A method for testing the posture of a suspended rotating body aircraft in a closed wind tunnel test

Abstract

This invention discloses a method for testing the attitude of a suspended spin-type aircraft in a closed wind tunnel test. The spin-type aircraft is suspended and supported in the test section of the closed wind tunnel by a cable traction system. The method includes: a calibration step, in which the spin-type aircraft is calibrated to a zero attitude; and a measurement step, in which the spin-type aircraft is subjected to wind tunnel testing with the zero attitude as the initial state. A calibration plate is fixed to the tail of the spin-type aircraft, and a camera captures images of the calibration plate at a preset frame rate. The attitude change data of the spin-type aircraft is obtained based on the attitude change of the calibration plate. The calibration plate has a calibration surface with N calibration points arranged in a preset manner. When the spin-type aircraft is in the zero attitude, the calibration surface faces the camera, and all calibration points on the calibration surface are located at the center of the camera's field of view. This testing method can improve the accuracy of the measurement data of the attitude and dynamic aerodynamic characteristics of the spin-type aircraft.

Description

Technical Field

[0001] This invention relates to the technical field of testing methods for spinning aircraft, specifically to a method for testing the attitude of a suspended spinning aircraft in a closed wind tunnel test. Background Technology

[0002] When a spinning projectile is in flight, its angular motion manifests in three forms: rotation around its axis, rotation around its velocity vector, and rotation around a spatial axis; precession and nutation. In wind tunnel testing, the main method for studying the dynamic aerodynamic characteristics of spinning projectiles is to simulate these three angular motions using a support mechanism and to measure the unsteady aerodynamic forces during these angular motions using a balance. Therefore, accurately calibrating the initial attitude of the spinning projectile and measuring the angular changes during single or coupled angular motions are crucial for understanding its precession, nutation characteristics, conical motion stability, and dynamic aerodynamic properties.

[0003] Existing wind tunnel testing setups for spinning aircraft typically employ either rigid strut supports or flexible rope supports. Rigid strut supports facilitate initial attitude calibration of the spinning aircraft, but vibrations in the support arms during testing can distort measurement data. Flexible rope supports, generally implemented via a rope traction system, are less convenient for initial attitude calibration, but the support system itself does not significantly impact the spinning aircraft. Furthermore, existing wind tunnel testing methods for spinning aircraft typically rely on sensors to measure attitude changes. This method is highly dependent on sensor accuracy and is susceptible to environmental influences, leading to inaccurate measurement data on the spinning aircraft model's attitude and dynamic aerodynamic characteristics. Summary of the Invention

[0004] The purpose of this invention is to overcome the aforementioned defects or problems in the prior art and to provide a method for testing the attitude of a suspended spinning aircraft in a closed wind tunnel test, which can improve the accuracy of the model attitude measurement data of the spinning aircraft during wind tunnel testing.

[0005] To achieve the above objectives, the present invention adopts the following technical solution:

[0006] Technical Solution 1: A method for testing the attitude of a suspended spin-type aircraft in a closed wind tunnel test, wherein the spin-type aircraft is suspended and supported in the test section of the closed wind tunnel by a rope traction system, comprising: a calibration step, in which the spin-type aircraft is calibrated to a zero attitude; and a measurement step, in which the spin-type aircraft is subjected to wind tunnel testing with the zero attitude as the initial state, a calibration plate is fixed at the tail of the spin-type aircraft, a camera captures the calibration plate at a preset frame rate, and the attitude change data of the spin-type aircraft is obtained based on the attitude change of the calibration plate; the calibration plate has a calibration surface, and the calibration surface has N calibration points arranged in a preset manner. When the spin-type aircraft is in a zero attitude, the calibration surface faces the camera, and all calibration points on the calibration surface are located at the center of the camera's field of view.

[0007] Technical Solution 2 is based on Technical Solution 1: The zero attitude refers to the deviation of the roll angle, pitch angle, and yaw angle of the spinning aircraft from the zero angle being less than a first threshold, and the deviation of its center of mass from the preset reference position in the forward and backward directions and the left and right directions being less than or equal to a second threshold; in the calibration step, the roll angle, pitch angle, and yaw angle of the spinning aircraft are first calibrated, and then the position of the center of mass of the spinning aircraft is calibrated.

