition corridor, but does not claim the mode switching itself; the roll recovery disposal chain and the wind suspension holding disposal chain are claimed by the related applications, and this invention claims the priority arbitration and suspension of the above disposal chains by the failure latch, as well as the graded emergency exit after failure; the return landing gear is claimed by the related applications, and this invention only refers to its high angle of attack controlled landing condition; the tail rotor dynamic horizontal stabilizer configuration is claimed by the related applications, and this invention reuses it in the cruise mode under the failure configuration of the single pitch trim function of the surviving side rotor, without repeatedly claiming the dynamic horizontal stabilizer configuration itself. Attached Figure Description

[0020] Figure 1 is a block diagram of the flight control system according to an embodiment of the present invention.

[0021] Figure 2 is a flowchart of single rotor failure detection according to an embodiment of the present invention.

[0022] Figure 3 is a block diagram of health perception control redistribution according to an embodiment of the present invention.

[0023] Figure 4 shows a hierarchical emergency exit decision tree according to an embodiment of the present invention.

[0024] Figure 5 is a priority arbitration timing diagram of an embodiment of the present invention.

[0025] Among them, 10 is the flight control computer; 20 is the coaxial upper rotor power unit; 21 is the upper rotor electronic speed controller; 22 is the coaxial lower rotor power unit; 23 is the lower rotor electronic speed controller; 30 is the left wingtip propeller power unit; 31 is the right wingtip propeller power unit; 40 is the roll control surface (in this embodiment, it is the slipstream inner aileron); 50 is the attitude inertial measurement unit; 60 is the wind speed estimation module; and 70 is the loop landing gear. Detailed Implementation

[0026] The present invention will be further described below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only for explaining the present invention and are not intended to limit the present invention. The control parameters in the embodiments are design values ​​verified by simulation at the prototype level, used to illustrate the relative relationship between each threshold, duration and gain, and not to limit the scope of protection of the present invention.

[0027] As shown in Figure 1, the flight control system of this embodiment includes a flight control computer 10, a coaxial upper rotor power unit 20, a coaxial lower rotor power unit 22, a left wingtip propeller power unit 30, a right wingtip propeller power unit 31, a slipstream inner aileron 40, an attitude inertial measurement unit 50, and a wind speed estimation module 60. The coaxial upper rotor power unit 20 and the coaxial lower rotor power unit 22 are coaxially arranged at the tail of the aircraft fuselage. Their rotation axes coincide and their rotation directions are opposite. Each includes a rotor, a drive motor, and an electronic speed controller. In this embodiment, the thrust readback device is implemented using telemetry readback from the electronic speed controller: both the upper rotor electronic speed controller 21 and the lower rotor electronic speed controller 23 have telemetry readback functionality. The telemetry readback content includes rotational speed and / or current, which can be used to calculate the actual thrust of the corresponding rotor. In this embodiment, the electronic speed controller is a 32-bit electronic speed controller (firmware is BLHeli32 or AM32) supporting the bidirectional DShot telemetry protocol. The telemetry readback refresh rate is not less than 100Hz, and the actual thrust signal output terminal is connected to the signal input terminal of the flight control computer 10 through a signal link. The thrust readback device can also be implemented by one or more of a current sensor, voltage sensor, rotational speed sensor, or thrust sensor, which respectively provide the actual thrust through a current-thrust model, a bus voltage-rotational speed model, a rotational speed-thrust coefficient model, or direct thrust measurement. The left wingtip propeller power unit 30 and the right wingtip propeller power unit 31 are respectively located at the left and right wingtips of the aircraft, and their thrust can be adjusted independently. Their differential thrust generates a yaw moment at the wingspan lever arm. In this embodiment, the roll control surface is an inner aileron 40 in the slipstream, which is set in the slipstream of the coaxial rotor assembly and serves as the roll channel; the roll control surface can also be a flap aileron or a left and right differential thrust unit. The attitude inertial measurement unit 50 (including a barometric altimeter) and the wind speed estimation module 60 provide the flight control computer 10 with attitude, angular velocity, altitude, and incoming wind speed estimates, respectively. Attitude and angular velocity constitute the airframe's kinematic characteristics, such as yaw rate, roll rate, and rate of change of attitude angle. Based on this, the wind speed estimation module 60 determines the wind suspension state (active / inactive): when the incoming wind speed estimated by the wind speed estimation module 60 is not less than the wind suspension threshold wind speed, the wind suspension state is activated and a wind suspension threshold group is selected; when it is lower than the wind suspension threshold wind speed, the wind suspension state is inactive and a calm wind threshold group is selected. The wind suspension threshold wind speed is the incoming wind speed upon which the wind suspension state changes from inactive to active; in this embodiment, a slow wind speed of 6 m / s is used (similar to the wind suspension entry / exit threshold). A loop-shaped landing gear 70 is provided on the lower part of the fuselage for controlled landing at high angles of attack.

