A control system and launch method for airborne folding-wing unmanned aerial vehicles (UAVs).

Abstract

This invention discloses a control system and launch method for an airborne folding-wing unmanned aerial vehicle (UAV). Before the launch of the sub-unit, the mother aircraft sends a voltage signal to the sub-unit, which is continuously monitored by the sub-unit's flight control system. Upon reaching the launch point, the mother aircraft is powered off, the unlocking device is activated, and the flight control system initiates control after the voltage monitoring disappears for a period of time. At this moment, the V-tail controls pitch and participates in roll control. After a certain period, the flight control system unlocks the locking servo, causing the wings to begin unfolding until they are fully extended. At this time, the ailerons participate in control, entering the unpowered dive-pull control phase. After a certain period, the wing control is locked. Subsequently, the flight control system determines that the attitude and airspeed are stable. Once stable, the engine is started, and the engine speed is determined. If the start fails, the start operation is repeated three times. After the engine starts, the flight path is executed normally, the ground station launch process is exited, and normal flight control logic is entered. This invention enables full-process control of an airborne folding-wing UAV.

Description

Technical Field

[0001] This invention belongs to the field of unmanned aerial vehicle (UAV) control, specifically relating to a control system and launch method for an airborne, mounted, and launched folding-wing UAV. Background Technology

[0002] Currently, drones are limited by their own software, hardware, and design constraints, resulting in limitations in certain application environments. For individual drones, small drones, due to insufficient fuel and short range, cannot travel thousands of kilometers to perform missions. Using large drones, on the other hand, is costly and risky. Therefore, the concept of mother-daughter drone collaboration has emerged. This model uses a large, long-endurance, long-range drone (or amphibious drone) as the mother unit, carrying 1-2 small drones as daughter units. When the mother unit, carrying the daughter units, reaches the daughter units' operational range, it opens its mounting system to launch the daughter units from the air. After launch, the daughter units initiate control, utilizing their folding mechanism to restore normal flight operation; they then activate normal flight control, maintaining attitude and stabilizing their flight path. After completing their mission, the daughter units are typically recovered via parachute, net capture, or disposable loitering munitions.

[0003] The mother-daughter aircraft dropping and daughter aircraft participating in combat significantly increases the operational range of small unmanned aerial vehicles (UAVs), making it possible for a large number of small UAVs to participate in long-range reconnaissance and strike missions, greatly reducing the cost of long-range operations. Simultaneously, small UAVs participating in combat missions also offer advantages such as high stealth. The technology for daughter aircraft mounting is also relatively mature. The mounted daughter aircraft are often designed with wing-folding mechanisms. When mounted, the integrated wing rotates 90° and is positioned parallel to the fuselage below it, resembling a normally mounted missile. This design also greatly reduces flight drag during mother-daughter aircraft flight. Since a single mother aircraft can carry multiple daughter aircraft, multiple mother aircraft can simultaneously release daughter aircraft during missions, forming a cluster communication among the daughter aircraft. Using a cluster approach for saturation strikes greatly improves mission success rates.

[0004] By closely combining medium and large-sized UAVs with small and micro-sized UAVs, the limitations of each type are overcome. This combination fully utilizes the powerful firepower and long endurance of medium and large-sized UAVs, while leveraging the flexibility and maneuverability of small and micro-sized UAVs. Medium and large-sized UAVs can carry more munitions and equipment, enabling them to perform more complex and decisive missions, while small and micro-sized UAVs can provide real-time intelligence on the battlefield and conduct reconnaissance and strikes against enemy targets. The advantage of this combination is that it leverages the complementary characteristics of various UAV types to accomplish a wider range of missions and operational needs.

[0005] The airborne launch control of the sub-aircraft can be divided into the control logic at the launch moment and the sub-aircraft control method after normal launch. After normal launch, the initial attitude of the aircraft may be quite complex, with irregular magnitudes of the three-axis angles and angular velocities, often requiring the use of classic PID control algorithms. In the launch moment control logic, to prevent collision between the sub-aircraft and the mother aircraft after unlocking, the sub-aircraft usually does not immediately launch. This allows the sub-aircraft and mother aircraft to maintain a safe altitude difference, and the sub-aircraft's wing folding mechanism can be restored during this period. It can be seen that constructing a reasonable sequence of logic—disengagement, sub-aircraft wing extension, sub-aircraft launch, unpowered gliding launch, and powered launch—is crucial for safely completing the sub-aircraft launch process. Among the existing drop control methods, there are control methods for dropping quadcopter UAVs. These methods employ PID control algorithms and, based on this technology, fuzzy control methods. Some existing methods also design a mission flow for the carrier platform. The UAV's mission execution process is mainly divided into the hang-up phase, descent and deployment phase, mission phase, and recovery / self-destruct phase. The key descent and deployment phase involves the UAV deploying its parachute and landing at the predetermined location after being dropped. After stabilizing its attitude, it deploys its rotors and begins controllable powered flight after the parachute is deactivated. Summary of the Invention

[0006] To address the aforementioned problems, this invention proposes a control system and launch method for an airborne folding-wing UAV. In addition to controlling the normal flight of the UAV, this invention designs a launch process for airborne UAV launch. Furthermore, for each stage of the airborne launch process, control timing and time intervals are designed to ensure the stability and safety of the airborne launch. Control methods and control laws for each stage are designed, and protection logic for the airborne launch process is designed and incorporated into the flight control hardware, thus completing the design of a full-process flight control system.