[0008] Technical solution three is based on technical solution two: when calibrating the roll angle of the spin-integrated aircraft, a horizontal support plate is fixedly installed on the spin-integrated aircraft, and an inclinometer is placed on the horizontal support plate in the left and right directions. The relevant parameters of the rope traction system are adjusted according to the angle displayed by the inclinometer until the deviation between the roll angle of the spin-integrated aircraft and the zero-degree angle is less than the first threshold.

[0009] Technical solution four is based on technical solution three: when calibrating the pitch angle of the spin-type aircraft, a horizontal support plate is fixedly installed on the spin-type aircraft, and an inclinometer is placed on the horizontal support plate along the front-back direction. The relevant parameters of the rope traction system are adjusted according to the angle displayed by the inclinometer until the deviation between the pitch angle of the spin-type aircraft and the zero-degree angle is less than a first threshold.

[0010] Technical solution five is based on technical solution four: When calibrating the yaw angle of the spin-type aircraft, a first measurement point and a second measurement point are determined on the spin-type aircraft at the same altitude but at different positions. The distance between the first measurement point and the second measurement point is measured. Then, the distance between the first measurement point and the second measurement point relative to the same side wall of the closed wind tunnel is measured using an altitude vernier caliper. The relevant parameters of the rope traction system are adjusted based on the yaw angle calculated from the above distance data until the deviation of the yaw angle of the spin-type aircraft from zero degrees is less than a first threshold.

[0011] Technical Solution Six, based on Technical Solution Five: When calibrating the center of mass position of the spinning aircraft, a reference position is determined according to the size of the closed wind tunnel. Then, the distance between the center of mass of the spinning aircraft and the side and bottom walls of the closed wind tunnel is measured by a height vernier caliper. Based on this distance, the relevant parameters of the rope traction system are adjusted until the deviation of the center of mass position of the spinning aircraft from the reference position in the front-back and left-right directions is less than a second threshold.

[0012] Technical solution seven, which is based on technical solutions one to six: In the measurement step, the attitude change data of the spinning aircraft obtained by calculating the attitude change based on the calibration plate includes: calibrating the camera intrinsic parameters and obtaining the camera coordinate system; calibrating the camera extrinsic parameters and obtaining the transformation relationship between the world coordinate system and the camera coordinate system; obtaining the transformation relationship between the projectile coordinate system and the camera coordinate system through the calibration plate image captured by the camera; transforming the projectile coordinate system to the world coordinate system, and using the rotation matrix method to calculate and obtain the attitude change data of the spinning aircraft.

[0013] Technical solution eight, based on technical solution seven: In the measurement step, obtaining the pose change data of the rotating aircraft based on the pose change calculation of the calibration plate includes: calibrating the camera intrinsic parameters and obtaining the camera coordinate system O. c -x c y c z c ; Calibrate camera extrinsic parameters and obtain world coordinate system O w -x w y w z w and camera coordinate system O c -x c y c z c The transformation relationship is expressed by Equation 1: The coordinate system O of the projectile is obtained from the image of the calibration plate captured by the camera. a -x a y a z a With camera coordinate system O c -x c y c z c The transformation relationship is expressed by Equation 2:

[0014] Transform the projectile coordinate system to the world coordinate system, as shown in Equation 3: (Equation 3); The attitude change data of the rotating body aircraft are obtained by solving the rotation matrix method, including: expressing the relationship between the rotation matrix and the angle as Equation 4:

[0015]

[0016] (Equation 4); By comparing the corresponding position values ​​of the matrices in Equations 3 and 4 using the method of undetermined coefficients, and combining them with Equation 3, the attitude change data of the rotating aircraft can be obtained, i.e., Equation 5:

[0017] Pose = [t x t y t z arctan(r5 / r8) arctanr2 arctan(-r1 / r0)] T (Equation 5); where T aw =[t x ,t y ,t z ] T The displacements of the rotating aircraft in the forward / backward direction (x-axis), the vertical direction (y-axis), and the horizontal direction (z-axis) are given. Let φ be the roll angle, θ be the yaw angle, and ψ be the pitch angle of the rotating aircraft; [r0 r3 r6] T For O w -x w Axis at O c -x c y c z c The direction vector of the coordinate system; [r1 r4 r7] T For O w -y w Axis at O c -x c y c z c The direction vector of the coordinate system; [r2 r5 r8] T For O w -z w Axis at O c -x c y c z c The direction vector of the coordinate system; [t x t y t z ] T For O w -x w y w z w The origin of the coordinate system is at O c -x c y c z c Position in the coordinate system.