[0028] As shown in Figure 2, the failure detection process consists of three levels: graded arming, failure characteristic judgment, and readback confirmation. Graded arming: Failure detection is not initiated when the ground clearance is below the arming height (3 m in this embodiment), and the turbulent flow field segment above ground is not diagnosed; ground-level failures are absorbed by the structure upon landing. After detection is initiated, failure detection thresholds are set. In this embodiment, these thresholds are set according to the wind suspension state—when the wind suspension is inactive, a still wind barrier threshold group is used (yaw and spin threshold 0.5 rad / s, confirmation window 0.10 s in this embodiment); when the wind suspension is active, a wind suspension barrier threshold group is used (yaw and spin threshold 1.8 rad / s, confirmation window 0.15 s in this embodiment, covering a gust sweep envelope of 1.47 rad / s with approximately 22% margin). The shared command asymmetric threshold for both levels is set to 10 in this embodiment. N; As a simplified variant, it is also possible to disregard the wind-suspended state classification and adopt a single fixed threshold group. In this case, the failure characteristic condition can be constituted only by the selected airframe kinematic characteristics exceeding the corresponding threshold without the additional direction locking condition. The readback confirmation can also be established when the actual thrust drops below the specified ratio without setting a separate duration requirement. Failure characteristic judgment: Airframe kinematic characteristics In this embodiment, the airframe yaw rate is taken as the kinematic characteristic. When the difference between the thrust commands of the upper rotor and the lower rotor exceeds the asymmetry threshold (the side with the larger thrust command is the side that is pumped up by the closed loop), and the direction of the airframe yaw rate is consistent with the spin direction caused by the residual anti-torque of the whole aircraft when the rotor on that side fails (in this embodiment, the lower rotor failure corresponds to the lower rotor command being pumped up and adding negative spin, and the upper rotor failure is the opposite), and when this characteristic is continuously satisfied for the confirmation window duration, it is determined that there is a suspected single rotor failure; the airframe kinematic characteristics can also be the roll rate or the rate of change of attitude angle, or a combination of at least two of them. Readback confirmation: The actual thrust on the pumped side read back by the thrust readback device is used as confirmation evidence. If the actual thrust drops below the specified proportion of its pull command (0.6 in this embodiment, i.e., 60%) and the drop continues for the confirmation window duration, the suspected failure is confirmed and the failed side is latched. If the actual thrust re-tracks the pull command within the confirmation window duration, the judgment of suspected single rotor failure is cancelled, and control is maintained according to the healthy configuration without entering latching. The ratio of actual thrust on the healthy side to the command is not less than 0.9, with sufficient two-way margin between the specified proportion and the actual thrust on the healthy side. Gust rejection mechanism: During gust impact or gust correction, both the upper and lower rotors are intact. The asymmetric command output by the closed loop is actually executed by both rotors. The thrust back reading device reads the actual thrust tracking thrust command. The back reading confirmation evidence is not satisfied, and the suspected failure judgment is structurally rejected. The failure and the gust are indistinguishable in terms of kinematic characteristics, but there is a structural difference in the causal relationship of "command - actual thrust". This embodiment replaces pure kinematic statistics with actuator back reading evidence, reducing the false lock rate to zero.

