Embedded conformal aerodynamic guidance system and control method for a low-spin aircraft

Abstract

This invention discloses an embedded conformal aerodynamic guidance system and control method for a low-spinning aircraft, belonging to the field of intelligent deformable structure design for aircraft. It includes: a navigation and attitude sensing module for acquiring navigation, attitude, and airflow attitude information of the aircraft body; a flight state calculation module for calculating the real-time flight Mach number based on navigation information; and a guidance correction decision module for determining whether safety conditions are met based on the real-time flight Mach number and whether phase synchronization conditions are met based on the roll phase angle and ideal aerodynamic torque direction angle in the attitude information. If both safety and phase synchronization conditions are met, an activation command is generated. A conformal aerodynamic actuator is used to generate local deformation in response to the activation command. The activation timing of the conformal aerodynamic actuator is precisely correlated with the roll phase of the aircraft body, activating and inflating at a specific phase to accurately correct trajectory deviations and ensure flight stability.

Description

Technical Field

[0001] This invention relates to the field of intelligent deformable structure design for aircraft, and more particularly to an embedded conformal aerodynamic guidance system and control method for a low-rotation aircraft. Background Technology

[0002] Currently, Precision Guidance Kits (PGKs) have broad prospects in the civil aerospace and aviation fields. For example, they are crucial for the precise recovery and trajectory control of reusable small spacecraft reentry capsules or high-speed logistics vehicles, playing a vital role in supporting the low-altitude economy. They can also be extended to many related fields such as precision firefighting and artificial rainmaking. The core technical feature of the aforementioned PGK system is the use of fixed canards or small folding rudders as aerodynamic actuators. Its working principle heavily relies on the high-speed spin of the aircraft itself (typically >5 rps). Under high-speed rotation, the gyroscopic effect averages the lateral force generated by the deflection of the control surfaces, which changes periodically and rapidly in the body coordinate system, into a directionally stable control force in inertial space, thereby effectively correcting the aircraft's flight trajectory in space.

[0003] However, when attempting to directly transplant such mature PGK systems to tail-stabilized low-speed rotating aircraft (such as fire extinguishing bombs, or simply bombs / projectiles), serious compatibility problems arose, leading to abnormal flight trajectories, nutation divergence, and even mid-flight crashes. Through in-depth theoretical analysis and experimental verification, the fundamental reason lies in the essential differences in the dynamic characteristics of the two types of flight platforms, specifically manifested in the following three core contradictions:

[0004] 1) Control mechanism failure due to roll characteristic mismatch

[0005] Tail-stabilized aircraft rely on a slight tilt angle of the tail fin to induce low-speed roll, achieving self-stability during flight. Their typical roll angular velocity is only 0.5–3 rps. Under these low-spin conditions, traditional fixed canards cannot generate an effective gyro-averaging effect. Instead, they continuously apply a lateral force whose direction changes periodically with the aircraft's roll. The frequency of this force's change is close to the aircraft's natural nutation frequency, easily generating strong conical motion.

[0006] 2) The abrupt amplification of aerodynamic disturbances during high-speed flight

[0007] These high-speed aircraft fly at extremely high speeds, with maximum Mach numbers reaching approximately Ma 2.0. Under these high-speed conditions, even the slightest abrupt change in the nose's shape can cause severe shock wave interference and pressure drag. Test data shows that, for example, installing a fixed rudder PGK on a standard test shape can increase the nose drag coefficient by more than 18%, directly resulting in a loss of over 20% in range or intended landing point accuracy. This will have an unacceptable negative impact on the aircraft's range, reentry corridor accuracy, and final landing point dispersion, severely reducing the success rate and reliability of high-precision recovery missions for such aircraft.

[0008] 3) Deterioration of static stability after adding components

[0009] The static stability of a primitive aircraft is achieved through careful design of the relative positions of its center of mass and center of pressure. Typically, the center of mass is required to be ahead of the center of pressure, maintaining a static stability margin of 0.5–1.0 times the characteristic diameter. However, after adding the PGK assembly to the nose, without systematic mass-aerodynamic optimization, the mass distribution and aerodynamic shape of the aircraft will inevitably change. This usually leads to a forward shift of the center of pressure or a rearward shift of the center of mass, severely compressing the static stability margin, and even causing the aircraft to become statically unstable under certain flight conditions, resulting in a complete loss of flight stability.

