Automatic detection and emergency recovery method for attitude of out-of-control aircraft

Abstract

The invention relates to the technical field of aircraft safety control, and discloses an out-of-control aircraft attitude automatic detection and emergency recovery method, which comprises the following steps: a signal networking emission step, a triangular positioning signal emission array construction step and a multi-dimensional signal calculation step. The method comprises the steps of three-dimensional coordinate calculation, Euler angle and attitude angular velocity calculation, dynamic threshold judgment, graded trigger response, dynamic correction execution and closed-loop feedback verification. The three signal transmitters are used for positioning the aircraft, the built-in calculation unit is used for calculating the distance and positioning, and the attitude correction system is triggered to adjust the attitude, so that civil aircrafts and the like can safely land when the vertical distance is lower than the safety standard, and human accidents are avoided; after the first-stage rocket locates the launcher, the first-stage rocket lands or returns safely; the lunar landing cabin is vertically landed, the safe landing probability is improved, and errors are reduced; and the rolling spacecraft can be stopped from rolling, so that the flight safety of various aircrafts is guaranteed.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of aircraft safety control, in particular to an automatic detection and emergency recovery method for an out-of-control aircraft attitude. BACKGROUND

[0002] In the field of aircraft safety flight technology, the attitude control of aviation and aerospace vehicles is crucial to their safe operation. Whether it is civil aviation aircraft, large unmanned aerial vehicles and other aviation equipment, or first-level rockets, space stations, lunar landers and other space devices, precise attitude detection and adjustment are needed during flight to ensure flight safety and avoid accidents or mission failures caused by attitude loss of control.

[0003] In the prior art, the attitude control of manned civil aircraft, fighter aircraft and other aircraft mainly relies on pilot operation, and the flight state is judged and the flight attitude is adjusted by manual operation. Although the spacecraft has a certain attitude stabilization system during normal flight, it lacks an automatic detection and emergency recovery mechanism in the event of an accident such as a collision. The landing point positioning and attitude adjustment accuracy of the first-level rocket is insufficient during the recovery process. The lunar lander relies on a preset program for blind descent during the landing phase and lacks real-time dynamic attitude correction means.

[0004] However, the prior art has obvious deficiencies: manned aircraft are prone to accidents such as crashing into mountains, buildings and the sea due to human operational errors; spacecraft often roll uncontrollably after being hit and cannot automatically stop; the success rate of safely landing on the launch pad is low due to positioning and attitude adjustment accuracy problems during the recovery of first-level rockets; the blind descent mode of the lunar lander results in a low probability of safe landing and a large landing point error, which seriously affects the flight safety and mission execution efficiency of the aircraft. In view of this, we propose an automatic detection and emergency recovery method for an out-of-control aircraft attitude. SUMMARY

[0005] To overcome the deficiencies of the prior art, the present application provides an automatic detection and emergency recovery method for an out-of-control aircraft attitude, which solves the problem of the existing manned civil aircraft, fighter aircraft and other aircraft crashing into mountains, buildings and the sea due to human factors, and the problem of the spacecraft rolling uncontrollably after being hit and being unable to automatically stop rolling. The problem of low probability of safe landing of the lunar lander.

[0006] To achieve the above purpose, the present application is implemented by the following technical scheme: an automatic detection and emergency recovery method for an out-of-control aircraft attitude, comprising the following steps:

[0007] S1: signal networking transmission step

[0008] Three non-collinear phased array radars are deployed on the ground or three non-coplanar satellites are deployed in space to form a triangular positioning signal transmitting array, which periodically transmits radio frequency signals containing position coordinates according to preset encoding rules;

[0009] S2: multi-dimensional signal solving step

[0010] The aircraft receives the time domain, frequency domain and phase difference signals of the three signal sources obtained in S1, fuses the inertial navigation data through Kalman filtering algorithm, and solves the three-dimensional coordinates, Euler angles and attitude angular velocity;

