Experimental method, device and medium for multi-body separation trajectory prediction under burst interference

Abstract

The application provides a test method, device and medium for multi-body separation trajectory prediction under sudden wind disturbance. The method according to the application comprises: initializing a test environment, determining a parallel mechanism configuration for driving six-degree-of-freedom motion of a secondary aircraft under a preset condition, and calibrating positioning accuracy and precision of a three-degree-of-freedom motion driving mechanism of a primary aircraft and a parallel mechanism of six-degree-of-freedom motion of the secondary aircraft; performing linkage test; performing test preparation work; driving the primary aircraft and the secondary aircraft to move according to the pose of the primary aircraft and the secondary aircraft at the next moment; determining a wind tunnel safety boundary, and giving coordinates of easy collision points of the secondary aircraft at each moment; calculating a safety margin according to the easy collision point coordinates of the safety boundary, predicting collision through the safety margin, and obtaining trajectory capture test data of all dangerous region separation states. The technical solution provided by the application can realize multi-body separation motion trajectory prediction under a changing wind field.

Description

Technical Field

[0001] This document relates to the technical field of multi-body separation trajectory prediction, and in particular to an experimental method, equipment and medium for multi-body separation trajectory prediction under gust interference. Background Art

[0002] Wind tunnel gust response technology uses an actuator to drive an airfoil structure within the wind tunnel, rotating it in the form of a mathematical function at a predetermined frequency, thereby generating a flow field with discrete gusts. Wind tunnel testing is crucial for gust response research, primarily through model testing to examine the effects of gusts and validate simulation results. Wind tunnel gust response testing typically employs two methods: discrete isolated gusts and continuous gusts. Discrete isolated gusts use high-pressure airflow to create localized gust conditions within the wind tunnel; continuous gusts use a mechanism to drive blades in a repetitive, regular motion, creating a continuous gust field. Discrete isolated gusts can generate significant loads, but their impact area is small, preventing the formation of a large gust field. Continuous gusts, on the other hand, deflect the wind tunnel airflow, resulting in a larger gust field with continuously and regularly varying wind speeds. Different gust response testing methods can be selected based on test requirements.

[0003] The Capture Trajectory Test (CTS) is a commonly used two-stage separation wind tunnel test method. Its advantages include high prediction accuracy, the ability to obtain test results that are essentially consistent with full-scale flight test data, and the ability to simulate complex external load separation conditions and special flight states, such as pitch, climb, or accelerated flight, through computer software. This allows for launches within certain aircraft envelope flight states, and the ability to directly generate external load separation trajectories during the test, allowing for timely assessment of separation characteristics. The test not only produces full-scale separation trajectories but also directly measures the aerodynamic loads on the external load at various measurement points along the separation trajectory, facilitating analysis and improvement of external load separation characteristics. Given these advantages, it is a mandatory aircraft / external load compatibility assessment and verification project prior to flight testing.

[0004] Conventional separation tests of aircraft in normal flow fields have become increasingly mature in China. However, in recent years, it has been discovered that even during the same separation state during field flight tests, the separated aircraft can have completely different positions due to atmospheric disturbances. This necessitates a wind tunnel capable of simulating trajectory capture tests under gust disturbances. Currently, there are no systems or methods in China capable of simulating this situation. Summary of the Invention

[0005] The embodiments of the present invention provide an experimental method, equipment and medium for predicting the trajectory of multi-body separation under gust interference, which solves the problem of possible collision when the two-stage aircraft are subjected to gust interference before, during and after separation, and the problem of dangerous blind spots in the separation of the two-stage aircraft; after the two-stage aircraft are safely separated, its displacement and attitude can no longer meet the requirements of the next flight mission.

[0006] According to an embodiment of the present invention, a test method for multi-body separation trajectory prediction under gust disturbance is provided, comprising:

[0007] According to an embodiment of the present invention, there is provided an electronic device, including:

[0008] processor; and,

[0009] A memory is arranged to store computer-executable instructions, which, when executed, cause the processor to perform the steps of the experimental method for multi-body separation trajectory prediction under gust disturbance as described above.

[0010] According to an embodiment of the present invention, a storage medium is provided for storing computer-executable instructions, which, when executed, implement the steps of the experimental method for predicting multi-body separation trajectories under gust disturbance as described above.