[0007] The control system for airborne launch of a folding-wing UAV of the present invention includes a folding-wing V-tail UAV, a carrier aircraft, a flight control and navigation system module, a ground station module, airborne launch flight control software, a pylon hook, etc.

[0008] The detailed steps of the control system and launch method for the airborne mounted folding-wing UAV described in this invention are as follows:

[0009] Step 1: The mother unit continuously supplies a voltage signal to the daughter unit as the signal voltage for airborne transmission.

[0010] Step 2: The ground unit confirms that the conditions for deployment are met and sends a deployment signal to the mother unit, thus entering the deployment ready state;

[0011] Step 3: The slave flight controller continuously monitors the voltage signal sent by the master unit in Step 1.

[0012] Step 4: The operator of the main machine cuts off the voltage signal, the locking device of the main machine unlocks, the hook opens, and the slave machine is deployed;

[0013] Step 5: The slave flight controller ADC detects the voltage signal. If the monitored voltage signal is 0 for a certain period of time, it means that it has separated from the mother machine. Record the current time as T0 and proceed to step 6.

[0014] Step Six: The control surfaces of the tail fin of the sub-aircraft begin to be controlled. The roll attitude of the folded aircraft is adjusted by deflecting the tail fin control surfaces to suppress the lateral roll of the folded body.

[0015] Step 7: After a period of time from T0, the time becomes T1. At T1, the flight controller sends a command to the wing lock servo, which opens the wing lock servo and controls the wings to begin rotating and deploying.

[0016] Step 8: After the wing lock servo is engaged, a period of time elapses until time T2. At this time, the wings are fully deployed and the ailerons intervene to control and stabilize the aircraft's attitude. After the wings are fully deployed, a period of time elapses until time T3'. At this time, the flight control system sends a command to the wing lock servo, which then locks the wings.

[0017] Step Nine: After the wings are fully deployed, the aircraft enters the unpowered gliding phase. At this time, the aircraft is in a high-speed nose-down dive. It is controlled by pulling up during the unpowered dive. After a period of time, the attitude stabilizes. The moment of attitude stabilization is T3. If the aircraft's altitude is lower than the minimum parachute opening altitude, the parachute is opened for landing. If the aircraft's altitude is not lower than the minimum parachute opening altitude after the attitude stabilizes, proceed to Step Ten.

[0018] Step 10: At time T3, the flight control system sends an engine start command and checks if the engine has started successfully. The time when the engine starts successfully is recorded as T4. If the engine is not detected to have started, it is restarted after a 2-second interval. If the engine fails to start three times and the aircraft's altitude is below the minimum parachute deployment altitude, the aircraft deploys its parachute for landing. If the altitude is above the minimum parachute deployment altitude, the aircraft continues to glide without power. During gliding, the ground station starts the engine. If the engine fails to start and the aircraft's altitude is below the minimum parachute deployment altitude, the aircraft deploys its parachute for landing.

[0019] Step 11: After the engine starts, exit the air-based launch procedure and the aircraft begins the mission phase.

[0020] Step 12: Perform the normal flight path mission through the conventional control law unit in the air-launched flight control module until the parachute is recovered.

[0021] In the aforementioned launch method for air-launched folding-wing UAVs, protection logic is designed to ensure the safety of air-launch during the drop and pull-up phases of the sub-unit. Specifically, this includes:

[0022] (1) Engine start failure protection

[0023] If the engine fails to start after three attempts, the engine start failure protection logic is activated. At this point, the drone stops attempting to start the engine and begins unpowered gliding towards its home point, which is located at the initial drop point but at an altitude of 150m. Upon reaching the home point, the drone begins circling it, awaiting further instructions from the ground station. During this protection period, manual ignition commands can be issued via the ground station, or the drone can be switched to semi-autonomous mode and manually controlled remotely. If the engine still fails to start successfully, and the drone falls below the altitude protection line during circling and descent, altitude protection is triggered, and the drone deploys its parachute for descent.

[0024] (2) Forced pull-up

[0025] When the drone's altitude drops to 800m during the pull-up control process, and the flight control determines that the drone's attitude has not yet stabilized, a forced pull-up is initiated. The engine starts, and the pitch channel control law is switched to the normal flight state, with a target pitch angle of 0°.

[0026] (3) Attitude protection

[0027] If the drone pitches more than 70° or rolls more than 90° during normal flight and fails to recover within 4 seconds, the drone's attitude protection will be triggered. The flight controller will assume that the drone is out of control at this time and will issue a parachute deployment command.