[0010] To address the aforementioned issues, academia and industry have proposed an improved approach using conformal aerodynamic control surfaces. This involves embedding deformable structures into the aircraft surface, which are flush with the overall shape under normal conditions and bulge locally to generate control force during corrections. For example, CN116659315A describes a "current-temperature-controlled flexible composite material deformable actuating control surface" that uses shape memory alloy (SMA) to drive skin deformation. In principle, this approach can eliminate exposed structures and reduce drag. However, after in-depth analysis, existing conformal control technologies, represented by CN116659315A, have the following fundamental limitations when applied to the specific scenario of low-spinning aircraft:

[0011] 1) Lack of platform adaptability: Existing conformal control technologies are designed for general-purpose aircraft platforms and do not specifically consider the low-spin, high dynamic pressure, and intense thermal environment of tail-stabilized aircraft. In addition, the control logic and other components in existing conformal control technologies are not optimized for these harsh operating conditions.

[0012] 2) Timing mismatch issue: Existing conformal control technology has not resolved the contradiction between the response lag of the flexible actuator (typical thermal cycling response time of SMA > 200ms) and the body's roll dynamics. For a body with a roll speed of 3 rps, it has already rotated more than 216° within 200ms. If the bulge is activated at any phase, the body has already deviated significantly from its initial position, causing the direction of the generated control force to deviate severely from the required correction direction, and may even produce a negative correction effect, exacerbating the flight trajectory deviation.

[0013] 3) Lack of a safe activation mechanism: Existing conformal control technologies lack a safe activation threshold linked to flight conditions (such as Mach number). Activation at high speeds (such as Ma>1.8) may result in the bulging structure being suppressed by the strong airflow and unable to deform effectively, or severe aerodynamic heating may seriously interfere with the performance consistency of actuators such as SMA, leading to control failure. Summary of the Invention

[0014] The purpose of this invention is to overcome the problems of the prior art and provide an embedded conformal aerodynamic guidance system and control method for low-rotation aircraft.

[0015] The objective of this invention is achieved through the following technical solution: an embedded conformal aerodynamic guidance system for a low-rotation aircraft, comprising:

[0016] The navigation and attitude sensing module is used to acquire motion state information of the aircraft body; the motion state information includes navigation information, attitude information, and airflow attitude information.

[0017] The flight status calculation module, connected to the navigation and attitude sensing module, is used to calculate the real-time flight Mach number based on navigation information;

[0018] The guidance correction decision module, connected to the navigation and attitude sensing module and the flight state calculation module, is used to receive motion state information and real-time flight Mach number, and determine whether the safety conditions are met based on the real-time flight Mach number; calculate the flight trajectory deviation based on navigation information and the preset target flight trajectory; calculate the ideal aerodynamic moment direction angle based on the pitch angle, yaw angle and airflow attitude information in the attitude information; and determine whether the phase synchronization condition is met based on the roll phase angle and ideal aerodynamic moment direction angle in the attitude information. If both the safety condition and the phase synchronization condition are met, an activation command is generated.

[0019] The conformal aerodynamic actuator is located at the nose of the aircraft body. Its outer surface conforms to the aerodynamic shape of the nose of the aircraft body. It is connected to the guidance correction decision module and is used to generate local deformation in response to activation commands.

[0020] The energy management module provides power to the navigation and attitude sensing module, flight status calculation module, guidance correction decision module, and conformal aerodynamic actuator.

[0021] In one example, the navigation information includes three-dimensional position, three-dimensional velocity vector and angular velocity; attitude information includes roll phase angle, pitch angle and yaw angle; airflow attitude information includes angle of attack and sideslip angle.

[0022] In one example, the outer surface of the conformal pneumatic actuator is divided circumferentially into at least four independently controllable action quadrants.

[0023] In one example, the conformal pneumatic actuator is driven by a shape memory alloy driver or a piezoelectric ceramic driver, thereby producing localized bulging deformation.

[0024] It should be further noted that the technical features corresponding to the above system examples can be combined or replaced to form new technical solutions.

[0025] This invention also includes an embedded conformal aerodynamic guidance and control method for a low-spinning aircraft, the method comprising the following steps:

[0026] Acquire motion state information of the aircraft body; the motion state information includes navigation information, attitude information, and airflow attitude information;

[0027] Calculate the real-time flight Mach number based on navigation information;

[0028] The system determines whether the safety conditions are met based on the real-time flight Mach number; calculates the flight trajectory deviation based on the navigation information and the preset target flight trajectory; calculates the ideal aerodynamic moment direction angle based on the pitch angle, yaw angle and airflow attitude information in the attitude information; and determines whether the phase synchronization conditions are met based on the roll phase angle and ideal aerodynamic moment direction angle in the attitude information. If both the safety conditions and the phase synchronization conditions are met, an activation command is generated.

[0029] In response to the activation command, the conformal aerodynamic actuator at the nose of the aircraft body undergoes local deformation.

