104 results about "Level flight" patented technology

The invention discloses a flight trainee attention and error behavior detection method based on a multi-modal physiological signal, and the method comprises the steps: obtaining the electrocardiogram data, eye movement data and electroencephalogram data of a flight trainee in a flight training process, and subjective evaluation data; based on subjective evaluation data, dividing flight student attention distribution conditions into a high grade, a middle grade and a low grade by adopting a clustering algorithm; performing feature optimization processing on the physiological data, and constructing a physiological feature attention level flight behavior associated data set; constructing an attention distribution level judgment model and a flight error behavior detection and early warning model; and outputting an attention level judgment result and a flight error behavior detection and early warning result. According to the method, association mapping of the multi-modal physiological signals and flight performance is realized, an attention grading standard for a flight training specific scene is established, and conversion from passive response to active early warning is realized.

A vertical take-off and landing aircraft which uses fixed rotors for both VTOL and forward flight operations. Some of the rotors are forward canted, while some of the rotors are forward facing. The forward canted rotors are tilted forward slightly from vertical and provide some forward propulsion during horizontal flight. The forward facing rotors may be fixed in a forward facing configuration. Forward swept wings allow for rotor placement which brings efficiency during hover operations.

A flight mode transitioning device, method and system configured to facilitate transitioning between vertical flight mode and horizontal flight mode in a vertical take-off and landing (VTOL) aircraft, the device comprising at least one actuator mediated member capable of limiting the movement the aircraft wing so as to enable both forward fight and vertical flight.

1. Name of the product in this design: Graphical User Interface for Adding Flight Information to Electronic Devices. 2. Intended use of this design: for use in an electronic device. 3. The key design feature of this product is its graphical user interface. 4. The image or photo that best illustrates the design points: Design 1 Interface Change State Diagram. 5. Other views are omitted as they are standard design. 6. Design 1 is designated as the basic design. 7. Purpose of the graphical user interface: This interface is used to add flight information in city-level flight service systems, such as operator information or aircraft information. Design 1: The main view is the initial interface. Figure 1 shows the interface that appears after clicking the "+" icon in the upper right corner of the main view of Design 1. Design 2's interaction is consistent with Design 1.

The invention discloses a fixed wing patrol aircraft flight attitude control method, and belongs to the technical field of unmanned aerial vehicles. Aiming at the technical contradiction that high-speed flight maneuverability and weapon shooting stability cannot be compatible, the method achieves breakthrough through the following innovative schemes that a four-wing propeller collaborative framework is designed, and each propeller is provided with a horizontal linear driver to achieve front-back telescopic displacement in the axial direction of a fuselage; in the horizontal flight stage, based on the feedback of a gyro stabilizer, pitching / rolling double-axis attitude accurate control is realized through the cooperative adjustment of propeller position stretching and rotation speed difference; triggering a layered compensation strategy by an ignition event, and synchronously adjusting the horizontal position and the rotating speed of a propeller according to the recoil force direction; and the tail vane intervenes in fine adjustment only when the yaw is out of tolerance. By combining the seamless conversion design of the vertical take-off and landing stage, the three technical problems of height loss without initial-speed deployment, low aerodynamic efficiency of high-speed flight and insufficient shooting precision are solved.

The invention provides a method and a device for judging the flight stage of a transport aircraft conforming to the operation expectation of a pilot, and the method comprises the steps: determining a guided flight stage of a current period according to a preset period, a working mode of an autopilot, a preselected height, a horizontal flight plan and a vertical flight plan; the guided flight stage comprises climbing CLM, cruising CRS and descending DES; the guide flight stage is used for generating a vertical guide instruction of the flight management system; and determining a man-machine interface flight stage according to the working mode of the autopilot, the preselected height, the horizontal flight plan and the vertical flight plan based on the judgment result of the guided flight stage. The method accords with the operation intention of the unit, enhances the vertical situation awareness capability of the unit, and provides a precondition for the output of the vertical guide instruction combination.

