683 results about "Angle of attack" patented technology

An attitude display system for use in an aircraft has at least one display unit visible by a pilot within a cockpit. A symbology generator is configured for receiving input from at least one attitude sensor and generating symbology image data representative of a pear-shaped attitude indicator, distorted vertically to represent a pitch of the aircraft and rotated to represent a roll angle of the aircraft. The symbology generator may further be configured for adding a horizontal line that represents a horizon, a regular trapezoid representing a downward gravity vector at a base of the trapezoid, a pitch angle of the aircraft, split wings representing an angle of attack of the aircraft, and a stall warning in a sawtooth crack shape. A tail portion of a T-shape with the curved top edge is itself curved in a direction opposite that of an aircraft spin direction.

The invention relates to the technical field of flight control, in particular to a fixed-wing unmanned aerial vehicle attitude correction and precise flight control method fused with visual navigation, which comprises the following steps: extracting ORB feature points and constructing a dynamic feature point cloud; simultaneously acquiring angular velocity and acceleration data of the inertial measurement unit; performing tight coupling fusion processing on the dynamic feature point cloud and the inertial measurement data to generate real-time pose data; inputting the real-time pose data into an attitude correction parameter calculation module, and generating an attitude correction parameter through a dynamic weight distribution algorithm in combination with a preset flight path planning model and GPS positioning data; and a control surface deflection instruction and an accelerator control instruction are generated through the multi-constraint optimization model, and unmanned aerial vehicle flight attitude closed-loop correction is completed. According to the method, aerodynamic parameters such as the air speed, the wind field lateral disturbance and the attack angle are introduced to serve as coupling input, the control surface deflection amount and limiting conditions are dynamically adjusted, and self-adaptive modeling and control over the nonlinear flight characteristics are achieved.

This disclosure provides an adaptive STSMC hierarchical control method for depth tracking of an underwater vehicle. The method includes: using an adaptive line-of-sight guidance algorithm to estimate the vehicle's angle of attack in real time and obtain an expected pitch angle; employing a nonlinear interference observer based on a sliding mode surface error to detect unknown interference; and applying an STSMC controller with adaptive saturation compensation to determine the expected control rudder angle from the sliding mode surface error, the expected pitch angle, and the observed interference. The adaptive saturation compensation corrects the rudder angle based on the saturation deviation, which is determined by the previously calculated expected control rudder angle and the actual output rudder angle at the previous moment. By introducing an adaptive saturation compensator in the dynamic control layer, the issue of rudder angle saturation is mitigated, thereby enhancing both control performance and stability.

The invention provides a multi-airfoil combined axial flow fan blade design method and system based on CST parameterization, and relates to the technical field of axial flow fan blade optimizing.The multi-airfoil combined axial flow fan blade design method comprises the following steps that S1, a plurality of characteristic sections are arranged in the radial direction of a blade, and the airfoil corresponding to each characteristic section is determined; s2, geometric modeling is conducted on the airfoils of all the characteristic sections through a CST parameterization method, and description parameter sets corresponding to all the airfoils are obtained; s3, performing aerodynamic performance simulation on the airfoils through the description parameter group to determine the attack angle of each airfoil; s4, the airfoils are combined at the corresponding attack angles, smooth transition cambered surfaces are generated between the airfoils, and an initial blade model is obtained; and S5, optimizing the initial blade model to obtain a target blade model by adjusting the description parameter group of each airfoil by taking the maximization of the lift-drag ratio as a target. Compared with a traditional single arc thin plate blade profile design, the pneumatic performance of the blade is more optimized, and the problem that a single blade profile is difficult to cope with complex pneumatic conditions where different sections are located is effectively solved.

