Test method and device for predicting multi-body separation trajectory under gust interference and medium

Abstract

The invention provides a test method and device for predicting a multi-body separation trajectory under gust interference and a medium, and the method comprises the steps: initializing a test environment, and determining a parallel mechanism configuration for driving a two-stage aircraft to move at six degrees of freedom under a preset condition, calibrating the positioning precision and accuracy of the three-degree-of-freedom motion driving mechanism of the first-stage aircraft and the six-degree-of-freedom motion parallel mechanism of the second-stage aircraft; carrying out a linkage test; carrying out preparation work before the test; driving the first-stage aircraft and the second-stage aircraft to move according to the poses of the first-stage aircraft and the second-stage aircraft at the next moment; determining a wind tunnel safety boundary, giving out coordinates of easy-to-collide points of a secondary aircraft at each moment, calculating safety margin according to the coordinates of the easy-to-collide points of the safety boundary, predicting collision through the safety margin, and obtaining trajectory capture test data of all dangerous areas in a separated state; according to the technical scheme provided by the invention, multi-body separation motion trail prediction in a variable wind field can be realized.

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 is to drive the airfoil mechanism in the wind tunnel through an actuator to rotate in a certain mathematical function at a given frequency, thereby generating a flow field with discrete gusts. Wind tunnel tests are of great significance to the study of gust response, mainly using model test methods to test the impact of gusts and verify the simulation results. Wind tunnel gust response tests usually use two methods, one is discrete isolated gusts and the other is continuous gusts. Discrete isolated gusts use high-pressure blowing to generate local gust conditions in the wind tunnel; continuous gusts use mechanisms to drive blades to swing repeatedly and regularly to form a continuous gust field. Discrete isolated gusts can generate large loads, but the affected area is small and cannot form a large gust field. The principle of continuous gusts is to deflect the wind tunnel airflow, so it can form a larger gust field, and the wind speed is continuous and regularly changing. Different gust response test methods can be selected according to the requirements of the test.

[0003] The Capture Trajectory Test (CTS) is a commonly used two-stage separation wind tunnel test method. Its advantages are high prediction accuracy, the ability to obtain test results that are basically consistent with full-scale flight test data, the ability to simulate complex separation conditions and special flight states of external attachments through computer software, such as pitch, climb or accelerated flight, etc., and the ability to adapt to launches under certain envelope flight states of the aircraft. The external attachment separation trajectory can be directly given in the test, and the separation characteristics can be evaluated in a timely manner. The test can not only obtain the full-scale condition separation trajectory, but also directly measure the aerodynamic loads of the external attachment at each measurement point on the separation trajectory, which is conducive to the analysis and improvement of the external attachment separation characteristics. In view of this advantage, it is an aircraft / external attachment compatibility assessment and verification project that must be carried out before conducting flight tests.

[0004] The test of conventional separation of aircraft in normal flow field has become increasingly mature in China. However, in recent years, it has been found that when aircraft are tested in the field, even in the same separation state, the posture of the separated aircraft is completely different due to the influence of atmospheric disturbance. This requires that the wind tunnel can simulate the trajectory capture test under the interference of gusts. At present, there is no system or method to simulate this situation in China. Summary of the invention

[0005] The embodiments of the present invention provide an experimental method, equipment and medium for predicting multi-body separation trajectories under gust interference, which solve the problem of collision between two-stage aircraft before, during and after separation due to gust interference; and the problem of dangerous blind spots in the separation of two-stage aircraft; after the two-stage aircraft is safely separated, its displacement and posture 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 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) The existing wind tunnel gust response simulation technology cannot simulate the posture movement of the aircraft, let alone analyze the posture movement under the gust disturbance when the first and second stage aircraft separate. The present application realizes the trajectory prediction capability of multi-body separation under gust disturbance.

[0013] (2) Achieve integrated test technology capabilities for gust disturbance aerodynamic analysis - multi-body separation trajectory prediction - grid force measurement to obtain aerodynamic characteristics of strong disturbance danger 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 drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. 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 creative labor.