The present application relates to the technical field of aircraft operating mechanism, in particular to a drive-control integrated aircraft side stick operating mechanism, which comprises a box body, a controller and a driver, a roll transmission shaft is rotatably supported in the box body, a U-shaped decoupling mounting seat is fixedly connected to one end of the roll transmission shaft, a pitch shaft is horizontally supported in the U-shaped decoupling mounting seat, a base is fixedly sleeved on the pitch shaft, a handle connecting rod and a handle assembly are sequentially connected to the upper end of the base, a roll motor drives a rocker arm I through a rocker arm II and a transmission connecting rod I, the rocker arm I is hinged to the roll transmission shaft, a pitch motor drives a fixed head through a rocker arm III and a transmission connecting rod II, and the fixed head is fixedly arranged at the end of the pitch shaft. The present application provides a drive-control integrated aircraft side stick operating mechanism with high integration, ultra-small volume, double-channel synchronous force sensing, microsecond-level real-time bus communication and plug-and-play, which is particularly suitable for the demand of high-fidelity, high-reliability and small-sized operating device of high-level flight simulator.

This invention discloses a blown-wing vertical takeoff and landing (VTOL) fixed-wing aircraft and its control method, belonging to the field of aircraft technology. It includes a fuselage, main wing, retractable wing, tail-mounted tilting electric duct, and a control system. The fuselage has a built-in pressurized duct and an air intake at the front. An array of nozzles is arranged on the upper surface of the main wing and connected to the pressurized duct. The retractable wing is mounted on the trailing edge of the main wing. Symmetrically arranged tail-mounted electric ducts capable of vertical / horizontal tilting are located at the tail. The entire aircraft has no exposed rotating blades. During VTOL, the retractable wing deploys, and the pressurized duct supplies air to the array of nozzles, generating primary lift through the Coanda effect. The retractable wing guides air to generate secondary lift, and the tail-mounted duct provides vertical thrust for auxiliary lift and attitude adjustment. During level flight, the retractable wing retracts, and the tail-mounted duct provides horizontal rearward cruise thrust. This invention achieves smooth switching between VTOL, transitional flight, and level flight cruise through its control method, making it suitable for emergency rescue and low-altitude operations in complex terrain scenarios.

The application belongs to the technical field of overall design of vertical take-off and landing aircraft, and particularly relates to a single-blade rotor type vertical take-off and landing fixed-wing aircraft power layout structure. The main body of the power layout structure is a rotor with only one blade, one side of the rotor rotating shaft is a load cabin for loading and counterweight, and the other side is a fixed pitch rotor blade, and the rotor blade can swing; the rotor blade is provided with a driving propeller for driving the rotor to rotate or generating a level flight thrust. The single-blade rotor is a main component for generating lift in vertical take-off (rotor mode) and cruising flight (fixed-wing mode). The single-blade rotor system provided by the application can be used as a power system of a vertical take-off and landing aircraft, the power layout can consider cruising level flight efficiency and vertical take-off and landing hovering efficiency, and static balance and dynamic balance can be realized in a steady state hovering state.

The invention discloses a fan blade inspection local dynamic path planning method and system based on an unmanned aerial vehicle, and relates to the technical field of unmanned aerial vehicles, and the method comprises the following steps: S1, a template generation stage: based on a blade three-dimensional model, generating a standard inspection template, the standard routing inspection template defines a global optimal routing inspection sequence covering the surface of the blade and a starting point of the global optimal routing inspection sequence; s2, a path mapping stage: mapping the standard inspection template to an actual blade, and generating a three-dimensional space path which keeps a safe distance from the curved surface of the blade based on a preset rule; and S3, an online execution stage: the unmanned aerial vehicle executes the three-dimensional space path. The method solves the problem that the conventional unmanned aerial vehicle path planning method does not fully consider the physical limitation that the unmanned aerial vehicle must keep approximately horizontal flight in the inspection process, so that the unmanned aerial vehicle is forced to frequently adjust the attitude, and the inspection efficiency is greatly reduced.

Symmetrical balance of the helicopter is achieved by selecting a symmetrically-designed novel piston type engine, a fixed horizontal umbrella wing is arranged on the top, and when the helicopter flies normally and horizontally, linear resistance is small; the damping effect of avoiding too high speed is achieved during up-and-down rising and falling. The helicopter is completely balanced in a static state, and when the helicopter flies, the thrust formed by airflow entering the upper portion of the main rotor to the small louver fan wings with a certain inclination angle on the upper portion of the main rotor and the thrust formed by airflow flowing out of the lower portion of the main rotor to the balance wings with the adjustable inclination angle are utilized. And the resultant force equivalently counteracts the counter-acting force of the engine by adjusting the inclination angle, so that the aircraft is stable in direction. A novel piston type engine is adopted, due to the unique structural design, the rotating speed and variable work of the engine can be accurately adjusted in a high-precision mode, and a power shaft of the engine vertically operates. A small-diameter turbofan'main rotor 'with a fixed attack angle is adopted, so that conversion of horizontal power into vertical power through a bevel gear is omitted. The aircraft flies at a low speed through forward or backward force generated by downwash airflow of the main rotor to the angle-adjustable driving wings on the lower portion of the main rotor.