[0028] (4) High level of protection

[0029] When the drone's altitude drops below 100m, altitude protection is triggered, and the drone deploys its parachute for a landing.

[0030] The timing logic of the above-mentioned launch method for air-launched folding-wing UAVs is as follows:

[0031] Timing 1: When the disengagement signal is detected, the ADC detects that the voltage has dropped to 0V. The disengagement signal is continuously detected for 200ms to prevent false judgments caused by voltage instability. After the disengagement signal is continuously detected for 200ms, the current time of detection is recorded as T0. At this time, control is activated to adjust the aircraft's attitude in the folded state via the tail rudder.

[0032] Timing 2: After the wings are deployed and the detachment is detected, 300ms after the detachment is time T1. At this time, the flight controller sends a wing lock servo command, the servo is released, and the wings begin to rotate and deploy. T1 = T0 + 300ms.

[0033] Timing 3: Wing lock. After 1300ms, the servo is engaged, which is time T2. At this time, the wing is fully deployed, T2 = T1 + 1300ms. After the wing is fully deployed, it takes 500ms to reach time T3', T3' = T2 + 500ms. The flight control system sends a wing lock command to the servo, and the servo locks the wing.

[0034] Sequence 4: After the wings are fully deployed, the aircraft enters the unpowered gliding phase. After Nms, the attitude is considered stable, N = 5s to 20s. At this time, if the altitude is lower than the minimum parachute deployment altitude, the parachute will be deployed for landing.

[0035] Timing 5: At time T3, the flight controller sends an engine start command, T3 = T2 + Nms; after 3.3s, it checks whether the engine has started successfully, i.e., the speed is >2000rpm for 3 seconds to prevent misjudgment caused by the wind blowing the propeller blades at high speed and instantaneous rotation. The start-up completion time is recorded as T4; if it is detected that the engine has not started, it will restart after an interval of 2s.

[0036] Timing 6: Engine start-up complete. After the aircraft's attitude stabilizes, it enters the mission phase. If the engine starts successfully on the first attempt, then T4 = T3 + 3300ms; otherwise, T4 is calculated based on the actual number of ignitions and the time taken.

[0037] The advantages of this invention are:

[0038] (1) The present invention provides a control system and launch method for airborne folding-wing UAVs, which provides a larger combat radius for loitering drones. Compared with the traditional drop method, the airborne launch process and control method can stably and safely complete the airdrop mission, and has been tested and verified.

[0039] (2) The present invention provides a control system and launch method for an airborne folding-wing UAV, which designs a reasonable airborne launch process and timing logic, which greatly ensures the safety of the mother and daughter aircraft and the attitude stability of the daughter aircraft during the launch process.

[0040] (3) The present invention provides a control system and launch method for an airborne folding-wing UAV, which has two control phases designed specifically: folding tail wing attitude control and gliding pull-up control, making the flight during the launch phase safer and more stable. The modified protection logic also allows the UAV to adapt to the launch process. Attached Figure Description

[0041] Figure 1 This is a flowchart of the airborne launch method for a folding-wing UAV according to the present invention.

[0042] Figure 2 This is a schematic diagram of the mother and daughter aircraft during the launch process of the airborne folding-wing UAV of the present invention.

[0043] Figure 3 This is a launch timing logic diagram of an airborne launch method for a folding-wing UAV according to the present invention. Detailed Implementation

[0044] The present invention will now be described in further detail with reference to the accompanying drawings.

[0045] The control system for airborne launch of a folding-wing unmanned aerial vehicle (UAV) of the present invention includes a folding-wing V-tail UAV, a carrier aircraft, a flight control and navigation module, a ground station module, an airborne launch flight control module, and hooks and hoops.

[0046] The mother aircraft is equipped with a hook on its belly, and the daughter aircraft has two longitudinally arranged hooks on its upper fuselage. On the ground, the hooks and hooks are connected, locking the mother and daughter aircraft. Upon reaching the air-to-ground launch site, the mother aircraft sends a release command (voltage signal) to the daughter aircraft's flight control system. Simultaneously, the hooks on the mother aircraft's belly release, launching the daughter aircraft in a projectile-like motion. The flight control and navigation module is installed on the daughter aircraft, responsible for communication between the daughter and mother aircraft, and between the daughter aircraft and the ground station. The daughter aircraft carries an air-to-ground launch flight control module to ensure successful air-to-ground launch and stable, normal flight. The ground station module is responsible for communication with the flight control system, sending command signals, and monitoring the aircraft's status.

[0047] The aforementioned airborne launch flight control module is the core of the patent design, including the launch control part of the airborne launch UAV, the launch timing logic part of the airborne launch UAV, the launch process control law part of the airborne launch UAV, the traditional control law part, and the protection logic part.

