A catapult type reconnaissance and attack integrated patrol aircraft attitude control system and method

Abstract

The present application relates to the technical field of non-electric variable control, and discloses a posture control system and method for a catapult type reconnaissance and attack integrated cruise vehicle, comprising a high dynamic inertia measurement module, a pipe transient transition smoothing filter module, a state response interval solving module, a dynamic saturation limiting amplitude adaptive distribution module, a nonlinear adaptive damping solving module and a control loop arbitration distribution module. The present application generates a rigid locking instruction by the state response interval solving module when the real-time flight speed of the cruise vehicle and the height relative to the launch point are in a preset launch transition interval, fixes the time domain dynamic gain compensation coefficient, and suspends the decay logic adjusted by the time constant decay factor. By high-bit proportional overload correction, the angular displacement deviation is corrected, the limit correction overload coverage boundary is adaptively extended with the physical phase change state, the angular velocity suppression time divergence problem caused by the open-loop time axis misalignment is solved, the physical saturation of the rudder is prevented, and a high-damping priority control response loop is constructed.

Description

Technical Field

[0001] This invention relates to an attitude control system and method for a catapult-launched reconnaissance and strike integrated loitering drone, belonging to the field of non-electric variable control technology. Background Technology

[0002] In the current overall design architecture of tube-launched high-dynamic loitering weapons, the control loop typically employs a time-axis gain scheduling strategy to suppress initial attitude jumps. This architecture relies on the repeatability of the aircraft's state during the tube exit transition phase. By attenuating the pre-stage correction control power through a fixed time-domain open-loop curve, when the tube-launched system applies initial propulsion, the flow field on the surface of the launcher and the folded aerodynamic control surfaces generate an asymmetric transient impact at the moment the constraint at the tube opening is released. The axial and radial spatial angular velocity vectors acquired by the control loop subsequently undergo a jump, causing an instantaneous breakdown of the momentum balance within the system.

[0003] The displacement deviation during the rollout caused by changes in operating conditions leads to random spatiotemporal drift in the phase of high-dynamic impact. Simply relying on optimizing the launch tube constraints or the design of hardware physical structures such as folding wings is limited by the inherent constraints of the overall weight and aerodynamic shape of the launch vehicle, and cannot eliminate multi-dimensional coupled high-dynamic disturbances. Coordinated optimization of software dimensions such as control methods becomes the key to ensuring attitude stability in the initial stage of launch. For example, Chinese invention patent application CN115682832A discloses a loitering munition launch method and system. When the axial acceleration is greater than a set value, a fixed delay period is used to wait for the projectile to exit the tube and the wings to unfold, and the motor and control surface deflection are controlled to correct the flight attitude. The technical solution still implicitly relies on the absolute ideal consistency of the launch conditions at the underlying logic, and the open-loop time axis scheduling strategy adopted cannot be compatible with actual conditions. During launch, random spatiotemporal drift can occur due to factors such as displacement deviation during tube gliding and turbulent flow. When the actual physical tube release time and the theoretically preset time axis intersection point are misaligned by milliseconds, the control loop is prone to premature attenuation of the suppression gain during the severe overload window of the preceding stage, resulting in a control weight deficit. This open-loop scheduling mode makes the control loop dependent on the consistency of the launch environment and difficult to be compatible with the nonlinear extension of external disturbance energy on the spatiotemporal axis, which in turn leads to initial attitude divergence. Increasing the proportional gain of the control loop or introducing external sensing hardware is not enough to eliminate spatiotemporal displacement fluctuations. Increasing the proportional gain can easily cause the servo ESC to fall into the physical saturation dead zone during large overload correction, causing the differential and damping control terms to fail, the loop to degenerate into an open-loop state, and induce attitude instability. Adding sensors increases the mass of the device and introduces signal delay.

[0004] Therefore, how to accurately capture the launch transition phase using the characteristics of spatial kinematics intervals, and construct a high-damping priority control response loop while effectively preventing physical saturation of the servo motor, has become the technical problem to be solved by this invention. Summary of the Invention

[0005] To address the problems in the background art, the technical solution of the present invention is as follows: A catapult-launched reconnaissance and strike integrated loitering rovers attitude control system, comprising:

[0006] The high-dynamic inertial measurement module collects angular velocity and axial acceleration data.

[0007] The tube-out transient transition smoothing filter module establishes a spatiotemporal correction anchor point and calculates the time constant decay factor when the axial acceleration data suddenly drops to a preset acceleration threshold.

[0008] The control loop arbitration and allocation module combines the time-domain dynamic gain compensation coefficient and the square-term nonlinear damping term to synthesize the final servo control command.

[0009] The state response interval calculation module generates a rigid lock command when the real-time flight speed of the loitering drone and the altitude relative to the launch point are within the preset launch transition interval. This command locks the time-domain dynamic gain compensation coefficient at the preset maximum gain limit value and suspends the attenuation logic. The attenuation rate of the attenuation logic is adjusted by the time constant attenuation factor.

[0010] The dynamic saturation limiting adaptive allocation module generates a negative feedback penalty factor when the absolute value of the final servo control command exceeds the physical saturation threshold. This factor is then input to the control loop arbitration allocation module to reduce the weight of the proportional control term. The nonlinear adaptive damping solution module outputs a squared nonlinear damping term to reduce the transient divergence rate scalar.

[0011] Preferably, the tube exit transient transition smoothing filter module further includes: establishing a spatiotemporal correction anchor point when the axial acceleration data suddenly drops to between 0.5 and 0.8 times the preset gravitational acceleration; calculating the actual time deviation between the spatiotemporal correction anchor point and the preset theoretical tube exit time point; if the actual time deviation is greater than 0, reducing the time constant attenuation factor; if the actual time deviation is less than or equal to 0, increasing the time constant attenuation factor.

[0012] Preferably, the state response interval calculation module further includes: retrieving the real-time flight speed of the loitering drone and its altitude relative to the launch point; when the real-time flight speed is between 14m / s and 35m / s and the altitude relative to the launch point is between 0.5m and 10m, the loitering drone is determined to be in the preset launch transition interval and a rigid lock command is generated.

[0013] Preferably, the maximum gain limit value set in the state response interval solution module is between 3.5 and 5.5, and during the period when the rigid locking command is in effect, the state response interval solution module locks the time-domain dynamic gain compensation coefficient at the maximum gain limit value until the nonlinear adaptive damping solution module reduces the transient divergence rate scalar to below the preset safe divergence threshold.