The application belongs to the field of aerospace technology, and particularly relates to a vertical take-off and landing horizontal flight combined power unmanned aerial vehicle, which comprises a fuselage, main wings, rotors, horizontal tail wings and movable vertical tail wings; the main wings are symmetrically connected to the middle part of the fuselage, and the rotors are symmetrically connected to the main wings on the two sides; the horizontal tail wings are symmetrically connected to the tail part of the fuselage, and the movable vertical tail wings are connected to the upper part of the tail part of the fuselage; a pressure measuring device is arranged on the movable vertical tail wings; the pressure measuring device collects the pressure on the two sides of the movable vertical tail wings to a flight control system, and the flight control system calculates the pressure difference between the left and right sides to control the deflection of the movable vertical tail wings. The cruising efficiency of the unmanned aerial vehicle is improved, the driving device and structure thereof are relatively simple and light compared with the full-motion vertical tail wing. Meanwhile, the unmanned aerial vehicle also has the cruising efficiency of a fixed-wing aircraft and the low requirement for a take-off and landing site of a rotor vertical take-off and landing aircraft.

The application discloses a heavy-load long-endurance tilting rotor unmanned aerial vehicle. The unmanned aerial vehicle comprises a wing frame for supporting and tilting, unmanned wings installed on symmetrical two sides of the wing frame, a cargo hold installed on the lower part of the wing frame and used for loading goods, and an unmanned aerial vehicle hydraulic system installed on the cargo hold and used for controlling the unmanned wings. The unmanned aerial vehicle can realize vertical take-off and landing, hovering and fast level flight, is easy to control the moving direction and has strong environmental adaptability. The unmanned aerial vehicle has simple structure, low take-off requirement, small wind resistance, fast level flight speed, low energy consumption and long endurance, so that the voyage and load can be increased. The wing frame overall tilting system has low structural complexity, airflow disturbance is reduced, and the safety of flight mode conversion can be improved.

The invention provides an unmanned aerial vehicle airborne antenna layout optimization method, and relates to the technical field of unmanned aerial vehicle communication, and the method comprises the steps: obtaining an attitude angle time sequence of typical attitudes including level flight, climbing, descending, turning, hovering and the like according to a task demand, discretizing a pitch angle and a roll angle into a plurality of attitude units, and carrying out the optimization of the attitude units; counting the duration of each attitude unit in combination with historical data, and calculating the comprehensive weight of each attitude unit; a plurality of alternative antenna layout schemes are generated under the condition that geometric, engineering and dynamic constraint conditions are met, communication performance evaluation values under all attitude units are obtained through electromagnetic simulation, and the comprehensive communication performance of the schemes is obtained through weighting according to the weights of the attitude units; and finally, with maximization of comprehensive communication performance and minimization of engineering complexity as targets, multi-target sorting is performed by using a TOPSIS method, and an optimal antenna layout scheme is determined in combination with a Pareto solution set, so that collaborative optimization of communication performance and engineering feasibility of the unmanned aerial vehicle in a dynamic flight task is realized.

This invention discloses a thin-walled, narrow-edge ducted fan that reduces crosswind interference. The fan comprises a duct shell, a multi-faceted frame, a trailing edge fairing, and a leading edge fairing. A drive motor is housed within the trailing edge fairing, and the bottom of the fairing is connected to the drive motor's base via screws. A motor rod extends from the drive motor, passing through the multi-faceted frame. The duct shell and leading edge fairing are fixedly connected by screws, and the motor rod is rotatably connected to the hub frame of the multi-faceted frame via bearings. The duct shell features a thin-walled, narrow-edge design, with the ratio of the duct shell's side length L to its inner diameter d being L / d ≤ 1 / 6. The interior of the duct shell is filled with foam, and the maximum wall thickness h of the outer wall satisfies either h = 1.5 mm, L ≤ 20 mm or 1.5 mm ≤ h ≤ 0.075L, L ≥ 20 mm. This design significantly reduces the weight of the shell compared to ordinary ducted fans, and when mounted on drones, it can reduce drag during high-speed level flight and mitigate crosswind effects during hovering and vertical takeoff and landing.