The invention provides a wind tunnel experiment-based wing high-attack-angle flow separation combined intelligent control method, which comprises the following steps of: arranging active flow control devices at 10% chord length and trailing edge flap of a wing to inhibit wing high-attack-angle flow separation; the active flow control device at the 10% chord length of the wing adopts direct jet flow, the jet flow direction is perpendicular to the chord line of the wing, three jet flow exciters are uniformly arranged along the spanwise direction, the active flow control device at the trailing edge flap adopts sweeping jet flow, the jet flow direction is parallel to the surface of the wing at the position, and 18 sweeping jet flow actuators are uniformly arranged along the spanwise direction; the jet flow of the two sets of flow control devices is adjusted through corresponding mass flow controllers, jet flow momentum distribution is dynamically optimized in combination with a reinforcement learning algorithm, the lift coefficient is remarkably increased, and energy consumption is reduced. The flow field state is fed back in real time through a pressure measuring system and a force measuring system, a closed-loop control system is constructed, and an execution action is output according to the flow field information around the wing to control an active flow control device arranged on the wing so as to inhibit flow separation on the surface of the wing, so that the wing gets rid of the stall state and the lift force is improved.

The invention belongs to the technical field of aircraft design, and particularly relates to a static pressure source position error correction method, which comprises the following steps: performing theoretical appearance static pressure source position error correction analysis based on current numerical simulation and flight test state points to obtain a static pressure source pressure coefficient Cp, a flight test static pressure source pressure coefficient Cp0 and an actually measured static pressure source pressure coefficient Cp1; static pressure source position error correction data optimization is carried out, a static pressure source position error correction relation Cp'during real flight of the theoretical appearance is obtained, and static pressure source position error correction is carried out; appearance measurement, CFD simulation and test flight data comprehensive application are realized; correcting the position error of the static pressure source into a function of Mach number, angle of attack and height; and not only can the requirement of the CCAR25 part for the air pressure height be met, but also the requirement of RVSM operation in AC-21-13 for the error of a height measurement system can be met. The method can be used for atmospheric data correction of transportation aircrafts, and atmospheric data correction of other aircrafts can be used for reference.

Disclosed in the present invention are a method and system for estimating an airspeed, an angle of attack, and a sideslip angle of an aircraft. The method comprises: modeling the full six-degree-of-freedom dynamics, the aerodynamics, a force, and a moment of an aircraft in a nonlinear state-space form; using an inertial measurement unit to measure the body angular rate and body acceleration of the aircraft; using a global navigation satellite system to measure a ground speed; using data of systems of the aircraft other than an air data system to determine air data measurement; and using a plurality of Kalman filters to estimate the airspeed and / or the angle of attack and / or the sideslip angle from all measured data. The method and the system provided by the present invention solve the problem in the prior art of inability to improve the safety and reliability by installing a redundant air data computer / sensor on a size-, weight-, and power-constrained aircraft, and the problem in the prior art of being not independent of a conventional air data system at the measurement principle level.

The invention provides a displacement vortex generator and a position regulation and control method, and belongs to the field of active flow control equipment. The VG comprises a belt, a belt guide rail, a belt bearing, a sliding block, a supporting column, a triangular VG body, a guide rail supporting block, a limiting groove, a body connecting plate and a motor. The belt is arranged on the belt guide rail and is driven by the motor; the sliding block is arranged on the belt, is in an inverted concave shape and is connected with the body connecting plate through the supporting column, and the VG bodies are fixed to the two ends of the body connecting plate. The belt guide rail is arranged on the guide rail supporting block, and the sliding block can penetrate through the limiting groove. The motor is controlled to drive the belt to drive the sliding block and the VG body to move, and accurate adjustment of the position of the vortex generator body is achieved. The position of the vortex generator is automatically adjusted according to changes of the incoming flow wind speed and the attack angle, the optimal pneumatic effect is kept under different working conditions, the pneumatic performance and the power generation efficiency of the wind turbine blade are remarkably improved, and the negative influence of the front edge roughness can be effectively weakened.