[0048] A launch control method for airborne folding-wing UAVs based on an airborne launch flight module, such as... Figure 1 , Figure 2 As shown, the details are as follows:

[0049] Step 1: Preparations before disembarkation:

[0050] The mother aircraft continuously supplies a DC 28V voltage signal to the daughter aircraft as the signal voltage for airborne transmission. During the flight phase, the daughter aircraft must manually press the release button on the ground station to send a release-ready signal to the mother aircraft, entering the release-ready state. Only then can the mother aircraft proceed with the subsequent drop-off process, avoiding misjudgments caused by voltage signal fluctuations. When not in the release-ready state, the daughter aircraft's flight control system will not monitor the voltage or enter the drop-off control process.

[0051] Step 2: Deployment Confirmation

[0052] Once the ground station confirms that the deployment conditions are met (all equipment is in normal condition and the separation signal is normal), manually click the deployment button in the ground station to send a deployment signal to the host unit and enter the deployment state.

[0053] Step 3: Voltage Monitoring

[0054] The flight controller of the sub-aircraft continuously monitors the DC28V voltage (range 0-36V) and does not initiate the deployment process unless an unlock command is received.

[0055] Step 4: Sub-machine deployment:

[0056] When the operator of the main machine presses the release button (cuts off the DC28V voltage), the locking device of the main machine is unlocked, the hook is opened, and the slave machine is deployed.

[0057] Step 5: Detect the detachment signal:

[0058] The slave unit's flight control ADC detects the disconnection signal, i.e., the voltage changes from 28V to 0V (from present to absent). It continues to detect the disconnection signal for a period of time to confirm that the voltage signal has been cut off and the slave unit has disconnected from the mother unit.

[0059] Step Six: Folded Tail Wing Attitude Control

[0060] When the departure signal is detected continuously until the set time, the current time is recorded as T0. At this time, the control surfaces of the tail fin of the aircraft begin to control the aircraft, adjusting the roll attitude of the folded aircraft by deflecting the tail fin control surfaces, and suppressing the lateral roll of the folded body.

[0061] Step 7: Deploy the wings:

[0062] After the slave unit is confirmed to have disengaged at time T0, a period of time is observed until time T1. At this time, the flight controller sends a command to the wing lock servo, which then opens the wing lock servo and controls the wing to begin rotating and deploying.

[0063] Step 8: Lock the wings:

[0064] After a period of time following the engagement of the wing-lock servo, at time T2, the wings are fully deployed, and the ailerons intervene to stabilize the aircraft's attitude. After a period of time following the full deployment of the wings, at time T3', the flight control system sends a command to the wing-lock servo, which then locks the wings.

[0065] Step Nine, Unpowered Gliding:

[0066] After the wings are fully deployed, the aircraft enters the unpowered gliding phase. During this phase, the aircraft is in a high-speed, nose-down dive, controlled by pulling up during the unpowered dive. After a period of time, the attitude stabilizes at point T3. The stabilization conditions are determined by the flight control system, primarily based on three-axis attitude and airspeed. If the aircraft's altitude is below the minimum parachute deployment altitude, it will deploy and land. If, after stabilization, the aircraft's altitude is not lower than the minimum parachute deployment altitude, then proceed to step ten.

[0067] Step 10: Start the engine:

[0068] At time T3, the flight controller sends an engine start command and checks if the engine has started successfully (speed > 2000 rpm). The time when the engine starts successfully is recorded as T4. If the engine is not detected to have started, it is restarted after a 2-second interval. If the engine fails to start three times and the aircraft's altitude is below the minimum parachute deployment altitude, the aircraft will deploy and land. If the aircraft's altitude is above the minimum parachute deployment altitude, the aircraft will continue to glide without power. During the glide, the ground station will attempt to start the engine. If the engine fails to start and the aircraft's altitude is below the minimum parachute deployment altitude, the aircraft will deploy and land.

[0069] Step 11: After the engine starts, exit the air-based launch procedure and the aircraft begins the mission phase.

[0070] Step 12: Perform the normal flight path mission through the conventional control law unit in the air-launched flight control module until the parachute is recovered.

[0071] During the aforementioned control method, from the start of the airborne launch of the drone to the full deployment of its wings, the drone is in a folded wing state or a partially unfolded "semi-folded state." Due to the lack of lateral roll damping provided by the wings, the drone's lateral stability is poor. Once crosswinds cause lateral forces and moments, the drone will experience large roll angles and roll angular velocities, making it difficult to control after the wings are deployed and potentially causing the drone to enter an unstable mode or attitude divergence. Since this invention uses a folded-wing V-tail drone, the tail fin adopts a "V-tail" configuration with a 30° angle to the horizontal. When the two tail fins move differentially, the lift on one tail fin increases while the lift on the other decreases, distributing to the horizontal direction, similar to an aileron. This provides a certain roll torque to suppress the drone's roll in the "folded state" and maintain stability in the roll direction. Therefore, this invention designs a folded-wing attitude control method, in which the folded tail fin participates in attitude control. By allocating a portion of the "V-tail" stroke, additional roll torque is differentially generated to suppress lateral roll.