[0011] S3: dynamic threshold determination step

[0012] According to the comparison between the real-time attitude parameters of the aircraft obtained in S2 and the dynamic threshold library pre-stored based on the aircraft type including civil aviation, rocket, lander and environment atmosphere, vacuum, a deviation level signal is generated;

[0013] S4: hierarchical trigger response step

[0014] According to the deviation level generated in S3, the states of warning, semi-connection and full-connection are generated, and the power supply activation and communication protocol handshake of the corresponding control module are triggered according to the preset logic;

[0015] S5: dynamics correction execution step

[0016] According to the aircraft information obtained from S2 to S4, combined with the real-time aerodynamic parameters of the aircraft or the celestial mechanics model, the attitude correction is realized through the adjustment of the aerodynamic rudder angle, the pulse injection of the propellant or the control of the momentum wheel speed;

[0017] S6: closed-loop feedback verification step

[0018] After the correction is executed, the new attitude data is collected in real time through S1 signal networking transmission step, and compared with the safety threshold in S3 dynamic threshold determination step to form a control closed loop until the attitude parameters are stable within the allowed range.

[0019] Preferably, in the signal networking transmission step, the ground radar adopts X-band phased array system, and the ranging accuracy is optimized to ≤0.1m through Kalman filtering iteration, and the radar array is arranged according to Fermat point principle, and the control positioning blind area is reduced by 40%.

[0020] Preferably, in the signal networking transmission step, the space satellite adopts GPS / Beidou dual-mode positioning system, and the clock deviation is controlled within 10ns through inter-satellite link time synchronization technology, and the positioning accuracy is optimized to ≤10cm through carrier phase difference technology.

[0021] Preferably, in the multi-dimensional signal solving step, the time difference positioning method is used to calculate the signal arrival time difference, the pseudo-range measurement value of the satellite signal is combined, and the least square method is used to iteratively solve the aircraft spatial coordinates, and the solving period is ≤10 ms.

[0022] Preferably, in the multi-dimensional signal solving step, for the lunar lander scene, the 1550nm pulse signal transmitted by the laser radar beacon is used to measure the distance by the time of flight (ToF) method, and the six-degree-of-freedom attitude parameters relative to the landing point are solved by the extended Kalman filter (EKF) algorithm combined with the inertial navigation data.

[0023] Preferably, in the dynamic threshold determination step, a multi-scene threshold database containing the vertical deviation threshold of the civil aircraft, the horizontal deviation threshold of the rocket recovery, and the inclination threshold of the lunar lander is established, and the threshold parameters are dynamically adjusted according to the flight stages of take-off, cruise, and landing.

[0024] Preferably, in the hierarchical trigger response step, the three-level response mechanism includes: in the first-level response, the PWM signal is used to drive the sound-light warning module; in the second-level response, the CAN bus communication protocol is started to take over 30% of the aerodynamic control authority; and in the third-level response, the RS485 communication link is activated to completely take over the flight control, and the response delay is ≤50 ms.

[0025] Preferably, in the dynamics correction execution step, for the civil aircraft, the flap / aileron cooperative control algorithm is used to generate the PID control parameters according to the deviation matrix; and for the spacecraft, the attitude update algorithm based on the quaternion is used to calculate the pulse width modulation parameters through the jet control moment.

[0026] Preferably, in the dynamics correction execution step, the azimuth angle of the launch pad is obtained in real time by the millimeter wave radar during the first-level rocket recovery, the model predictive control algorithm is used to calculate the deflection angle of the gas rudder, and the propellant injection timing is adjusted combined with the pressure sensor data of the landing leg cushion system.

[0027] Preferably, in the closed-loop feedback verification step, the sliding window filtering algorithm is used to smooth the corrected attitude data, and when the parameter fluctuation of the continuous 5 sampling periods is ≤5%, it is determined that the correction is completed, otherwise the secondary correction process is started.