[0014] Preferably, the dynamic saturation limiting adaptive allocation module further includes: retrieving the final servo control command and comparing the absolute value of the final servo control command with the difference between the final servo control command and the physical saturation threshold, wherein the physical saturation threshold corresponds to 90% of the standard drive voltage amplitude of the servo execution module; generating a negative feedback penalty factor when the absolute value of the final servo control command exceeds the physical saturation threshold; and introducing the negative feedback penalty factor into the proportional control term in the control loop arbitration allocation module to reduce the weight of the proportional control term allocation.

[0015] Preferably, the nonlinear adaptive damping solution module further includes: retrieving the measured angular velocity data of the current cycle, calculating the product of the measured angular velocity data and its own absolute value to generate a squared nonlinear damping term; and inputting the squared nonlinear damping term into the attitude control loop to suppress the difference slope between the amplitude and the angular velocity data, thereby providing a damping overload with high-frequency disturbance limiting cutoff characteristics.

[0016] Preferably, the control loop arbitration allocation module is connected to the dynamic saturation limiting adaptive allocation module to collect and synthesize the final servo control command containing the proportional control term and the square term nonlinear damping term, and to receive the negative feedback penalty factor output by the dynamic saturation limiting adaptive allocation module when the absolute value of the final servo control command exceeds the physical saturation threshold.

[0017] Preferably, the high dynamic inertial measurement module includes a high dynamic angular velocity detection module and an axial acceleration detection module. The angular velocity acquisition range of the high dynamic angular velocity detection module is not less than 400 rad / s, so as to achieve angular velocity range coverage of the initial momentum disturbance caused by catapult impact and the extreme transient asymmetric aerodynamic impact torque generated when the fore and aft wings of the tandem wing are asynchronously deployed.

[0018] Preferably, the system calibrates the physical saturation threshold by adjusting the power supply pulse width modulation ratio of the standard drive voltage of the servo execution module, and introduces a negative feedback penalty factor by adjusting the proportional gain matrix in the control loop arbitration allocation module. Under the hierarchical amplitude control state of the state response interval solution module and the nonlinear adaptive damping solution module, the pitch angle divergence and yaw angle divergence of the loitering aircraft within a second (e.g., 2s to 5s) after leaving the control tower are both less than 2°.

[0019] A method for attitude control of a catapult-launched reconnaissance and strike loitering munition, used to operate an attitude control system for a catapult-launched reconnaissance and strike loitering munition, includes the following steps:

[0020] Step S1: Acquire angular velocity and axial acceleration data using a high dynamic inertial measurement module;

[0021] Step S2: When the axial acceleration data suddenly drops to a preset acceleration threshold, the spatiotemporal correction anchor point is established and the time constant decay factor is calculated using the tube exit transient transition smoothing filter module.

[0022] Step S3: The final servo control command is synthesized by combining the time-domain dynamic gain compensation coefficient and the square-term nonlinear damping term with the control loop arbitration allocation module.

[0023] Step S4: When the real-time flight speed of the loitering vehicle and the altitude relative to the launch point are in the preset launch transition range, the state response interval calculation module generates a rigid locking command to lock the time-domain dynamic gain compensation coefficient at the preset maximum gain limit value and suspend the attenuation logic. The attenuation rate of the attenuation logic is adjusted by the time constant attenuation factor.

[0024] Step S5: When the absolute value of the final servo control command exceeds the physical saturation threshold, the dynamic saturation limit adaptive allocation module generates a negative feedback penalty factor and inputs it to the control loop arbitration allocation module to reduce the weight of the proportional control term. The nonlinear adaptive damping solution module outputs a squared nonlinear damping term to reduce the transient divergence rate scalar.

[0025] Compared with the prior art, the beneficial effects of the present invention are:

[0026] 1. In the attitude control of the catapult-launched reconnaissance and strike loitering pod, the state response interval operator obtains the real-time flight speed and relative launch point altitude of the loitering pod in real time. When the real-time motion state of the loitering pod enters the launch transition interval composed of the preset flight speed and relative altitude, the operator generates a rigid locking command, which forcibly locks the time-domain dynamic gain compensation coefficient in the loop to the preset maximum gain limit value, suspends the attenuation program that decreases monotonically with time axis, and uses high-level proportional overload to correct angular displacement deviation, so that the coverage boundary of the limit correction overload adaptively extends with the physical phase change state, solving the problem of angular velocity suppression time divergence caused by open-loop time axis misalignment.

[0027] 2. The dynamic saturation limiting adaptive allocation unit retrieves the servo control command synthesized by the loop on a cycle-by-cycle basis. When the absolute value of the command exceeds the physical saturation threshold, it introduces a negative feedback penalty factor that is positively correlated with the saturation excess in the arbitration allocation loop, and reduces the allocation weight of the proportional control term in situ. This leaves priority control capacity for the nonlinear adaptive damping solution unit, preventing the servo mechanism from falling into the open-loop dead zone under large-range attitude divergence conditions. In addition, the damping suppression effect generated by the nonlinear adaptive damping solution unit reduces the angular velocity and its differential slope, reduces the transient divergence rate scalar, and determines whether the state response interval operator continues to maintain the locked state.

[0028] 3. The transient transition smoothing filter unit monitors the changes in the rigid body's axial acceleration and establishes the transient spatiotemporal correction anchor point at the instant the axial acceleration suddenly drops to a set threshold. It then adjusts the time constant decay factor based on the deviation between the spatiotemporal correction anchor point and the theoretical exit time. If the exit is earlier, the time constant decay factor is smoothly reduced to extend the high-gain maintenance phase; if the exit is later, the gain decay is accelerated. This avoids aerodynamic overload divergence caused by the strong airflow after exiting the tube, eliminates the dependence of the open-loop time axis scheduling on the absolute consistency of the external physical environment, and allows the parameters in the control loop to adaptively evolve with the actual physical phase transition state. Attached Figure Description

[0029] Figure 1 This is a flowchart of the execution steps of the attitude control method for a catapult-launched reconnaissance and strike integrated loitering drone of the present invention;

[0030] Figure 2 This is a hardware architecture diagram of the attitude control system for a catapult-launched reconnaissance and strike integrated loitering drone according to the present invention.

[0031] The objectives, features, and advantages of this invention will be further explained in conjunction with the embodiments and with reference to the accompanying drawings. Detailed Implementation

[0032] The technical solutions of the embodiments of this application will be clearly described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of this application, not all embodiments. All other embodiments obtained by those skilled in the art based on the embodiments of this application are within the scope of protection of this application.