[0072] The tail control surfaces of the folding-wing V-tail UAV have a travel of ±15°. During tail attitude control, pitch control is also required to maintain stability in the pitch direction. However, after being launched, the UAV initially adopts a nose-down flight pattern, waiting for the wings to fully extend before pulling up. Therefore, minimal pitch torque is needed to maintain level flight. This invention employs a 5° V-tail control surface travel for pitch control, using the normal control logic. A 10° V-tail control surface travel is used to control roll, employing inner-loop PID control. The target control variable is the UAV's target roll angle Φ. a With the current roll angle Φ of the drone c The difference between them is due to the target roll angle Φ of the UAV at this stage. a The angle is always 0°, therefore the target control variable is also the UAV roll angle Φ.c The control quantity is processed using a PID controller to ultimately obtain the V-tail differential steering quantity δ. r1 The control law is as follows:

[0073]

[0074] in, These are the control parameters for the gain circuit. For the control parameters of the integral element, This represents the gain of the differential element.

[0075] The initiation time of the aforementioned folded tail attitude control is the launch time of the UAV slave unit, i.e., the moment the flight controller receives the disappearance of the AD28V voltage at T0. At this time, the slave unit separates from the mother unit, and the slave unit performs a projectile-like motion. The V-tail begins to control and suppress roll. The end time of the folded tail attitude control is the moment the wings are fully deployed, i.e., time T2. At this time, the slave unit's ailerons initiate control, using conventional control methods to control the UAV's lateral stability. Simultaneously, the V-tail control logic switches to normal control logic to control heading and pitch.

[0076] In the above control method, the pitch channel control method used in the unpowered dive-pull-up control phase is a pitch-airspeed closed-loop control, while the roll channel remains at 0°. Since the initial velocity of the UAV is 55 m / s, and its cruising speed is 50 m / s, after air-launch, the UAV will dive for a short period. At the start of the unpowered dive-pull-up control, the UAV's flight state is at a large angle and high speed with its nose down. Therefore, pitch-airspeed closed-loop control is used, controlling the elevator surfaces to control airspeed, causing the UAV to pitch up and decelerate, ultimately achieving stable pitch attitude and airspeed. This invention designs a set of attitude stability judgment logic. After stabilization, the flight controller automatically determines whether a stable state has been reached. Once a stable state is reached, the flight controller sends an engine ignition command. The start time of the unpowered dive-pull-up control phase is when the wings are fully deployed and the ailerons are activated; the end time is when the flight controller determines that the UAV's attitude is stable and the engine is started for the first time. The dive-pull-up control in this invention differs from the pitch channel control in normal control mode, and employs an independent control law and entirely new control parameters, specifically:

[0077] The physical quantities and control parameters for the unpowered dive-pull-up control phase are as follows:

[0078] Pull-up phase flag temp: temp displays 2 for pull-up preparation; 1 for pull-up in progress; and 0 for pull-up completion.

[0079] Pitch control airspeed proportional coefficient The default value is 2;

[0080] Pitch control airspeed integral coefficient Default is 0.1;

[0081] The initial value of the drone's speed, V0, measured at the moment the control is pulled up;

[0082] The speed V of the drone measured during the pull-up phase;

[0083] Given the final value of the target velocity V c The cruising speed is 50 m / s.

[0084] Given an initial velocity bias ΔV;

[0085] Exponential transition time constant T: defaults to 2s;

[0086] The target speed V during the pull-up control phase tar A function that changes exponentially with time T, with an initial time of V0 + ΔV and a final time of V. c ;

[0087] The table below summarizes the state variables and control parameters during the pull-up control phase:

[0088] Table 1: State variables and control parameters during the pull-up control phase

[0089]

[0090]

[0091] The above pitch-airspeed closed-loop control uses a PI controller to stabilize the UAV's airspeed, and the control law is as follows:

[0092]

[0093] The elevator value δ of the UAV during the pull-up control phase is obtained. e_pull It is used to control speed stability and pitch angle stability.

[0094] In the aforementioned control method, this invention also incorporates flight control attitude stability judgment logic, which serves as the command signal for engine start-up. The flight control attitude stability judgment logic is as follows: the pull-up is considered complete only when all three conditions are met: current pitch is within ±12 degrees, roll angle is within ±10 degrees, and current airspeed is within ±3 degrees of the set airspeed (i.e., cutoff airspeed) value. Engine start-up then begins. In this invention, during the airborne launch phase of the UAV sub-unit, the flight state and flight logic differ from normal flight, rendering some conventional protection logic inapplicable. Therefore, this invention also incorporates a series of protection logics during the sub-unit drop and pull-up phases, such as altitude protection and forced pull-up, to ensure the safety of airborne launch.