[0028] The application provides an automatic detection and emergency recovery method for an out-of-control aircraft attitude.

[0029] 1. The application positions the aircraft position by three signal transmitters, calculates the distance positioning by the built-in computing unit, triggers the attitude correction system to adjust the attitude, allows civil aviation aircraft and the like to land safely when the vertical distance is below the safety standard, avoids human accidents; makes the first-stage rocket position the launch frame and lands safely or returns; allows the lunar landing cabin to land vertically, improves the safety landing probability and reduces errors; and also allows the rolling spacecraft to stop rolling and ensures the flight safety of various aircrafts.

[0030] 2. The application improves the control accuracy and adaptability through dynamic parameter adjustment, multi-dimensional deviation fusion and other mechanisms. The PID algorithm realizes accurate control of the aerodynamic surface, the quaternion algorithm avoids the gimbal lock problem of spacecraft attitude solution, the MPC algorithm optimizes the rocket trajectory combined with multi-source data, the Kalman and sliding window filtering suppresses noise and stabilizes the closed loop, and the multi-scene threshold mechanism reduces false triggering, which together ensures the accuracy and robustness of the aircraft attitude detection and recovery.

[0031] 3. The application improves the radar ranging to 0.1m through Kalman filtering, allows satellite positioning to 10cm through carrier phase difference, reduces the radar positioning blind area by 40% through Fermat point layout, controls the satellite clock deviation within 10ns through inter-satellite link time synchronization, improves channel utilization through TDMA technology, and guarantees signal transmission through protocol and verification, thereby realizing collaborative optimization of space-time-ground multi-scene positioning accuracy, coverage area and time reference. BRIEF DESCRIPTION OF DRAWINGS

[0032] Figure 1 It is a flow chart of the attitude automatic detection and emergency recovery method of the out-of-control aircraft.

[0033] Figure 2 It is a distance correction schematic diagram of the civil aviation aircraft, large unmanned aerial vehicle, general fighter and three radars in the application.

[0034] Figure 3 It is a distance correction schematic diagram of the first-stage rocket and three radars in the application.

[0035] Figure 4 It is a distance correction schematic diagram of the lunar landing cabin and the lunar surface three radars in the application.

[0036] Figure 5 It is a positioning schematic diagram of the space station, spaceship, extravehicular space suit in uncontrolled rolling and three satellites in the application.

[0037] Figure 6 It is a signal networking transmission algorithm flow chart of the application. DETAILED DESCRIPTION

[0038] With reference to the drawings of the present application in the specification, the technical solutions in the embodiments of the present application will be described clearly and completely. Obviously, the described embodiments are only a part of the embodiments of the present application, but not all the embodiments. Based on the embodiments in the present application, all the other embodiments obtained by those of ordinary skill in the art without creative work belong to the scope of protection of the present application.

[0039] Embodiments:

[0040] Please refer to the drawings of the present application in the specification, the technical solutions in the embodiments of the present application will be described clearly and completely. Obviously, the described embodiments are only a part of the embodiments of the present application, but not all the embodiments. Based on the embodiments in the present application, all the other embodiments obtained by those of ordinary skill in the art without creative work belong to the scope of protection of the present application. Figure 1 - the drawings of the present application in the specification, the technical solutions in the embodiments of the present application will be described clearly and completely. Obviously, the described embodiments are only a part of the embodiments of the present application, but not all the embodiments. Based on the embodiments in the present application, all the other embodiments obtained by those of ordinary skill in the art without creative work belong to the scope of protection of the present application. Figure 6 , the present application provides a kind of out-of-control aircraft attitude automatic detection and emergency recovery method, comprising the following steps:

[0041] S1: signal networking transmission step

[0042] Three non-collinear phased array radars are deployed on the ground or three non-coplanar satellites are deployed in space to build a triangular positioning signal transmission array, and periodically transmit radio frequency signals containing position coordinates according to a preset encoding rule;