[0018] As can be seen from the above description of the present invention, compared with the prior art, the present invention has the following beneficial effects:

[0019] This invention provides a method for measuring the attitude of a suspended spin-type aircraft in a closed wind tunnel test. The method utilizes a cable traction system to suspend and support the spin-type aircraft, providing a larger dynamic attitude testing space within the relatively confined space of a closed wind tunnel. Based on this, the attitude calibration and measurement of the spin-type aircraft are carried out. First, the attitude of the spin-type aircraft is calibrated through a calibration process, ensuring it is in the zero-position for testing and measurement. Then, the wind tunnel test begins, with the spin-type aircraft moving according to the desired attitude driven by the cable traction system. During this process, a camera captures images of a calibration plate on the spin-type aircraft at a preset frame rate. By calculating the attitude of the calibration plate, the attitude changes of the spin-type aircraft can be determined. To achieve this, several calibration points are set on the calibration plate, arranged in a preset manner and in sufficient quantity. When a change in attitude occurs, the positions of the calibration points on the calibration plate captured by the camera in the image will also change accordingly. By observing the changes in the positions of the calibration points in the image and the transformation relationships between relevant coordinate systems, the attitude change of the spinning vehicle in space can be obtained accordingly. Furthermore, since the calibration plate is fixed to the tail of the spinning vehicle, the coordinate system of the calibration plate coincides with the coordinate system of the projectile (the three-dimensional positional relationship between the origins of the coordinate systems is known). The attitude change of the spinning vehicle can be inferred from the attitude of the calibration plate. Therefore, the camera only needs to capture the calibration plate at the tail of the projectile. Compared with the method of arranging calibration points on the body of the spinning vehicle to calculate the attitude, the method of fixing the calibration plate at the tail has relatively concentrated calibration points. The camera can capture a larger range of projectile attitude changes, providing more attitude measurement data for the aforementioned "larger dynamic attitude test workspace". In addition, using the calibration plate to calculate the attitude can significantly reduce the amount of calculation and improve the calculation efficiency. Meanwhile, aerodynamic forces are closely related to attitude angles. After static calibration of the position and attitude, the error between the actual attitude angle and the zero position and attitude can be reduced. During the dynamic measurement of the position and attitude, visual measurement improves the measurement accuracy. Through the test method provided by this invention, the accuracy of measuring the position and attitude and dynamic aerodynamic characteristics data of a rotating aircraft can be improved. Attached Figure Description

[0020] To more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings used in the following description of the embodiments are briefly introduced. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0021] Figure 1 A schematic diagram of the state of a spin-type aircraft in a closed wind tunnel, as provided in an embodiment of the present invention.

[0022] Figure 2A schematic diagram illustrating the calibration of the roll angle of a spinning aircraft according to an embodiment of the present invention;

[0023] Figure 3 A schematic diagram illustrating the calibration of the pitch angle of a rotating aircraft according to an embodiment of the present invention;

[0024] Figure 4 A schematic diagram illustrating the calibration of the yaw angle of a spinning aircraft according to an embodiment of the present invention;

[0025] Figure 5 A schematic diagram illustrating the vertical position of the center of mass of a spinning aircraft as provided in an embodiment of the present invention;

[0026] Figure 6 This is a schematic diagram illustrating the calibration of the center of mass position of a spinning aircraft in the left-right direction, provided in an embodiment of the present invention.

[0027] Figure 7 A schematic diagram of the structure of the calibration plate provided in the embodiment of the present invention for a rotating body aircraft;

[0028] Figure 8 A schematic diagram of a calibration plate on a spinning aircraft provided in an embodiment of the present invention;

[0029] Figure 9 This is a schematic diagram of camera imaging provided in an embodiment of the present invention.

[0030] Explanation of key figure labels:

[0031] 1. Rotating aircraft; 2. Closed wind tunnel; 3. Rope traction system; 4. Inclinometer; 5. Altitude vernier caliper; 6. Calibration plate; 7. Calibration point; 8. Horizontal bearing plate; 9. Projectile body; 10. Fixing block; 11. Tail rod. Detailed Implementation

[0032] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are preferred embodiments of the present invention and should not be considered as excluding other embodiments. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without creative effort are within the scope of protection of the present invention.

[0033] Unless otherwise expressly defined, the use of terms such as "first," "second," or "third" in the claims, description, and accompanying drawings of this invention is for distinguishing different objects and not for describing a specific order.