The present invention discloses a dynamic composite distortion simulator and a working method thereof in the technical field of aero-engine air inlet and engine matching. The simulator comprises a process air inlet, a spoiler column, a rotary valve and a high-speed motor. The spoiler column is installed inside the process air inlet, and the rotary valve is installed in the cavity of the spoiler column. One end of the rotary valve extending out of the spoiler column is connected to the high-speed motor. Dynamic distortion vortices are formed by utilizing the flow around the column and the rotary valve, and parameters such as the intensity and scale frequency of the vortex change with changes in the motor speed and the angle of attack and height of the spoiler column. This can effectively simulate the dynamic distortion flow field structure generated in the air inlet due to various reasons, overcome the shortcomings of existing simulators in simulating dynamic distortion, form a flow structure similar to the distortion flow field at the outlet of the real air inlet, so as to more realistically simulate the air inlet matching process, significantly improve the experimental efficiency and accuracy, and do not require an external air source.

The invention belongs to the technical field of aerodynamic parameter identification, and relates to an intelligent fitting optimization method for aerodynamic parameters of a fixed-wing aircraft based on direct lift force. According to the invention, a BP network is used to carry out fitting of a lift coefficient and an aerodynamic derivative of the lift coefficient to an attack angle on noiseless data. The input of the BP network comprises an attack angle, an elevator, a flap, a speed and a pitch angle speed, and is input into the BP network to obtain a prediction result when no noise influence exists, a loss function value is calculated, a gradient is calculated according to the loss function value, and then network parameters are updated according to a gradient descent algorithm to reduce the loss function value. When the loss function value is small, the difference between the prediction result of the network and the actual result is small, that is, the lift coefficient identification value output by the network is accurate in the BP network. The aerodynamic derivative obtained through calculation can be applied to a controller and has good control quality.

The intelligent bionic ornithopter with the body and the self-adaptive attack angle adjusting joints comprises an ornithopter body, wings and an empennage, the wings and the empennage are arranged on the ornithopter body, driving motors are symmetrically arranged on the flapping portions of the wings, and each driving motor is provided with a limiting joint; the limiting joint comprises an outer ring, an inner ring and an elastic cushion block, the outer ring is composed of a plane end and first protruding parts, the plane end and the driving motor are connected and rotate synchronously, the first protruding parts are located on the two sides of the plane end and provided with penetrating holes, the inner ring is a cylinder, and the inner ring is arranged between the first protruding parts and swings with the middle of the plane end as the center. The inner ring and the penetrating hole are matched and jointly sleeved and fixedly connected with the edge of the wing, a second protruding part is arranged on the inner ring, and elastic cushion blocks are arranged at the upper end and the lower end of the second protruding part. The invention aims to provide the bionic ornithopter which can fully simulate flapping wing flapping actions of flying organisms and can adaptively adjust the attack angle in the flapping process so as to improve the flapping lift force and reduce the resistance.

The invention relates to the technical field of aircraft control, and discloses a flight program planning method and device for an aircraft, equipment and a medium, and the method comprises the steps: obtaining a nominal flight program and a corresponding flight envelope function which is generated in a preset profile domain based on a process constraint condition and a terminal constraint condition of the aircraft; key nodes are selected in the speed domain of the nominal flight program, random quantities are overlaid on all calculation nodes based on a random walk strategy, and an overlaid result is generated; performing constraint correction on the superposition result through a flight envelope function to generate a corrected attack angle node value; and performing polynomial interpolation fitting on the corrected attack angle node value to generate a random walk flight program. According to the method, the random walk flight program with randomness and constraint conformity can be generated, so that the track unpredictability of the hypersonic gliding aircraft is enhanced, and diversified combat requirements of the aircraft such as high maneuverability and strong penetration are met.

The invention relates to the technical field of flight safety of civil aircrafts, in particular to a flight trajectory prediction method under the condition of air high-speed target hood perforation, which comprises the following steps: acquiring calculation coefficient parameter data of a target flight dynamics model, establishing a trajectory prediction framework based on a three-degree-of-freedom target flight dynamics model and target real-time flight data; the method comprises the following steps: performing numerical simulation on a flow field of a hood perforation area through fluid mechanics analysis software to obtain speed field and pressure field data, establishing a hood perforation flow field pneumatic analysis model by adopting a high-order polynomial fitting method, and calculating Mach number and pressure distribution after hood perforation; dynamically updating aerodynamic coefficient data and an aerodynamic parameter database; the speed and pressure distribution data are introduced into the three-degree-of-freedom target flight dynamics model for registration updating, a prediction track is obtained, the prediction precision is improved by 30%-50%, and especially under the condition of a large angle of attack, the prediction error is controlled to be within 5% from 20% or above.