[0011] By adopting the embodiments of the present invention, the following beneficial effects are achieved:

[0012] (1) Existing wind tunnel gust response simulation technology cannot simulate the posture motion of the aircraft, let alone analyze the posture motion under gust interference when the first and second stage aircraft separate. This application realizes the trajectory prediction capability of multi-body separation under gust interference.

[0013] (2) Achieve integrated test technology capabilities for gust interference aerodynamic analysis - multi-body separation trajectory prediction - grid force measurement to obtain aerodynamic characteristics of strong interference dangerous areas - and design multi-body separation posture control laws. BRIEF DESCRIPTION OF THE DRAWINGS

[0014] In order to more clearly illustrate one or more embodiments of this specification or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are only some embodiments recorded in this specification. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0015] Figure 1 This is a flow chart of an experimental method for predicting multi-body separation trajectories under gust disturbances according to an embodiment of the present invention;

[0016] Figure 2A schematic diagram of a wind tunnel burst response driving mechanism of an embodiment of the present application;

[0017] Figure 3 A structural diagram of an LQ-080 pulsating pressure sensor of an embodiment of the present application;

[0018] Figure 4 A schematic diagram of positions of installing pulsating pressure sensors on a primary aircraft and a secondary aircraft of an embodiment of the present application;

[0019] Figure 5 A schematic diagram of a three-degree-of-freedom angular displacement driving mechanism of a primary aircraft of an embodiment of the present application;

[0020] Figure 6 A schematic diagram of trajectory lines in a separated state of a dangerous area of an embodiment of the present application;

[0021] Figure 7 A schematic diagram of a grid force measurement station site of an embodiment of the present application;

[0022] Figure 8 A pretest flow of capturing a multi-body separation trajectory under burst interference based on an embodiment of the present application. DETAILED DESCRIPTION

[0023] In order for those skilled in the art to better understand the technical solutions in the one or more embodiments of the present specification, the technical solutions in the one or more embodiments of the present specification will be described clearly and completely in conjunction with the drawings in the one or more embodiments of the present specification. Obviously, the described embodiments are only a part of the embodiments of the present specification, rather than all the embodiments. Based on the one or more embodiments of the present specification, all other embodiments obtained by those skilled in the art without creative labor should belong to the protection scope of the present document.

[0024] Method embodiments

[0025] According to an embodiment of the present application, an experimental method for predicting a multi-body separation trajectory under burst interference is provided, Figure 1 A flowchart of the experimental method for predicting a multi-body separation trajectory under burst interference of an embodiment of the present application, according to Figure 1 The experimental method for predicting a multi-body separation trajectory under burst interference of an embodiment of the present application includes:

[0026] S1, initializing a test environment, including: determining wind tunnel model parameters of an aircraft in the test, determining a burst interference mathematical model and a corresponding motion response actuator, and determining monitoring flow field related parameters, wherein a preset pulsating pressure sensor is arranged in the aircraft;

[0027] Determining a wind tunnel model scale ratio and separation parameters of a primary aircraft and a secondary aircraft, including:

[0028] The ratio of the first-stage aircraft to the full-scale first-stage aircraft, the flight altitude, flight Mach number, and flight overload of the full-scale first-stage aircraft, the mass, inertia, and center of mass position of the first-stage aircraft, and the angle of attack, sideslip angle, and roll angle of the first-stage aircraft in maneuvering or level flight;

[0029] The ratio of the secondary aircraft model to the full-scale secondary aircraft, the flight speed, flight altitude, flight overload of the secondary aircraft, and the separation position of the secondary aircraft from the first aircraft, the mass, inertia and center of mass position of the secondary aircraft, the separation speed and separation angular velocity of the secondary aircraft, and the attitude angle of the secondary aircraft relative to the first aircraft

[0030] Determining the mathematical model of gust disturbance and the corresponding motion response actuator includes:

[0031] Gust interference mathematical models include: longitudinal slope gust model, step gust model, full-wavelength (1-consine) discrete gust model, half-wavelength (1-consine) discrete gust model, full-wavelength (1-consine) continuous gust model, and half-wavelength (1-consine) continuous gust model. The mathematical form of the (1-consine) gust model is shown in Formula 1:

[0032]

[0033] In Formula 1, s represents the flight distance, Uds represents the maximum gust speed, and H represents the gust gradient. For the specific gradient form, please refer to CCAR-25: PART C.25.645.