[0030] In one example, determining whether the safety conditions are met based on the real-time flight Mach number includes:

[0031] Determine whether the real-time flight Mach number is less than the preset Mach number safety threshold. If it is less, the safety condition is met; otherwise, the safety condition is not met.

[0032] In one example, the Mach number safety threshold ranges from 1.5 to 2.0.

[0033] In one example, determining whether the phase synchronization condition is met based on the roll phase angle and ideal aerodynamic moment direction angle in the attitude information includes:

[0034] Calculate the absolute difference between the current roll phase angle and the ideal aerodynamic moment direction angle;

[0035] Determine whether the absolute difference is less than the preset synchronization angle threshold. If it is less, the phase synchronization condition is met; otherwise, the phase synchronization condition is not met.

[0036] In one example, the synchronization angle threshold ranges from 10 degrees to 30 degrees.

[0037] In one example, before or after determining whether the safety conditions and phase synchronization conditions are met, it is determined whether the number of corrections performed in this flight mission is less than the preset maximum allowed number. When the safety conditions, phase synchronization conditions, and correction number conditions are all met, an activation command is generated.

[0038] It should be further noted that the technical features corresponding to the above examples can be combined or replaced to form new technical solutions.

[0039] Compared with the prior art, the beneficial effects of the present invention are:

[0040] 1. Based on the real-time flight Mach number, it determines whether the safety conditions are met, realizing the activation of the safety threshold linked to the flight status to avoid control failure; based on the roll phase angle and flight trajectory deviation, it determines whether the phase synchronization conditions are met, precisely linking the activation timing of the conformal aerodynamic actuator with the roll phase of the missile body, activating the bulge at a specific phase to accurately correct the trajectory deviation. This overcomes the problems of traditional fixed control surfaces failing to generate effective average control force under low-spin conditions, instead triggering conical motion and causing nutation divergence, and low reliability at high speeds due to lack of environmental adaptation, thus ensuring flight stability.

[0041] 2. Thanks to its conformal integrated design, the conformal aerodynamic actuator is completely integrated with the projectile's shape under normal conditions, completely eliminating shock wave interference and pressure drag caused by traditional control surfaces. Compared to the serious consequence of adding fixed control surfaces, which results in a range loss of over 20%, this invention adds almost no additional drag, maximizing the success rate and reliability of high-precision recovery missions for aircraft.

[0042] 3. By determining the safe activation threshold based on the real-time flight Mach number, the system proactively avoids adverse operating conditions such as high dynamic pressure and strong aerodynamic heat, preventing actuator failure or performance interference. Simultaneously, by limiting the maximum number of activations per flight, the system effectively avoids the risks of energy depletion and actuator thermal accumulation failure, significantly improving its robustness and reliability in complex flight environments.

[0043] 4. The entire system is integrated into the aircraft's control cabin in a modular form, enabling plug-and-play upgrades without any radial changes and with strong compatibility with existing platforms. At the same time, the modular design of the system reduces system complexity, and the failure of a single module will not affect the overall aerodynamic shape, thus possessing high fault tolerance.

[0044] 5. The system of this invention is not an improvement on a single component, but a complete closed-loop system solution integrating perception, decision-making, execution, and energy. It successfully integrates conformal aerodynamic control, embedded guidance, and intelligent decision-making technologies, providing a low-cost, high-efficiency precision guidance upgrade path for a large number of existing low-spin tail stabilized aircraft. Attached Figure Description

[0045] The specific embodiments of the present invention will be further described in detail below with reference to the accompanying drawings. The accompanying drawings are provided to provide a further understanding of the present application and constitute a part of the present application. The same reference numerals are used in these drawings to denote the same or similar parts. The illustrative embodiments of the present application and their descriptions are used to explain the present application and do not constitute an improper limitation of the present application.

[0046] Figure 1 A system block diagram provided as an example of the present invention;

[0047] Figure 2 This is a schematic diagram of the quadrant distribution of a flexible skin actuator provided as an example of the present invention;

[0048] Figure 3 This is a schematic diagram illustrating the activation of a flexible skin actuator bulge, as provided in an example of the present invention.

[0049] Figure 4 This is a control flowchart provided as an example of the present invention.

[0050] In the diagram: 1 - First quadrant; 2 - Second quadrant; 3 - Third quadrant; 4 - Fourth quadrant; 5 - Inactive bulging area; 6 - Active bulging area. Detailed Implementation

[0051] The technical solution of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0052] Furthermore, the technical features involved in the different embodiments of the present invention described below can be combined with each other as long as they do not conflict with each other.