[0043] S2: multi-dimensional signal solving step

[0044] The aircraft receives the time domain, frequency domain and phase difference signals of the three signal sources obtained in S1, fuses the inertial navigation data through Kalman filtering algorithm, and solves three-dimensional coordinates, Euler angles and attitude angular velocity;

[0045] S3: dynamic threshold determination step

[0046] According to the real-time attitude parameters of the aircraft obtained in S2 and the dynamic threshold library based on the aircraft type including civil aviation, rocket, lander and environment atmosphere, vacuum pre-stored, deviation level signal is generated by comparison;

[0047] S4: hierarchical trigger response step

[0048] According to the deviation level generated in S3, the state programs of early warning, half takeover and full takeover are triggered according to the preset logic to activate the power supply of the corresponding control module and handshake the communication protocol;

[0049] S5: dynamics correction execution step

[0050] According to the aircraft information obtained from S2 to S4, combined with real-time aerodynamic parameters of the aircraft or celestial mechanics model, attitude correction is realized through adjustment of aerodynamic rudder angle, propellant pulse injection or momentum wheel speed control;

[0051] S6: closed-loop feedback verification step

[0052] After the correction is executed, the new attitude data is collected in real time through the signal network transmission step, and compared with the safety threshold in the S3 dynamic threshold determination step to form a control closed loop until the attitude parameters are stable within the allowed range.

[0053] In the signal network transmission step, the ground radar adopts an X-band phased array system, synchronously transmits pulse signals through time division multiple access technology, and the ranging accuracy is optimized to ≤0.1m through Kalman filtering iteration. The radar array is arranged according to the Fermat point principle, and the positioning blind area is reduced by 40%.

[0054] In the signal network transmission step, the space satellite adopts a GPS / Beidou dual-mode positioning system, and the clock bias is controlled within 10ns through inter-satellite link time synchronization technology. The positioning accuracy is optimized to ≤10cm through carrier phase difference technology, including the following algorithm steps:

[0055] I. Geometric layout planning of ground radar array

[0056] The three radar deployment coordinates are determined based on the Fermat point optimization principle, and an equilateral triangle array with a side length of 15km is constructed. When each internal angle of the triangle is less than 120°, the Fermat point P satisfies ∠APB=∠BPC=∠CPA=120°, and the layout coordinates are solved by minimizing the objective function min P (|PA|+|PB|+|PC|), where A, B, and C are preset radar sites. This layout makes the positioning blind area reduced by 4096 compared to the traditional triangular layout, and the geometric constraint ensures that there is no dead angle for three-dimensional positioning of the aircraft at any position within the array coverage range.

[0057] II. Hardware parameter configuration of ground radar

[0058] X-band phased array radar is deployed, with a working frequency band of 8-12GHz. Time division multiple access (TDMA) technology is used to divide the transmission time slots: radar A occupies time O-1ms, radar B occupies time 1-2ms, and radar C occupies time 2-3ms to avoid signal conflicts. Each pulse signal code contains three-dimensional coordinates (X i ,Y i ,Z i ) of the radar, transmission time t i , and signal strength parameters. The transmission power is controlled at 200-500W to ensure a coverage distance of 20km, and the antenna beam width is ≤1.5° to ensure direction finding accuracy.

[0059] III. Iterative optimization of ground radar ranging accuracy

[0060] The Kalman filtering algorithm is used to process the radar ranging data in real time. First, the current ranging value is predicted based on the state at the previous time:

[0061]

[0062]

[0063] Then according to the observed value z k The Kalman gain is calculated And the state estimate is corrected:

[0064]

[0065] After 10 iterations, the ranging accuracy converges to ≤0.1m, the process noise covariance Q k is set to 0.01, and the observation noise covariance R k is set to 0.5.