The invention discloses a compact omnidirectional vector wind-resistant four-rotor aircraft and a control method. The aircraft comprises a fuselage, four vector machine arms distributed on the front, back, left and right sides of the fuselage in a cross shape, a flight controller, thrust motors and propellers. The vector machine arm comprises a machine arm body, an axial tilting steering engine, an axial rotating assembly, a tangential tilting steering engine and a limiting motor base. The axial tilting steering engine drives the axial rotating assembly to continuously rotate around the axis of the aircraft arm within the range of 90 degrees left and right, the tangential tilting steering engine drives the limiting motor base to continuously rotate relative to the axial rotating assembly within the range of 40 degrees inside and outside, and the limiting motor base is used for fixing the thrust motor. After being combined with the propellers, the thrust motor can generate three-dimensional vector thrust along with the two-degree-of-freedom movement. The defect that position control and attitude control of a traditional multi-rotor aircraft are mutually coupled is overcome, and high-wind-resistance-level flight under the condition that spot hover in any attitude at any position in the air is kept and the fuselage is horizontal is achieved.

This invention relates to the field of unmanned aerial vehicle (UAV) control technology, specifically to an integrated gun-barrel top-attack UAV and its control method. The UAV includes a main body, wings, a recoilless gun mounted obliquely on the fuselage, a reconnaissance module, a flight control module, and a communication module. The control method includes: receiving waypoint instructions; the waypoint manager determining a flight sequence including takeoff, landing, hovering, level flight, and mode transitions based on preset decision logic; executing the sequence and generating control commands through a flight mode controller; and achieving attitude stabilization through an attitude controller. This invention achieves deep integration of the UAV platform and weapon, possessing rapid deployment, flexible multi-mode flight, intelligent route optimization, and precise top-attack capabilities, significantly improving the UAV's combat effectiveness and survivability in complex battlefield environments.

The invention provides an electric variable-pitch air propeller system suitable for a tilting unmanned aerial vehicle and a design method. The electric variable-pitch air propeller system comprises a propeller root assembly, a propeller hub and a variable-pitch mechanism. The pitch changing mechanism and the propeller root assembly are installed in the propeller hub and are in transmission connection, the steering engine is connected with the pitch changing mechanism, the propeller blades are connected with the propeller root assembly, and the steering engine drives the pitch changing mechanism to control the propeller root assembly to carry the propeller blades to rotate. The defects of an existing design method are overcome, multi-objective optimization design of the propeller is completed by coupling multiple disciplines based on an intelligent optimization algorithm, and the tilting propeller is compatible with a vertical lifting working condition and a flat flying working condition. A propeller pitch changing mechanism suitable for the tilting transition process of the unmanned aerial vehicle and the multi-working-condition pneumatic pushing and twisting characteristic is reasonably designed on the basis of the electric steering engine, the pitch angle of the blades is adjusted according to the actual working operation requirement of the propeller, and it is guaranteed that the propeller is in the working state with the optimal pneumatic performance under all the using working conditions.

The application discloses a method for determining the scale ratio of a simplified airplane scale model for airplane lightning partition test, which comprises the following steps: simplifying an airplane model, processing different scale simplified airplane scale models, and performing a simplified lightning attachment test on the simplified airplane scale models, calculating the attachment probability of lightning attachment points to vertical tails and back of the simplified airplane scale models in a positive polarity discharge state in a level flight attitude, and judging the coincidence of the lightning partition test result obtained in a specific scale test with the result of the initial lightning attachment point of a real airplane, so as to determine the best scale ratio of the scale model used in the lightning partition test. The application can accurately determine a suitable scale ratio, and provides favorable support for quickly judging the coincidence of the lightning partition test result obtained in a specific scale test with the result of the initial lightning attachment point of a real airplane.

The present application relates to a drone receiving system (100) comprising a receiving device (200) for receiving a drone (300), the receiving device (200) being designed to be arranged on a structure (400) formed almost vertically, and a drone (300), the drone (300) comprising a holding mechanism (310), wherein the holding mechanism (310) is designed to arrange the drone (300) on the receiving device (200). The present application also relates to a method for landing (100) a drone (300) using a drone receiving system (100) according to one of the preceding embodiments, the method having the following steps: placing the holding mechanism (310) of the drone (300) on a target surface (230) while the drone is in a horizontal flight state; converting a locking unit (210) from an unlocked position to a locked position; and pivoting the drone (300) from the horizontal flight state to a vertical position on the structure (400) formed almost vertically.