[0034] Unless otherwise expressly defined, in the claims, description, and accompanying drawings of this invention, the use of directional terms such as "center," "lateral," "longitudinal," "horizontal," "vertical," "top," "bottom," "inner," "outer," "upper," "lower," "front," "rear," "left," "right," "clockwise," and "counterclockwise" to indicate orientation or positional relationships is based on the orientation and positional relationships shown in the accompanying drawings and is only for the convenience of describing the invention and simplifying the description, and is not intended to indicate or imply that the device or element referred to must have a specific orientation or be constructed and operated in a specific orientation, and therefore should not be construed as limiting the specific scope of protection of this invention.

[0035] Unless otherwise expressly defined, the terms "fixed connection" or "fixed connection" used in the claims, description and drawings of this invention should be interpreted broadly to refer to any connection in which there is no displacement or relative rotation relationship between the two parties, including non-removable fixed connection, detachable fixed connection, integral connection and fixed connection by other means or components.

[0036] In the claims, description and accompanying drawings of this invention, the terms "comprising," "having," and variations thereof are used to mean "including but not limited to."

[0037] This invention provides a method for testing the attitude of a suspended spinning aircraft in a closed wind tunnel. The method includes two main steps: a calibration step and a measurement step. In the measurement step, the spinning aircraft 1 undergoes a simulated flight test according to conventional testing procedures. In this invention, the spinning aircraft 1 is tested in a closed wind tunnel 2, referring to… Figure 1 The closed wind tunnel 2 is roughly a rectangular metal structure with a cavity, surrounded by wind tunnel walls, with the front end open to receive airflow.

[0038] In this embodiment, the spin-type aircraft 1 is suspended and supported in the test section of the closed wind tunnel 2 by a rope traction system 3. Specifically, the rope traction system 3 mainly consists of multiple drive motors, traction cables, and steering pulleys. Typically, the drive motors in the rope traction system 3 can be fixedly installed on the outside of the wind tunnel wall of the closed wind tunnel 2, or fixedly installed on a fixed base. The output shaft of the drive motor is connected to the traction cable, which can extend or shorten the traction cable. The steering pulley is fixedly installed on the wind tunnel wall, either on the outside, inside, or across the wind tunnel wall. The other end of the traction cable connected to the drive motor passes around the steering pulley and enters the interior of the closed wind tunnel 2, connecting to a predetermined position on the spin-type aircraft 1, thus suspending and supporting the spin-type aircraft 1. When adjusting the model's attitude, corresponding control signals can be input to the drive motors to adjust the length of the corresponding traction cable, thereby adjusting the position of the spin-type aircraft 1 in the closed wind tunnel 2, as well as its roll, pitch, and yaw angles. It should be understood that the applicant has already filed multiple patent applications for the rope traction system 3. Therefore, the rope traction system 3 used will not be described in detail in this specification. Those skilled in the art can choose the rope traction system 3 with different application scenarios according to actual needs.

[0039] In the test method provided in this embodiment, the rotating body aircraft 1 is first calibrated to zero attitude in the calibration step.

[0040] Reference Figure 1 and Figure 7 The main body of the spin-integrated aircraft 1 is a conventional projectile body 9. A connecting rod extends rearward from the tail of the projectile body 9. A fixing block 10 is provided at the rear end of the connecting rod. A calibration plate 6 is provided on the fixing block 10. Two tail rods 11 extend to the left and right sides from the fixing block 10. The difference between this spin-integrated aircraft 1 and existing spin-integrated aircraft is that a calibration plate 6 is provided at the tail of the spin-integrated aircraft 1. The function of the calibration plate 6 will be described in detail below.

[0041] Reference Figure 1 In this specification and claims, six directions are defined: up, down, left, right, front, and back, as follows: Figure 1 The coordinate system is shown in the figure. In this coordinate system, the front end of the spinning aircraft 1 is front, the rear end of the spinning aircraft 1 is rear, the top of the spinning aircraft 1 is up, the bottom of the spinning aircraft 1 is down, the left hand direction when facing the spinning aircraft 1 is left, and the right hand direction when facing the spinning aircraft 1 is right.

[0042] In this embodiment, the zero attitude of the spinning aircraft 1 refers to the deviation of its roll angle, pitch angle, and yaw angle from the zero-degree angle being less than a first threshold, and the deviation of its center of mass from the preset reference position in the forward and backward directions and the left and right directions being less than or equal to a second threshold. To ensure calibration accuracy, the first threshold range cannot exceed 0.1°; in this embodiment, 0.05° is selected. The second threshold cannot exceed 0.05mm; in this embodiment, 0.03mm is selected.