[0033] An attitude control system for a catapult-launched reconnaissance and strike loitering rovers includes:

[0034] The high-dynamic inertial measurement module collects angular velocity and axial acceleration data.

[0035] The tube-out transient transition smoothing filter module establishes a spatiotemporal correction anchor point and calculates the time constant decay factor when the axial acceleration data suddenly drops to a preset acceleration threshold.

[0036] The control loop arbitration and allocation module combines the time-domain dynamic gain compensation coefficient and the square-term nonlinear damping term to synthesize the final servo control command.

[0037] The state response interval calculation module generates a rigid lock command when the real-time flight speed of the loitering drone and the altitude relative to the launch point are within the preset launch transition interval. This command locks the time-domain dynamic gain compensation coefficient at the preset maximum gain limit value and suspends the attenuation logic. The attenuation rate of the attenuation logic is adjusted by the time constant attenuation factor.

[0038] The dynamic saturation limiting adaptive allocation module generates a negative feedback penalty factor when the absolute value of the final servo control command exceeds the physical saturation threshold. This factor is then input to the control loop arbitration allocation module to reduce the weight of the proportional control term. The nonlinear adaptive damping solution module outputs a squared nonlinear damping term to reduce the transient divergence rate scalar.

[0039] Preferably, the tube exit transient transition smoothing filter module further includes: establishing a spatiotemporal correction anchor point when the axial acceleration data suddenly drops to between 0.5 and 0.8 times the preset gravitational acceleration; calculating the actual time deviation between the spatiotemporal correction anchor point and the preset theoretical tube exit time point; if the actual time deviation is greater than 0, reducing the time constant attenuation factor; if the actual time deviation is less than or equal to 0, increasing the time constant attenuation factor.

[0040] Preferably, the state response interval calculation module further includes: retrieving the real-time flight speed of the loitering drone and its altitude relative to the launch point; when the real-time flight speed is between 14m / s and 35m / s and the altitude relative to the launch point is between 0.5m and 10m, the loitering drone is determined to be in the preset launch transition interval and a rigid lock command is generated.

[0041] Preferably, the maximum gain limit value set in the state response interval solution module is between 3.5 and 5.5, and during the period when the rigid locking command is in effect, the state response interval solution module locks the time-domain dynamic gain compensation coefficient at the maximum gain limit value until the nonlinear adaptive damping solution module reduces the transient divergence rate scalar to below the preset safe divergence threshold.

[0042] Preferably, the dynamic saturation limiting adaptive allocation module further includes: retrieving the final servo control command and comparing the absolute value of the final servo control command with the difference between the final servo control command and the physical saturation threshold, wherein the physical saturation threshold corresponds to 90% of the standard drive voltage amplitude of the servo execution module; generating a negative feedback penalty factor when the absolute value of the final servo control command exceeds the physical saturation threshold; and introducing the negative feedback penalty factor into the proportional control term in the control loop arbitration allocation module to reduce the weight of the proportional control term allocation.

[0043] Preferably, the nonlinear adaptive damping solution module further includes: retrieving the measured angular velocity data of the current cycle, calculating the product of the measured angular velocity data and its own absolute value to generate a squared nonlinear damping term; and inputting the squared nonlinear damping term into the attitude control loop to suppress the difference slope between the amplitude and the angular velocity data, thereby providing a damping overload with high-frequency disturbance limiting cutoff characteristics.

[0044] Preferably, the control loop arbitration allocation module is connected to the dynamic saturation limiting adaptive allocation module to collect and synthesize the final servo control command containing the proportional control term and the square term nonlinear damping term, and to receive the negative feedback penalty factor output by the dynamic saturation limiting adaptive allocation module when the absolute value of the final servo control command exceeds the physical saturation threshold.

[0045] Preferably, the high dynamic inertial measurement module includes a high dynamic angular velocity detection module and an axial acceleration detection module. The angular velocity acquisition range of the high dynamic angular velocity detection module is not less than 400 rad / s, so as to achieve angular velocity range coverage of the initial momentum disturbance caused by catapult impact and the extreme transient asymmetric aerodynamic impact torque generated when the fore and aft wings of the tandem wing are asynchronously deployed.

[0046] Preferably, the system calibrates the physical saturation threshold by adjusting the power supply pulse width modulation ratio of the standard drive voltage of the servo execution module, and introduces a negative feedback penalty factor by adjusting the proportional gain matrix in the control loop arbitration allocation module. Under the hierarchical amplitude control state of the state response interval solution module and the nonlinear adaptive damping solution module, the pitch angle divergence and yaw angle divergence of the loitering aircraft within a second (e.g., 2s to 5s) after leaving the control tower are both less than 2°.

[0047] A method for implementing the attitude control system of a catapult-launched reconnaissance and strike loitering rovers includes the following steps:

[0048] Step S1: Acquire angular velocity and axial acceleration data using a high dynamic inertial measurement module;

[0049] Step S2: When the axial acceleration data suddenly drops to a preset acceleration threshold, the spatiotemporal correction anchor point is established and the time constant decay factor is calculated using the tube exit transient transition smoothing filter module.

[0050] Step S3: The final servo control command is synthesized by combining the time-domain dynamic gain compensation coefficient and the square-term nonlinear damping term with the control loop arbitration allocation module.

[0051] Step S4: When the real-time flight speed of the loitering vehicle and the altitude relative to the launch point are in the preset launch transition range, the state response interval calculation module generates a rigid locking command to lock the time-domain dynamic gain compensation coefficient at the preset maximum gain limit value and suspend the attenuation logic. The attenuation rate of the attenuation logic is adjusted by the time constant attenuation factor.

[0052] Step S5: When the absolute value of the final servo control command exceeds the physical saturation threshold, the dynamic saturation limit adaptive allocation module generates a negative feedback penalty factor and inputs it to the control loop arbitration allocation module to reduce the weight of the proportional control term. The nonlinear adaptive damping solution module outputs a squared nonlinear damping term to reduce the transient divergence rate scalar.