[0029] As shown in Figure 3, after latching the failed side, the flight control computer 10 performs health-aware control redistribution: the thrust command of the failed rotor is cleared to zero, and the rotor on the healthy side bears the combined thrust requirement of the upper and lower rotors (in this embodiment, the hovering requirement is about 27 N, which is about 40% of the physical upper limit of 63.2 N for a single rotor, with sufficient margin); the yaw channel is switched from coaxial differential speed to differential thrust of the left wingtip propeller power unit 30 and the right wingtip propeller power unit 31. The differential thrust generates yaw moment at the wingspan lever arm, and the anti-torque of the healthy side rotor is included as a feedforward term in the differential distribution (in this embodiment, the yaw differential gain is 0.44 N·m / N, including the wingspan lever arm and the anti-torque canard component); the roll channel is undertaken by the roll control surface. In this embodiment, the roll control surface is the slipstream inner aileron 40, which is differentially undertaken by the slipstream inner aileron 40. Zeroing the failed side applies to both vertical takeoff and landing mode and fixed-wing cruise mode: In cruise mode, the coaxial rotor assembly does not bear the main lift. After zeroing the failed side, the surviving rotor takes over the pitch trim function of the dynamic horizontal stabilizer, and the whole aircraft retains complete fixed-wing cruise capability.

[0030] As shown in Figure 4, the graded emergency exits are diverted according to the latching conditions. Latching along the action: The heading target is frozen to the current heading at the time of latching—emergency response is not concerned with the heading, because the yaw requirement to return to the old heading will occupy the propeller differential control authority for a long time, and its roll side effect is equivalent to the aileron control authority, which will trigger critical oscillations; the control gain switches to the emergency response group; if in the transition corridor, the transition is immediately terminated and the aircraft returns to hovering mode. Stabilization confirmation: The aircraft's attitude angular velocity is lower than the specified angular velocity value (0.3 rad / s in this embodiment), the attitude deviation is lower than the specified angle value (12° in this embodiment), and this condition persists for the specified duration (0.3 s in this embodiment). Calm wind exit (slow wind speed below 6 m / s): After stabilization confirmation, the aircraft lands vertically in place. Windshield exit (slow wind speed not less than 6 m / s): First, stop the attack to maintain attitude—suspend the windshield maintenance chain and let it drift with the wind, with position maintenance giving way to attitude maintenance; after stabilization is confirmed, the flow is divided according to the altitude: when the altitude is not lower than the specified altitude value (15 m in this embodiment), directly enter the forward transition acceleration and switch to fixed-wing cruise with the wing bearing the load—climbing is the weakest segment of the failure configuration, and only the acceleration upper wing can fundamentally prevent sinking; when the altitude is insufficient, land in place under controlled conditions, in conjunction with the high angle of attack landing condition of the 70° angle of attack of the loop landing gear.

[0031] As shown in Figure 5, the latch of single rotor failure is the highest priority disability event: during the latch period, the roll recovery handling chain and the wind suspension holding handling chain are suspended and rearmapping is prohibited (including immediate reset when there is competition with the failure latch in the same control phase), and the redistributed control configuration takes full control; if the handling chain is not suspended, the handling chain tuned according to the healthy configuration and the failure handling chain interfere with each other. Simulation shows that the interaction between the two chains will lead to a crash.

[0032] The main control parameters of this embodiment are shown in the table below: Armed Altitude >3 m Turbulent flow fields above ground are not diagnosed; failures near ground are absorbed by the structure. Instruction asymmetric threshold 10 N The still wind deflector and the wind suspension deflector are shared. Still air intake self-rotating threshold / confirmation window 0.5 rad / s / 0.10 s Windshield Rotary Threshold / Confirmation Window 1.8 rad / s / 0.15 s Covers a gust sweep envelope of 1.47 rad / s with approximately 22% margin. Reread confirmation ratio 0.6 Actual thrust < 0.6 × command thrust indicates failure; healthy thrust ≥ 0.9 indicates sufficient two-way margin. Stabilization Confirmation Angular velocity < 0.3 rad / s, attitude deviation < 12°, duration 0.3 s Wind-suspended / calm wind boundary Slow wind speed 6 m / s The same meaning as the wind-suspended threshold. High diversion threshold 15 m The same semantic meaning as the low-altitude disposal threshold of the same species Yaw differential gain 0.44 N·m / N Wingtip propeller differential (wing span lever arm + anti-torque canard component) The maximum thrust a single rotor can withstand 63.2 N (hovering requirement approximately 27 N) The physical feasibility basis of redistribution

