A flight control and power control integrated system of a UAV

Abstract

The application discloses a flight control and power control integrated system of an unmanned plane, and relates to the field of unmanned plane control.The system comprises a first microprocessor, a second microprocessor, and a communication connection between the two for data interaction and a transient hard interrupt response mechanism.The first microprocessor runs a real-time operating system to perform attitude calculation and generate an expected throttle value.The second microprocessor runs an independent firmware to perform power device bottom layer state machine control based on power sensor parameters.The second microprocessor reversely transmits full parameter state data sets of the power device to the first microprocessor for adaptive attitude compensation, whole machine level space-time collaborative collection, and software task level cascade scheduling.The system reduces electromagnetic interference through function partition layout.The application realizes high integration of the flight control and power control, improves real-time performance, reliability and anti-interference capability, reduces volume and weight, and lowers power consumption and cost, and is suitable for power control of small and medium-sized unmanned planes.

Description

Technical Field

[0001] This invention relates to the field of unmanned aerial vehicle (UAV) control, specifically to an integrated system for flight control and power control of a UAV. Background Technology

[0002] Currently, the power control systems of small and medium-sized UAVs typically employ a control architecture with separate flight control computers and engine electronic control units (ECUs). The control process generally involves the flight control computer generating control commands such as engine throttle and target speed based on the flight mission and control strategy. These commands are then sent to the ECU via communication interfaces such as UART and CAN, outputting recognizable PWM, analog, or communication control signals. The ECU, combining its own collected engine speed, temperature, and pressure parameters, calculates control quantities such as fuel injection pulse width and ignition timing, and controls engine actuators such as fuel injectors and ignition coils through a power drive circuit. Simultaneously, engine operating status information is collected by the ECU and forwarded to the flight control computer via the flight control board for status monitoring and closed-loop control.

[0003] In the aforementioned serial control architecture of "flight control computer → ECU → engine", the flight control computer and ECU typically exist as independent hardware units. Data transmission between these units requires wiring harnesses, connectors, and multi-level communication interfaces. While this architecture can achieve engine control functions, it suffers from problems such as long control links, the need for multi-level forwarding of commands and status data, increased communication delays, and increased risk of anomalies. Furthermore, multiple independent control boards are configured with processors, power supplies, communication interfaces, and signal conditioning circuits, resulting in redundant hardware resource configurations and increasing system size, weight, power consumption, and installation and wiring complexity. In addition, the large number of inter-board connectors and wiring harnesses increases the risk of connection failures under vibration, shock, and complex environmental conditions, which is detrimental to the miniaturization, lightweighting, and high reliability design of UAV power control systems.

[0004] Therefore, it is necessary to integrate the control and interface functions of the flight control computer and the ECU onto the same control board, and to use multiple microcontrollers to complete tasks such as flight control, interface processing and engine control, thereby shortening the control link, reducing inter-board communication and external connections, reducing hardware redundancy, and improving system integration and operational reliability while retaining functional division of labor. Summary of the Invention

[0005] The purpose of this invention is to provide an integrated flight control and power control system for unmanned aerial vehicles (UAVs). This system integrates flight control, interface communication and signal processing, and power unit management functions onto a single control circuit board. Multiple microcontrollers are used to perform corresponding control tasks, enabling in-board coordination of flight control command processing, power unit control calculation, and status feedback. This reduces external inter-board communication, protocol conversion, and wiring harness connection links in the original system, shortens the power unit control and status feedback links, thereby reducing control transmission delay and improving the system's real-time response capability, thus solving the problems mentioned in the background art.

[0006] To achieve the above objectives, the present invention provides the following technical solution: an integrated system for flight control and power control of an unmanned aerial vehicle (UAV), comprising:

[0007] The first microprocessor runs a lightweight real-time operating system to perform flight control tasks and calculates the airframe attitude based on the onboard attitude sensor array to generate the desired throttle value.

[0008] The second microprocessor runs independent power unit management firmware, which is used to independently execute the underlying state machine control based on the real-time operating parameters of the power unit collected by the power sensor group. The first microprocessor and the second microprocessor are connected in communication.

[0009] The isolated intelligent power management module receives the airborne total power input and physically divides it into an electrically isolated flight control power supply branch and an ECU power supply branch on the printed circuit board, which respectively supply power to the first microprocessor and the second microprocessor. Through the electrical isolation of the physical domain, the reverse electromagnetic interference generated when the power load on the second microprocessor side is activated is blocked from coupling to the flight control power supply branch, so as to maintain the deterministic timing state of the first microprocessor in performing attitude calculation and task scheduling.

[0010] In the deterministic timing state, the first microprocessor writes the generated expected throttle value into a designated receive buffer on the second microprocessor side in real time;

[0011] The second microprocessor cyclically reads the instructions in the receiving buffer and performs multi-track comparison with the real-time parameters collected by the power sensor group to calculate the underlying power execution quantity to drive the power unit. At the same time, it transmits the calculated full parameter status dataset of the power unit back to the first microprocessor, which then performs adaptive flight attitude compensation based on the power loss status, thereby constructing a dynamic functional closed-loop interactive network of electromechanical control on a single printed circuit board.

[0012] As a preferred option, it also includes a hardware interruption interaction mechanism for emergencies, specifically including:

[0013] The first microprocessor and the second microprocessor are also configured with bidirectional hardware interrupt-triggered communication independent of their own communication.