[0053] Example 1: In the tube-launched catapult launch of a tube-launched loitering glider, the rigid body coordinate system experiences divergent motion variables due to turbulent flow at the tube inlet. A high-dynamic inertial measurement module periodically collects angular velocity and axial acceleration data at a sampling frequency of 200Hz to 400Hz. The state response interval calculation module obtains the real-time flight velocity and relative altitude of the loitering glider to the launch point. When the loitering glider is in the physical phase transition stage of cold gas ejection from the tube and the self-locking deployment of the front and rear folding wings, if the real-time flight velocity is within the launch transition zone of 14m / s to 35m / s... Within a stress window of 0.5m to 10m relative height, the state response interval solution module immediately generates a rigid locking command, forcibly locking the time-domain dynamic gain compensation coefficient in the control loop at a preset maximum gain limit value, and rigidly suspending the attenuation program that monotonically decreases with time. Simultaneously, the state response interval solution module continuously reads angular velocity data within two control sampling cycles and calculates its discrete difference slope to calculate the transient divergence rate scalar in each axis of the rigid body in real time. The formula for calculating the difference slope of the angular velocity data is as follows: ,in, The slope of the difference in the angular velocity data of the roll axis. The roll axis angular velocity in the current k-th control cycle. The roll axis angular velocity is the value of the previous (k-1)th control cycle, with a constant control cycle. The corresponding time is between 2.5ms and 5ms.

[0054] During the process of keeping the state response interval solution module locked, the formula for determining its time-domain dynamic gain compensation coefficient is as follows: ,in, For time-domain dynamic gain compensation coefficients, The preset maximum gain limit value. The decay factor is the time constant. As a discrete-time quantity, the dynamic saturation limiting adaptive allocation module retrieves the final servo control command synthesized by the control loop arbitration allocation module cycle by cycle. If the absolute value of the final servo control command exceeds the preset physical saturation threshold due to proportional correction, the dynamic saturation limiting adaptive allocation module introduces a negative feedback penalty factor positively correlated with the saturation exceedance in the control loop arbitration allocation module, thereby reducing the allocation weight of the proportional control term in situ. This allows for priority action capacity for the square-term nonlinear damping term output by the nonlinear adaptive damping solution module, preventing the actuator from falling into the saturation dead zone. The calculation formula for the square-term nonlinear damping term is as follows: ,in, The nonlinear damping term is the square term of the roll axis. The preset basic damping coefficient, To obtain the measured angular velocity data, the control loop arbitration and allocation module combines the time-domain dynamic gain compensation coefficient and the square-term nonlinear damping term to synthesize the final servo control command. The calculation formula for the final servo control command is as follows: ,in, For the final servo control command, As the reference proportional gain, This refers to the attitude angle deviation of the roll axis. This is the proportional control term, i.e., the reference proportional gain. Attitude angle deviation from the roll axis The product term; the control loop arbitration allocation module corrects the reference proportional gain based on the difference between the expected absolute value of the final servo control command and the physical saturation threshold. When the difference is greater than zero, the ratio of the difference to the physical saturation threshold is multiplied by a calibration constant of 0.25 to generate a negative feedback penalty factor with an output upper limit of 0.40. In this closed-loop interaction, in order to eliminate the dimensional structure difference between the single-valued scalar form of the negative feedback penalty factor and the matrix form of the gain, the system establishes a normalization transformation path. By directly mapping the calculated scalar penalty factor as a diagonal operator into a third-order identity matrix, the system derives... An array architecture with equivalent algebraic dimensions is used to bridge subsequent matrix-based gain reduction operations. The baseline gain is reduced proportionally using an attenuation matrix. The nonlinear adaptive damping solution module reserves at least 35% of the drive capacity in its square-term nonlinear damping term output. Specifically, to provide damped overload with high-frequency disturbance limiting cutoff characteristics, the nonlinear adaptive damping solution module monitors the absolute value of the first derivative of the angular velocity data in real time during operation. When this absolute value is less than or equal to a preset angular velocity change rate threshold, the module normally performs the rolling calculation of the square-term nonlinear damping term; once the monitoring... When the structural vibration caused by the high elasticity of the folding wing causes a sudden increase in the absolute value of the first derivative of the angular velocity data and exceeds the threshold of the rate of change of angular velocity, the embedded high-frequency disturbance limiting and truncation operator will be activated. This will hard-freeze and lock the amplitude output of the nonlinear adaptive damping term in the current control cycle, stopping the further multiplication and amplification of the damping term with the fluctuating measured angular velocity data, until the absolute value of the first derivative falls back below the threshold, thus smoothly outputting the control commands to the servo mechanism. When the tube exit transient transition smoothing filter module detects a sudden drop in axial acceleration data to a preset acceleration threshold, the tube exit transient transition smoothing filter module establishes... The spatiotemporal correction anchor point is calculated and the time constant decay factor is calculated. Based on the actual time deviation between the spatiotemporal correction anchor point and the theoretical exit time, the value of the time constant decay factor is smoothly corrected. Thus, an adaptive parameter evolution channel is established at the moment when the loitering vehicle leaves the tube opening constraint, so that the control gain curve is aligned with the physical change node in the time domain. The control commands are smoothly output to the servo mechanism, enabling the loitering vehicle to generate a high-damping priority control response loop under the structural vibration conditions caused by the tube opening area and the unfolding of the folding wings. This achieves attitude recovery with pitch angle divergence and yaw angle divergence of less than 2° at the second level (e.g., 2s to 5s) after exiting the tube.