The invention discloses a large flexible wing static aeroelasticity analysis method based on a physical information neural network, and belongs to the field of aircraft design. The method comprises the following steps: for the wing with the super-high aspect ratio, firstly, dispersing the wing into n spanwise sections based on a strip theory, and initializing a reference angle-of-attack parameter of each spanwise section; then, a physical information neural network (PINN) is built, unsteady flow field solving is carried out on each section, for the current ith iteration, the physical information neural network is used for sequentially traversing and extracting pressure distribution of the airfoil surface of each section, respective aerodynamic load distribution is calculated, and the generalized force of the airfoil is obtained; and solving a structural dynamics equation based on the generalized force and structural information of the wing, calculating the deformation of the large-deformation wing by using a modal rotation method, and finally, carrying out solving iterative circulation on aerodynamic force and structural deformation until the attack angle is converged, so as to obtain the actual lift force L of the wing. According to the method, the calculation precision and efficiency are remarkably improved, and different flight conditions and wing design can be adapted.

The invention discloses a fixed-wing aircraft large-angle-of-attack landing control method based on a dynamic reference trajectory, and belongs to the technical field of fixed-wing aircraft autonomous landing, and the method comprises the following steps: establishing a landing control strategy based on a kinematic model and an attitude angle conversion model in the fixed-wing aircraft autonomous landing process, designing two-stage annular auxiliary line radiuses of an air flight stage of the fixed-wing aircraft, designing an expected track yaw angle, an expected track inclination angle and an expected roll angle of the fixed-wing aircraft, solving the expected yaw angle and the expected pitch angle, and constructing a control law of an attack angle when the fixed-wing aircraft is about to land; aiming at a flight stage and a taxiing stage of fixed-wing aircraft landing, a proper angular velocity controller and an air speed controller are designed based on a dynamic reference trajectory, so that the pose of the fixed-wing aircraft and a large attack angle when the fixed-wing aircraft is about to land are controlled, the dangerous conditions of collision and stall are avoided, and the safety of the fixed-wing aircraft is improved. And the accuracy and safety of autonomous landing control are improved.

The present invention relates to a method and system for generating a real-time penetration trajectory of an aircraft. The present invention solves the penetration trajectory optimization problem by using a second-order cone programming method to obtain the penetration trajectory under different initial conditions, and after generating a data set, uses the generated data set to train a deep neural network to obtain a trained deep neural network model, and then uses the trained deep neural network model as a real-time controller of the angle of attack and the roll angle to quickly generate a penetration trajectory based on the current state parameters of the hypersonic gliding aircraft. In the actual application process, the penetration trajectory can be obtained by simply inputting the obtained current state parameters, so as to improve the solution efficiency of the convex optimization problem, thereby meeting the real-time requirements of the highly dynamic environment of the HGV.

Systems and methods for mechanically rotating an aircraft about its center-of-gravity (CG) are disclosed. The system can enable the rear, or main, landing gear to squat, while the nose landing gear raises to generate a positive angle of attack for the aircraft for takeoff or landing. The system can also enable the nose gear and main gear to return to a relatively level fuselage attitude for ground operations. The system can include one or more hydraulically linked hydraulic cylinders to control the overall height of the nose gear and the main gear. Because the hydraulic cylinders are linked, a change on the length of the nose cylinder generates a proportional, and opposite, change in the length of the main cylinder, and vice-versa. A method and control system for monitoring and controlling the relative positions of the nose gear and main gear is also disclosed.