[0034] The driving mechanism of the motion response to gust disturbance is composed of cam mechanism, gear rack mechanism and crank rocker mechanism, all of which can realize swing motion. However, considering the aerodynamic characteristics, dynamic performance, economy and processability, the driving mechanism of gust response adopts the combination of crank connecting rod and quadrilateral connecting rod mechanism to realize the reciprocating swing of the cascade. Figure 1 As shown in the figure, the mechanism consists of a crank, a main connecting rod and a quadrilateral connecting rod mechanism, and the quadrilateral connecting rod mechanism consists of a rocker, a long connecting rod and a slave connecting rod. Figure 2 As shown;

[0035] The motion response actuator corresponding to gust disturbance is composed of a combination of crank-connecting rod-polygonal connecting rod mechanism, flywheel mechanism, synchronous belt mechanism and other mechanisms. It has the characteristics of high operating frequency and large blade swing amplitude. According to the mathematical model of gust disturbance, the blade swing changes according to a cosine law.

[0036] The blade cascade swing angle is shown in Formula 2:

[0037] θ = A·cos(2πft) Formula 2;

[0038] Therefore, the cascade swing angular velocity is shown as formula 3:

[0039]

[0040] The angular velocity is shown as formula 4:

[0041]

[0042] In formula 2-4, A represents the swing amplitude of the cascade mechanism, and f represents the frequency (0-15 Hz) of the cascade mechanism swing.

[0043] Selection of preset pulsating pressure sensor in aircraft:

[0044] LQ-080 series, using frequency 0-2 kHz, the pulsating pressure sensor appearance diagram is shown as Figure 3 The micro pulsating pressure sensor is a small disc shape, the disc diameter is 4.1 mm, the length is 9.6 mm, and the pressure range is selected as 50 psi.

[0045] The schematic diagram of the point position of the pulsating pressure sensor installed in the model is shown as Figure 4 After the pulsating pressure sensor is installed, the connecting line is routed from the inside of the model, passes through the support mechanism of the first and second aircraft models, and is led out to the wall of the wind tunnel.

[0046] S2, determine the parallel mechanism configuration of driving the six-degree-of-freedom motion of the second aircraft under the preset condition, calibrate the positioning accuracy and accuracy of the three-degree-of-freedom motion driving mechanism of the first aircraft, and calibrate the positioning accuracy and accuracy of the six-degree-of-freedom motion parallel mechanism of the second aircraft;

[0047] Figure 5 The schematic diagram of the three-degree-of-freedom angular displacement driving mechanism of the first aircraft in the embodiment of the application is shown, and formula 5 gives the inverse solution motion mathematical model of the three-degree-of-freedom angular displacement driving of the first aircraft:

[0048]

[0049] α1=arctan(tan(θ m )cos(γ1))

[0050]

[0051] α0=-(α1+α2)

[0052]

[0053] In formula 5, θ mis the preset angle between the main shaft and the tail shaft, a0 is the attack angle device rotation angle, g1 is the main shaft rotation angle around its own axis, g2 is the tail shaft rotation angle around its own axis, a is the tail shaft attack angle, b is the tail shaft sideslip angle, and g is the tail shaft roll angle.

[0054] The three-degree-of-freedom angular displacement driving mechanism of the primary aircraft performs position inverse solution according to the inverse kinematics model of formula 5, gives motor driving instructions, completes three-degree-of-freedom attitude angle movement, and then uses the absolute measurement wall to build a spatial coordinate system to complete the actual measurement of the attitude angle movement of the primary aircraft model in the wind tunnel, and finally measures 100 groups of actual attitude angles. The actual measurement attitude angle and a0, g1 and g2 obtained by inverse solution of the inverse kinematics mathematical model are used as inputs, and the nonlinear least squares method is used to obtain the error parameters of the three-degree-of-freedom angular displacement driving mechanism of the primary aircraft. The error parameters are substituted into formula 6, and a group of command attitude angles is given again to drive the primary motion mechanism to move. When the attitude angle movement is completed, the absolute measurement wall is used to measure the attitude angle again. If the difference between the actual attitude angle and the command attitude angle meets the positioning accuracy requirement, the calibration is completed. If it does not meet the requirement, another group of command attitude angles needs to be measured until the accuracy requirement is met, that is, the calibration is completed.