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

[0016] Figure 2A schematic diagram of a wind tunnel gust response driving mechanism according to an embodiment of the present invention;

[0017] Figure 3 This is a structural diagram of the LQ-080 pulsating pressure sensor according to an embodiment of the present invention;

[0018] Figure 4 A schematic diagram of the locations of the pulsating pressure sensors installed in the first and second stage aircrafts according to an embodiment of the present invention;

[0019] Figure 5 It is a schematic diagram of a three-degree-of-freedom angular displacement driving mechanism of a first-stage aircraft according to an embodiment of the present invention;

[0020] Figure 6 A schematic diagram of a trajectory line of a dangerous area separation state according to an embodiment of the present invention;

[0021] Figure 7 A schematic diagram of a grid force measurement station location according to an embodiment of the present invention;

[0022] Figure 8 This is a multi-body separation trajectory capture prediction test process based on gust interference according to an embodiment of the present invention. DETAILED DESCRIPTION

[0023] In order to enable those skilled in the art to better understand the technical solutions in one or more embodiments of this specification, the technical solutions in one or more embodiments of this specification will be clearly and completely described below in conjunction with the drawings in one or more embodiments of this specification. Obviously, the described embodiments are only part of the embodiments of this specification, not all of them. Based on one or more embodiments of this specification, all other embodiments obtained by ordinary technicians in this field without creative work should fall within the scope of protection of this document.

[0024] Method Embodiment

[0025] According to an embodiment of the present invention, a test method for multi-body separation trajectory prediction under gust disturbance is provided. Figure 1 Flow chart of the experimental method for predicting multi-body separation trajectories under gust disturbance according to an embodiment of the present invention. Figure 1 As shown, the experimental method for predicting multi-body separation trajectories under gust disturbance in an embodiment of the present invention includes:

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

[0027] Determine the scale and separation parameters of the first and second stage aircraft wind tunnel models, including:

[0028] The ratio of the first-stage aircraft to the full-size first-stage aircraft, the flight altitude, flight Mach number, and flight overload of the full-size 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-size 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] The mathematical model of gust disturbance and the corresponding motion response actuators are determined as follows:

[0031] Gust disturbance mathematical models: 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, 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 of 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, such as 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 the gust disturbance is composed of a combination of crank-connecting rod-polygonal connecting rod mechanism, flywheel mechanism, synchronous belt mechanism and other mechanisms, and 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] It can be seen that the blade cascade swing angular velocity is as shown in formula 3:

[0039]

[0040] The angular velocity is shown in Formula 4:

[0041]

[0042] In Formula 2 to Formula 4, A represents the amplitude of the oscillation of the cascade mechanism, and f represents the frequency (0 to 15 Hz) of the oscillation of the cascade mechanism.

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

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

[0045] The schematic diagram of the location of the pulsating pressure sensor installed on the model is as follows Figure 4 As shown, after the pulsating pressure sensor is installed, its connecting line is routed from the inside of the model, passes through the supporting mechanisms of the first and second aircraft models, and is led out to the outside of the wind tunnel wall.

[0046] S2. Determine the configuration of the parallel mechanism driving 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 driving 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;

[0047] Figure 5 Schematic diagram of the three-degree-of-freedom angular displacement drive mechanism of the first-stage aircraft according to an embodiment of the present invention. Formula 5 gives the inverse motion mathematical model of the three-degree-of-freedom angular displacement drive of the first-stage 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, α0 is the rotation angle of the angle of attack device, γ1 is the rotation angle of the main shaft around its own axis, γ2 is the rotation angle of the tail shaft around its own axis, α is the tail shaft attack angle, β is the tail shaft sideslip angle, and γ is the tail shaft roll angle.