The invention relates to the technical field of aircraft control, in particular to an aircraft airspeed attack angle estimation method based on augmented generalized Kalman filtering, which comprises the following steps: acquiring sensor data and correcting acceleration, calculating initial estimation of an airspeed attack angle, constructing an error matrix and scoring, extracting variation trend adjustment parameters, and calculating a state variation coefficient. And multi-source information is fused to output the combined airspeed angle of attack quantity. According to the method, through acceleration component body coordinate system projection and back projection correction, an airspeed initial estimation value is constructed in combination with a height difference value, a physical matching degree is enhanced, a slope score and trend change extraction are performed to establish a dynamic compensation mechanism, and the adaptive capacity to a non-stationary flight state is improved; the introduction of the pitch angle speed, the control surface deflection and the aerodynamic coefficient bias makes the state evolution more structural, a change coefficient is established through ratio calculation, the coupling of state prediction and maneuvering dynamic is enhanced, the weighted fusion of an observation residual error and a predicted value is realized, and the convergence capability of an estimated value to complex aerodynamic interference is improved.

The invention discloses a terminal course angle and time constraint iterative reentry guidance method, and belongs to the technical field of hypersonic flight vehicle gliding section trajectory planing.The method comprises the steps that firstly, a resistance acceleration profile is planned according to a reentry corridor, and the size of a heeling angle is obtained; a heeling angle symbol is obtained through a heeling reversal strategy based on an SA-SQP optimization method; and finally, on-line correction is carried out on the basis of the standard attack angle profile, the interference error of longitudinal and lateral guidance is weakened, the guidance precision is improved, and a three-dimensional trajectory considering time and terminal course angle constraints is finally obtained. According to the method, the three core modules of two-parameter resistance acceleration profile design, heeling inversion search based on the SA-SQP optimization method and attack angle profile online correction are combined, the coupling interference error between longitudinal guidance and lateral guidance is weakened, and therefore high-precision reentry guidance under multiple constraints is achieved. According to the method, the reentry guidance precision can be effectively improved, the task adaptability is high, and flight time and terminal course angle constraints can be strictly met.

The invention belongs to the field of embedded atmosphere data system fault identification and elimination calculation, and particularly relates to a flying wing aircraft seven-pressure-measuring-point embedded atmosphere data system fault identification and calculation method, which comprises the following steps: step 1, pressure measuring points are grouped: each group of pressure measuring points comprises nose cone lower surface pressure measuring points, left and right symmetrical pressure measuring points and upper and lower symmetrical pressure measuring points, the pressure measuring points are used for solving the attack angle alpha and the sideslip angle beta, one pressure measuring point is reused, the pressure measuring points are divided into four groups under different attack angles, and each group of pressure measuring points is composed of four pressure measuring points; step 2, identifying the fault type of the pressure measuring point and the fault point: according to the attack angles alpha 1, alpha 2, alpha 3 and alpha 4, the sideslip angles beta 1, beta 2, beta 3 and beta 4 obtained by resolving the four groups of pressure measuring points under different attack angles, and the pressure measuring point used for resolving the attack angle alpha and the sideslip angle beta, identifying the fault type of the pressure measuring point, and further identifying the fault point; and step 3, removing fault points, and calculating and outputting atmospheric parameters including an attack angle alpha and a sideslip angle beta.

The invention discloses a roll angle measuring method for a continuous variable roll test of a transient wind tunnel, and belongs to the technical field of wind tunnel test. Comprising the following steps: step S1, early-stage preparation: selecting two angle measurement sensors to carry out measurement work, and installing the two angle measurement sensors and an installation platform according to a preset angle; s2, carrying out field calibration operation; carrying out field calibration on the coefficient of the angle measurement sensor by adopting a fixed attack angle and variable roll angle mode; and S3, data interception, discretization and application during a step variable attack angle continuous variable roll angle test: in a wind tunnel test process, continuously collecting output voltage values of the two angle measurement sensors, calculating roll angles corresponding to the output voltage values, intercepting roll angle range values corresponding to the two angle measurement sensors according to needs, and calculating the roll angle range values of the two angle measurement sensors according to the intercepted roll angle range values. Splicing into a target measurement range; compared with a traditional step variable rolling and fixed rolling variable attack angle test mode, the scheme has the advantage that the test efficiency is greatly improved.