[0055]

[0056] a1 = arctan(tan(0 m + a2 m )cos(g1))

[0057]

[0058] a0 = -(a1 + a2)

[0059]

[0060] The positioning accuracy and accuracy of the six-degree-of-freedom parallel mechanism of the secondary aircraft axial displacement, lateral displacement, longitudinal displacement, pitch angle, yaw angle and roll angle are calibrated, including:

[0061] First, the theoretical position value is given, and the motor code disc value is obtained by inverse solution according to the six-degree-of-freedom parallel mechanism kinematics mathematical model (formula 7). The six-rod static spherical hinge of the parallel mechanism is driven to move linearly by the motor code disc rotation, and finally the actual position movement is completed by the secondary aircraft model at the end of the parallel mechanism. At the same time, 100 groups of position commands are measured by using the absolute measurement wall, and the unknown parameters in the kinematics mathematical model are obtained by the nonlinear least squares method. Then 20 groups of command attitude angles are given again. If the error between the position command and the actual measurement position does not meet the accuracy requirement, 20 groups of command attitude angles need to be measured again until the accuracy requirement is met.

[0062]

[0063] In formula 7, A ix is the x-coordinate of the spherical joint of the i-th branch of the six-degree-of-freedom parallel mechanism after static coordinate transformation in the wind tunnel coordinate system;

[0064] A iy is the y coordinate of the spherical joint of the i-th branch of the six-degree-of-freedom parallel mechanism after static coordinate transformation in the wind tunnel coordinate system;

[0065] A iz is the z coordinate of the spherical joint of the i-th branch of the six-degree-of-freedom parallel mechanism after static coordinate transformation in the wind tunnel coordinate system;

[0066] x ia , x ib are the slope and intercept of the x-coordinate of the static spherical joint of the i-th branch of the six-degree-of-freedom parallel mechanism with respect to the motor encoder value;

[0067] y ia ,y ib is the slope and intercept of the y-coordinate of the static spherical joint of the i-th branch of the six-degree-of-freedom parallel mechanism with respect to the motor encoder value;

[0068] z ia , z ib are the slope and intercept of the z-coordinate of the static spherical joint of the i-th branch of the six-degree-of-freedom parallel mechanism with respect to the motor encoder value;

[0069] h i is the encoder value of the i-th branch motor;

[0070] L i is the length of the i-th branch chain;

[0071]

[0072] In formula 8, a ix is the x-coordinate of the spherical joint of the i-th branch of the six-degree-of-freedom parallel mechanism in the model coordinate system;

[0073] a iy is the y-coordinate of the spherical joint of the i-th branch of the six-degree-of-freedom parallel mechanism in the model coordinate system;

[0074] a iz is the z coordinate of the spherical joint of the i-th branch of the six-degree-of-freedom parallel mechanism in the model coordinate system;

[0075] T wm is the transformation matrix from the secondary aircraft model coordinate system to the wind tunnel coordinate system.

[0076] S3. Perform a linkage test on the gust response actuator, the pulsating pressure sensor, the first-stage aircraft motion drive mechanism, and the six-degree-of-freedom parallel mechanism, driving the six-degree-of-freedom parallel mechanism equipped with the second-stage aircraft model to move to a hanging position within a preset range of the first-stage aircraft model;

[0077] S4. Carry out pre-test preparations;

[0078] S4 specifically includes:

[0079] Start the pulsating pressure sensor and signal acquisition equipment;

[0080] Start the wind tunnel and wait for the wind tunnel flow field to stabilize. Then start the ejection and separation of the first and second stage aircraft models. When the ejection process is completed, the second stage aircraft model has moved to the separation position, including:

[0081] The ejection process described above is based on a pre-existing mathematical model, which provides a mathematical relationship between aerodynamic forces or moments and position from the moment of separation to the moment of separation completion. This mathematical relationship is used to simulate the physical process of ejection and separation between the primary and secondary vehicles.