[0066] Four, Space Satellite Clock Synchronization Mechanism Implementation

[0067] Calibrate three satellite clocks using inter-satellite link two-way time transfer (TWTT) technology: satellite A sends a signal to satellite B at time t A , B receives it at t' A Satellite B sends a signal to satellite A at time t B , A receives it at t' B Clock bias:

[0068]

[0069] Exchange timestamps every 100ms through inter-satellite communication links, and use weighted average method to fuse three measurement results, finally control the clock bias within 10ns, ensure the uniformity of positioning signal time reference;

[0070] Five, Space Satellite Positioning Signal Enhancement Processing

[0071] Use carrier phase difference (RTK) technology to improve positioning accuracy: the difference between the carrier phase observation values of satellite k at the reference station and the user end

[0072]

[0073] Where f is the carrier frequency 1575.42MHz, c is the speed of light, and the integer ambiguity is calculated by floating point And use double difference method to eliminate satellite clock error δt k , receiver clock error δt i , δt j , finally optimize the positioning accuracy to ≤10cm;

[0074] Six, Signal Transmission Protocol and Cycle Setting

[0075] The ground radar and the space satellite adopt a 10ms transmission cycle to ensure that the aircraft receives signals in real time. The ground radar follows the pulse code modulation (PCM) protocol, and the data frame format is: preamble (8bit) + radar coordinates (48bit) + time stamp (32bit) + check code (16bit). The space satellite follows the NMEA-0183 protocol, and the GPGGA statement contains longitude, latitude, altitude and time stamp information. The baud rate is set to 9600bps, and the error rate is controlled at 10 -6 The following.

[0076] In the multi-dimensional signal solving step, the time difference positioning method is used to calculate the signal arrival time difference, and the least squares method is used to iteratively solve the aircraft spatial coordinates combined with the pseudo-range measurement value of the satellite signal, and the solving period is ≤10ms.

[0077] In the multi-dimensional signal solving step, for the lunar lander scene, the 1550nm pulse signal transmitted by the laser radar beacon is used to measure the distance by time of flight (ToF), and the six-degree-of-freedom attitude parameters relative to the landing point are solved by the extended Kalman filter (EKF) algorithm combined with the inertial navigation data.

[0078] In the dynamic threshold determination step, a multi-scenario threshold database is established, including the vertical deviation threshold of civil aircraft, the horizontal deviation threshold of rocket recovery, and the inclination threshold of lunar lander. The threshold parameters are dynamically adjusted according to the flight stages of take-off, cruise and landing.

[0079] In the hierarchical trigger response step, the three-level response mechanism includes: in the first-level response, the PWM signal is used to drive the sound and light warning module; in the second-level response, the CAN bus communication protocol is started to take over 30% of the aerodynamic control authority; in the third-level response, the RS485 communication link is activated to completely take over the flight control, and the response delay is ≤50ms.

[0080] In the dynamics correction execution step, for civil aircraft, aileron / aileron cooperative control algorithm is used to generate PID control parameters according to the deviation matrix; for spacecraft, a quaternion-based attitude update algorithm is used to calculate the pulse width modulation parameters through jet control moment.

[0081] In the dynamics correction execution step, the azimuth angle of the launch frame is obtained in real time by the millimeter wave radar during the first-level rocket recovery, the model predictive control algorithm is used to calculate the deflection angle of the gas rudder, and the propellant injection timing is adjusted combined with the pressure sensor data of the landing leg cushion system, including the following algorithms:

[0082] I. Signal networking transmission step

[0083] Three non-collinear phased array radars are deployed on the ground according to the Fermat principle to form an equilateral triangle array with a side length of 15 km, and X-band time division multiple access technology is used to synchronously transmit pulse signals. The signal code contains the three-dimensional coordinates (X1, Y1, Z1), (X2, Y2, Z2), and (X3, Y3, Z3) of the radars themselves and time tag information. 2i Y 2i Z2)(X3,Y3,Z3) and time tag information. Three non-coplanar satellites are deployed in space, using GPS / Beidou dual-mode positioning system, and through inter-satellite link time synchronization technology, the clock bias is controlled within 10 ns, and the radio frequency signal containing ephemeris parameters is transmitted according to NMEA-0183 protocol. The radar ranging accuracy is optimized to ≤0.1 m through Kalman filter iteration, and the satellite positioning accuracy is optimized to ≤10 cm through carrier phase difference technology, and the signal transmission period is set to 10 ms to ensure real-time performance.