[0053] To address the three core problems of traditional PGK schemes—nutation divergence, surge in aerodynamic drag, and deterioration of static stability—and to overcome the shortcomings of existing conformal aerodynamic control schemes in terms of platform adaptability, environmental robustness, and system integration, this invention proposes an embedded conformal aerodynamic guidance system suitable for low-rotation aircraft.

[0054] In one example, such as Figure 1As shown, an embedded conformal aerodynamic guidance system for a low-spinning aircraft includes a navigation and attitude sensing module, a flight state calculation module, a guidance correction decision module, and a conformal aerodynamic actuator connected in sequence. The navigation and attitude sensing module is connected to the guidance correction decision module. The system also includes an energy management module to provide stable and reliable power to the entire system and manage the high pulse current demand when the actuator is activated. The entire system is encapsulated in a compact control cabin whose dimensions are fully compatible with standard fuses, allowing for direct replacement installation.

[0055] Specifically, the navigation and attitude sensing module is the system's perception front end, continuously collecting motion state information of the aircraft and transmitting it to the flight state calculation module and guidance correction decision module via an internal high-speed data bus (such as SPI or CAN). This provides foundational data for subsequent trajectory calculation and control decisions. The motion state information includes navigation information, attitude information, and airflow attitude information. Navigation information includes three-dimensional position (longitude, latitude, altitude), three-dimensional velocity vector, and angular velocity, used for real-time calculation of the deviation vector between the current flight trajectory and the preset target flight trajectory. Attitude information includes roll phase angle. Pitch angle, yaw angle, roll phase angle Used to identify the aircraft's current roll attitude, it is a key input for phase-synchronous control and can be obtained through geomagnetic or gyroscope data. Airflow attitude information, including angle of attack α and sideslip angle β, is used for accurate aerodynamic modeling and correction, improving control precision.

[0056] Optionally, the navigation and attitude sensing module employs a combination of a Micro-Electro-Mechanical Systems (MEMS) inertial measurement unit with high overload resistance and a satellite navigation receiver, or other satellite navigation and geomagnetic solutions, to ensure reliable operation even under high overload launch shocks. Navigation information can be acquired via a satellite navigation receiver; airflow attitude information (angle of attack, sideslip angle) can be directly measured via pitot tubes and anemometer sensors, or indirectly calculated based on an aerodynamic model combined with flight parameters (velocity vector, attitude angle, angular velocity); attitude information can be acquired by the inertial measurement unit by acquiring raw angular velocity and acceleration data, combined with information from auxiliary sensors such as magnetometers, and calculated in real time using an embedded sensor fusion algorithm (such as a Kalman filter). In high-speed rotation scenarios, the roll phase angle primarily relies on gyroscope integration to ensure dynamic response accuracy; in low-speed or static scenarios, the roll phase angle is corrected using absolute reference values ​​provided by magnetic sensors, thereby ensuring measurement stability and accuracy under all operating conditions.

[0057] Optionally, the navigation and attitude sensing module calculates the flight trajectory deviation based on the navigation information and the preset target flight trajectory: it establishes a motion-aerodynamic coupling model of the aircraft based on an extended Kalman filter, uses three-dimensional position, three-dimensional velocity, and angular velocity as system observations, and estimates the aircraft state variables in real time, including three-dimensional position, three-dimensional velocity vector, pitch angle, yaw angle, and roll phase angle; it compares the three-dimensional position and three-dimensional velocity in the state variables with the expected state of the target flight trajectory, calculates the flight trajectory deviation (vector), and sends the flight trajectory deviation to the guidance correction decision module.

[0058] The flight status calculation module receives the three-dimensional velocity vector from the navigation and attitude sensing modules, combines it with a pre-stored standard atmospheric model, calculates the current Mach number in real time, and feeds the result back to the guidance correction decision module. Specifically, the calculation of the flight status calculation module follows the basic principles of compressible fluid dynamics and uses the following formula for calculation:

[0059]

[0060] in, For real-time flight Mach number; The magnitude of the three-dimensional velocity vector; The specific heat ratio of air; It is the gas constant; The atmospheric temperature at the current altitude is obtained by looking up a table using the altitude model. The real-time Mach number output by the flight status calculation module is a key environmental criterion for the guidance decision module to determine safe activation.