[0054] The first-stage aircraft three-degree-of-freedom angular displacement drive mechanism performs position inverse solution according to the motion inverse solution model described in Formula 5, gives motor drive instructions, completes the three-degree-of-freedom attitude angle movement, and then uses the space coordinate system built by the absolute measurement wall to complete the actual measurement of the attitude angle movement of the first-stage aircraft model in the wind tunnel, and finally measures 100 sets of actual attitude angles. The actual measured attitude angle and α0, γ1, and γ2 obtained by inverse solution of the inverse motion mathematical model are used as inputs, and the error parameters of the first-stage aircraft three-degree-of-freedom angular displacement drive mechanism are obtained by nonlinear least squares method. The error parameters are substituted into Formula 6, and a set of command attitude angles are given again to drive the first-stage motion mechanism to move. After the attitude angle movement is completed, the attitude angle is measured again using the absolute measurement wall. If the difference between the actual attitude angle and the command attitude angle meets the positioning accuracy requirement, the calibration is completed. If not, it is necessary to measure another set of command attitude angles until the accuracy requirement is met, that is, the calibration is completed.

[0055]

[0056] α1=arctan(tan(θ m +Δθ m )cos(γ1))

[0057]

[0058] α0=-(α1+α2)

[0059]

[0060] Calibrate the positioning accuracy and precision of the six-degree-of-freedom motion parallel mechanism of the secondary aircraft's axial displacement, lateral displacement, longitudinal displacement, pitch angle, yaw angle and roll angle, including:

[0061] First, given the theoretical posture value, according to the six-degree-of-freedom parallel mechanism motion mathematical model (Formula 7), the motor encoder value is obtained by inverse solution. The static ball joints of the six rods of the parallel mechanism are driven to perform linear motion by the rotation of the motor encoder, and finally the actual posture movement is completed by the secondary aircraft model at the end of the parallel mechanism. At the same time, the 100 sets of posture instructions given are actually measured using an absolute measurement wall, and the unknown parameters in the motion mathematical model are obtained by the nonlinear least squares method. Then 20 sets of instruction attitude angles are re-given. If the error between the posture instruction and the actual measured posture does not meet the accuracy requirements, the group needs to re-measure 20 sets of instruction attitude angles until they meet the accuracy requirements.

[0062]

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

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

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

[0066] x ia , x ib is the slope and intercept of the x-coordinate of the static spherical joint point 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 point 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 is the slope and intercept of the z coordinate of the static spherical joint point 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 moving ball 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 moving ball 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 moving ball 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, performing 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, and 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 preparatory work before the test;

[0078] S4 specifically includes:

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

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

[0081] The above ejection process is a given ejection mathematical model, which gives a mathematical relationship between the aerodynamic force or aerodynamic torque and the posture during the time period from the separation action to the separation completion time. The mathematical relationship is used to simulate the physical process of ejection separation of the first and second stage aircraft.

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

[0083] S5. Monitor the pulsating pressure sensor until the fluctuating flow field tends to be stable, calculate the aerodynamic coefficient according to the built-in balance of the first-stage aircraft and the second-stage aircraft, calculate the posture 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 posture 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 solved 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 angle motion of pitch angle, yaw angle and roll angle;

[0087] At the same time, the aerodynamic coefficients of the secondary aircraft at the same time, including aerodynamic force and aerodynamic moment, are calculated through the built-in balance of the secondary aircraft. The displacement and attitude relative to the ground are solved according to the six-degree-of-freedom dynamic equation, and the secondary 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 angle motion relative to the ground, including:

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

[0089] The built-in balance is directly connected to the acquisition device of the balance measurement signal to form a closed loop. When the aerodynamic force and aerodynamic moment act on the model, the built-in balance will deform, and then the resistance value will change, generating electrical signal fluctuations. The electrical signal fluctuation value is read by the balance acquisition device, which is called the code value. The aerodynamic force and aerodynamic moment coefficients are then calculated from the code value. The calculation process is shown in Formula 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 give the coordinates of the collision-prone points of the secondary aircraft at each moment. Calculate the safety margin based on the coordinates of the collision-prone points of the safety boundary. Predict the collision through the safety margin and obtain the trajectory capture test data of the separation state of all dangerous areas.