[0014] When the first microprocessor senses a change in the body's attitude angular velocity that triggers a preset threshold, it bypasses the regular bus periodic communication flow and directly applies a thrust compensation hard interrupt signal to the second microprocessor through the bidirectional hardware interrupt trigger communication. This forces the second microprocessor to interrupt its current regular state machine execution sequence and complete the adjustment in the next working stroke immediately following the power unit, thereby achieving cross-domain transient function decoupling and interaction between attitude control and thrust output.

[0015] Preferably, a ground potential balance mechanism is established between the isolated intelligent power management module and the control drive circuit of the power device, specifically including:

[0016] The control drive circuit of the power unit adopts a low-side switching power direct drive structure. The power control square wave directly generated by the advanced timer of the second microprocessor is output to the on-board gate drive circuit for transient current amplification, so as to control the conduction and turn-off of the power MOSFET, thereby driving the power unit.

[0017] The source ground loop of the power MOSFET is single-point connected to the dedicated power ground of the ECU power supply branch. The isolated intelligent power management module limits the propagation of high current transient noise on the power ground to the signal ground of the flight control power supply branch through its internal isolation components, ensuring a low bit error rate for cross-core bus data transmission.

[0018] Preferably, a digital twin prediction interaction mechanism for the power system is established between the first microprocessor and the second microprocessor, specifically including:

[0019] The second microprocessor captures power parameters of the power unit in real time, such as micro-phase timing fluctuations, fuel pressure pulsation transients, and cylinder block temperature gradient data.

[0020] The first microprocessor acquires the power parameters, performs spatiotemporal alignment and data fusion of the power parameters with the current three-axis acceleration and airspeed of the machine body, and drives the power unit state degradation model in the background RTOS thread to output predictive maintenance instructions for spark plug carbon buildup or filter blockage before latent faults appear.

[0021] Preferably, a cross-domain security monitoring mechanism is established between the first microprocessor and the second microprocessor, specifically including:

[0022] The first microprocessor performs real-time window monitoring of the running heartbeat status word of the second microprocessor;

[0023] When it is determined that the second microprocessor is deadlocked due to a sudden electromagnetic pulse and the heartbeat timeout has not been refreshed, the first microprocessor instantly takes over the control of the whole machine and controls the flight control actuator to switch to a safe gliding attitude. Simultaneously, a millisecond-level low-level reset pulse is applied to the second microprocessor through the hardware reset pin directly connected to the board to force it to hard restart.

[0024] After restarting, the second microprocessor reads the residual parameters of the unpowered state and restores the phase of the power unit operating state machine before the restart within microseconds, thus restoring ignition and fuel injection control and completing cross-domain fault-tolerant self-healing.

[0025] Preferably, the input terminal of the isolated intelligent power management module is equipped with a bus voltage high-frequency sampling unit. A programmable electronic switch controlled by the second microprocessor is connected in series on the key inductive load circuit of the ECU power supply branch. When the bus voltage high-frequency sampling unit detects that the bus voltage drops below the safety threshold, the first microprocessor sends a highest priority energy degradation command to the second microprocessor. The second microprocessor responds to the command and instantly disconnects the programmable electronic switch to cut off the auxiliary power supply of the power unit, and routes the remaining electrical energy to the flight control power supply branch and the direct drive circuit used to maintain the operation of the power unit.

[0026] Preferably, the onboard attitude sensor group and the power sensor group are connected by a machine-wide spatiotemporal collaborative acquisition network, specifically including:

[0027] The inertial measurement unit, barometer, and magnetometer located in the flight control power supply branch interact with the first microprocessor at high speed via a short-distance SPI bus on the board.

[0028] The power parameter acquisition sensor of the power unit connected to the input terminal of the second microprocessor and the conditioning circuit performs local low-level signal capture;

[0029] The first microprocessor uses its on-chip hardware timer to periodically send clock synchronization timestamps to the second microprocessor, forcing the second microprocessor to mark and align the acquisition time of the power sensor data, ensuring that the full parameter status dataset uploaded to the first microprocessor has complete timing consistency with the attitude dataset on the flight control side.

[0030] Preferably, a software task-level cascaded scheduling and anti-collision interaction mechanism is also established between the first microprocessor and the second microprocessor, specifically including:

[0031] The first microprocessor does not participate in any real-time drive control of the underlying power unit, but is specifically responsible for the high-level attitude computing power iteration and communication with the external long-distance data link, and only updates the expected throttle value to the bus in the low-priority periodic thread of the RTOS.

[0032] The second microprocessor uses an interrupt preemption mechanism to prioritize responding to the refresh event of the desired throttle value, thereby achieving parallel decoupling of the deterministic execution of the flight control algorithm and the underlying real-time performance of the power unit control on the circuit board software architecture, avoiding cross-domain task scheduling conflicts.

[0033] Preferably, the printed circuit board is spatially divided into a flight control unit area far from the high-current load of the power unit and an ECU control unit area close to the physical interface of the actuator. The isolated intelligent power management module is located on the electrical boundary line between the two control unit areas. The first microprocessor and the onboard attitude sensor group are both located in the flight control unit area. The high-power MOSFET switch, gate drive chip and freewheeling protection network are all located in the ECU control unit area. By combining the distance attenuation of physical space with the channel isolation of the power module, the electromagnetic interference damage of power switch radiated noise to flight control sensitive signals is dynamically reduced.

[0034] As a preferred option, this also includes a dynamic coordination mechanism among the implementing agencies, specifically including:

[0035] The first microprocessor's general-purpose timer independently generates multiple high-frequency PWM signals, which are then amplified by the onboard level conversion stage and directly drive the aileron servo, elevator servo, and directional servo of the UAV to adjust its aerodynamic shape.