[0061] The guidance correction decision module is the control core of the system. It can be a low-power, high-performance 32-bit microcontroller, such as an ARM Cortex-M4 core MCU (e.g., the STM32F405 series). These chips have sufficient computing power to run the aforementioned control logic in real time, and their power consumption is extremely low. Furthermore, the guidance correction decision module incorporates an intelligent decision-making logic based on flight state and roll phase to generate the optimal activation command. In this example, the guidance correction decision module receives motion state information and real-time flight Mach number in real time, and determines whether the safety conditions are met based on the real-time flight Mach number. Preferably, the Mach number safety threshold ranges from 1.5 to 2.0, with a typical value of 1.8. When the real-time flight Mach number is less than the Mach number safety threshold, the safety conditions are met. At the same time, the flight trajectory deviation is calculated based on the navigation information and the preset target flight trajectory: a motion-aerodynamic coupling model of the aircraft is established based on the extended Kalman filter. The three-dimensional position, three-dimensional velocity, and angular velocity output by the navigation and attitude sensing modules are used as system observations to estimate the aircraft state variables in real time, including three-dimensional position, three-dimensional velocity vector, pitch angle, yaw angle, and roll phase angle. The three-dimensional position and three-dimensional velocity in the state variables are compared with the expected state of the target flight trajectory to calculate the flight trajectory deviation (vector). Simultaneously, the ideal aerodynamic moment direction angle is calculated based on the pitch angle, yaw angle, and airflow attitude information in the attitude information: A body coordinate system is established based on the pitch and yaw angles. The flight trajectory deviation vector is projected onto the desired correction direction determined by the aircraft cross-section in the body coordinate system. The direction is corrected by combining the aerodynamic characteristics corresponding to the angle of attack and sideslip angle, thus calculating the ideal aerodynamic moment direction angle within the aircraft cross-section. Simultaneously, the roll phase angle and ideal aerodynamic moment direction angle in the attitude information are used to determine whether the phase synchronization condition is met: 1) The absolute difference between the current roll phase angle and the ideal aerodynamic moment direction angle is calculated; 2) The absolute difference is determined to be less than a preset synchronization angle threshold. If less, the phase synchronization condition is met; otherwise, it is not. If both the safety condition and the phase synchronization condition are met, an activation command is generated and sent to the conformal aerodynamic actuator.

[0062] Conformal aerodynamic actuators are the end effectors of the system. They can be flexible skin actuators, such as shape memory alloy actuators or piezoelectric ceramic actuators, capable of generating localized bulging deformation and serving as a conformal disturbance source for corrective aerodynamic forces. The flexible skin actuator is located on the outer surface of the aircraft's nose cone. Its coverage area is aerodynamically optimized, and its outer surface conforms to the aerodynamic shape of the aircraft's nose. In the non-operating state, its outer surface is completely conformal to the projectile surface, with continuous curvature, forming a smooth, continuous aerodynamic shape without any steps or gaps. This achieves the design goal of no radial dimension increase, thus ensuring the conformal characteristics of the system. If necessary, the distance to the flexible skin actuator can be defined as 0.3–0.6 times the projectile diameter backward from the apex. This region exhibits high aerodynamic control efficiency at subsonic and transonic speeds. The flexible skin actuator responds to activation commands by generating localized bulging deformation at a specified location. After the actuation, the flexible skin actuator automatically resets. Taking a flexible skin actuator driven by a shape memory alloy as an example, the working principle is as follows: a shape memory alloy mesh or sheet is embedded under the flexible silicone skin. When the activation command (high current pulse) arrives in the corresponding area, the shape memory alloy is energized and heats up, undergoing an austenitic phase transformation, generating a contraction force that pulls the skin to bulge locally. After the power is turned off, the shape memory alloy cools down and reverts to the martensitic phase. The skin returns to its flat state under its own elasticity or the action of a pre-set reset spring, thus generating local bulging deformation and producing corrective aerodynamic force.

[0063] Preferably, the outer surface of the flexible skin actuator is divided circumferentially into at least four independently controllable action quadrants, such as... Figure 2 As shown, the skin surface of this example flexible skin actuator is divided circumferentially into four independent controllable quadrants: Quadrant 1 (0° to 90°), Quadrant 2 (90° to 180°), Quadrant 3 (180° to 270°), and Quadrant 4 (270° to 360°). Each quadrant can independently receive activation commands and generate bulging deformation. The inactive bulging region 5 and the activated bulging region 6 are shown below. Figure 2 As shown, the independent activation control achievable in each quadrant provides the physical basis for vector control in the roll coordinate system. For example, when the guidance correction decision module calculates that a correction force pointing towards 0° is needed, it will be applied to the roll phase in the first quadrant. When the temperature approaches 0°, an activation command is sent to the flexible skin actuator, thereby applying corrective control at the most appropriate time. Upon receiving the activation command, the flexible skin actuator causes the skin structure in the specified quadrant to bulge locally under the drive of an electrical signal, as shown below. Figure 3 As shown, the height of the activated bulge area 6 is controlled within the range of 0.3–10 mm. After a duration of Δt, it naturally returns to flatness by its own elasticity or by the reset mechanism.