[0094] Determine the wind tunnel safety boundary and give 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 of the safety boundary and predict collisions through the safety margin, including:

[0095] The first-stage aircraft model and the center of mass coordinates of the first-stage aircraft model in the wind tunnel are imported into the mechanism dynamics simulation software to obtain the coordinates of the first-stage aircraft outer surface, and the first-stage aircraft shape coordinates are translated 10mm along the vertical outward direction of the model outer surface as a safety margin. At the same time, the second-stage aircraft model is regarded as an envelope, and the coordinates of the easy-to-collision points of the second-stage aircraft model when performing posture movement are calculated through the mechanism dynamics simulation software. Therefore, the safety margin Δh = safety margin-easy-to-collision point coordinates, and the safety margin is used to predict the collision, and the condition for judging whether there is a collision 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 aircrafts 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 also includes:

[0099] After obtaining the trajectory capture test data of all dangerous area separation states, the force measuring stations of the secondary aircraft's pitch angle, yaw angle and roll angle combination sequence are designed in the Cartesian coordinate system according to the separation displacement of the secondary aircraft and the first aircraft, namely the grid force measuring points, including:

[0100] The trajectory lines of all dangerous areas separation states are as follows Figure 6 As shown in Figure 2, and the grid force measurement station locations are as follows Figure 7 shown.

[0101] According to each grid force measuring station, the first-stage aircraft motion mechanism drives the first-stage aircraft model to perform attitude angle movement, and the second-stage aircraft 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 through the built-in balance of the first and second-stage models to obtain aerodynamic force and aerodynamic moment data;

[0102] The grid force measurement point data is made into a database as the input of the PID parameter adjustment process, and a control strategy that meets the requirements of the first and second level aircraft maneuvers to get out of the dangerous area is designed, and then the rudder deflection control law of the first and second level aircraft is generated;

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

[0104] Conduct experiments based on multi-body separation trajectory prediction under gust interference to obtain trajectory lines and aerodynamic data of separation states 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 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 dangerous area of ​​the separation boundary, the second-stage aircraft model is moved to the separation position through the parallel six-degree-of-freedom mechanism. The pulsating pressure sensor is turned on and the wind tunnel is started. After the flow field of the wind tunnel itself is stable, the gust response actuator is started and the pulsating pressure sensor is observed. 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, at this time, the first and second stage aircraft use the internal five / six-component force balance to calculate the aerodynamic force and aerodynamic torque of the first and second stage aircraft models at the current moment. 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 command 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 move to the specified displacement and attitude angle according to the position command. As Figure 7 As shown, multi-body separation collision detection is performed. If it is judged that there is no collision, the above process is repeated until the simulation time or simulation stroke reaches the motion space threshold and the entire trajectory test is completed. If it is judged that a collision is about to occur, the primary and secondary aircraft drive mechanisms will stop and return to the response safety position respectively, then the wind tunnel will be shut down and a multi-body separation capture trajectory based on gust interference will be generated.

[0107] After obtaining all separation trajectories that are prone to collision in all gust responses, select the working conditions that do not meet the flight safety standards and design the two-stage separation grid force measurement station sites (i.e., in the Cartesian coordinate system, based on the separation displacement of the secondary aircraft and the first aircraft, design the force measurement station sites of the second-stage aircraft's pitch angle, yaw angle and roll angle combination sequence). Grid force measurement tests are carried out by performing conventional force measurements using the built-in balances of the first and second stage models to obtain aerodynamic force and aerodynamic 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 correct out of the dangerous area when the two stages collide is obtained.