[0095] (1) Engine start failure protection

[0096] In the launch control method of this invention, after the flight control system determines that the UAV's attitude is stable, it will initiate engine ignition. After ignition, the flight control system will determine the propeller speed. If ignition fails, a second ignition will be attempted, for a total of three ignition attempts. Due to the influence of factors such as altitude, temperature, and airspeed, the success rate of high-altitude ignition is lower than that of ground ignition. If all three ignition attempts fail, the engine start failure protection logic will be activated. At this time, the UAV will no longer attempt engine start and will begin unpowered gliding. The gliding target waypoint is the "home" point, which is located at the initial drop point but at an altitude of 150m. After reaching the "home" point, the UAV will begin circling around it, awaiting further instructions from the ground station. During the entire protection period, manual ignition commands can be issued using the ground station, or the system can switch to semi-autonomous mode and manually control the UAV using remote control. If the engine still fails to start successfully, and the UAV falls below the altitude protection line during circling and descent, altitude protection will be triggered, and the UAV will deploy its parachute for descent.

[0097] (2) Forced pull-up

[0098] After the wings are fully extended, the aircraft initially assumes a nose-down dive attitude. At this point, the engines are not started, and the aircraft begins unpowered gliding. Simultaneously, the elevators control the drone to slowly pull up. The control law for this pull-up process is independent of the control law for the pitch channel under normal flight conditions. If the pull-up process is too slow, it will affect the subsequent engine start-up sequence; therefore, a forced pull-up protection mechanism is implemented. When the drone's altitude drop reaches 800m during the pull-up control process, and the flight control determines that the drone's attitude has not yet stabilized, a forced pull-up is initiated. The engines are started, and the aircraft switches to the pitch channel control law under normal flight conditions, with a target pitch angle of 0°.

[0099] (3) Attitude protection

[0100] Under normal flight conditions, the drone will not exhibit extreme attitude situations. Therefore, when the drone pitches more than 70° and rolls more than 90° and fails to recover within 4 seconds, the drone's attitude protection will be triggered. The flight controller will assume that the drone is out of control at this time and will issue a parachute deployment command. During the drone's air-based launch phase, especially before the wings are deployed, it is normal for the drone to exhibit large attitudes in the pitch and roll directions for extended periods. In this case, it is not necessary to trigger attitude protection. Instead, it is necessary to wait for the drone's wings to deploy and then slowly adjust the drone's attitude. Therefore, the attitude protection logic for the drone during the airdrop phase is canceled. Once the drone airdrop process is successfully completed, i.e., the engine ignites successfully, the attitude protection logic returns to the normal flight state. (4) Altitude Protection

[0101] Similar to normal flight, the launch control method of this invention also includes altitude protection to prevent a high-speed crash of the drone due to a process error during air-based launch. When the drone's altitude drops below 100m, altitude protection is triggered, and the drone deploys its parachute for a descent.

[0102] In addition to the aforementioned launch control method flow design, it is necessary to design the time intervals between each process, also known as the airborne launch timing logic design. A suitable airborne launch timing logic can make the control process of airborne launch of folding-wing UAVs more reasonable, more stable, and safer. It can also reduce the risk of collision between the slave unit and the mother unit, reduce the large attitude changes of the slave unit during airborne launch, and ensure the success of the airborne launch mission. Therefore, the timing logic design in this invention is as follows: Figure 3 As shown, the details are as follows:

[0103] Timing 1: When the disengagement signal is detected, the ADC detects a voltage change to 0V (from present to absent). The disengagement signal is continuously detected for 200ms to prevent false alarms due to voltage instability. After 200ms of continuous detection, the current time of detection is recorded as T0. At this point, control is activated, and the aircraft's attitude in the folded state is adjusted via the tail rudder.

[0104] Timing 2: After the wings are deployed and the detachment is detected, 300ms after the detachment is time T1. At this time, the flight controller sends a wing lock servo command, the servo is released, and the wings begin to rotate and deploy. T1 = T0 + 300ms.

[0105] Timing 3: Wing lock. After 1300ms, the servo is engaged, which is time T2. At this time, the wing is fully deployed, T2 = T1 + 1300ms. After the wing is fully deployed, it takes 500ms to reach time T3', T3' = T2 + 500ms. The flight control system sends a wing lock command to the servo, and the servo locks the wing.

[0106] Time Sequence 4: After the wings are fully deployed, the aircraft enters the unpowered gliding phase. Attitude is considered stable after Nms (N = 5s–20s). If the altitude is below the minimum parachute deployment altitude at this time, the aircraft will deploy and land. Time Sequence 5: At time T3, the flight control system issues an engine start command (T3 = T2 + Nms). After 3.3s, it checks if the engine has started successfully (i.e., engine speed > 2000 rpm for 3 seconds) to prevent misjudgment caused by wind-induced high-speed instantaneous rotation of the propeller blades. The start-up completion time is recorded as T4. If no start is detected, the engine will restart after a 2-second interval. If three attempts to start the engine fail, and the altitude is above the minimum parachute deployment altitude, unpowered gliding continues. If the altitude is below the minimum parachute deployment altitude, the aircraft will deploy and land.