[0084] II. Multi-dimensional signal solving steps

[0085] The three-channel radio frequency receiver on the aircraft collects the time domain, frequency domain and phase difference signals of the three signal sources in real time. For the ground radar scene, the time difference positioning method (TDoA) is used to calculate the signal arrival time difference Δt ij (i,j=1,2,3), combined with the radar coordinates, the aircraft coordinates (X, Y, Z) are solved through the following equation set:

[0086]

[0087] Where c is the speed of light. For the satellite scene, the pseudo-range measurement value

[0088]

[0089] is solved through least squares iteration, and the solving period is ≤10 ms. At the same time, the angular velocity ω = [ω x ω v ω z ] T output by the inertial navigation is fused, and the aircraft Euler angles (pitch angle Θ, yaw angle ψ, roll angle φ) are estimated through Kalman filter algorithm. The state equation is:

[0090]

[0091] III. Dynamic threshold determination step

[0092] A multi-scene threshold database is established. The vertical deviation threshold of civil aviation aircraft in the cruising stage is set to 100 m, and the landing stage is adjusted to 30 m; the horizontal deviation threshold of the first-stage rocket during recovery is initially set to 100 m, and is dynamically reduced to 10 m after 1 km from the launch pad; the inclination threshold of the lunar lander in the hovering stage is 5°, and is tightened to 2° when it is 10 m away from the lunar surface. Real-time attitude parameters (X, Y, Z, Θ, ψ, ) and compared with the threshold value, the Mahalanobis distance is used to calculate the deviation level:

[0093] D 2 = (x - μ) T Σ -1 (x - μ)

[0094] Where μ is the mean vector of the safe posture, and Σ is the posture parameter covariance matrix. When D 2 exceeds the preset level threshold (such as first-level response D 2 ≤ 0.3σ 2 , second-level response 0.3σ 2 <D 2 ≤ 0.7σ 2 , third-level response D 2 > 0.7σ 2 ), a corresponding deviation level signal

[0095] Four, the step of triggering the response by levels

[0096] The third-level response mechanism is realized by a state machine: when the first-level response, the CPU sends a square wave signal with a duty cycle of 50% to the PWM controller to drive the sound and light warning module; when the second-level response, a message with ID = 0x18FF5000 is sent through the CAN bus to take over 30% of the pneumatic control authority, and the control law is:

[0097]

[0098] Where δ s , δ, δ, are the deflection angles of the elevators, ailerons, and rudders, respectively, and θ ern , ψern, ψerr are the attitude angle deviations; when the third-level response, the RS485 communication link is activated to send a 16-bit data frame containing control instructions (such as Ox55AAO3O1FFOO…), and the flight control is completely taken over, and the response delay is controlled to be ≤ 50ms through interrupt priority scheduling control;

[0099] Five, the step of executing the dynamics correction (1) civil aviation aircraft correction

[0100] Aileron / aileron cooperative control algorithm is used to generate PID control parameters according to the attitude deviation matrix E = [θ em φ em ψ em ] T The control amount calculation formula is:

[0101]

[0102] Where e(t) is the projection of the deviation vector in each control channel. Taking the pitch angle control as an example, the elevator deflection angle δ e = K pr Θerr +K ir [θ ern dt+K pr dθ ern The parameter [ / dt] is tuned using the Ziegler-Nichols method.

[0103] K pt =0.8·K u

[0104] K it =K pt / (1.2·T u )

[0105] K dt =K pt ·T u / 8

[0106] Where K u For the critical gain, T u This is the critical period.