[0033] Simulation results show that in the worst-case configuration (sudden rotor failure under main lift), without baseline intervention, failure to latch within 0.25 s results in attitude overshoot and grounding within 0.4 s (baseline probe confirmation). In this case, the latching time for this configuration is 0.21~0.22 s, meeting the intervention window. The latching time for upper rotor failure is 0.32 s, and for half-descent slow-launch failure it is 0.86 s. Due to the low attitude deterioration rate, both resulted in safe landings. In wind-suspended conditions with a force 9 wind (22 m / s plus gusts), the latching time is approximately 0.40 s, and after stabilization, the aircraft accelerates back to cruise. Zero false locks were observed throughout the 3~22 m / s compression gust profile sweep and in the force 9 gust impact test scenario (the original area of ​​severe false locks). Zero false locks were also observed in calm high / low altitude, half-descent, transition corridor, and force 9 wind conditions. All seven types of failures during suspension in winds of force 9 were handled safely (successful landing or escape), with no ground contact or divergence; in the worst case (complete failure of the main rotor at an altitude of 40 m), the aircraft lost altitude and converged to 6.25 m before landing normally; in the case of failure during suspension in winds of force 9, the aircraft converged and lost altitude by about 20 m before accelerating and escaping to cruise; the failure configuration still retains complete fixed-wing cruise capability.

[0034] Alternative implementation methods: The thrust readback device can be any form other than telemetry readback from an electronic speed controller, including current sensors, voltage sensors, speed sensors, and thrust sensors, as well as combinations thereof; the electronic speed controller can be any other electronic speed controller with speed and / or current telemetry readback functions, not limited to the bidirectional DShot protocol and BLHeli32, AM32 firmware; the actual thrust can be calculated from the readback speed using the thrust coefficient, from the readback current using the motor model, from the bus voltage using the voltage-speed model, from the speed sensor signal using the speed-thrust coefficient model, or directly measured by the thrust sensor, or a combination of multiple methods can be used; the roll control surface can be any form other than the slipstream inner aileron, including flaps and left / right differential thrust units; the readback confirmation ratio can be 40%~80%. The internal calibration is based on the distributed nature of the power system; the failure detection threshold can be graded according to the wind spur state, using multiple levels and continuously interpolated according to wind speed, or it can be a simplified variant that is not graded according to the wind spur state and uses a single fixed threshold group; the airframe kinematic characteristics on which the failure characteristic conditions are based can be any one of the airframe yaw rate, roll rate and attitude angle change rate, and the failure characteristic conditions can be accompanied by or without the direction locking condition; the duration requirement for readback confirmation can also be omitted, and the actual thrust falling below a specified ratio is the necessary and sufficient evidence for confirmation; the actuator for yaw differential can be a thrust unit arranged symmetrically on both sides other than the wingtip propeller; the altitude split threshold and stabilization criterion can be scaled according to the aircraft model; all the above variants can realize the function of confirming failure by reading back the actual thrust collapse and locking, redistributing and then classifying the exit.

[0035] The specific embodiments of the present invention have been described above with reference to the accompanying drawings. However, these descriptions should not be construed as limiting the scope of the present invention. The scope of protection of the present invention is defined by the appended claims. Any modifications based on the claims of the present invention are within the scope of protection of the present invention.