[0043] In the calibration process, the roll angle, pitch angle, and yaw angle of the spinning aircraft 1 are first calibrated, and then the position of the center of mass of the spinning aircraft 1 is calibrated. It is easy to understand that the roll angle refers to the angle by which the spinning aircraft rolls relative to its own axis; the pitch angle refers to the angle by which the spinning aircraft changes relative to the horizontal plane; and the yaw angle refers to the angle by which the spinning aircraft changes relative to the vertical plane.

[0044] The first step is to calibrate the roll angle. (Refer to...) Figure 2 When calibrating the roll angle of the spinning aircraft 1, a horizontal support plate 8 is fixedly installed on the spinning aircraft 1, and an inclinometer 4 is placed on the horizontal support plate 8 along the left-right direction. The relevant parameters of the cable traction system 3 are adjusted according to the angle displayed by the inclinometer 4 until the deviation between the roll angle of the spinning aircraft 1 and the zero-degree angle is less than a first threshold. The horizontal support plate 8 is located at the top of the spinning aircraft 1 and is set according to the center of mass position of the spinning aircraft 1, so it will not affect the attitude calibration of the spinning aircraft 1. A horizontal support surface is formed on the top of the horizontal support plate 8, on which the inclinometer 4 can be placed. The inclinometer 4 is an instrument that can display the angle with the horizontal direction. It should be noted that when calibrating the roll angle, the inclinometer 4 is placed in the left-right direction, at which time the inclinometer 4 can display the current roll angle of the spinning aircraft 1. When the deviation between the current roll angle of the spinning aircraft 1 and the zero-degree angle is greater than the first threshold, the relevant parameters of the rope traction system 3 can be adjusted, and the corresponding drive motor can be controlled to adjust the length of the corresponding traction cable, thereby adjusting the roll angle of the spinning aircraft 1 until its deviation from the zero-degree angle is less than the first threshold.

[0045] Next, the pitch angle is calibrated. (Refer to...) Figure 3When calibrating the pitch angle of the spinning aircraft 1, a horizontal support plate 8 is fixedly installed on the spinning aircraft 1, and an inclinometer 4 is placed on the horizontal support plate 8 along the front-to-back direction. The relevant parameters of the cable traction system 3 are adjusted according to the angle displayed by the inclinometer 4 until the deviation between the pitch angle of the spinning aircraft 1 and the zero-degree angle is less than a first threshold. Similar to calibrating the roll angle, the inclinometer 4 needs to be placed on the horizontal support plate 8 along the front-to-back direction. At this time, the inclinometer 4 can display the current pitch angle of the spinning aircraft 1. When the deviation between the current pitch angle of the spinning aircraft 1 and the zero-degree angle is greater than the first threshold, the relevant parameters of the cable traction system 3 can be adjusted, controlling the corresponding drive motor to adjust the length of the corresponding traction cable, thereby adjusting the pitch angle of the spinning aircraft 1 until its deviation from the zero-degree angle is less than the first threshold.

[0046] Next, the yaw angle is calibrated. It should be noted that because the horizontal support plate 8 is horizontally positioned, the yaw angle of the rotating aircraft 1 cannot be directly measured by placing the inclinometer 4 on the horizontal support plate 8. Therefore, an altitude vernier caliper 5 is used to indirectly measure the yaw angle of the rotating aircraft 1. (Refer to...) Figure 4 When calibrating the yaw angle of the spinning aircraft 1, a first measuring point and a second measuring point are determined on the spinning aircraft 1 at the same altitude but at different positions. The distance between the first measuring point and the second measuring point is measured. Then, the distance between the first measuring point and the second measuring point relative to the same sidewall of the closed wind tunnel 2 is measured using an altitude vernier caliper 5. Based on the yaw angle calculated from the above distance data, the relevant parameters of the rope traction system 3 are adjusted until the deviation of the yaw angle of the spinning aircraft 1 from zero degrees is less than a first threshold. Specifically, the altitude vernier caliper 5 can be an altitude ruler with a base. First, the first measuring point and the second measuring point are selected on the body of the spinning aircraft 1. The first measuring point and the second measuring point are at the same altitude, and the distance between them on the spinning aircraft 1 can be measured. At this point, the altimeter is placed horizontally, with a leveling block placed below the far end of the handle to ensure the altimeter is horizontally level. Then, the base of the altimeter is pressed against one side of the wind tunnel wall. The scale on the upper end of the handle is slid to align with the first measuring point, and the second measuring point is aligned in the same way. The yaw distances of the first and second measuring points relative to the wind tunnel wall against which the altimeter is pressed are obtained. The yaw angle can then be calculated using the arcsine function. When the deviation between the yaw angle of the spinning aircraft 1 and zero degrees exceeds a first threshold, the relevant parameters of the rope traction system 3 are adjusted, controlling the corresponding drive motor to adjust the length of the corresponding traction cable, thereby adjusting the yaw angle of the spinning aircraft 1 until its deviation from zero degrees is less than the first threshold.