The application discloses a kind of aircraft level flight acceleration and deceleration aerodynamic parameter identification and its performance evaluation method, comprising: S1, based on the mass m of aircraft, thrust T and aerodynamic resistance D establishes the dynamics model of aircraft level flight acceleration and deceleration, determines the flight parameter required for aerodynamic parameter identification;S2, pre-processing is carried out to flight parameter;S3, construct aerodynamic derivative influence matrix and resistance coefficient vector;S4, based on aerodynamic derivative influence matrix and resistance coefficient vector, combined with bayesian identification module, obtains aerodynamic resistance coefficient;S5, based on the aerodynamic resistance coefficient, combined with dynamics model, evaluates flight performance.The application carries out sampling rate correction, outlier rejection and filtering processing to the flight parameter required for aerodynamic parameter identification, improves the accuracy of flight parameter, and then uses bayesian parameter identification method to obtain aerodynamic resistance, calculates the flight performance of aircraft level flight acceleration and deceleration, and improves the accuracy of flight performance calculation.

The application discloses a fan blade inspection local dynamic path planning method and system based on a UAV, relates to the technical field of unmanned aerial vehicles, and comprises the following steps: S1, a template generation stage: generating a standard inspection template based on a three-dimensional model of a blade, wherein the standard inspection template defines a globally optimal inspection sequence covering the surface of the blade and a starting point thereof; S2, a path mapping stage: mapping the standard inspection template to an actual blade and generating a three-dimensional space path maintaining a safe distance from the curved surface of the blade based on a preset rule; and S3, an online execution stage: a UAV executing the three-dimensional space path.The application solves the problem that existing UAV path planning methods do not fully consider the physical limitation that the UAV must maintain approximate horizontal flight during the inspection process, thereby causing the UAV to be forced to frequently adjust the attitude and substantially reducing the inspection efficiency.

This invention relates to the field of aircraft technology, specifically to a vertical takeoff and landing (VTOL) aircraft and its control method. The VTOL aircraft includes a tailless flying wing fuselage. Multiple drag rudders are mounted on the trailing edges of the left and right wing surfaces of the fuselage. These drag rudders can deflect vertically, generating differential drag through asymmetric deflection to provide yaw control torque for the fuselage. A tiltrotor quadcopter system is also mounted on the fuselage. The left and right propellers of the tiltrotor quadcopter system can differentially adjust their speed to further assist in providing yaw control torque. This invention significantly improves the aircraft's crosswind resistance during VTOL by eliminating the vertical takeoff and landing phase, while also achieving better aerodynamic efficiency during level flight. The integrated drag rudders and tiltrotor quadcopter system on the fuselage provide yaw control torque, thereby improving the aircraft's directional stability.

The system comprises an ultrasonic anemometer, a redundant GPS (Global Positioning System), a height accelerometer, a high-precision accelerometer with three course angles, a CCD (Charge Coupled Device) industrial camera, a camera optical center, a high-precision geographic identifier, a ground host, an airborne host and an unmanned aerial vehicle platform, a high-precision geographic identification database is input into the system; the unmanned aerial vehicle platform carries an airborne host, an ultrasonic anemometer, a redundant GPS, a height accelerometer, a high-precision accelerometer and a CCD industrial camera. The unmanned aerial vehicle climbs and flies to the position above the identification, a CCD industrial camera is used for photographing the position above the identification, two or more geographic identifications with the enough distance are shot, and the unmanned aerial vehicle flies horizontally as much as possible; iteratively solving the current optical center position of the unmanned aerial vehicle; solving the navigational speed of the unmanned aerial vehicle; solving the navigational speed of the ultrasonic anemometer; the wind speed is solved, and wind following sailing is used for reducing measurement errors. According to the invention, high-precision and high-temporal-spatial-resolution wind field measurement can be realized.

The invention provides a multi-rotor single aircraft and a vertical take-off and landing control method thereof, relates to the technical field of aircrafts, and has the advantages of multi-rotor vertical take-off and landing and fixed-wing efficient level flight. Comprising a fuselage frame, a main wing, a main power unit, a full-motion front canard, a front power unit, a servo actuating mechanism and a directional rear empennage. According to the invention, the full-motion front canard is driven to rotate through the servo actuating mechanism, so that the conversion of the aircraft between a vertical take-off and landing mode and a level flight mode is ingeniously realized; during vertical take-off and landing, the servo mechanism drives the canard wings to upwards deflect to the maximum angle, so that the planes of propellers of the front power units on the canard wings are turned to be horizontal, the front power units and the main power units on the main wings jointly provide vertical lift force, and helicopter type in-situ take-off and landing are achieved; when the aircraft is converted into level flight, the servo mechanisms drive the canards to deflect downwards, change the thrust direction and provide pneumatic control torque, so that the aircraft body is stably transited into a horizontal posture from a vertical posture, and the dual advantages of a multi-rotor aircraft and a fixed-wing aircraft are organically combined.