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

[0110] It can realize the control technology of the first and second stage aircraft to resist aerodynamic interference under gust interference. The test method can realize the prediction of multi-body separation motion trajectory under changing wind field, and innovatively propose to combine wind tunnel gust response technology with trajectory capture test technology to realize the superposition of aerodynamic interference area of ​​the first and second stage aircraft under mutual motion and aerodynamic interference area under the change of flow field itself, and then provide a more comprehensive design basis for multi-body separation safety analysis and maneuver control technology iteration.

[0111] Device Example 1

[0112] According to an embodiment of the present invention, there is provided an electronic device, including:

[0113] processor; and,

[0114] A memory is arranged to store computer executable instructions, which when executed cause the processor to perform the steps of the above method embodiments.

[0115] Device Example 2

[0116] According to an embodiment of the present invention, a storage medium is provided for storing computer executable instructions, wherein the computer executable instructions implement the steps of the above method embodiment as claimed when executed.

[0117] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, rather than to limit it. Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or replace some or all of the technical features therein by equivalents. However, these modifications or replacements do not cause the essence of the corresponding technical solutions to deviate from the scope of the technical solutions of the embodiments of the present invention.

Claims

1. An experimental method for multi-body separation trajectory prediction under gust disturbance, characterized in that include: S1, initializing the test environment, including: determining the wind tunnel model parameters of the aircraft under test, determining the mathematical model of gust interference and the corresponding motion response actuator, and determining the parameters related to the monitoring flow field, wherein a pulsating pressure sensor is preset in the aircraft; S2. Determine the configuration of the parallel mechanism driving 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 driving 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, performing 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, and 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 preparatory work before the test; S5. Monitor the pulsating pressure sensor until the fluctuating flow field tends to be stable, calculate the aerodynamic coefficient according to the built-in balance of the first-stage aircraft and the second-stage aircraft, calculate the posture 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 posture of the first-stage aircraft and the second-stage aircraft at the next moment; S6. Determine the wind tunnel safety boundary and give the coordinates of the collision-prone points of the secondary aircraft at each moment. Calculate the safety margin based on the coordinates of the collision-prone points of the safety boundary. Predict the collision through the safety margin and obtain the trajectory capture test data of the separation state of all dangerous areas.

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-size first-stage aircraft, the flight altitude, flight Mach number, and flight overload of the full-size 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, 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 the gust disturbance 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, characterized in that: The monitoring flow field related parameters include: Determine the model of the pulsation pressure sensor for monitoring the flow field, the frequency of use of the pulsation pressure sensor, the size of the pulsation pressure sensor, the signal acquisition equipment and the high-performance computer.

5. The method according to claim 1, characterized in that The S4 specifically includes: Start the wind tunnel, and after the wind tunnel flow field is stable, 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 specified 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 to monitor 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, and the position and attitude at the next moment relative to the ground 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 angle movement of the 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, and the displacement and attitude at the next moment relative to the ground 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 movement relative to the first-stage aircraft model and attitude angle movement relative to the ground.

7. The method according to claim 1, characterized in that The S6 specifically includes: If the collision does not occur, S5 is executed. If the 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 is completed, and trajectory data is generated; 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, and 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, according to the separation displacement of the secondary aircraft and the first-stage aircraft, the force measuring station positions of the second-stage aircraft pitch angle, yaw angle and roll angle combination sequence are designed. According to the force measuring station positions, the first-stage aircraft motion mechanism drives the first-stage aircraft model to perform attitude angle movement, and the second-stage aircraft 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 through the built-in balances of the first and second-stage models to obtain aerodynamic force and aerodynamic moment data; The grid force measurement point data is made into a database as the input of the PID parameter adjustment process, and a control strategy that meets the requirements of the first and second level aircraft maneuvers to get out of the dangerous area is designed, and then the rudder deflection control law of the first and second level aircraft is generated; The first and second stage aircraft with rudder control law will be subjected to trajectory capture test according to all separation states in the danger zone to verify the control capability of the first and second stage aircraft under aerodynamic disturbance 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, which, when executed, cause the processor to perform 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.

10. A storage medium 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 in any one of claims 1 to 8.

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