[0047] Next, the center of mass position of the spinning-integrated aircraft 1 is calibrated. First, the center of mass position of the spinning-integrated aircraft 1 should be determined; this process can be calibrated on the spinning-integrated aircraft 1 during manufacturing. When calibrating the center of mass position of the spinning-integrated aircraft 1, a reference position is determined based on the dimensions of the closed wind tunnel 2. Then, the distance between the center of mass of the spinning-integrated aircraft 1 and the side and bottom walls of the closed wind tunnel 2 is measured using a height vernier caliper 5. Based on this distance, the relevant parameters of the rope traction system 3 are adjusted until the deviation of the center of mass position of the spinning-integrated aircraft 1 from the reference position in both the front-to-back and left-to-right directions is less than a second threshold. The reference position is determined based on the dimensions of the closed wind tunnel 2 and the actual experimental conditions, and can be predetermined before the experiment. Then, referring to… Figure 5 and Figure 6 The distances between the center of mass of the spinning aircraft 1 and the side and bottom walls of the closed wind tunnel 2 are measured respectively. When the deviation of the center of mass of the spinning aircraft 1 from the reference position is greater than the second threshold, the relevant parameters of the rope traction system 3 can be adjusted, and the corresponding drive motor can be controlled to adjust the length of the corresponding traction cable, thereby adjusting the deviation of the center of mass of the spinning aircraft 1 from the reference position in the front-back and left-right directions to be less than the second threshold.

[0048] After calibrating the attitude of the spinning vehicle 1 to zero attitude, a closed wind tunnel simulation flight test of the spinning vehicle 1 can be conducted. During the test, the attitude changes of the spinning vehicle 1 can be measured. In the measurement step, the spinning vehicle 1 is tested in the wind tunnel with zero attitude as the initial state. A calibration plate 6 is fixed at the tail of the spinning vehicle 1. A camera captures images of the calibration plate 6 at a preset frame rate, and the attitude change data of the spinning vehicle 1 is obtained based on the attitude change of the calibration plate 6. The calibration plate 6 has a calibration surface with N calibration points 7 arranged in a preset manner. When the spinning vehicle 1 is in zero attitude, the calibration surface faces the camera, and all calibration points 7 on the calibration surface fall within the camera's field of view.

[0049] Reference Figure 7 and Figure 8 The calibration plate 6 is located at the tail of the rotating integrated aircraft 1. During attitude calibration of the rotating integrated aircraft 1, the calibration plate 6 needs to be pre-installed on the aircraft 1. Calibration points 7 are formed on the calibration surface of the calibration plate 6. In this embodiment, there are nine calibration points 7 arranged in a square. In other embodiments, the number of calibration points 7 may be more or less, but should not be less than six.

[0050] During the measurement process, the rope traction system 3 adjusts the attitude of the spinning aircraft 1 in the airflow by adjusting the length of the traction cable. Since the calibration plate 6 is fixed on the spinning aircraft 1, the attitude change of the spinning aircraft 1 will cause the calibration plate 6 to change its attitude synchronously. Furthermore, the spatial relative relationship between the calibration plate 6 and the spinning aircraft 1 is fixed. Therefore, the attitude change of the spinning aircraft 1 can be inferred by measuring the attitude change of the calibration plate 6. Based on this principle, during the measurement process, the camera does not need to photograph the spinning aircraft 1, but only the calibration plate 6. Since the shape of the calibration plate 6 is fixed and the arrangement of the calibration points 7 has a strong regularity, the calculation efficiency is higher and the accuracy is higher than directly calculating the attitude change data through the motion changes of the spinning aircraft 1. At the same time, compared with setting sensors on the spinning aircraft 1 to detect the attitude change of the spinning aircraft 1, there is no need to worry about the sensor being affected by external factors and causing the measured data to be distorted.