[0082] The wind tunnel response actuator is started and moves to a designated position according to the selected mathematical model of gust disturbance;

[0083] S5. Monitor the pulsating pressure sensor until the fluctuating flow field tends to a stable state, calculate the aerodynamic coefficients based on the built-in balances of the first-stage aircraft and the second-stage aircraft, calculate the postures of the first-stage aircraft and the second-stage aircraft at the next moment, and drive the first-stage aircraft and the second-stage aircraft to move according to the postures of the first-stage aircraft and the second-stage aircraft at the next moment;

[0084] S5 specifically includes:

[0085] A high-performance computer that observes and processes pulsating pressure signals monitors whether the fluctuating flow field has stabilized.

[0086] If it is stable, the aerodynamic coefficients of the first-stage aircraft at this moment, including aerodynamic force and aerodynamic moment, are calculated through the built-in balance of the first-stage aircraft. The displacement relative to the ground is calculated according to the six-degree-of-freedom dynamic equation, and the first-stage aircraft motion mechanism is driven to complete the three-degree-of-freedom attitude angular motion of pitch angle, yaw angle and roll angle.

[0087] At the same time, the aerodynamic coefficients of the secondary aircraft at the same moment, including aerodynamic force and aerodynamic torque, are calculated through the built-in balance of the secondary aircraft. The displacement and attitude relative to the ground are calculated according to the six-degree-of-freedom dynamic equation. The six-degree-of-freedom parallel mechanism is used to drive the secondary aircraft model to perform displacement motion relative to the first-stage aircraft model and attitude angular motion relative to the ground, including:

[0088] Both the first and second level aircraft models mentioned above require a built-in balance to obtain the aerodynamic force and aerodynamic moment coefficients. The process of measuring and calculating the aerodynamic force and aerodynamic moment coefficients using the built-in balance is as follows:

[0089] The built-in balance is directly connected to the equipment that collects the balance's measurement signals, forming a closed loop. When aerodynamic forces and moments act on the model, the built-in balance deforms, causing changes in resistance and generating electrical signal fluctuations. The electrical signal fluctuations are read by the balance acquisition equipment and are called code values. The code values ​​are then used to calculate the aerodynamic force and moment coefficients, as shown in Equation 9:

[0090]

[0091] In formula 9, A ω is the aerodynamic load of the balance, at the initial moment A ω is zero, Tp is the six-component code value of the balance, PK is the interference term, F is the aerodynamic coefficient, at this time i=2,4,6,j=2,4,6;

[0092] M is the aerodynamic coefficient, at this time i=1,3,5,j=1,3,5,ρ v is the wind tunnel dynamic pressure, S r is the model reference area, L r is the model reference length;

[0093] S6. Determine the wind tunnel safety boundary and provide the coordinates of the secondary aircraft's collision-prone points at each moment. Calculate the safety margin based on the coordinates of the collision-prone points within the safety boundary. Use the safety margin to predict collisions and obtain trajectory capture test data for all hazardous area separation states.

[0094] Determine the wind tunnel safety boundary and provide the coordinates of the collision-prone points of the secondary aircraft model at each moment. Calculate the safety margin based on the coordinates of the collision-prone points in the safety boundary and predict collisions using the safety margin, including:

[0095] The coordinates of the center of mass of the first-stage aircraft model and the first-stage aircraft model in the wind tunnel were imported into the mechanical dynamics simulation software to obtain the coordinates of the first-stage aircraft's outer surface. The coordinates of the first-stage aircraft's outer surface were then translated 10mm perpendicularly outward from the model's outer surface to serve as a safety margin. Simultaneously, the second-stage aircraft model was treated as an envelope, and the mechanical dynamics simulation software was used to calculate the coordinates of the vulnerable points of the second-stage aircraft model during its posture motion. Therefore, the safety margin Δh = safety margin - vulnerable point coordinates. This safety margin is used to predict collisions, and the condition for determining whether a collision occurs is Δh ≤ 10mm.

[0096] If the predicted collision does not occur, the calculation iteration of the next moment begins, that is, return to step 13, and the first and second level aircraft perform separation movement at the next moment until the trajectory capture test is completed, and data is generated at the same time; if the predicted collision occurs, the six-degree-of-freedom parallel mechanism is driven to move the second level aircraft model to a safe position, the trajectory capture test is completed, and data is generated.

[0097] According to the flight mission requirements of the first and second level aircraft, design and carry out test conditions prone to collision when encountering gust response, and conduct trajectory capture tests for each test condition prone to collision, repeating S5-S6.