The invention relates to the field of wind driven generators, and particularly discloses a floating type vertical axis wind turbine multi-partition variable pitch control strategy which comprises the steps that S1, basic parameters and operating condition parameters of a unit are obtained; s2, establishing an aerodynamic model based on a double-disk multi-flow management theory and a superposition wind speed method; s3, measuring parameters through a sensor, and calculating a theoretical attack angle of the blade; s4, establishing a multi-partition pitch angle change function based on the theoretical attack angle; s5, constructing an NSGA-II multi-target optimization model, and taking maximization of an average power coefficient and minimization of a blade thrust standard deviation as targets; s6, substituting a function to obtain an optimal variable pitch control factor and an optimal pitch angle; and S7, the optimal pitch angle is converted into a variable pitch actuator instruction. The problems that an existing strategy is limited in partition, the wind direction influence is ignored, and optimal parameters for working condition adaptation are lacked are solved, the wind energy capture efficiency is improved, the load fluctuation is reduced, and the variable pitch control accuracy is improved.

The invention belongs to the technical field of hypersonic flight vehicle control, and particularly relates to an INDI-based flight and thrust integrated control method for a hypersonic flight vehicle. The method comprises the following specific steps: designing a vertical-direction-controlled flight-thrust integrated longitudinal dynamic model of the hypersonic aircraft, designing a trajectory guiding outer ring to realize height and speed tracking control, designing an attitude stabilizing outer ring to realize attack angle tracking control, constructing an attitude control inner ring to realize flight-thrust control of an acceleration level, and designing a flight-thrust model of the hypersonic aircraft. And designing an engine compensation layer to process dynamic delay of an actuator. Through the MIMO-INDI mechanism, the pneumatic / propulsion coupling effect is actively utilized, collaborative distribution of the fuel equivalence ratio of the elevator and the engine is achieved, and the comprehensive control performance of the aircraft is improved. The method is a flight / propulsion integrated control method and has a wide application prospect.

The invention discloses a wind turbine airfoil full-attack-angle pneumatic measurement and prediction method for a wind tunnel experiment in the field of wind tunnel experiments, and the method comprises the steps: 1) carrying out the modeling of a wind turbine airfoil, designing a wind turbine airfoil model of the wind tunnel experiment, and building a test platform; (2) checking the safety of equipment and an experimental environment; 3) opening a wind tunnel to obtain a set wind speed; 4) sequentially changing the attack angle, obtaining the pressure data of the wind turbine airfoil at the expected attack angle moment, and closing the wind tunnel after the measurement is finished; 5) adjusting the pitot tube to be right in front of the wind turbine airfoil, repeating the step 3), and collecting incoming flow pressure right in front of the wind turbine airfoil when the attack angles are 90 degrees and 270 degrees; (6) data processing is conducted, and correction coefficients related to the incoming flow wind speed are obtained; and 7) calculating a lift coefficient and a resistance coefficient, obtaining a resistance coefficient shrinkage ratio, and predicting the lift coefficient and the resistance coefficient of the wind turbine airfoil under the full-attack-angle working condition. The method can quickly predict the aerodynamic change condition of the wind turbine airfoil under the expected working condition.

The invention discloses an aircraft attack angle protection control method based on dynamic instruction correction, belongs to the field of flight control methods, and is used for solving the tracking control problem under the condition that the aircraft attack angle is limited. Firstly, based on an aircraft dynamics model, a backstepping method is utilized to design an adaptive controller to ensure stable tracking of a system, a virtual attack angle instruction part needs to be limited by adopting a saturation function, and a limited boundary value is provided by a subsequent protection system. And then, designing a prediction model representing correlation between the attack angle and the attack angle instruction and dynamics based on the designed closed-loop control system, reversely deducing a complex function relationship between the attack angle instruction and the system state and dynamics through the prediction model, and constructing an attack angle protection system. And finally, calculating an attack angle instruction boundary value through the prediction model and a system motion state trend under the current input, and correcting an original virtual attack angle instruction boundary so as to realize attack angle protection. According to the method, stable tracking of the system under attack angle constraint can be ensured, the residual attack angle in the envelope can be fully utilized, the maneuverability of the aircraft is improved, and the method has an engineering application reference value.