[0051] Specifically, in the measurement steps, the attitude change data of the spinning aircraft 1 obtained based on the attitude change calculation of the calibration plate 6 includes: calibrating the camera intrinsic parameters and obtaining the camera coordinate system; calibrating the camera extrinsic parameters and obtaining the transformation relationship between the world coordinate system and the camera coordinate system; obtaining the transformation relationship between the projectile coordinate system and the camera coordinate system through the image of the calibration plate 6 captured by the camera; transforming the projectile coordinate system to the world coordinate system, and using the rotation matrix method to calculate and obtain the attitude change data of the spinning aircraft 1.

[0052] Reference Figure 9 It shows the position and meaning of each coordinate system in the camera imaging model.

[0053] The calibration of camera intrinsic parameters can be achieved using existing camera intrinsic parameter calibration software, such as Halcon. After this step, the camera coordinate system O... c -x c y c z c It can be determined.

[0054] Next, the camera extrinsic parameters were calibrated, and the world coordinate system O was obtained. w -x w y w z w and camera coordinate system O c -x c y c z c The transformation relationship is expressed by Equation 1:

[0055]

[0056] Specifically, a calibration plate is typically used to determine the world coordinate system. In this embodiment, the calibration plate 6 can be the calibration plate 6 on the rotating body aircraft 1. In Equation 1, c H w Let R be the pose matrix of the camera coordinate system relative to the world coordinate system, where the initial R is... cw T cw It can be obtained by least squares of the camera coordinate system and world coordinate system matrices.

[0057] Then, the coordinate system O of the projectile was obtained from the image of calibration board 6 taken by the camera. a -x a y a z a With camera coordinate system O c -x c y c z c The transformation relationship is expressed by Equation 2:

[0058]

[0059] The projectile coordinate system is used to represent the pose change of the calibration plate 6. Initially, the spinning aircraft 1 is in zero pose, at which point the projectile coordinate system coincides with the world coordinate system. When the spinning aircraft 1 begins to move, the camera captures images at a preset frequency, and the projectile coordinate system changes with the pose change of the calibration plate 6 in the images. c H a This is the pose matrix of the camera coordinate system relative to the projectile coordinate system.

[0060] Next, the projectile coordinate system is transformed to the world coordinate system, as shown in Equation 3:

[0061]

[0062] in, a H w Let be the pose matrix of the projectile coordinate system relative to the world coordinate system.

[0063] Subsequently, the attitude change data of the rotating body aircraft 1 was obtained by solving the rotation matrix method, including:

[0064] The relationship between the rotation matrix and the angle is expressed as Equation 4:

[0065]

[0066] By comparing specific coefficients in Equations 3 and 4, the attitude change data of the rotating body aircraft 1 can be obtained, namely Equation 5:

[0067] Pose = [t x t yt z arctan(r5 / r8) arctanr2 arctan(-r1 / 0)] T (Equation 5);

[0068] Among them, T aw =[t x ,t y ,t z ] T The displacements of the rotating aircraft 1 in the forward / backward direction (x-axis), the vertical direction (y-axis), and the horizontal direction (z-axis) are given. Let φ be the roll angle, θ be the yaw angle, and ψ be the pitch angle of the rotating aircraft 1; [r0 r3 r6] T For O w -x w Axis at O c -x c y c z c The direction vector of the coordinate system; [r1 r4 r7] T For O w -y w Axis at O c -x c y c z c The direction vector of the coordinate system; [r2 r5 r8] T For O w -z w Axis at O c -x c y c z c The direction vector of the coordinate system; [t x t y t z ] T For O w -x w y w z w The origin of the coordinate system is at O c -x c y c z c Position in the coordinate system.

[0069] In the above formula, x, y, and z in the projectile coordinate system represent the front-back direction with the airflow direction in the closed wind tunnel 2 as the front-back direction, the y-axis as the up-down direction, and the z-axis as the left-right direction, respectively. The same applies to other coordinate systems.

[0070] Through the above process, the attitude change data of the rotating body aircraft 1 can be calculated based on the image of the calibration board 6 captured by the camera.