[0098] The embodiment of the present invention further includes:

[0099] After obtaining the trajectory capture test data of all hazardous area separation states, design the force measurement station locations of the secondary aircraft's pitch angle, yaw angle, and roll angle combination sequence based on the separation displacement of the secondary aircraft and the first aircraft in the Cartesian coordinate system, namely the grid force measurement points, including:

[0100] The trajectory lines of all dangerous area separation states are as follows Figure 6 As shown, and the grid force measurement station locations are as follows Figure 7 shown.

[0101] At each grid force measurement station, the first-stage aircraft motion mechanism drives the first-stage aircraft model to perform attitude angular motion, while the second-stage aircraft parallel six-degree-of-freedom motion mechanism drives the second-stage aircraft model to perform displacement and attitude angular motion. After the displacement and attitude angles of the first and second-stage models are in place, conventional force measurements are performed using the built-in balances of the first and second-stage models to obtain aerodynamic force and aerodynamic torque data.

[0102] The grid force measurement point data is stored in a database and used as input for the PID parameter adjustment process to design a control strategy that meets the requirements for maneuvering out of dangerous areas for the first and second level aircraft, and then generate the rudder deflection control law for the first and second level aircraft.

[0103] The first and second stage aircraft with rudder control law will be subjected to trajectory capture tests according to all separation states in the danger zone to verify the control capabilities of the first and second stage aircraft under aerodynamic interference in the danger zone and their ability to maneuver out of the danger zone where a collision is about to occur.

[0104] Carry out experiments based on multi-body separation trajectory prediction under gust interference to obtain trajectory lines and aerodynamic data of separation state in dangerous areas. The specific process is as follows: Figure 8 shown.

[0105] According to the mission requirements of the first and second stage aircraft, a separation state that is prone to collision when responding to gusts of wind is designed, and a trajectory capture test is carried out for each separation condition.

[0106] At the beginning of the test, first of all, according to the test conditions at the separation boundary danger zone, the secondary aircraft model is moved to the separation position through the parallel six-degree-of-freedom mechanism. Turn on the pulsating pressure sensor, start the wind tunnel, and after the wind tunnel's own flow field is stable, start the gust response actuator and observe the pulsating pressure sensor. The pulsating pressure sensor is used to observe whether the flow field fluctuation is stable. If the fluctuating flow field cannot approach a certain gust mathematical model form for 15 seconds, the wind tunnel will be shut down; if the fluctuating flow field can approach a certain gust function form, then the first and second stage aircraft calculate the aerodynamic force and aerodynamic torque of the first and second stage aircraft models at the current moment through the internal five / six-component force balance. Subsequently, the trajectory solving software receives the current posture, separation parameters, aerodynamic force and torque, and solves the posture of the first and second stage aircraft at the next moment through the flight dynamics equation. Using the calculated displacement and attitude angle, the displacement instruction of the motion mechanism control system is solved through the kinematic inverse formula of the two-stage aircraft. The motion mechanism control system drives the two-stage aircraft to the specified displacement and attitude angle according to the position instruction. As Figure 7 As shown, multi-body separation collision detection is performed. If no collision is detected, the above process is repeated until the simulation time or simulated travel reaches the motion space threshold, completing the entire trajectory test. If a collision is detected as imminent, the primary and secondary vehicle drive mechanisms will cease motion and return to their respective safe response positions. The wind tunnel is then shut down, and a multi-body separation capture trajectory based on gust interference is generated.

[0107] After obtaining all separation trajectories prone to collision in response to all gusts, operating conditions that do not meet flight safety standards are selected and grid force measurement station locations for two-stage separation are designed (i.e., based on the separation displacement of the secondary vehicle from the first vehicle in the Cartesian coordinate system, force measurement station locations for a combined sequence of the second vehicle's pitch, yaw, and roll angles are designed). Grid force measurement tests are then conducted using conventional force measurements using the built-in balances of the first and second stage models to obtain aerodynamic force and moment data.

[0108] After obtaining the aerodynamic data of all grid force measurement points, the design of the rudder control law is carried out in the form of a database. Through the PID parameter adjustment method, the ability of the second-stage aircraft to autonomously return to the dangerous area when a collision occurs between the two stages is obtained.