[0064] The energy management module provides a stable power supply for the navigation and attitude sensing module, flight status calculation module, guidance correction decision module, and conformal aerodynamic actuators. It includes a high-density energy storage unit, an energy storage unit, and a power conversion unit. The high-density energy storage unit uses high-energy-density batteries such as thermal batteries and lithium batteries, offering high energy density and a storage life of up to 10 years. The energy storage unit employs a supercapacitor array for instantaneous high-power discharge. Normally, it is trickle-charged by the main energy storage unit and releases a high-pulse current instantaneously when the actuator is activated. The power conversion unit supports low-power standby in standby mode and provides sufficient peak drive power during activation. Simultaneously, the energy management module integrates power management circuitry, intelligently allocating energy and prioritizing continuous power supply to the navigation and decision module, releasing high-pulse current to the conformal aerodynamic actuators only during the window period when the decision module issues an activation command.

[0065] When the system of this invention is running, the modules work together to form an efficient and adaptive closed-loop control process:

[0066] 1) Continuous sensing: The navigation and attitude sensing module continuously collects and outputs the aircraft's motion state information, including three-dimensional position, three-dimensional velocity vector, and roll phase angle. Pitch angle, yaw angle, angle of attack α, and sideslip angle β.

[0067] 2) Status Update: The flight status calculation module updates the current flight Mach number in real time.

[0068] 3) Intelligent decision-making: The guidance correction decision module executes the above decision logic in a loop and judges in real time whether an activation window that meets the safety conditions and phase synchronization conditions has appeared.

[0069] 4) Precise execution: When the safety conditions and phase synchronization conditions are met, the guidance correction decision module sends a precise activation command to the corresponding quadrant of the flexible skin actuator.

[0070] 5) Track correction: The flexible skin actuator inflates within a specified area and for a specified duration, generating a transient, directionally controllable aerodynamic torque, thereby precisely correcting the flight trajectory.

[0071] 6) Reset and standby: After the activation command ends, the flexible skin actuator automatically resets, and the system as a whole returns to the monitoring state, waiting for the next correction window.

[0072] The system of this invention has a clearly defined scope of application and physical integration: 1) Applicable platform: specifically designed for low-rotation aircraft, but can be extended to other aircraft. 2) Integration method: the entire system is completely embedded in the standard guidance component segment of the aircraft's nose. 3) Physical constraints: after installation, the radial dimensions of the system remain completely consistent with the original component, achieving conformal integration without outward protrusion or increased diameter. The technical solution of this invention consists of five functionally tightly coupled and logically coordinated modules, constructing a highly integrated, intelligent decision-making closed-loop control system of perception-decision-execution. Under the strict spatial constraints of the guidance component, it achieves high-precision and high-reliability correction of the flight trajectory of low-rotation aircraft, while maintaining the integrity and stealth of the projectile's aerodynamic shape.

[0073] This invention also includes an embedded conformal aerodynamic guidance and control method for a low-spinning aircraft, the method comprising the following steps:

[0074] S1: Obtain motion state information of the aircraft body.

[0075] The motion status information includes navigation information, attitude information, and airflow attitude information. The navigation information includes three-dimensional position, three-dimensional velocity vector, and angular velocity; the attitude information includes roll phase angle, pitch angle, and yaw angle; and the airflow attitude information includes angle of attack and sideslip angle.

[0076] S2: Calculate the real-time flight Mach number based on navigation information.

[0077] The expression for calculating the real-time flight Mach number is as follows:

[0078]

[0079] in, For real-time flight Mach number; The magnitude of the three-dimensional velocity vector; The specific heat ratio of air; It is the gas constant; The atmospheric temperature at the current altitude is obtained by looking up a table using the altitude model.

[0080] S3: Determine whether the safety conditions are met based on the real-time flight Mach number; calculate the flight trajectory deviation based on the navigation information and the preset target flight trajectory; calculate the ideal aerodynamic moment direction angle based on the pitch angle, yaw angle and airflow attitude information in the attitude information; determine whether the phase synchronization conditions are met based on the roll phase angle and ideal aerodynamic moment direction angle in the attitude information; if both the safety conditions and the phase synchronization conditions are met, generate an activation command.

[0081] Among these, determining whether the safety conditions are met is based on the real-time flight Mach number, i.e., judging the real-time flight Mach number. Is it less than the preset Mach number safety threshold? The Mach number safety threshold ranges from 1.5 to 2.0, with a typical value of 1.8. If... If the safety conditions are met; This indicates that the current flight environment is too harsh (high dynamic pressure, strong aerodynamic heat), and activating the actuator may lead to control failure or structural damage. At this time, any activation commands must be prohibited, and the system returns to monitoring status.