Claims

1. A method for detecting and handling single rotor failure in a tail-seat coaxial twin-rotor aircraft, wherein the tail-seat coaxial twin-rotor aircraft includes an upper rotor and a lower rotor arranged coaxially, a left propeller and a right propeller respectively disposed at the left and right wingtips, and a roll control surface, wherein the upper rotor and the lower rotor are each driven by a drive motor, the aircraft is equipped with a thrust feedback device, the thrust feedback device being used to feedback the actual thrust of the upper rotor and the lower rotor, the thrust feedback device being at least one of the following: telemetry feedback of the electronic speed controller of the drive motor, a current sensor, a voltage sensor, a speed sensor, and a thrust sensor, and the roll control surface being at least one of the following: an aileron, a flap aileron, and a left and right differential thrust unit located within the slipstream of the upper rotor and the lower rotor, characterized in that... Includes the following steps: Step 1: Real-time acquisition of the thrust commands of the upper rotor and the lower rotor, the actual thrust read back by the thrust feedback device, the airframe kinematic characteristics, and the wind suspension status. The airframe kinematic characteristics include the airframe yaw rate, and at least one of the roll rate and attitude angle change rate. Step 2: Set failure detection thresholds according to the wind spool state classification. When the wind spool is inactive, the still wind barrier threshold group is used. When the wind spool is active, the wind spool barrier threshold group is used. The yaw and spin threshold of the wind spool barrier threshold group is higher than the yaw and spin threshold of the still wind barrier threshold group, and the confirmation window duration of the wind spool barrier threshold group is not shorter than the confirmation window duration of the still wind barrier threshold group. When the difference between the thrust commands of the upper rotor and the lower rotor exceeds the asymmetry threshold, and at least one of the body kinematic characteristics satisfies the failure characteristic condition and continues to meet the confirmation window duration, a suspected single rotor failure is determined. The side with the larger thrust command between the upper rotor and the lower rotor is the side that was pumped up. The failure characteristic condition includes a direction locking condition, which requires that the direction of the body yaw rate be consistent with the spin direction caused by the residual counter-torque of the entire machine when the rotor on the pumped-up side fails. Step 3: Using the actual thrust read back by the thrust readback device as confirmation evidence, when the actual thrust of the pumped rotor falls below a specified proportion of its thrust command, and this fall continues for the duration of the confirmation window, the suspected single rotor failure is confirmed, the rotor on that side is determined to be faulty, and the faulty side is locked; when the actual thrust tracks the thrust command within the confirmation window duration, the determination of the suspected single rotor failure is revoked, and control is maintained according to the healthy configuration, thus structurally rejecting the gust impact and gust correction conditions; Step 4: Perform health-sensing control reallocation according to the latched failure side: clear the thrust command of the failure side rotor to zero, and let the healthy side rotor bear the total thrust demand of the upper rotor and the lower rotor. Switch the yaw channel from coaxial differential speed to differential thrust of the left propeller and the right propeller, and let the roll channel be borne by the roll control surface. Step 5: Set the latch of the single rotor failure as the highest priority disabling event. During the latch's active period, suspend other emergency response chains and prohibit the suspended other emergency response chains from being rearmed. When the latch and other emergency response chains compete within the same control frame, the latch takes priority and the other emergency response chains are immediately reset and taken over by the redistributed control configuration. Step 6: Implement graded emergency exits: Freeze the heading target and switch to emergency control gain, and handle according to wind conditions—in calm wind conditions, land on the spot after the attitude stabilizes; in wind-suspended conditions, maintain the attitude first, and after the attitude stabilizes, divert according to the altitude, and execute acceleration to switch to fixed-wing cruise or land on the spot under controlled conditions.

2. The detection and treatment method according to claim 1, characterized in that, The tail-seat coaxial dual-rotor aircraft also includes a wind speed estimation module. The wind suspension state is determined by the output of the wind speed estimation module. The wind suspension state is activated when the incoming wind speed estimated by the wind speed estimation module is not less than the wind suspension boundary wind speed. The calm wind barrier threshold group and the wind suspension barrier threshold group are selected by the activation and deactivation of the wind suspension state, respectively.

3. The detection and treatment method according to claim 1 or 2, characterized in that, When the ground clearance is lower than the prescribed armed clearance, the failure detection in steps 2 and 3 will not be initiated.

4. The detection and treatment method according to claim 1, characterized in that, In step 2, the kinematic characteristic of the aircraft is the yaw rate of the aircraft, and the failure characteristic condition is the direction lock condition.

5. The detection and treatment method according to claim 1, characterized in that, In step 3, the specified ratio is 40% to 80%.

6. The detection and treatment method according to claim 1, characterized in that, In step 4, when the yaw channel is switched to the differential thrust, the counter-torque of the surviving rotor is included as a feedforward term in the distribution of the differential thrust; the roll control surface is an aileron located in the slipstream of the upper rotor and the lower rotor; when the single rotor fails in the fixed-wing cruise mode, the thrust command of the failed rotor is cleared to zero, and the surviving rotor takes on the pitch trim function of the dynamic horizontal stabilizer.

7. The detection and treatment method according to claim 1, characterized in that, In step 5, the other emergency response chains that are suspended include the roll-out response chain and the wind suspension response chain.