[0055] Example 2: In a hardware-in-the-loop simulation test, the flight control test system connects a high-dynamic flight turntable to a real-time computing platform. The dynamic control speed change rate threshold of the high-dynamic flight turntable is used as the dynamic control speed characteristic threshold for the speed adaptive response in the system algorithm, which is greater than or equal to 10000° / s. 2The angular velocity measurement resolution is 0.01° / s. Under millisecond-level deterministic task scheduling, the real-time computing platform's overall control response and the semi-physical simulation time window completely cover the second level, thus reproducing the high-dynamic rigid body boundary conditions of the tube-launched loitering pod at the instant of ejection from the launch tube at room temperature; and addressing the constant control period in the loop. The settings were analyzed to determine the natural frequencies of the structural vibration during the folding wing deployment transient and the maximum computational load of the data processing unit. Since the fundamental frequency of the folding wing's main structure is 50Hz, to avoid spectral aliasing in the attitude sampling signal and maintain a phase margin of more than four times that under the Nyquist sampling theorem, the minimum refresh frequency of the control loop must be greater than or equal to 200Hz. Simultaneously, to ensure the data processing unit's full load rate remains below the safety limit of 80%, a constant control period was maintained. With defined boundary constraints, a constant control period is achieved by embedding a time-series hard real-time scheduler into the real-time computing platform. The values ​​are adjusted and locked between 2.5ms and 5ms to provide a refresh rate of 200Hz to 400Hz for attitude algorithm control. To test the anti-saturation divergence performance of the attitude control system in an environment filled with non-ideal dust, turbulence, and uncertain electromechanical disturbances, a power spectral density of 0.05 rad is superimposed on the original input stream of the high dynamic inertial measurement module from the signal generator of the real-time computing platform. 2 / s 2 Gaussian white noise is used to construct a non-ideal semi-physical test condition containing measurement noise. The high-dynamic inertial measurement module collects the roll axis angular velocity data of the rigid body coordinate system under simulated crosswind conditions. The state response interval solution module calculates the differential slope of the angular velocity data based on the angular velocity data within a continuous control cycle. The formula for calculating the differential slope of the angular velocity data is as follows: ,in, The slope of the difference in the angular velocity data of the roll axis. The roll axis angular velocity in the current k-th control cycle. The angular velocity of the roll axis in the previous (k-1)th control cycle. To maintain a constant control cycle, when the real-time flight speed of the loitering rovers and the altitude relative to the launch point are within the preset launch transition range, the state response interval calculation module generates a rigid lock command, forcibly locking the time-domain dynamic gain compensation coefficient in the control loop at the preset maximum gain limit value, and suspending a monotonically decreasing attenuation program. The formula for calculating the time-domain dynamic gain compensation coefficient is as follows: ,in, For time-domain dynamic gain compensation coefficients, The preset maximum gain limit value. The decay factor is the time constant. It is a discrete-time quantity.

[0056] The control loop arbitration allocation module combines the time-domain dynamic gain compensation coefficient with the squared nonlinear damping term calculated by the nonlinear adaptive damping solution module to synthesize the final servo control command. When the strong proportional correction causes the absolute value of the final servo control command to exceed the preset physical saturation threshold, the dynamic saturation limiting adaptive allocation module introduces a negative feedback penalty factor in the control loop arbitration allocation module to reduce the allocation weight of the proportional control term, reserving priority action capacity for the squared nonlinear damping term. The calculation formulas for the squared nonlinear damping term and the final servo control command are as follows: , ,in, The nonlinear damping term is the square term of the roll axis. The preset basic damping coefficient, These are measured angular velocity data. For the final servo control command, As the reference proportional gain, This refers to the attitude angle deviation of the roll axis. This is the proportional control term, i.e., the reference proportional gain. Attitude angle deviation from the roll axis The product term; to verify the complete support of this technical solution for the boundaries defined by this invention, the sample group of this invention using the complete attitude control system of this invention, the partially missing control group with the dynamic saturation limiting adaptive allocation module removed, and the deviation from the maximum gain limiting value were run in the above-mentioned semi-physical simulation environment respectively. An out-of-range control group was established, and crosswind speeds of 15 m / s, 20 m / s, and 25 m / s were set to construct a stepped-enhanced problem variable interference. Under a crosswind of 15 m / s, the high-dynamic inertial measurement module outputs the original roll axis angular velocity including background noise, and the initial value of the differential slope of the calculated angular velocity data reaches 1250 rad / s². During the initial stage of launch transition, the spacecraft experiences a strong interference window of acceleration from rest to launch and the popping of folded wings. During the extremely short period of time during which its height relative to the launch point rises from 0.5 m to 10 m, its real-time flight speed naturally falls within the launch transition speed range of 14 m / s to 35 m / s. The state response interval solution module accurately captures this kinematic characteristic. The system triggers a rigid locking command immediately, enabling the high-gain register to complete hardware latching. This ensures that the system remains in the locked maximum gain state even when the differential slope continues to rise to 1250 rad / s² or higher step stress conditions, thus guaranteeing the adaptive extension of the correction overload boundary. Under this condition, the control loop of some missing control groups, lacking a dynamic saturation distribution mechanism, continuously triggers the 9.0V saturation voltage dead zone in its synthesized control command. This causes the square term nonlinear damping term to be limited and squeezed, failing to suppress the transient asymmetric flow field torque at the outlet pipe. This expands the maximum attitude angle deviation of its roll axis to 8.4°, and delays the angular velocity convergence time to 0.142 seconds. In contrast, the sample of this invention, at the maximum gain limit value... When the absolute lower limit is set to 3.5, the state response interval calculation module locks the time-domain dynamic gain, and the peak value of the synthesized final servo control command is 7.2V, which is within the physical saturation threshold. Its final damping correction response limits the maximum roll attitude angle deviation to 1.8°, and completes angle convergence within 0.045 seconds after exiting the tube; the maximum gain limit value is set to... When adjusted to the normal median value of 4.5, the sample group of the present invention outputs proportional overload under the same divergent stress, and the measured maximum roll attitude angle deviation is reduced to 1.1°.

[0057] When the crosswind speed is increased to the limiting perturbation gradient of 25 m / s and the gain is greatly limited... When the absolute upper limit is set to 5.5, the differential slope of the roll axis angular velocity data increases to 4200 rad / s² (this differential slope serves as a characterization quantity, corresponding to the first velocity change rate threshold preset by the trigger system). At this time, the initial command amplitude calculated by the control loop arbitration allocation module is expected to reach 11.4V. The dynamic saturation limiting adaptive allocation module detects that this amplitude exceeds the limit in real time. By applying a negative feedback penalty factor of 0.26 in the arbitration loop, the weight of the proportional control term is reduced in place, thereby reserving 35% of the output drive capacity for the nonlinear adaptive damping solution module. The nonlinear adaptive damping solution module generates a 3.1V square term nonlinear damping term based on the current measured angular velocity data of 45 rad / s, which smooths the pitch angle divergence and yaw angle divergence within the second-level window of the tube exit to a stable state of 1.4° and 1.6°, respectively. As a reverse empirical test of the boundary control effectiveness, when the out-of-range control group greatly increases the gain limiting value... When set to the lower limit of 3.0, due to insufficient proportional overload expansion boundary, the system control quantity cannot counteract the 25m / s wind drag moment, causing the pitch angle deviation to diverge to 12.3° and the angular displacement to exceed the tolerance. However, when the maximum gain limit is set... When the gain is set to the upper limit of 6.0, the excessively high gain configuration causes the negative feedback penalty factor to jump at a high frequency of 65Hz between 0.0 and 0.8, inducing the servo motor to vibrate the transmission mechanism, resulting in a final yaw angle divergence of 4.8°. The measured parameter range of 3.5 to 5.5 blocks the attitude under-control divergence at the lower limit and converges the nonlinear self-resonance at the upper limit, forming a control gain working window with specific boundary stability characteristics. The spatial kinematic interval characteristics constructed by the real-time flight speed and relative altitude of the loitering glider during the launch transition phase are directly mapped to the adaptive phase transition node of the control system. With the boundary limit of the final servo control command voltage amplitude by the dynamic saturation limit adaptive allocation module, the damping term and proportional term allocation weight of the servo mechanism are adjusted without relying on external timing open-loop scheduling. This suppresses the multidimensional nonlinear divergence of attitude variables caused by the asymmetric flow field disturbance at the moment of catapult launch, and realizes the dynamic tracking and convergence of the attitude angle of the loitering glider after it is freed from the constraints of the tubular tube.