[0036] The high-resolution PWM signal is directly output by the local advanced timer of the second microprocessor. After power amplification, it drives the damper servo to adjust the air intake and coordinates with the power MOSFET directly driven power unit to control the chemical energy release power.

[0037] The changes in aerodynamic shape and the amplitude of chemical energy release power are globally unified and scheduled by the first microprocessor based on the uploaded full-parameter state dataset, forming a dynamic collaborative control chain of thrust and drag at the whole machine level.

[0038] Preferably, the power unit can be one of a piston engine, a turbojet engine, a turbofan engine, a turboshaft engine, and an electric ducted propulsion system.

[0039] In summary, the beneficial effects of this invention are:

[0040] 1. In the prior art, flight control computers and ECUs are usually set up as independent devices, each with its own circuit board, power supply interface, communication interface, connector, mounting structure and protective housing. The devices also need to be connected by wiring harnesses. The above-mentioned discrete setup not only occupies a large amount of internal installation space, but also increases the weight of the whole aircraft and the difficulty of layout.

[0041] This application integrates flight control, interface processing, and power unit management circuits onto a single control circuit board, transforming previously dispersed control units into a unified control assembly. By reducing the number of independent circuit boards, housings, inter-board connectors, and external wiring harnesses, the space occupied by the power control system and the weight of the equipment can be effectively reduced. For small and medium-sized UAVs with limited internal space and sensitive to payload weight, this effect helps optimize the layout of internal equipment, reduce the proportion of non-missionary payloads, and reserve more installation space for fuel, payloads, or other mission equipment.

[0042] 2. In existing discrete solutions, the power plant control commands generated by the flight control computer typically need to pass through the flight control output interface, ECU input parsing, and control calculations before finally forming fuel injection, ignition, or other power plant execution control signals; power plant status parameters also need to be transmitted back step by step in the opposite direction. Due to the multi-level communication, data forwarding, interface conversion, and signal conditioning processes, both control information and status feedback information may incur additional transmission and processing delays.

[0043] This application integrates the aforementioned functions onto a single control circuit board, enabling the flight control computer and the power plant management module to exchange control commands and status data via an on-board communication interface or a high-speed internal bus. This reduces data forwarding links between external boards and unnecessary protocol conversion processes. Consequently, the control path from the flight control computer generating the signal output to the power plant is shorter, as is the feedback path from the power plant status back to the flight control logic. This improves the power plant's response speed to changes in throttle, flight status, and control commands.

[0044] Especially in situations where UAVs are accelerating, climbing, adjusting their attitude, or experiencing rapid changes in power status, shorter control and feedback links help the flight control system obtain power status information in a timely manner and adjust the power unit control inputs accordingly, thereby improving the coordination between flight control and power output.

[0045] 3. In the existing architecture, flight control computer commands may need to pass through the transmit / receive buffers, protocol encapsulation, protocol parsing, level conversion, and forwarding scheduling of different control units in sequence. Different devices have their own independent task cycles, communication cycles, and processing sequences, which can easily lead to problems such as waiting, asynchrony, or inconsistent update cycles during the transmission of control data.

[0046] This application integrates flight control computer processing, interface processing, and power plant control processing onto a single circuit board, enabling control data to be transmitted and scheduled within a unified on-board architecture, thereby reducing uncertain delays caused by cross-device transmission. Because the communication cycles, data formats, task scheduling relationships, and fault handling strategies of each control module on the board can be designed uniformly, it is easier to ensure the timing consistency between throttle commands, speed information, power plant status information, and executed control outputs, thus improving the predictability and repeatability of the control process.

[0047] This effect not only improves the control accuracy of the power unit, but also facilitates subsequent control parameter calibration, fault reproduction, and system testing.

[0048] 4. During operation, UAVs may be affected by power unit vibration, airframe vibration, impact loads, and temperature changes. In existing discrete solutions, there are many external connectors, terminals, and wiring harnesses between the flight control computer and the ECU. Loose connectors, poor terminal contact, worn wiring harnesses, incorrect insertion, and momentary disconnections caused by vibration can all lead to loss of control commands, abnormal status feedback, and even power unit control failure.

[0049] This application integrates multiple control functions onto a single control circuit board, reducing the external signal connections between the flight control computer and the ECU. This allows some control signals and status data that previously required external wiring harnesses to be transmitted within the board. The reduced number of external connection points lowers the probability of mechanical failures caused by connectors, wiring harnesses, and mounting structures, thus improving the system's connection reliability in environments subject to continuous vibration and shock.

[0050] In addition, reducing the number of wire harnesses simplifies internal wiring and reduces the risks of wire harness tangling, wear, assembly compression, and mis-insertion or omission during maintenance.

[0051] 5. In existing discrete solutions, power unit control signals may need to be transmitted through long inter-board wiring harnesses. The UAV contains multiple electrical devices, including ignition, fuel injection, power conversion, motor or servo drive, and data communication equipment. Some of these devices may generate strong pulse interference or electromagnetic noise during operation. Long external signal harnesses are more susceptible to interference coupling from nearby power lines or pulse signals, thus affecting the stability of control and acquisition signals.