[0109] By adopting the embodiments of the present invention, the following beneficial effects are achieved:

[0110] The control technology of the first and second aircrafts resisting aerodynamic interference under the sudden wind interference can be realized. By using the test method, the multi-body separation motion trajectory prediction under the changing wind field can be realized. The wind tunnel sudden wind response technology and the trajectory capture test technology are combined innovatively, the superposition of the aerodynamic interference zone of the first and second aircrafts under the mutual motion and the aerodynamic interference zone under the change of the flow field itself is realized, and then more comprehensive design basis for the multi-body separation safety analysis and the maneuvering control technology iteration is provided.

[0111] Device embodiment one

[0112] According to the embodiment of the present application, an electronic device is provided, comprising:

[0113] a processor; and

[0114] a memory arranged to store computer-executable instructions that, when executed, cause the processor to perform the steps of the above method embodiments.

[0115] Device embodiment two

[0116] According to the embodiment of the present application, a storage medium is provided for storing computer-executable instructions that, when executed, implement the steps of the above method embodiments.

[0117] Finally, it should be noted that: the above embodiments are only used to illustrate the technical solutions of the present application, and not to limit them; although the present application has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand: it can still modify the technical solutions recorded in the foregoing embodiments, or make equivalent replacement for part or all of the technical features; and these modifications or replacements do not make the essence of the corresponding technical solutions deviate from the scope of the technical solutions of the embodiments of the present application.

Claims

1. An experimental method for multi-body separation trajectory prediction under gust disturbance, characterized by include: S1. Initializing the test environment, including: determining wind tunnel model parameters of the aircraft under test, determining a mathematical model of gust interference and a corresponding motion response actuator, and determining parameters related to monitoring the flow field, wherein a pulsating pressure sensor is pre-set in the aircraft; S2. Determine the configuration of the parallel mechanism that drives the six-degree-of-freedom motion of the secondary aircraft under preset conditions, calibrate the positioning accuracy and precision of the three-degree-of-freedom motion drive mechanism of the primary aircraft, and calibrate the positioning accuracy and precision of the six-degree-of-freedom motion parallel mechanism of the secondary aircraft; S3. Perform a linkage test on the gust response actuator, the pulsating pressure sensor, the first-stage aircraft motion drive mechanism, and the six-degree-of-freedom parallel mechanism, driving the six-degree-of-freedom parallel mechanism equipped with the second-stage aircraft model to move to a hanging position within a preset range of the first-stage aircraft model; S4. Carry out pre-test preparations; S5. Monitor the pulsating pressure sensor until the fluctuating flow field tends to a stable state, calculate the aerodynamic coefficients based on the built-in balances of the first-stage aircraft and the second-stage aircraft, calculate the postures of the first-stage aircraft and the second-stage aircraft at the next moment, and drive the first-stage aircraft and the second-stage aircraft to move according to the postures of the first-stage aircraft and the second-stage aircraft at the next moment; S6. Determine the wind tunnel safety boundary and provide the coordinates of the secondary aircraft's collision-prone points at each moment. Calculate the safety margin based on the coordinates of the collision-prone points within the safety boundary. Use the safety margin to predict collisions and obtain trajectory capture test data for all hazardous area separation states.

2. The method according to claim 1, characterized in that The wind tunnel model parameters of the aircraft under test are specifically determined as follows: Determine the wind tunnel model scale and separation parameters of the first-stage and second-stage aircraft, including: The ratio of the first-stage aircraft to the full-scale first-stage aircraft, the flight altitude, flight Mach number, and flight overload of the full-scale first-stage aircraft, the mass, inertia, and center of mass position of the first-stage aircraft, and the angle of attack, sideslip angle, and roll angle of the first-stage aircraft in maneuvering or level flight; The ratio of the secondary aircraft model to the full-size secondary aircraft, the flight speed, flight altitude, flight overload of the secondary aircraft and the separation position of the secondary aircraft and the first-stage aircraft, the mass, inertia and center of mass position of the secondary aircraft, the separation speed and separation angular velocity of the secondary aircraft and the attitude angle of the secondary aircraft relative to the first-stage aircraft.