[0058] Example 3: In the tube-launched loitering pod's tube-type catapult launch, the rigid body coordinate system experiences spatial non-electrical motion state variables diverging due to turbulent flow at the tube inlet. The system's high-dynamic inertial measurement module collects triaxial angular velocity and axial acceleration data at a sampling frequency of 200Hz to 400Hz. At this time, the transmission and actuation mechanisms inside the loitering pod generate dynamic impedance fluctuations and zero-bias temperature drift, causing the time-domain gain switching critical point to misalign under fixed-time constraints. This leads to divergent instability of the attitude control loop under a wide range of uncertain external stress impacts. To address the aforementioned parameter control deviations, the flight control system establishes initial state definition conditions, clarifying that the core object of the high-dynamic inertial measurement module is the rigid body loitering pod with a folding wing structure, whose initial angular velocity components at the tube exit transient are in the range of 0 rad / s to 5 rad / s. The calculation... The processing chip provides a 120MHz operating frequency, and the data temporary storage register allocates 32-bit floating-point storage space as a functional specification for the enabling environment. The state response interval solution module executes a quantization program during operation, directly receiving the roll axis angular velocity, pitch axis angular velocity, and yaw axis angular velocity collected by the high-dynamic inertial measurement module. It continuously reads the three-axis angular velocity data within the current control cycle and calculates their discrete difference slope. To demodulate the three-axis attitude angular velocity divergence caused by multidimensional non-equilibrium energy dissipation into a single-valued scalar form, the state response interval solution module converts the angular velocity change rates along the roll axis, pitch axis, and yaw axis into single-valued quantities through square root operations, establishing a transient divergence rate envelope index that reflects the severity of the total motion divergence in rigid body space. The mathematical calculation formula for the transient divergence rate envelope index is as follows: ,in, The transient divergence rate envelope index, The slope of the difference in the angular velocity data of the roll axis. The slope of the differential angular velocity data for the pitch axis. The differential slope of the angular velocity data for the yaw axis.

[0059] In the specific numerical parameter adjustment operation, when the differential slope of the roll axis angular velocity data measured by the high dynamic inertial measurement module... The differential slope of the pitch axis angular velocity data at 1200 rad / s² Differential slope for yaw axis angular velocity data of 800 rad / s² When the value is 600 rad / s², the transient divergence rate envelope index calculation unit in the state response interval solution module performs the above square root operation, and outputs the transient divergence rate envelope index. The state response interval calculation module presets a suppression critical threshold of 1500 rad / s², which is derived from offline statistical data of multi-attitude coupled ground catapult taxiing tests. The module employs a multi-dimensional composite decision boundary, setting the flight speed range of 14 m / s to 35 m / s and the relative altitude range of 0.5 m to 10 m as the kinematic trigger thresholds for the launch transition phase. The transient divergence rate envelope index, calculated based on the spatial three-axis angular velocity discrete differential trajectory, serves as the intensity index for evaluating the energy of multi-axis motion cross-coupling. When the system detects that the real-time flight speed and relative altitude have entered the launch transition interval threshold, it spontaneously generates a rigid locking command, thereby precisely covering the abnormal movement window caused by the asynchronous ejection of tandem wings and the asymmetric flow field at the source of spatial kinematics. This addresses the parameter mutations during the transition from constrained taxiing within the tube to the strong impact phase of the asymmetric flow field at the tube opening. During this process, the preset safe divergence threshold is specifically calibrated as follows: This serves as a safety convergence benchmark for determining whether the multi-axis attitude divergence kinetic energy has been sufficiently dissipated and allows for the release of rigid lock; simultaneously, the velocity change rate threshold for damping adaptive activation in the nonlinear adaptive damping solution module, which is mapped to the dynamic velocity change rate threshold of the damped response in the control chain, is set as follows. It is specifically designed to decouple and isolate the structural flutter noise introduced during the second-level tube ejection transition period caused by the high elasticity of the folding wing from the large overload of the macroscopic rigid body attitude divergence of the device in bandwidth. When the system detects that the real-time flight speed and relative altitude are entering the launch transition interval threshold, causing the rigid lock command to take effect, the time-domain dynamic gain compensation coefficient in the control loop is forcibly locked at the preset maximum gain limit value, and the transient divergence rate envelope index is calculated by the state response interval solution module. The system continuously evaluates the dissipation state of the cross-coupling energy of multi-axis motions and simultaneously calculates a second-order dissipation compensation operator to offset transmission losses. This second-order dissipation compensation operator is output to the control loop arbitration and allocation module to correct the control input to the servo mechanism. The second-order dissipation compensation operator, run by the state response interval solution module, includes a first-order stiffness dissipation component and a second-order damping dissipation component. The first-order stiffness dissipation component is generated by multiplying the differential slope of the angular velocity data by a first dissipation constant, and the second-order damping dissipation component is generated by multiplying the time derivative of the differential slope by a second dissipation constant. The two components are then added together and output to the control loop arbitration and allocation module to offset the back electromotive force loss of the actuator. The first and second dissipation constants are determined based on test results of a loitering pod fixed to a high-dynamic three-axis turntable and inputting a 50Hz sinusoidal sweep frequency signal. The second dissipation constant is incrementally adjusted... A dissipation constant reduces the transmission hysteresis phase shift of the turntable feedback to within 3°. Adjusting the second dissipation constant maintains the feedback current overshoot peak of the servo amplifier below 85% of the rated protection current. Specifically, the first-order stiffness dissipation component is obtained by multiplying the differential slope of the angular velocity data of the roll axis, pitch axis, or yaw axis by the first dissipation constant. The second-order damping dissipation component is obtained by multiplying the time derivative of the differential slope by the second dissipation constant. The two are algebraically added to construct the second-order dissipation compensation operator terms for each corresponding axis. After receiving this operator term, the control loop arbitration and allocation module directly adds it in-situ as an independent linear weighted superposition term to the final servo control command, which includes the time-domain dynamic gain compensation coefficient, reference proportional gain, attitude angle deviation, and square-term nonlinear damping term. The formula for calculating the corrected final servo control command is as follows: ,in, This is the proportional control term, i.e., the reference proportional gain. Attitude angle deviation from the roll axis The product term, Let be the first dissipation constant. Let be the second dissipation constant. To correspond to the time derivative of the axial differential slope, the corrected final servo control command is superimposed with a dynamic level compensation for the reverse electromotive force of the actuating mechanism and the transmission hysteresis, on the basis of the original proportional correction voltage and the square term nonlinear damping voltage. This completes the algebraic correction of the amplitude of the drive voltage output to the high-speed servo amplifier. Since the final servo control command contains a reverse correction component composed of a second-order dissipation compensation operator, the physical time delay lag and energy transmission loss of the high-speed servo amplifier are canceled in situ. The total control energy gain of the control loop is spontaneously fixed at the corresponding physical position with the sudden change of non-electric variables, blocking the saturation instability state caused by the cross-coupling of multi-dimensional actions. The loitering glider achieves smooth convergence of the spatial three-axis attitude within the second time of exiting the tube under the condition of large sideslip at the tube opening, eliminating the spatiotemporal divergence of non-electric motion state parameters.