[0052] This application places flight control computer interface processing and power plant management-related functions on the same circuit board, transforming the transmission path of critical control signals and status feedback signals from long-distance inter-board transmission to short-distance intra-board transmission, thereby reducing the interference of the external environment on the signal link. Simultaneously, on the same circuit board, a holistic layout and routing design can be implemented for control circuits, sampling circuits, and power drive circuits. This includes functional zoning, rationally setting ground plane return paths, isolating high-noise drive areas from sensitive signal areas, and filtering and shielding critical signals.

[0053] Therefore, this application can not only reduce the risk of interference coupling caused by long wire harnesses, but also transform electromagnetic compatibility design from passive coordination between multiple independent devices to unified design and unified control within the circuit board scope, which is conducive to improving the stability of control signals and reducing the possibility of the power control system interfering with navigation, communication and other airborne electronic equipment.

[0054] 6. When the existing flight control computer and ECU are set up separately, each device usually needs to be configured with its own power supply circuit, interface protection circuit, communication transceiver circuit, mounting structure, connectors, and housing, etc., resulting in some functions being redundantly configured. In addition, multiple circuit boards need to be processed, mounted, tested, assembled, and wired together separately, resulting in a large number of material types, a long assembly process, and high production management complexity.

[0055] By adopting a circuit board integrated structure, this application allows for the overall planning of resources such as power input, interface protection, communication connections, mounting structures, and housings, while meeting functional isolation and reliability requirements. This reduces unnecessary duplicate components and structures. Furthermore, the reduced number of independent boards and external connections correspondingly decreases the number of PCBs, housings, connectors, wire harnesses, and assembly steps required.

[0056] Therefore, this application is beneficial to reducing material costs, processing costs, assembly costs and subsequent maintenance costs, and is also beneficial to improving the consistency and assembly efficiency of products in mass production.

[0057] 7. In the existing multi-device discrete structure, the flight control computer and ECU, as well as the wiring harnesses between them, need to be installed, fixed, and connected separately during system installation. During commissioning, the power supply status, communication status, interface configuration, and signal matching relationships between multiple devices also need to be checked separately. Once a fault occurs, it is necessary to check each segment of the multi-level control link, making the fault location process quite complex.

[0058] By integrating relevant control functions onto a single circuit board, this application simplifies the connections between external devices, reduces the number of components requiring fixing and connection during installation, and lowers the operational complexity for assembly personnel. Since the on-board communication and control logic can be uniformly designed, tested, and have its status information recorded on the same hardware platform, it is easier to obtain complete control process data and quickly determine whether the fault originates from input acquisition, control calculation, drive output, or an external actuator during system debugging, parameter calibration, and fault diagnosis.

[0059] Therefore, this application can shorten the system integration and debugging cycle, improve maintenance efficiency, and reduce the risk of failure caused by manual wiring errors or inconsistent interface configurations.

[0060] 8. In existing technologies, the flight control computer and the ECU operate separately, each possessing different information. The flight control computer primarily obtains flight attitude, navigation, and mission control information, while the ECU primarily obtains information such as engine speed, temperature, pressure, and operational status from the power plant. Because data exchange between the two requires intermediate equipment and communication links, under certain abnormal operating conditions, problems may arise such as untimely acquisition of status information or uncoordinated fault handling.

[0061] This application integrates flight control processing, interface processing, and power plant management functions onto a single circuit board, enabling more timely sharing of flight status information and power plant operating status information within the board. This facilitates the establishment of a unified status monitoring and fault handling mechanism. For example, when the power plant experiences abnormal engine speed, temperature, pressure, fuel injection drive, or communication, the control system can promptly take appropriate protective measures based on the current flight status, such as limiting throttle, switching control states, recording fault information, or outputting alarms.

[0062] This technology can improve the timeliness of handling abnormal states in the UAV power system and help prevent power unit abnormalities from further affecting flight safety.

[0063] 9. In existing discrete solutions, the flight control computer and ECU typically only transmit limited throttle commands and engine status parameters of the power unit. Furthermore, the coupling between power unit control and flight control is limited due to the influence of communication cycles and data forwarding processes.

[0064] By adopting a monolithic integrated structure, this application allows the flight control module to obtain operational parameters from the power unit, such as engine speed, temperature, pressure, and execution status, more promptly. The power unit management module can also obtain information such as throttle demand, flight phase, and control mode more quickly. Based on this information sharing, coordinated control between engine output and flight control requirements can be further achieved. For example, appropriate power control strategies can be adopted in different flight phases such as takeoff, climb, cruise, descent, or anomaly handling, or power commands can be limited and protected based on the power unit status.

[0065] Therefore, this application not only enables the integrated replacement of existing engine control links, but also provides a hardware foundation for the subsequent realization of higher-level flight control and power management collaborative strategies.

[0066] 10. Although this application integrates multiple control functions onto the same circuit board, flight control, interface processing, and power plant management tasks can still be handled separately by multiple microcontrollers or multiple functional areas, thus maintaining relative functional independence while achieving physical integration. Compared to simply centralizing all functions onto a single processor, this structure facilitates separate software development, testing and verification, fault isolation, and interface expansion for different functions.

[0067] Therefore, this application can achieve the advantages of circuit board integration in terms of size, weight, connectivity and cost, while maintaining clear boundaries between various control functions. This facilitates functional configuration and adaptation according to different engine models, different flight control requirements or different UAV platforms, thereby improving the product's versatility and engineering promotion value. Attached Figure Description

[0068] To more clearly illustrate the technical solutions in the embodiments of the invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0069] Figure 1 This is a schematic diagram of the microcontroller layout framework in an integrated flight control and power control system for an unmanned aerial vehicle (UAV) according to the present invention.