3. The method according to claim 1, characterized in that The determination of the gust disturbance mathematical model and the corresponding motion response execution mechanism includes: The mathematical models of gust disturbance include: longitudinal slope gust model, step gust model, full wavelength discrete gust model, half wavelength discrete gust model, full wavelength continuous gust model, and half wavelength continuous gust model; The driving mechanism of gust response adopts a combination of crank-connecting rod and quadrilateral connecting rod mechanism to realize the reciprocating swing of the blade. The motion response actuator corresponding to gust interference is composed of a crank-connecting rod and polygonal connecting rod mechanism, a flywheel mechanism and a synchronous belt mechanism.

4. The method according to claim 1, wherein The monitoring flow field related parameters include: Determine the model of the pulsating pressure sensor for monitoring the flow field, the frequency of use of the pulsating pressure sensor, the size of the pulsating pressure sensor, the signal acquisition equipment and the high-performance computer.

5. The method according to claim 1, wherein The S4 specifically includes: Start the wind tunnel, and after the wind tunnel flow field stabilizes, eject and separate the first and second stage aircraft models, so that the second stage aircraft model moves to the separation position; The wind tunnel response actuator is started and moves to a designated position according to a predetermined gust disturbance mathematical model.

6. The method according to claim 1, characterized in that The S5 specifically includes: Monitor whether the fluctuating flow field of the pulsating pressure sensor tends to be stable. If it is not stable, continue monitoring for a preset time. If it still does not tend to be stable after the preset time, shut down the wind tunnel. If it is stable, the aerodynamic coefficients of the first-stage aircraft at this moment, including aerodynamic force and aerodynamic moment, are calculated through the built-in balance of the first-stage aircraft. The position and attitude relative to the ground at the next moment are solved according to the flight dynamics equation, and the first-stage aircraft motion mechanism is driven to complete the three-degree-of-freedom attitude angular motion of pitch angle, yaw angle and roll angle; at the same time, the aerodynamic coefficients of the second-stage aircraft at the same moment, including aerodynamic force and aerodynamic moment, are calculated through the built-in balance of the second-stage aircraft. The displacement and attitude relative to the ground at the next moment are solved according to the flight dynamics equation, and the second-stage aircraft model is driven by the six-degree-of-freedom parallel mechanism to perform displacement motion relative to the first-stage aircraft model and attitude angular motion relative to the ground.

7. The method according to claim 1, characterized in that The S6 specifically includes: If no collision occurs, S5 is executed. If a collision occurs, the six-degree-of-freedom parallel mechanism is driven to move the secondary aircraft model to a safe position. The trajectory capture test ends and trajectory data is generated. Design and carry out test conditions prone to collision when responding to gusts, and conduct trajectory capture tests for each test condition prone to collision to obtain trajectory capture test data for all hazardous area separation conditions.

8. The method according to claim 1, characterized in that The method further comprises: In the Cartesian coordinate system, based on the separation displacement of the secondary aircraft and the first-stage aircraft, the force measurement station locations of the second-stage aircraft's pitch angle, yaw angle, and roll angle combination sequence are designed. According to the force measurement station locations, the first-stage aircraft's motion mechanism drives the first-stage aircraft model to perform attitude angle movement, and the second-stage aircraft's parallel six-degree-of-freedom motion mechanism drives the second-stage aircraft model to perform displacement and attitude angle movement. After the displacement and attitude angle of the first and second-stage models are in place, conventional force measurement is performed using the built-in balances of the first and second-stage models to obtain aerodynamic force and aerodynamic torque data; The grid force measurement point data is stored in a database and used as input for the PID parameter adjustment process to design a control strategy that meets the requirements for maneuvering out of dangerous areas for the first and second level aircraft, and then generate the rudder deflection control law for the first and second level aircraft. The first and second stage aircraft with rudder control law will be subjected to trajectory capture tests according to all separation states in the danger zone to verify the control capabilities of the first and second stage aircraft under aerodynamic interference in the danger zone and their ability to maneuver out of the danger zone where a collision is about to occur.

9. An electronic device comprising: processor; as well as, A memory arranged to store computer-executable instructions, wherein when the computer-executable instructions are executed, the processor is caused to perform the steps of the experimental method for multi-body separation trajectory prediction under gust disturbance according to any one of claims 1 to 8.

10. A storage medium for storing computer-executable instructions, wherein the computer-executable instructions, when executed, implement the steps of the experimental method for multi-body separation trajectory prediction under gust disturbance as claimed in any one of claims 1 to 8.

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