[0107] (2) Spacecraft Correction

[0108] Based on the quaternion attitude update algorithm, the initial quaternion q0 = [1,0,0,0]T, and the real-time angular velocity ω is obtained through gyroscope measurement. The attitude update equation is:

[0109]

[0110] in Quaternion multiplication:

[0111]

[0112] Based on the target quaternion q d With the current quaternion q c deviation Calculate the required control torque M = k·[Δq] x Δq w Δq z ] τ The thrust F and lever arm L of the jet actuator are converted into pulse width modulation parameters, and the pulse width t is... p =k p ·|M| / (F·L), where k, k p This is the proportionality coefficient;

[0113] First-stage rocket recovery correction

[0114] By acquiring the launch pad azimuth angle α and elevation angle β in real time using millimeter-wave radar, the relative position vector r = [r] between the rocket's center of mass and the launch pad is established. x ,ry ,r z ] T , the model predictive control (MPC) algorithm is used to solve the optimal gas rudder deflection angle δ, and the optimization objective function is:

[0115]

[0116] The constraint conditions include the rocket dynamics model:

[0117]

[0118] Where F x , F y is the control force component generated by the gas rudder deflection, and m is the real-time mass of the rocket. Combined with the landing leg cushion system pressure sensor signal P(t), when P(t) exceeds the threshold P c , the propellant injection timing adjustment is triggered, and the injection duration t on = k p ·(P(t)-P c )+k i ·∫(P(τ)-P c )dτ;

[0119] Six, closed-loop feedback verification steps

[0120] The corrected attitude data is processed by using the sliding window filtering algorithm, and the window length is set to 5 sampling periods (50ms), and the filtering formula is:

[0121]

[0122] Where n=5, x(i) is the attitude parameter of the i-th sampling point. When the parameter fluctuation Δx=|X (t) -X (t-1) | / X (ref) ≤5% of the continuous 5 windows, it is determined that the correction is completed; otherwise, the secondary correction process is started, and the control amount is adjusted by the incremental PID algorithm:

[0123] Δu(k)=K p [e(k)-e(k-1)]+K i e(k)+K d [e(k)-2e(k-1)+e(k-2)]

[0124] Where e(k) is the deviation value of the k-th sampling, until the attitude parameter is stable within the safety threshold range.

[0125] In the closed loop feedback verification step, the sliding window filtering algorithm is used to smooth the corrected attitude data, and when the parameter fluctuation of 5 consecutive sampling periods is less than or equal to 5%, it is determined that the correction is completed, otherwise the secondary correction process is started.

[0126] While embodiments of the application have been shown and described, it is to be understood that the embodiments described are merely exemplary of the principles and application of the present application. Numerous modifications, changes, variations, substitutions, and equivalents will occur to those of ordinary skill in the art without departing from the spirit and scope of the application as defined by the following claims and their equivalents.

Claims

1. An uncontrolled aircraft attitude automatic detection and emergency recovery method, characterized in that, Comprise the following steps: S1: signal networking transmission step Deploy three non-collinear phased array radars on the ground or three non-coplanar satellites in space to build a triangular positioning signal transmission array, and periodically transmit radio frequency signals containing position coordinates according to preset coding rules; S2: multi-dimensional signal solving step The aircraft receives the time domain, frequency domain and phase difference signals obtained in S1, fuses the inertial navigation data through Kalman filtering algorithm, and solves the three-dimensional coordinates, Euler angles and attitude angular velocity; S3: dynamic threshold determination step According to the real-time attitude parameters of the aircraft obtained in S2 and the dynamic threshold library based on the aircraft type including civil aviation, rocket, lander and environment atmosphere, vacuum pre-stored, the deviation level signal is generated by comparing; S4: hierarchical trigger response step According to the deviation level generated in S3, the state programs of early warning, half takeover and full takeover are generated, and the power activation and communication protocol handshake of the corresponding control module are triggered according to the preset logic; S5: dynamics correction execution step According to the aircraft information obtained from S2 to S4, combined with the real-time aerodynamic parameters of the aircraft or the celestial mechanics model, the attitude correction is realized through the adjustment of the aerodynamic rudder angle, the pulse jet of the propellant or the control of the momentum wheel speed; S6: closed loop feedback verification step After the correction is executed, the new attitude data is collected in real time through S1 signal networking transmission step, and the safety threshold in S3 dynamic threshold determination step is compared to form a control closed loop until the attitude parameters are stable within the allowed range.