The invention provides a method for automatically identifying and testing objective subjects of a D-level flight simulator, which comprises the following steps of: processing and intercepting flight test data: respectively carrying out flight test data processing and intercepting work according to flight performance and flight characteristic subjects specified by a simulator objective test standard; making and generating a test configuration file, extracting test flight condition information from the intercepted flight test data, and making a configuration file; automatic identification testing: loading the configuration file and the intercepted flight test data into an existing flight simulation program of the flight simulator, respectively executing automatic identification testing according to subjects specified by a standard, and recording a test result; and automatically evaluating a test result, drawing an intercepted flight test data curve and a tolerance specified by a standard, drawing a test result curve, identifying a tolerance conforming part and an out-of-tolerance part, and giving conformity evaluation. The method ensures the legitimacy and validity of the test process and result, and meets the objective test requirements of the D-level analog machine specified by the standard.

An aircraft has a boom, a propulsion assembly coupled to a first end of the boom, and a first wing coupled to a second end of the boom. The first wing can be coupled to the boom by a (first) articulated joint. The propulsion assembly can be coupled to the boom by a (second) articulated joint. The articulated joints can provide various adjustments between the wing(s) and the boom (e.g., stowed, deployed, anhedral, dihedral, sweep, angle of attack, collective pitch control, the cyclic pitch control, etc.) and between the propulsion system and the boom (e.g., to move between vertical flight and horizontal flight). A payload can be connected to the boom with one of the articulated joints whereby the articulated joint adjusts a position of the payload during flight to maintain desired flight characteristics and stabilize the payload.

The present application relates to the field of aircraft technology, and discloses a kind of airship and airship transport method, the method comprises: from the ground of departure point, airship is lifted to the first preset height, and make airship realize force balance in vertical direction;Airship is lowered to the ground of departure point at the first preset height with rope, and cargo is bound with rope;Airship lifts cargo through rope, airship falls a preset distance to the second preset height along vertical direction under the action of gravity of cargo, make airship and cargo realize force balance in vertical direction;Airship keeps the second preset height level flight to destination, then unloads cargo.This airship transport method, when airship is in cargo state, the distance that needs to run in vertical direction is only this section of preset distance, greatly saves the energy of airship, thereby guarantees the endurance and loading capacity of airship;And this airship transport method does not need to use slide bubble take-off and landing mode, greatly reduces the requirement for landing site.

The invention provides a method and a device for determining a power supply scheme of a carrier rocket servo system. The method comprises the following steps: acquiring a first-stage flight stage parameter and a second-stage flight stage parameter of a carrier rocket servo system; according to the first-level flight stage parameter, obtaining a first-level flight stage electric quantity; obtaining secondary flight stage parameter electric quantity according to the secondary flight stage parameters; according to the first-level flight stage electric quantity and the second-level flight stage parameter electric quantity, obtaining a total electric quantity; obtaining an initial lithium battery capacity according to the total electric quantity and a preset proportion; and adjusting the initial lithium battery capacity according to the obtained battery health state parameters to obtain a target lithium battery capacity, and controlling a lithium battery with the target lithium battery capacity and a standby battery to supply power to the carrier rocket servo system. According to the invention, the accuracy of the lithium battery capacity is improved, the cost is reduced, and the power utilization safety of the carrier rocket servo system can be improved.

The invention relates to the technical field of aircraft aerodynamic model modeling, in particular to a flight test method for fixed-wing aircraft global dynamic aerodynamic modeling, which comprises the following steps: arranging a sensor on an aircraft, manually controlling the aircraft to fly, and optimizing control parameters according to sensor response data. Enabling the aircraft to autonomously fly by adopting the optimized control parameters; determining whether the autonomous flight of the aircraft meets an identification test condition based on the preset trajectory and the actual trajectory of the autonomous flight; when identification test conditions are met, the aircraft is controlled to carry out multiple times of alternate level flight and excitation flight, and sensor response data are obtained to serve as identification test data; determining an independent variable and a to-be-identified parameter, and based on the identification test data, establishing a nonlinear aerodynamic model by constructing a regression quantity, iteratively screening the regression quantity and a least square method; errors in pneumatic modeling can be reduced, and the test efficiency is improved.