8. The detection and treatment method according to claim 1, characterized in that, In step 6, freezing the heading target means freezing the heading target to the current heading at the time of latching; the criterion for attitude stabilization is: the aircraft's attitude angular velocity is lower than the specified angular velocity value, the attitude deviation is lower than the specified angle value and continues for a specified duration; under calm wind conditions, after attitude stabilization is confirmed, a vertical landing is performed in place.

9. The detection and treatment method according to claim 1, characterized in that, In step 6, under wind suspension conditions, the wind suspension holding chain is first engaged to maintain the attitude and allow it to drift with the wind. After the attitude stabilizes and is confirmed: when the ground clearance is not lower than the specified altitude value, it enters the forward transition acceleration and switches to fixed-wing cruise supported by the wing; when the ground clearance is lower than the specified altitude value, it performs a controlled landing in place.

10. A flight control system for handling single rotor failure in a tail-seat coaxial twin-rotor aircraft, characterized in that, include: A flight control computer; a coaxial upper rotor power unit and a coaxial lower rotor power unit, each including a rotor, a drive motor, and an electronic speed controller; a thrust readback device for reading back the actual thrust of the coaxial upper rotor power unit and the coaxial lower rotor power unit, the thrust readback device being at least one of the following: telemetry readback of the electronic speed controller of the drive motor, a current sensor, a voltage sensor, a speed sensor, and a thrust sensor, the signal output terminal of the thrust readback device being connected to the signal input terminal of the flight control computer; a left propeller power unit and a right propeller power unit respectively located at the left and right wingtips; a roll control surface, the roll control surface being at least one of the following: an aileron, a flap aileron, and a left and right differential thrust unit located within the slipstream of the coaxial upper rotor power unit and the coaxial lower rotor power unit; and an attitude inertial measurement unit connected to the flight control computer; the flight control computer is configured to perform the detection and handling method according to any one of claims 1 to 9.

11. A method for detecting and handling single rotor failure in a tail-seat coaxial twin-rotor aircraft, wherein the tail-seat coaxial twin-rotor aircraft includes an upper rotor and a lower rotor arranged coaxially, a left propeller and a right propeller respectively disposed at the left and right wingtips, and roll control surfaces, wherein the upper rotor and the lower rotor are each driven by a drive motor, and the aircraft is equipped with a thrust feedback device for feedback of the actual thrust of the upper rotor and the lower rotor, characterized in that... Includes the following steps: Step 1: Real-time acquisition of the thrust commands of the upper rotor and the lower rotor, the actual thrust read back by the thrust feedback device, and the kinematic characteristics of the airframe, wherein the kinematic characteristics of the airframe include at least one of the airframe yaw rate, roll rate, and attitude angle change rate; Step 2: When the difference between the thrust commands of the upper rotor and the lower rotor exceeds the asymmetry threshold, and at least one of the body kinematic characteristics meets the failure characteristic condition and continues to meet the confirmation window duration, it is determined that there is a suspected single rotor failure, and the side with the larger thrust command between the upper rotor and the lower rotor is the side that is pumped up. Step 3: Using the actual thrust read back by the thrust readback device as confirmation evidence, when the actual thrust of the pumped rotor falls below a specified proportion of its thrust command, the suspected single rotor failure is confirmed, the rotor on that side is determined to be faulty, and the faulty side is locked; when the actual thrust tracks the thrust command, the determination of the suspected single rotor failure is revoked, and control is maintained according to the healthy configuration; Step 4: Perform health-sensing control reallocation according to the latched failure side: clear the thrust command of the failure side rotor to zero, and let the healthy side rotor bear the total thrust demand of the upper rotor and the lower rotor. Switch the yaw channel from coaxial differential speed to differential thrust of the left propeller and the right propeller, and let the roll channel be borne by the roll control surface. Step 5: Set the latch of the single rotor failure as the highest priority disabling event, suspend other emergency response chains, and let the redistributed control configuration take over. Step 6: Implement graded emergency exits: Freeze the heading target and switch to emergency control gain, and handle according to wind conditions—in calm wind conditions, land on the spot after the attitude stabilizes; in wind-suspended conditions, maintain the attitude first, and after the attitude stabilizes, divert according to the altitude, and execute acceleration to switch to fixed-wing cruise or land on the spot under controlled conditions.

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