[0107] Timing 6: Engine start-up complete. After the aircraft's attitude stabilizes, it enters the mission phase. If the engine starts successfully on the first attempt, then T4 = T3 + 3300ms; otherwise, T4 is calculated based on the actual number of ignitions and the time taken.

Claims

1. A method for launching an airborne folding-wing unmanned aerial vehicle (UAV), characterized in that: The specific steps are as follows: Step 1: The mother unit continuously supplies a voltage signal to the daughter unit as the signal voltage for airborne transmission; Step 2: The ground unit confirms that the conditions for deployment are met and sends a deployment signal to the mother unit, thus entering the deployment ready state; Step 3: The slave flight controller continuously monitors the voltage signal sent by the master unit in Step 1; Step 4: The operator of the main machine cuts off the voltage signal, the locking device of the main machine unlocks, the hook opens, and the slave machine is deployed; Step 5: The slave flight controller ADC detects the voltage signal. If the monitored voltage signal is 0 for a certain period of time, it means that it has separated from the mother machine. Record the current time as T0 and proceed to step 6. Step 6: The control surfaces of the tail fin of the sub-aircraft begin to be controlled. The roll attitude of the folded aircraft is adjusted by deflecting the tail fin control surfaces to suppress the lateral roll of the folded body. Step 7: After a period of time from T0, it becomes T1; at T1, the flight controller sends a command to the wing lock servo, the wing lock servo opens, and the wing begins to rotate and deploy. Step 8: After the wing-locking servo is engaged, a period of time elapses, which is time T2. At this point, the wings are fully deployed, and the ailerons intervene to stabilize the aircraft's attitude. After the wings are fully deployed, a period of time elapses, which is time T3. ， At this moment, the flight control system sends a command to the wing-locking servo, which then locks the wing. Step Nine: After the wings are fully deployed, the aircraft enters the unpowered gliding phase. At this time, the aircraft is in a high-speed nose-down dive. It is controlled by pulling up during the unpowered dive. After a period of time, the attitude stabilizes. The moment of attitude stabilization is T3. If the aircraft's altitude is lower than the minimum parachute opening altitude, the parachute is opened for landing. If the aircraft's altitude is not lower than the minimum parachute opening altitude after the attitude stabilizes, proceed to Step Ten. Step 10: At time T3, the flight control sends an engine start command and determines whether the engine has started successfully. The time when the engine starts successfully is recorded as T4. If the engine is not detected to have started, it is restarted after an interval of 2 seconds. If the engine fails to start 3 times and the altitude of the aircraft is below the minimum parachute opening altitude, the aircraft will deploy its parachute and land. If the altitude is above the minimum parachute opening altitude, the aircraft will continue to glide without power. During the glide, the ground station will start the engine. If the engine fails to start and the altitude of the aircraft is below the minimum parachute opening altitude, the aircraft will deploy its parachute and land. Step 11: After the engine starts, exit the air-based launch procedure and the aircraft begins the mission phase; Step 12: Perform the normal flight path mission through the conventional control law unit in the air-launched flight control module until the parachute is recovered.

2. The launch method for an airborne mounted folding-wing UAV as described in claim 1, characterized in that: In step six, the folded tail fin attitude control method is used, with the folded tail fin participating in the attitude control. A 5° V-tail control surface stroke is allocated to control pitch, and the control logic is the normal control logic; a 10° V-tail control surface stroke controls roll, using inner-loop PID control, with the target control variable being the UAV target roll angle Φ. a With the current roll angle Φ of the drone c The difference between them is due to the target roll angle Φ of the UAV at this stage. a The angle is always 0°, therefore the target control variable is also the UAV roll angle Φ. c The control quantity is processed using a PID controller to ultimately obtain the V-tail differential steering quantity δ. r1 The control law is: in, These are the control parameters for the gain circuit. For the control parameters of the integral element, This represents the gain of the differential element.

3. The launch method for an airborne mounted folding-wing UAV as described in claim 1, characterized in that: In step nine, the pitch channel control method used for the unpowered dive-pull-up control is pitch-airspeed closed-loop control. The various physical quantities and control parameters in the unpowered dive-pull-up control stage are as follows: Pull-up phase flag temp: temp displays 2 for pull-up preparation; 1 for pull-up in progress; and 0 for pull-up completion. Pitch control airspeed proportional coefficient The default value is 2; Pitch control airspeed integral coefficient Default is 0.1; The initial value of the drone's speed, V0, measured at the moment the control is pulled up; The speed V of the drone measured during the pull-up phase; Given the final value of the target velocity V c The cruising speed is 50 m / s. Given an initial velocity bias ΔV; Exponential transition time constant T: defaults to 2s; The target speed V during the pull-up control phase tar A function that changes exponentially with time T, with an initial time of V0 + ΔV and a final time of V. c ; A PI controller is used to stabilize the airspeed of the UAV. The control law is as follows: The elevator value δ of the UAV during the pull-up control phase is obtained. e_pull It is used to control speed stability and pitch angle stability.