[0070] Figure 2 This is a schematic diagram of the process framework structure of an integrated flight control and power control system for an unmanned aerial vehicle (UAV) according to the present invention.

[0071] Figure 3 This is a schematic diagram of the instruction flow of an integrated flight control and power control system for an unmanned aerial vehicle (UAV) according to the present invention.

[0072] Figure 4 This is a schematic diagram of the engine control process of an integrated flight control and power control system for an unmanned aerial vehicle (UAV) according to the present invention.

[0073] Figure 5 This is a schematic diagram of the first part of the flight control circuit of an integrated flight control and power control system for an unmanned aerial vehicle (UAV) according to the present invention.

[0074] Figure 6 This is a schematic diagram of the second part of the flight control circuit of an integrated flight control and power control system for an unmanned aerial vehicle (UAV) according to the present invention.

[0075] Figure 7 This is a schematic diagram of the third part of the flight control circuit of an integrated flight control and power control system for an unmanned aerial vehicle (UAV) according to the present invention.

[0076] Figure 8 This is a schematic diagram of the fourth circuit of the flight control system of an integrated flight control and power control system for an unmanned aerial vehicle (UAV) according to the present invention.

[0077] Figure 9 This is a schematic diagram of the first part of the ECU control circuit of the integrated flight control and power control system for a drone according to the present invention.

[0078] Figure 10 This is a schematic diagram of the first part of the ECU control circuit of an integrated flight control and power control system for a drone according to the present invention. Detailed Implementation

[0079] The present invention will now be described in further detail with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without creative effort are within the scope of protection of the present invention. These drawings are simplified schematic diagrams, which are only used to illustrate the basic structure of the present invention in a schematic manner, and therefore only show the components related to the present invention.

[0080] To facilitate understanding of the present invention, a more complete description of the invention will be given below with reference to the accompanying drawings, which illustrate several embodiments of the invention. However, the invention can be implemented in many different forms and is not limited to the embodiments described herein. Rather, these embodiments are provided so that the disclosure of the invention will be more thorough and complete.

[0081] All features disclosed in this specification, or all steps in all disclosed methods or processes, may be combined in any way, except for mutually exclusive features and / or steps.

[0082] Any feature disclosed in this specification (including any appended claims, abstract, and drawings) may be replaced by other equivalent or similar features, unless specifically stated otherwise. That is, unless specifically stated otherwise, each feature is merely one example of a series of equivalent or similar features.

[0083] In this invention, unless otherwise explicitly specified and limited, the terms "installation," "connection," "linking," and "fixing," etc., should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral part; they can refer to a mechanical connection, a direct connection, or an indirect connection through an intermediate medium; they can refer to the internal communication of at least two elements or the interaction relationship of at least two elements, unless otherwise explicitly limited. Those skilled in the art can understand the specific meaning of the above terms in this invention according to the specific circumstances.

[0084] Please see Figure 1 - Figure 10 This invention provides an embodiment of an integrated flight control and power control system for unmanned aerial vehicles (UAVs). Its core lies in the use of two high-performance 32-bit microcontrollers: one is the computing core, managing the flight control algorithm and sensor data fusion; the other is used for power unit control, receiving control commands from the computing core and collecting and feeding back real-time status information of the power unit to the computing core. This solution deeply integrates the traditional distributed flight control system and the power unit control module in both software and hardware, placing them on the same PCB board, directly driving the power unit and collecting sensor signals. (See attached diagram.) Figure 1 .

[0085] It should be noted that the power unit can be one of the following: piston engine, turbojet engine, turbofan engine, gas turbine engine, turboshaft engine, and electric ducted propulsion. In this embodiment, a piston engine is used as the power unit.

[0086] Specifically, the system mainly includes a first microprocessor, a second microprocessor, and an isolated intelligent power management module. Both the first and second microprocessors use 32-bit ARM Cortex-M series chips with high clock speeds (e.g., ≥168MHz).

[0087] The first microprocessor (flight control computing side): Internally running a lightweight real-time operating system (RTOS), it serves as the high-level computing hub of the system. It does not directly participate in the microsecond-level actions such as fuel injection and ignition at the lower level. Instead, it utilizes the thread scheduling mechanism of the RTOS to perform three-dimensional attitude calculation of the aircraft and navigation path planning based on the onboard attitude sensor group (inertial measurement unit, barometer, magnetometer), and finally generates the desired throttle value. At the same time, the general-purpose timer of the first microprocessor independently generates multiple high-frequency PWM signals, which are amplified by the onboard level conversion stage and directly drive the external aileron servo, elevator servo, and directional servo to adjust the aerodynamic shape of the UAV.

[0088] The second microprocessor (power unit control side): As the underlying power execution center, it runs a completely independent power unit management firmware. This firmware is based on hardware interrupts and state machines and does not rely on RTOS, thereby ensuring microsecond-level real-time response capability. It receives the expected throttle command from the first microprocessor and independently calculates the target fuel injection frequency, throttle opening degree and other power execution quantities in a closed loop by combining the real-time parameters collected by the power sensor group.