Principles described herein allow for wings to be optimized for aerodynamic and structural efficiency while reducing planform compromises required to address wing tip stall problems. All-moving outboard wings allow the outboard wing local angle of attack to be changed relative to the inner wing region to avoid tip stall at high aircraft angles of attack, such as that which is required for low-speed flight. During high-speed cruise flight, at low aircraft angle of attack, the outboard wing local angle of attack can be increased to achieve an ideal elliptical lift distribution.

The invention relates to the technical field of aircraft electric heating deicing, in particular to a pulse heating aircraft deicing method and system.The method comprises the following steps that the ice layer thickness, the ice layer type, the airspeed and the attack angle of each aircraft surface area are collected, the inherent pneumatic sensitivity coefficient of each area is called, the aerodynamic penalty index of each area is established, and the aerodynamic penalty index of each area is calculated; and according to the values of the aerodynamic penalty indexes of the areas, performing descending order arrangement to generate a deicing area priority sequence. Through multi-parameter joint evaluation of the ice layer thickness, the ice layer type, the airspeed and the incidence angle, the heating resource distribution sequence is optimized, energy waste is avoided, in the local heating execution stage, temperature change slope real-time monitoring is adopted to replace fixed time control, the actual thermal response behavior of the ice layer can be dynamically responded, and the real-time thermal response performance of the ice layer is improved. The initial phase change process of the ice crystal is identified through a phase change starting slope threshold value, and then a medium-power pulse control mode is switched, so that a steady-state regulation heating path is formed, and the phase change control continuity is strengthened.

The invention provides a high-speed aircraft online real-time energy management guidance method and system and a medium, and relates to the technical field of high-speed aircraft guidance, and the method comprises the steps: determining a flight sub-stage of an energy management stage; determining an energy corridor, and planning a low-energy nominal longitudinal track based on a Gaussian pseudo-spectral method; when the aircraft enters the energy management section, judging whether the current actual energy exceeds the energy upper boundary of the energy corridor, adding the S turning section if the current actual energy exceeds the energy upper boundary, otherwise, entering the capture section; predicting the to-fly distance of the aircraft, and obtaining an expected energy state by interpolating the energy profile of the longitudinal trajectory; obtaining an attack angle and a speed brake guidance instruction through a longitudinal guidance law according to the deviation between the current energy state and the expected energy state; outputting a heeling angle guidance instruction according to the flight sub-stage through a transverse lateral guidance law; and circularly executing until the aircraft successfully enters the pre-landing section. According to the strategy, online energy management is realized, and the aircraft is ensured to safely enter the landing section.

The present invention discloses a device for measuring the lift of an airfoil, which is used to be placed in a wind tunnel. The device includes: two first columns, which are arranged vertically and oppositely; two upper flat plates, which are arranged horizontally and are respectively fixed at the upper ends of the corresponding first columns; two lower flat plates, which are arranged horizontally and are respectively fixed at the lower ends of the corresponding first columns; two linear guides, which are arranged vertically and oppositely, and whose two ends are respectively fixed on the corresponding upper flat plates and lower flat plates; two bearings, which are arranged oppositely and are respectively sleeved on the two linear guides. After the airfoil receives the upward or downward lift in the wind tunnel, the airfoil drives the bearings to move upward or downward along the linear guides, thereby measuring the upward or downward displacement value of the airfoil and obtaining the magnitude of the lift of the airfoil. The present invention is simple, reliable and intuitive; the present invention does not require a wind tunnel balance, reducing the experimental cost; the present invention can change the angle of attack of the airfoil by setting a motor and can measure multiple groups of data.

A control system for rotating an airfoil along one axis that is offset from a perpendicular axis that varies the angle of attack (AOA), dihedral, and center of lift based on the position of the attachment to precisely maneuver an aircraft. This technique can eliminate the need for traditional control surfaces such as ailerons, elevators, and rudders. allowing for an aerodynamically advantaged shape.