4. The launch method for an airborne mounted folding-wing UAV as described in claim 1, characterized in that: In step nine, the attitude stability judgment logic is as follows: simultaneously satisfying that the current slave aircraft pitch is within ±12 degrees, roll angle is within ±10 degrees, and current slave aircraft airspeed is within the set airspeed value within ±3.

5. The launch method for an airborne mounted folding-wing UAV as described in claim 1, characterized in that: The design incorporates protective logic during the launch and recovery phases of the sub-unit to ensure the safety of air-based launches. This includes: (1) Engine start failure protection If the engine fails to start after three attempts, the engine start failure protection logic is activated. At this point, the drone stops attempting to start the engine and begins unpowered gliding towards its home point, which is located at the initial drop point but at an altitude of 150m. Upon reaching the home point, the drone begins circling it, awaiting further instructions from the ground station. During this protection period, manual ignition commands can be issued via the ground station, or the drone can be switched to semi-autonomous mode and manually controlled remotely. If the engine still fails to start successfully, and the drone falls below the altitude protection line during circling and descent, altitude protection is triggered, and the drone deploys its parachute for descent. (2) Forced pull-up When the drone's altitude drops to 800m during the pull-up control process, and the flight control determines that the drone's attitude has not yet reached stability, a forced pull-up is initiated, the engine starts, and the pitch channel control law of normal flight mode is switched, with a target pitch angle of 0°. (3) Attitude protection If the drone pitches more than 70° or rolls more than 90° during normal flight and fails to recover within 4 seconds, the drone's attitude protection will be triggered. The flight controller will assume that the drone is out of control at this time and will issue a parachute deployment command. (4) High level of protection When the drone's altitude drops below 100m, altitude protection is triggered, and the drone deploys its parachute for a landing.

6. The launch method for an airborne mounted folding-wing UAV as described in claim 1, characterized in that: The sequential logic is as follows: Timing 1: When detecting the disconnection signal, the ADC detects that the voltage has become 0V and continuously detects the disconnection signal for 200ms to prevent false judgments caused by voltage instability; If the disengagement signal is continuously detected for 200ms, record the current time of the disengagement detection as T0. At this time, activate the control system and adjust the aircraft's attitude in the folded state through the tail rudder. Timing 2: After the wings are deployed and the separation is detected, 300ms after the separation is T1. At this time, the flight controller sends a wing lock servo command, the servo is released, and the wings begin to rotate and deploy. T1 = T0 + 300ms. Timing 3: After locking the wings and engaging the servos, time T2 occurs 1300ms later. At this time, the wings are fully deployed, and T2 = T1 + 1300ms. After the wings are fully deployed, time T3 occurs 500ms later. ， At that moment, T3 ， =T2+500ms, the flight controller sends a command to the wing-lock servo, and the servo locks the wing; Sequence 4: After the wings are fully deployed, the aircraft enters the unpowered gliding phase. After Nms, the attitude is considered stable, N = 5s to 20s. At this time, if the altitude is lower than the minimum parachute deployment altitude, the parachute will be deployed for landing. Timing 5: At time T3, the flight controller sends an engine start command, T3 = T2 + Nms; after 3.3s, it checks whether the engine has started successfully, i.e., the speed is >2000rpm for 3 seconds to prevent false judgments caused by the wind blowing the propeller blades at high speed and instantaneous rotation. The start-up completion time is recorded as T4; if it is detected that the engine has not started, it will restart after an interval of 2s. Timing 6: The moment the engine starts successfully and the aircraft's attitude stabilizes, the mission phase begins. If the engine starts successfully on the first attempt, then T4 = T3 + 3300ms; otherwise, T4 is calculated based on the actual number of ignitions and the time taken.

7. A control system for air-launched folding-wing unmanned aerial vehicles (UAVs), the system implementing the launch method according to any one of claims 1 to 6, characterized in that: Includes a folding-wing V-tail UAV, a carrier aircraft, a flight control and navigation module, a ground station module, an air-launched flight control module, and hooks and hoops; The mother aircraft is equipped with a hook on its belly, and the daughter aircraft has two longitudinally arranged hooks on its upper fuselage. On the ground, the hooks and hooks are connected, locking the mother and daughter aircraft. Upon reaching the air-launch site, the mother aircraft sends a drop command to the daughter aircraft's flight control system, simultaneously releasing the hook on the mother aircraft's belly and launching the daughter aircraft in a projectile-like motion. A flight control and navigation module is installed on the daughter aircraft for communication between the daughter and mother aircraft, and between the daughter aircraft and the ground station. The daughter aircraft carries an air-launch flight control module to ensure successful air-launch and stable, normal flight. The ground station module is responsible for communication with the flight control system, sending command signals, and monitoring the aircraft's status. The air-launch flight control module includes the launch control section for the air-launched UAV, the launch timing logic section, the launch process control law section, the conventional control law section, and the protection logic section.

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