[0089] To ensure the stable operation of the system under the dual-core board architecture, this embodiment establishes a rigorous dynamic interaction relationship. First, the airborne main power supply, such as a 6S / 12S lithium battery or generator output, is connected to the isolated intelligent power management module. This module is divided into an electrically isolated flight control power supply branch and an ECU power supply branch at the PCB physical level. When the second microprocessor drives large inductive loads such as the ignition coil and fuel pump, the hundreds of volts of reverse electromotive force and high-frequency surge generated are completely blocked by the isolated DC-DC and high-speed optocoupler inside the isolation module, thereby providing an extremely clean power supply environment for the first microprocessor and maintaining the absolute determinism of its attitude calculation timing.

[0090] In this deterministic state, the first microprocessor calculates the desired throttle value, and the second microprocessor, using an interrupt preemption mechanism, prioritizes the throttle refresh to complete the underlying power output. Subsequently, the second microprocessor transmits a dataset containing full power parameter status, including engine speed, cylinder temperature, and actual injection pulse width, back to the first microprocessor. The first microprocessor then performs adaptive flight attitude compensation based on this power loss status, such as pre-adjusting the elevator surfaces to prevent stall when the engine speed drops, thus completing a complete dynamic closed loop of "sensing-computing-driving-feedback-compensation" on the circuit board.

[0091] It should be noted that existing drones experience a phase difference between attitude adjustment commands and thrust commands when encountering sudden strong crosswinds, resulting in asynchrony due to communication delays. To address this issue, this embodiment adds an independent bidirectional hardware interrupt-triggered communication in addition to the communication connection between the first and second microprocessors.

[0092] When the first microprocessor senses through the gyroscope that the rate of change of the body's attitude angular velocity exceeds the preset safety threshold, it will immediately determine that the regular periodic bus communication, which is usually refreshed once every tens of milliseconds, can no longer meet the transient response requirements. At this time, the first microprocessor directly pulls low or high the hardware interrupt trigger and sends a high-priority hardware interrupt signal to the second microprocessor, while placing the transient thrust compensation factor into the designated area.

[0093] Once the second microprocessor captures the hardware interrupt, it immediately suspends the current state machine's normal sequence and forcibly intervenes in the current or immediately following crankshaft working stroke, typically within milliseconds or even hundreds of microseconds. This can manifest as directly increasing the current single injection pulse width or advancing the ignition angle. Through this hardware-level decoupling interaction, microsecond-level absolute spatiotemporal synchronization between attitude control surface deflection and engine thrust is achieved.

[0094] It is worth mentioning that, in order to completely solve the spatial electromagnetic compatibility (EMC) problem caused by instantaneous high currents of tens of amperes on the circuit board, this embodiment has carried out functional interaction design in PCB wiring and grounding structure.

[0095] The PCB board is physically divided into a flight control area and an ECU control area. An isolated intelligent power management module is deployed across the electrical boundary between these two areas. All sensitive devices are placed in the flight control area, while all high-power MOSFET switches, gate drive chips, and freewheeling diode networks are placed in the ECU control area.

[0096] In terms of drive interaction, the engine drive adopts a low-side switching structure: the 3.3V square wave signal output by the advanced timer of the second microprocessor enters the gate drive chip, and after transient amplification, drives the gate of the power MOSFET; the drain of the MOSFET is connected to the high-impedance fuel injector, and the source is connected to the power ground.

[0097] Most importantly, the source ground loops of all power MOSFETs converge at a single point to the dedicated power ground (PGND) of the ECU power supply branch. This power ground is completely disconnected from the signal ground (SGND) of the flight control area in terms of physical copper pouring, and is only connected through optoelectronic / magnetic isolation devices in the isolation module. As a result, when the fuel injector or igniter is switched on and off at high frequency, the return transient noise of tens of amperes can only be discharged within PGND and is completely limited, thus ensuring zero bit error rate for high-frequency communication across the core bus.

[0098] In addition, to achieve advanced predictive maintenance of UAVs, this embodiment constructs a whole-machine-level spatiotemporal collaborative acquisition network and a high-frequency data fusion mechanism.

[0099] Sensors such as the IMU on the flight control side interact directly with the first microprocessor via short-range SPI; while the underlying signals such as the crankshaft Hall sensor, PT100 temperature sensor, and oil pressure sensor on the engine side are connected to the second microprocessor. Alternatively, the underlying signals such as the speed sensor (such as magnetoresistive or Hall type, used to measure the N1 / N2 rotor speed), exhaust temperature sensor (such as K-type thermocouple), and fuel / oil pressure sensor on the gas turbine engine side are connected to the second microprocessor.

[0100] To ensure that the two sets of data are comparable on the same timeline, the first microprocessor uses its hardware timer to periodically broadcast a high-precision "clock synchronization timestamp" to the second microprocessor via the bus. The second microprocessor automatically adds this timestamp to the underlying data when capturing the crankshaft's raw pulses or ADC sampling.

[0101] In practice, the second microprocessor captures the microscopic phase difference when the crankshaft passes the sensor at a high sampling rate of several kilohertz (kHz), reflecting power parameters such as instantaneous angular acceleration fluctuations and pressure pulsation details in the fuel passage. The first microprocessor acquires the power parameters and fuses the high-frequency power waveform with the time-stamped triaxial acceleration. For example, if it is found that the instantaneous angular acceleration of the crankshaft is consistently low during a specific stroke in a complete cycle, accompanied by slight oil pressure fluctuations, the degradation model in the first microprocessor can determine that the spark plug of a certain cylinder is carbon deposited or the fuel filter is slightly clogged, thereby issuing a predictive maintenance warning to the ground station before complete engine shutdown occurs.