[0082] Determine whether the phase synchronization condition is met based on the roll phase angle and ideal aerodynamic torque direction angle in the attitude information, including:

[0083] 1) Calculate the current roll phase angle With the ideal correction direction absolute difference And determine whether it is less than the preset synchronization angle threshold (phase tolerance window). Synchronization angle threshold The value ranges from 10° to 30°, with a typical value of 15°.

[0084] 2) Determine if the absolute difference is less than the preset synchronization angle threshold. If it is less, the phase synchronization condition is met; otherwise, the phase synchronization condition is not met. Specifically, if... This indicates that the target actuator quadrant has not yet rotated to the optimal force application position. Activating it at this time will cause the control force direction to deviate, or even produce a negative correction. The system continues to monitor and waits for the synchronization window. If This indicates that the synchronization window has appeared, and the track correction (activation) operation can be performed.

[0085] Furthermore, the activation command includes a region identifier (such as a target quadrant identifier) ​​and an activation pulse width. The activation command. The activation pulse width... The typical range is 50~200ms, which can be dynamically adjusted according to the current Mach number to match the thermal response characteristics of the driver under different dynamic pressure environments.

[0086] S4: In response to the activation command, the conformal aerodynamic actuator at the nose of the aircraft body undergoes local deformation.

[0087] Specifically, after receiving the activation command, the flexible skin actuator causes the skin structure in a specific area to bulge locally under the drive of an electrical signal. The bulge height is controlled within the range of 0.3–10 mm. After a duration of Δt, the skin structure naturally returns to flatness by its own elasticity or by a reset mechanism.

[0088] Preferably, before or after determining whether the safety conditions and phase synchronization conditions are met, it is determined whether the number of corrections performed in this flight mission is less than the preset maximum allowable number. When the safety conditions, phase synchronization conditions, and correction number conditions are all met, an activation command is generated.

[0089] Specifically, check the number of corrections performed during the current flight. Has the preset upper limit been reached? (Typical value is 5 flights). If To prevent energy depletion or driver failure due to heat buildup, the system will no longer be activated. If Then an activation command is generated.

[0090] Combining the above methods and applying them to a guidance system, the phase synchronization control logic flow running inside the guidance correction decision module is as follows: Figure 4 As shown, this process is executed cyclically at a high frequency (e.g., >100Hz):

[0091] S10: After the system is powered on, the guidance correction decision module enters the real-time monitoring loop, continuously acquiring motion state information and flight trajectory deviation vector from the navigation and attitude sensing modules, and the real-time flight Mach number from the flight state calculation module. ;

[0092] S20: Mach number determination: Determine the current... Is it less than the preset Mach number threshold? (Configured as 1.8 in this embodiment). If If, then return to S10; if Then proceed to S30.

[0093] S30: Calculate the correction direction: Calculate the ideal aerodynamic moment direction angle based on the pitch angle, yaw angle, and airflow attitude information in the attitude information. .

[0094] S40: Phase synchronization judgment: judgment Is it smaller than the preset phase tolerance window? (In this embodiment, the angle is configured as 15°); if If, then return to S10; if Then it enters S50;

[0095] S50: Activation Count Determination: Determines the number of corrections performed during the current flight. Has the preset maximum allowable value been reached? (In this embodiment, it is configured to be 5 times); if If, then return to S10; if Then proceed to S60.

[0096] S60: Generate and execute activation instructions: The guidance correction decision module, based on... The target quadrant is determined, and the required activation pulse width Δt is calculated based on the deviation. An activation command containing the target quadrant and Δt is generated and sent to the flexible skin actuator. Simultaneously, the activation count counter N is incremented by 1. The actuator automatically resets after completing the activation for duration Δt, and the system returns to S10.

[0097] This invention uses a unique phase synchronization control logic to precisely correlate the activation timing of the flexible skin actuator with the roll phase of the projectile, overcoming the fundamental defect that traditional fixed control surfaces cannot generate effective average control force under low-spin conditions and instead induce conical motion, thus ensuring flight stability.

[0098] The above detailed embodiments are a description of the present invention. It should not be considered that the specific embodiments of the present invention are limited to these descriptions. For those skilled in the art, several simple deductions and substitutions can be made without departing from the concept of the present invention, and all of these should be considered to fall within the protection scope of the present invention.