[0060] Example 4: When the system faces the pre-launch condition of installing the attitude control system on different batches of loitering pod shell structures, the high dynamic inertial measurement module is in a static constraint state before ejection. To eliminate the divergence of spatial non-electric motion state variables caused by the non-orthogonal residuals of the installation axis and the spatiotemporal drift of the sensor zero bias, the data processing unit calls the zero-point calibration program to continuously read the raw angular velocity data for 1024 control cycles in a static environment. The average value of the three-axis static zero bias is obtained by averaging and written into the configuration table in the memory. Before the attitude calculation begins, the static zero bias average value in the configuration table is used to subtract and offset the raw input stream, thereby blocking the zero bias drift before the attitude control and launch state response calculation. The calculation formula of the calibrated measured angular velocity data is as follows: ,in, The measured angular velocity data after calibration. This is the raw angular velocity data. The average value is the three-axis static zero bias. When the system is in the ground static simulation verification condition, the external test bench sends a preset simulation signal stream to the data processing unit through the bus interface to test the matching response of the state response interval calculation module to the preset launch transition interval. The data processing unit receives the simulation signal stream, calculates the simulated flight speed and simulated relative altitude under constant control period timing constraints, and continuously compares the simulated flight speed and simulated relative altitude with the launch transition interval threshold in the memory. When it is determined that the simulated flight speed at 5 consecutive sampling points is within 14m / sqm, the system will determine the target value. When the speed is between s and 35m / s and the simulated relative height is between 0.5m and 10m, the high-order gain register is automatically latched, and the time-domain dynamic gain compensation coefficient is locked at the preset maximum gain limit value. The control gain of the control loop is smoothly adjusted so that the feedback topology is tightened on the preset steady-state baseline before the aeroelastic stress wave peak caused by the full opening of the front and rear folding wings is fully released. This allows the control system to have a definite initial state when facing the large sideslip turbulence outside the tube and the strong impact of complex asymmetric aerodynamic flow, and drives the servo motor output shaft to deflect to the zero-position alignment state.

[0061] Example 5: When the system faces ground assembly and verification conditions where there are differences in dynamic response delays among ESCs from different production batches, the digital flight control calibration platform connects the ESCs to the real-time digital simulation main control unit. By inputting sinusoidal position sweep excitation signals with frequencies ranging from 1Hz to 100Hz to the ESCs, a high-precision position sensor synchronously acquires the actual deflection angle sequence of the servo's output shaft at a sampling frequency of 2000Hz. The real-time digital simulation main control unit compares the time-domain phase difference between the input sinusoidal position sweep excitation signal and the actual deflection angle sequence, extracting the rigid body physical response delay scalar of the current ESC at a cutoff frequency of 50Hz. This scalar is used to characterize the physical property residuals at the actuator component level. To establish a deterministic channel for the parameter evolution trajectory in the time-domain dynamic gain compensation coefficient, the calibration calculation unit calls the parameter matrix offline calibration program, using the measured rigid body physical response delay scalar as the independent variable input. The time constant attenuation factor required by the flight control loop is dynamically calculated using a reciprocal proportional relationship. The mathematical calculation formula is as follows: ,in, The decay factor is the time constant. For rigid body physical response time delay scalar, The preset calibration ratio is 0.05s.

[0062] In the specific quantization conversion process, when the high-precision position sensor measures the rigid body physical response delay scalar of the current batch of ESC servo motors... When the time is 0.0125 seconds, the time constant decay factor is output by running the above reciprocal proportional transformation through the calibration calculation unit. The corresponding value is 4.0, while in another test sample group, when the high-latency batch of servos was replaced and its rigid body physical response delay scalar was measured... When increased to 0.0156 seconds, the calculated time constant decay factor... The corresponding adaptive adjustment is set to 3.21. At this point, the central processing unit writes the calculated time constant attenuation factor into the designated address field of the non-volatile register. At the instant the loitering pod receives the ejection trigger signal, the ejection transient smoothing filter module directly reads the corresponding value from the non-volatile register and activates the time-domain dynamic gain compensation coefficient. The adaptive evolution attenuation logic enables the dissipation polarity of the control gain curve to form an end-to-end causal closed loop with the actual hardware response impedance of the actuator, overcoming the time axis misalignment problem caused by the production batch residual of the actuator, and realizing the multi-axis attitude angular velocity divergence convergence of each working condition sample group within a second-level spatial window that matches the preset launch transition interval and is free from the constraints of the launch tube pipeline.

[0063] It will be apparent to those skilled in the art that the present invention is not limited to the details of the exemplary embodiments described above, and that the present invention can be implemented in other specific forms without departing from the spirit or essential characteristics of the present invention.

[0064] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and are not intended to limit it. Although the present invention has been described in detail with reference to preferred embodiments, those skilled in the art should understand that modifications or equivalent substitutions can be made to the technical solutions of the present invention without departing from the spirit and scope of the technical solutions of the present invention.