[0102] If a gas turbine engine is used: when the instantaneous angular acceleration of the turbine speed is detected to have periodic high-frequency fluctuations within a specific speed range, and the outlet pressure pulsation frequency of the fuel metering unit is abnormal and accompanied by an exhaust gas temperature (EGT) rise with a slope deviating from the baseline curve, the degradation model in the first microprocessor can determine that there is coking in the combustion chamber nozzle or that the fuel filter bypass valve has begun to stick, thereby predicting a decrease in compressor efficiency or partial blockage of the turbine blade cooling channel; if the fused triaxial acceleration detects an increase in the sideband energy of the blade passing frequency (BPF) in the casing vibration spectrum, the model can further distinguish between combustion instability and rotor imbalance, thereby issuing a graded predictive maintenance warning to the ground station before severe performance deterioration or in-flight shutdown occurs.

[0103] This embodiment further illustrates the system's deep protection mechanism when encountering extreme external environments, such as lightning strikes causing system crashes or bus power outages.

[0104] Under strong electromagnetic pulse interference, the second microprocessor may experience program overrun or deadlock. To address this, the first microprocessor continuously monitors the "heartbeat status word" emitted by the second microprocessor via the bus. If the heartbeat is not refreshed within the set window period, the first microprocessor immediately takes over the entire system, controlling the servo to enter a "safe gliding attitude." Immediately afterwards, the first microprocessor outputs a millisecond-level low-level pulse through a control line directly connected to the hardware reset pin (RESET) of the second microprocessor, forcing the second microprocessor to undergo a hardware restart. The second microprocessor restarts and initializes in just a few microseconds, immediately retrieving the context status words such as "speed, temperature, and throttle" saved before the crash, seamlessly restoring the interrupted control state machine, and achieving rapid self-healing before the mid-air crash.

[0105] The isolated intelligent power management module is equipped with a high-precision bus voltage monitoring unit at the front end. In the ECU power branch, including non-core load circuits such as the auxiliary redundant oil pump and cylinder preheating circuit, programmable electronic switches controlled by the second microprocessor are connected in series.

[0106] When the main generator fails or the battery is severely damaged, causing the bus voltage to plummet below the safety threshold, the first microprocessor issues a "highest level energy degradation command." Upon receiving the command, the second microprocessor immediately and forcibly disconnects the programmable electronic switch, cutting off power to all non-lethal inductive loads. The system then routes the remaining, extremely precious electrical energy entirely to the lifeline systems: the first microprocessor on the flight control side for attitude calculations, and the power MOSFETs on the engine side for basic fuel injection direct drive. Through this intelligent power allocation network, the controllable flight and safe emergency landing time of the UAV are extended to the extreme.

[0107] The above description is merely a specific embodiment of the invention, but the scope of protection of the invention is not limited thereto. Any changes or substitutions conceived without inventive effort should be included within the scope of protection of the invention. Therefore, the scope of protection of the invention should be determined by the scope defined in the claims.

Claims

1. An integrated system for flight control and power control of an unmanned aerial vehicle (UAV), characterized in that: include: The first microprocessor runs a lightweight real-time operating system to perform flight control tasks and calculates the airframe attitude based on the onboard attitude sensor array to generate the desired throttle value. The second microprocessor is used to independently execute the underlying state machine control based on the real-time operating parameters of the power unit collected by the power sensor group. The first microprocessor and the second microprocessor are connected in communication. The isolated intelligent power management module receives the airborne total power input and physically divides it into an electrically isolated flight control power supply branch and an ECU power supply branch on the printed circuit board, which respectively supply power to the first microprocessor and the second microprocessor. Through the electrical isolation of the physical domain, the reverse electromagnetic interference generated when the power load on the second microprocessor side is activated is blocked from coupling to the flight control power supply branch, so as to maintain the deterministic timing state of the first microprocessor in performing attitude calculation and task scheduling. In the deterministic timing state, the first microprocessor writes the generated expected throttle value into a designated receive buffer on the second microprocessor side in real time; The second microprocessor reads the instructions in the receiving buffer in a loop and performs multi-track comparison with the real-time parameters collected by the power sensor group to calculate the underlying power execution amount to drive the power unit. At the same time, the calculated full parameter status dataset of the power unit is transmitted back to the first microprocessor, which performs adaptive flight attitude compensation based on the power loss status, thereby constructing a dynamic functional closed-loop interaction on a single printed circuit board.

2. The integrated flight control and power control system for an unmanned aerial vehicle (UAV) according to claim 1, characterized in that: It also includes a hardware interruption interaction mechanism for emergencies, specifically including: The first microprocessor and the second microprocessor are also configured with bidirectional hardware interrupt-triggered communication independent of their own communication. When the first microprocessor senses a change in the body's attitude angular velocity that triggers a preset threshold, it bypasses the regular bus periodic communication flow and directly applies a thrust compensation hard interrupt signal to the second microprocessor through the bidirectional hardware interrupt trigger communication. This forces the second microprocessor to interrupt its current regular state machine execution sequence and complete the adjustment in the next working stroke immediately following the power unit, thereby achieving cross-domain transient function decoupling and interaction between attitude control and thrust output.

3. The integrated flight control and power control system for a UAV according to claim 2, characterized in that: A ground potential balance mechanism is established between the isolated intelligent power management module and the control drive circuit of the power device, specifically including: The control drive circuit of the power unit adopts a low-side switching power direct drive structure. The power control square wave directly generated by the advanced timer of the second microprocessor is output to the on-board gate drive circuit for transient current amplification, so as to control the conduction and turn-off of the power MOSFET, thereby driving the power unit. The source ground loop of the power MOSFET is single-point connected to the dedicated power ground of the ECU power supply branch. The isolated intelligent power management module limits the propagation of high current transient noise on the power ground to the signal ground of the flight control power supply branch through its internal isolation components, ensuring a low bit error rate for cross-core bus data transmission.