Claims

1. An embedded conformal aerodynamic guidance system for a low-spin aircraft, comprising: include: The navigation and attitude sensing module is used to acquire motion state information of the aircraft body; the motion state information includes navigation information, attitude information, and airflow attitude information; the navigation information includes three-dimensional position, three-dimensional velocity vector, and angular velocity; the attitude information includes roll phase angle, pitch angle, and yaw angle; the airflow attitude information includes angle of attack and sideslip angle. The flight status calculation module, connected to the navigation and attitude sensing module, is used to calculate the real-time flight Mach number based on navigation information; The guidance correction decision module, connected to the navigation and attitude sensing module and the flight state calculation module, is used to receive motion state information and real-time flight Mach number, and determine whether the safety conditions are met based on the real-time flight Mach number; calculate the flight trajectory deviation based on navigation information and the preset target flight trajectory; calculate the ideal aerodynamic moment direction angle based on the pitch angle, yaw angle, airflow attitude information and flight trajectory deviation in the attitude information; and determine whether the phase synchronization condition is met based on the roll phase angle and ideal aerodynamic moment direction angle in the attitude information. If both the safety conditions and the phase synchronization condition are met, an activation command is generated. The conformal aerodynamic actuator is located at the nose of the aircraft body. Its outer surface conforms to the aerodynamic shape of the nose of the aircraft body. It is connected to the guidance correction decision module and is used to generate local deformation in response to activation commands. The energy management module provides power to the navigation and attitude sensing module, flight status calculation module, guidance correction decision module, and conformal aerodynamic actuator. The determination of whether safety conditions are met based on the real-time flight Mach number includes: Determine whether the real-time flight Mach number is less than the preset Mach number safety threshold. If it is less, the safety condition is met; otherwise, the safety condition is not met. The step of determining whether the phase synchronization condition is met based on the roll phase angle and ideal aerodynamic torque direction angle in the attitude information includes: Calculate the absolute difference between the current roll phase angle and the ideal aerodynamic moment direction angle; Determine whether the absolute difference is less than the preset synchronization angle threshold. If it is less, the phase synchronization condition is met; otherwise, the phase synchronization condition is not met.

2. The embedded conformal aerodynamic guidance system for low-rotation aircraft according to claim 1, characterized in that, The outer surface of the conformal pneumatic actuator is divided circumferentially into at least four independently controllable action quadrants.

3. The embedded conformal aerodynamic guidance system for low-rotation aircraft according to claim 1, characterized in that, The conformal pneumatic actuator is driven by a shape memory alloy driver or a piezoelectric ceramic driver, thereby generating local bulging deformation.

4. An embedded conformal aerodynamic guidance and control method for a low-spinning aircraft, characterized in that, Includes the following steps: Acquire motion state information of the aircraft body; the motion state information includes navigation information, attitude information combined with airflow attitude information; the navigation information includes three-dimensional position, three-dimensional velocity vector and angular velocity; the attitude information includes roll phase angle, pitch angle and yaw angle; the airflow attitude information includes angle of attack and sideslip angle. Calculate the real-time flight Mach number based on navigation information; The system determines whether the safety conditions are met based on the real-time flight Mach number; calculates the flight trajectory deviation based on navigation information and the preset target flight trajectory; calculates the ideal aerodynamic moment direction angle based on the pitch angle, yaw angle, airflow attitude information, and flight trajectory deviation in the attitude information; and determines whether the phase synchronization conditions are met based on the roll phase angle and ideal aerodynamic moment direction angle in the attitude information. If both the safety conditions and the phase synchronization conditions are met, an activation command is generated. In response to the activation command, the conformal aerodynamic actuator at the nose of the aircraft body undergoes local deformation; The determination of whether safety conditions are met based on the real-time flight Mach number includes: The system determines whether the real-time flight Mach number is less than a preset Mach number safety threshold. If it is less, the safety condition is met; otherwise, the safety condition is not met. The step of determining whether the phase synchronization condition is met based on the roll phase angle and ideal aerodynamic torque direction angle in the attitude information includes: Calculate the absolute difference between the current roll phase angle and the ideal aerodynamic moment direction angle; Determine whether the absolute difference is less than the preset synchronization angle threshold. If it is less, the phase synchronization condition is met; otherwise, the phase synchronization condition is not met.

5. The embedded conformal aerodynamic guidance and control method for a low-rotation aircraft according to claim 4, characterized in that, The Mach number safety threshold ranges from 1.5 to 2.

0.

6. The embedded conformal aerodynamic guidance and control method for a low-rotation aircraft according to claim 4, characterized in that, The range of the synchronization angle threshold is 10 degrees to 30 degrees.

7. The embedded conformal aerodynamic guidance and control method for a low-rotation aircraft according to claim 4, characterized in that, Before or after determining whether the safety conditions and phase synchronization conditions are met, it is determined whether the number of corrections performed in this flight mission is less than the preset maximum allowable number. When the safety conditions, phase synchronization conditions, and correction number conditions are all met, an activation command is generated.

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