Claims

1. An attitude control system for a catapult-launched reconnaissance and strike integrated loitering rovers, characterized in that, include: The high-dynamic inertial measurement module collects angular velocity and axial acceleration data. The tube-out transient transition smoothing filter module establishes a spatiotemporal correction anchor point and calculates the time constant decay factor when the axial acceleration data suddenly drops to a preset acceleration threshold. The control loop arbitration and allocation module combines the time-domain dynamic gain compensation coefficient and the square-term nonlinear damping term to synthesize the final servo control command. The state response interval calculation module generates a rigid lock command when the real-time flight speed of the loitering drone and the altitude relative to the launch point are within the preset launch transition interval. This command locks the time-domain dynamic gain compensation coefficient at the preset maximum gain limit value and suspends the attenuation logic. The attenuation rate of the attenuation logic is adjusted by the time constant attenuation factor. The dynamic saturation limiting adaptive allocation module generates a negative feedback penalty factor when the absolute value of the final servo control command exceeds the physical saturation threshold. This factor is then input to the control loop arbitration allocation module to reduce the weight of the proportional control term. The nonlinear adaptive damping solution module outputs a squared nonlinear damping term to reduce the transient divergence rate scalar.

2. The attitude control system for a catapult-launched reconnaissance and strike integrated loitering pod according to claim 1, characterized in that, The tube exit transient transition smoothing filter module also includes: when the axial acceleration data suddenly drops to between 0.5 and 0.8 times the preset gravitational acceleration, establishing a spatiotemporal correction anchor point; calculating the actual time deviation between the spatiotemporal correction anchor point and the preset theoretical tube exit time point; if the actual time deviation is greater than 0, reducing the time constant attenuation factor; if the actual time deviation is less than or equal to 0, increasing the time constant attenuation factor.

3. The attitude control system for a catapult-launched reconnaissance and strike integrated loitering pod according to claim 1, characterized in that, The state response interval calculation module also includes: retrieving the real-time flight speed of the loitering drone and its altitude relative to the launch point; when the real-time flight speed is between 14m / s and 35m / s and the altitude relative to the launch point is between 0.5m and 10m, the loitering drone is determined to be in the preset launch transition interval and a rigid lock command is generated.

4. The attitude control system for a catapult-launched reconnaissance and strike integrated loitering pod according to claim 1, characterized in that, The maximum gain limit value set in the state response interval solution module is between 3.5 and 5.

5. During the period when the rigid lock command is in effect, the state response interval solution module locks the time-domain dynamic gain compensation coefficient at the maximum gain limit value until the nonlinear adaptive damping solution module reduces the transient divergence rate scalar below the preset safe divergence threshold.

5. The attitude control system for a catapult-launched reconnaissance and strike integrated loitering pod according to claim 1, characterized in that, The dynamic saturation limiting adaptive allocation module also includes: retrieving the final servo control command and comparing the absolute value of the final servo control command with the difference between the physical saturation threshold, where the physical saturation threshold corresponds to 90% of the standard drive voltage amplitude of the servo execution module; generating a negative feedback penalty factor when the absolute value of the final servo control command exceeds the physical saturation threshold; and introducing the negative feedback penalty factor into the proportional control term in the control loop arbitration allocation module to reduce the weight of the proportional control term allocation.

6. The attitude control system for a catapult-launched reconnaissance and strike integrated loitering pod according to claim 1, characterized in that, The nonlinear adaptive damping solution module also includes: retrieving the measured angular velocity data of the current cycle, calculating the product of the measured angular velocity data and its own absolute value to generate a square-term nonlinear damping term; inputting the square-term nonlinear damping term into the attitude control loop to suppress the difference slope between the amplitude and the angular velocity data, and providing a damping overload with high-frequency disturbance limiting cutoff characteristics.

7. The attitude control system for a catapult-launched reconnaissance and strike integrated loitering pod according to claim 1, characterized in that, The control loop arbitration allocation module is connected to the dynamic saturation limiting adaptive allocation module. It is used to collect and synthesize the final servo control command, which includes a proportional control term and a square nonlinear damping term. When the absolute value of the final servo control command exceeds the physical saturation threshold, it receives the negative feedback penalty factor output by the dynamic saturation limiting adaptive allocation module.

8. The attitude control system for a catapult-launched reconnaissance and strike integrated loitering pod according to claim 1, characterized in that, The high dynamic inertial measurement module includes a high dynamic angular velocity detection module and an axial acceleration detection module. The angular velocity acquisition range of the high dynamic angular velocity detection module is not less than 400 rad / s, which is used to achieve angular velocity range coverage for the initial momentum disturbance caused by catapult impact and the extreme transient asymmetric aerodynamic impact torque generated when the fore and aft wings of the tandem wing are asynchronously deployed.

9. The attitude control system for a catapult-launched reconnaissance and strike integrated loitering pod according to claim 1, characterized in that, The system calibrates the physical saturation threshold by adjusting the power supply pulse width modulation ratio of the standard drive voltage of the servo execution module, and introduces a negative feedback penalty factor by adjusting the proportional gain matrix in the control loop arbitration allocation module. Under the hierarchical amplitude control state of the state response interval solution module and the nonlinear adaptive damping solution module, the pitch angle divergence and yaw angle divergence of the loitering aircraft are both less than 2° within 2s to 5s after leaving the control tube.

10. A method for attitude control of a catapult-launched reconnaissance and strike loitering munition, used to operate the attitude control system of the catapult-launched reconnaissance and strike loitering munition as described in claim 1, characterized in that, Includes the following steps: Step S1: Acquire angular velocity and axial acceleration data using a high dynamic inertial measurement module; Step S2: When the axial acceleration data suddenly drops to a preset acceleration threshold, the spatiotemporal correction anchor point is established and the time constant decay factor is calculated using the tube exit transient transition smoothing filter module. Step S3: The final servo control command is synthesized by combining the time-domain dynamic gain compensation coefficient and the square-term nonlinear damping term with the control loop arbitration allocation module. Step S4: When the real-time flight speed of the loitering vehicle and the altitude relative to the launch point are in the preset launch transition range, the state response interval calculation module generates a rigid locking command to lock the time-domain dynamic gain compensation coefficient at the preset maximum gain limit value and suspend the attenuation logic. The attenuation rate of the attenuation logic is adjusted by the time constant attenuation factor. Step S5: When the absolute value of the final servo control command exceeds the physical saturation threshold, the dynamic saturation limit adaptive allocation module generates a negative feedback penalty factor and inputs it to the control loop arbitration allocation module to reduce the weight of the proportional control term. The nonlinear adaptive damping solution module outputs a squared nonlinear damping term to reduce the transient divergence rate scalar.

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