4. The integrated flight control and power control system for a UAV according to claim 3, characterized in that: A dynamic system digital twin prediction interaction mechanism is established between the first microprocessor and the second microprocessor, specifically including: The second microprocessor captures the power parameters of the power unit in real time and transmits them to the first microprocessor. The power parameters are spatiotemporally aligned and fused with the current three-axis acceleration and airspeed of the machine body. The power unit state degradation model is driven in the background RTOS thread to output predictive maintenance instructions for the power unit before latent faults appear.

5. The integrated flight control and power control system for a UAV according to claim 2, characterized in that: A cross-domain security monitoring mechanism is established between the first microprocessor and the second microprocessor, specifically including: The first microprocessor performs real-time window monitoring of the running heartbeat status word of the second microprocessor; When it is determined that the second microprocessor is deadlocked due to a sudden electromagnetic pulse and the heartbeat timeout has not been refreshed, the first microprocessor instantly takes over the control of the whole machine and controls the flight control actuator to switch to a safe gliding attitude. Simultaneously, a low-level reset pulse is applied to the second microprocessor through the hardware reset pin directly connected to the board to force it to hard restart. After restarting, the second microprocessor reads the residual parameters of the unpowered state, restores the phase of the power unit's operating state machine before the restart, and resumes control to complete cross-domain fault-tolerant self-healing.

6. The integrated flight control and power control system for a UAV according to claim 5, characterized in that: The input terminal of the isolated intelligent power management module is equipped with a bus voltage high-frequency sampling unit. A programmable electronic switch controlled by the second microprocessor is connected in series on the key inductive load circuit of the ECU power supply branch. When the bus voltage high-frequency sampling unit detects that the bus voltage drops below the safety threshold, the first microprocessor sends a highest priority energy degradation command to the second microprocessor. The second microprocessor responds to the command and instantly disconnects the programmable electronic switch to cut off the auxiliary power supply of the power unit, and routes the remaining electrical energy to the flight control power supply branch and the direct drive circuit used to maintain the operation of the power unit.

7. The integrated flight control and power control system for an unmanned aerial vehicle (UAV) according to claim 6, characterized in that: The onboard attitude sensor group and the power sensor group are connected by a system-wide spatiotemporal collaborative acquisition network, specifically including: The inertial measurement unit, barometer, and magnetometer located in the flight control power supply branch interact with the first microprocessor at high speed via a short-distance SPI bus on the board. The power parameter acquisition sensor of the power unit connected to the input terminal of the second microprocessor and the conditioning circuit performs local low-level signal capture; The first microprocessor uses its on-chip hardware timer to periodically send clock synchronization timestamps to the second microprocessor, forcing the second microprocessor to mark and align the acquisition time of the power sensor data, ensuring that the full parameter status dataset uploaded to the first microprocessor has complete timing consistency with the attitude dataset on the flight control side.

8. The integrated flight control and power control system for an unmanned aerial vehicle (UAV) according to claim 7, characterized in that: A software task-level cascade scheduling and anti-collision interaction mechanism is also established between the first microprocessor and the second microprocessor, specifically including: The first microprocessor does not participate in any real-time drive control of the underlying power unit, but is specifically responsible for the high-level attitude computing power iteration and communication with the external long-distance data link, and only updates the expected throttle value to the bus in the low-priority periodic thread of the RTOS. The second microprocessor uses an interrupt preemption mechanism to prioritize responding to the refresh event of the desired throttle value, thereby achieving parallel decoupling of the deterministic execution of the flight control algorithm and the underlying real-time performance of the power unit control on the circuit board software architecture, avoiding cross-domain task scheduling conflicts.

9. The integrated flight control and power control system for an unmanned aerial vehicle (UAV) according to claim 8, characterized in that: The printed circuit board is spatially divided into a flight control unit area far from the high-current load of the power unit and an ECU control unit area close to the physical interface of the actuator. The isolated intelligent power management module is located on the electrical boundary line between the two control unit areas. The first microprocessor and the onboard attitude sensor group are both located in the flight control unit area. The high-power MOSFET switch, gate drive chip and freewheeling protection network are all located in the ECU control unit area. By combining the distance attenuation of physical space with the channel isolation of the power module, the electromagnetic interference damage of power switch radiated noise to flight control sensitive signals is dynamically reduced.

10. The integrated flight control and power control system for an unmanned aerial vehicle (UAV) according to claim 9, characterized in that: It also includes a dynamic coordination mechanism established between the various implementing agencies, specifically including: The first microprocessor's general-purpose timer independently generates multiple high-frequency PWM signals, which are then amplified by the onboard level conversion stage and directly drive the aileron servo, elevator servo, and directional servo of the UAV to adjust its aerodynamic shape. The high-resolution PWM signal is directly output by the local advanced timer of the second microprocessor. After power amplification, it drives the damper servo to adjust the air intake and coordinates with the power MOSFET directly driven power unit to control the chemical energy release power. The changes in aerodynamic shape and the amplitude of chemical energy release power are globally unified and scheduled by the first microprocessor based on the uploaded full parameter state dataset, forming a dynamic collaborative control chain of thrust and drag at the whole machine level.

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