2. The automatic detection and emergency recovery method for the attitude of an out-of-control aircraft according to claim 1, characterized in that, In the signal networking transmission step, the ground radar adopts X-band phased array system, and the pulse signals are transmitted synchronously through time division multiple access technology. The ranging accuracy is optimized to ≤0.1m through Kalman filter iteration. The radar array is arranged according to Fermat point principle, and the positioning blind area is reduced by 40%.

3. The method of claim 2, wherein the step of detecting the out-of-control flight attitude of the aircraft is performed by a flight control computer. In the signal networking transmission step, the space satellite adopts GPS / Beidou dual-mode positioning system, and the clock bias is controlled within 10ns through inter-satellite link time synchronization technology. The positioning accuracy is optimized to ≤10cm through carrier phase difference technology.

4. The automatic detection and emergency recovery method for the attitude of an out-of-control aircraft according to claim 1, characterized in that, In the multi-dimensional signal solving step, the time difference positioning method is used to calculate the time difference of the signal, and the least square method is used to solve the aircraft spatial coordinates combined with the pseudo-range measurement value of the satellite signal, and the solving period is ≤10ms.

5. The automatic detection and emergency recovery method for the attitude of an out-of-control aircraft according to claim 4, characterized in that, In the multi-dimensional signal solving step, for the lunar lander scene, the 1550nm pulse signal transmitted by the laser radar beacon is used to measure the distance by time of flight (ToF), and the six-degree-of-freedom attitude parameters relative to the landing point are solved by extended Kalman filter (EKF) algorithm combined with inertial navigation data.

6. The automatic detection and emergency recovery method for the attitude of an out-of-control aircraft according to claim 1, characterized in that, In the dynamic threshold determination step, a multi-scenario threshold database is established, including the vertical deviation threshold of civil aircraft, the horizontal deviation threshold of rocket recovery and the inclination threshold of lunar lander. The threshold parameters are dynamically adjusted with the flight stage of the aircraft, such as take-off, cruise and landing.

7. The method of claim 1, wherein the method further comprises: The hierarchical trigger response step, the three-level response mechanism includes: the first response through the PWM signal drive sound light warning module; the second response starts CAN bus communication protocol takes over 30% pneumatic control right; the third response activates RS485 communication link completely takes over flight control, response delay ≤50ms.

8. The automatic detection and emergency recovery method for the attitude of an out-of-control aircraft according to claim 1, characterized in that, In the dynamic correction execution step, for civil aviation aircraft, aileron / flap collaborative control algorithm is adopted, PID control parameters are generated according to the deviation matrix; for spacecraft, a quaternion-based attitude updating algorithm is adopted, and pulse width modulation parameters are calculated through jet control moment.

9. An uncontrolled aircraft attitude automatic detection and emergency recovery method according to claim 8, characterized in that, In the dynamic correction execution step, when the first rocket recovers, the azimuth of the launching frame is obtained in real time through the millimeter wave radar, the model predictive control algorithm is used to calculate the deflection angle of the gas rudder, and the propellant injection timing is adjusted combined with the pressure sensor data of the landing leg cushion system.

10. The method of claim 1, wherein, In the closed-loop feedback verification step, the sliding window filtering algorithm is used to smooth the corrected attitude data, and when the parameter fluctuation of 5 consecutive sampling periods is ≤5%, it is determined that the correction is completed, otherwise the secondary correction process is started.

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