[0071] The closed wind tunnel testing method for a spinning aircraft provided by this invention employs a rope traction system 3 to suspend and support the spinning aircraft 1. Within the relatively confined space of the closed wind tunnel 2, a larger dynamic attitude testing space can be provided for the spinning aircraft 1. Based on this, the attitude calibration and measurement of the spinning aircraft 1 are carried out. First, the attitude of the spinning aircraft 1 is calibrated through a calibration step, enabling the spinning aircraft 1 to be in the zero attitude for testing and measurement. Then, in the measurement step, the spinning aircraft 1... Wind tunnel testing commenced. During the wind tunnel test, a camera captured images of the calibration plate 6 on the spinning aircraft 1 at a fixed frequency. When the spinning aircraft 1 underwent attitude changes due to the cable traction system 3, since the calibration plate 6 was fixed to the spinning aircraft 1, the attitude changes of the spinning aircraft 1 could be calculated by analyzing the attitude changes of the calibration plate 6. To achieve this, several calibration points 7 were set on the calibration plate 6, and these calibration points 7 were arranged in a preset manner. With a sufficient number of calibration points, when the calibration plate 6 undergoes an attitude change, the position of the calibration points 7 on the calibration plate 6 captured by the camera will also change accordingly in the image. By observing the change in the position of the calibration points 7 in the image and the transformation relationship between the relevant coordinate systems, the attitude change of the spinning vehicle 1 in space can be obtained accordingly. Furthermore, since the calibration plate 6 is fixed to the tail of the spinning vehicle 1, the coordinate system of the calibration plate coincides with the coordinate system of the projectile (the three-dimensional positional relationship between the origins of the coordinate systems is known). The attitude change of the spinning vehicle 1 can be inferred from the attitude of the calibration plate 6. Therefore, the camera only needs to capture the calibration plate 6 at the tail of the spinning vehicle 1. Compared with the method of arranging calibration points on the body of the spinning vehicle 1 to calculate the attitude, the method of fixing the calibration plate 6 at the tail has relatively concentrated calibration points 7. The camera can capture a larger range of projectile attitude changes, providing more attitude measurement data for the aforementioned "larger dynamic attitude test workspace". In addition, using the calibration plate 6 to calculate the attitude can significantly reduce the amount of calculation and improve the calculation efficiency. Meanwhile, aerodynamic forces are closely related to attitude angles. After static calibration of the position and attitude, the error between the actual attitude angle and the zero position and attitude can be reduced. During the dynamic measurement of the position and attitude, visual measurement improves the measurement accuracy. Through the test method provided by this invention, the accuracy of the measurement data of the position and attitude and dynamic aerodynamic characteristics of the spinning aircraft can be improved.

[0072] The foregoing description of the specifications and embodiments is intended to explain the scope of protection of this invention, but does not constitute a limitation on the scope of protection of this invention. Modifications, equivalent substitutions, or other improvements to the embodiments of this invention or a portion thereof that can be obtained by those skilled in the art through logical analysis, reasoning, or limited experimentation, based on the teachings of this invention or the foregoing embodiments, in conjunction with common knowledge, general technical knowledge, and / or existing technology, should all be included within the scope of protection of this invention.

Claims

1. A method for testing the attitude of a suspended spin-type aircraft in a closed wind tunnel test, wherein the spin-type aircraft (1) is suspended and supported in the test section of a closed wind tunnel (2) by a rope traction system (3), characterized in that, include: The calibration step involves calibrating the spin-type aircraft (1) to a zero attitude. and Measurement steps: In this step, the spinning aircraft (1) is subjected to wind tunnel testing with zero attitude as the initial state. A calibration plate (6) is fixed at the tail of the spinning aircraft (1). The camera takes pictures of the calibration plate (6) at a preset frame rate and calculates the attitude change data of the spinning aircraft (1) based on the attitude change of the calibration plate (6). The calibration plate (6) has a calibration surface with N calibration points (7) arranged in a preset manner. When the spinning aircraft (1) is in zero attitude, the calibration surface faces the camera, and all calibration points (7) on the calibration surface are located at the center of the camera's field of view. The zero attitude refers to the fact that the deviation of the roll angle, pitch angle, and yaw angle of the spin-type aircraft (1) from the zero angle is less than the first threshold, and the deviation of its center of mass from the preset reference position in the forward and backward directions and the left and right directions is less than or equal to the second threshold. In the calibration step, the roll angle, pitch angle and yaw angle of the spin-type aircraft (1) are first calibrated, and then the position of the center of mass of the spin-type aircraft (1) is calibrated.

2. The method for testing the attitude of a suspended rotating aircraft in a closed wind tunnel test as described in claim 1, characterized in that, in When calibrating the roll angle of the spin-integrated aircraft (1), a horizontal support plate is fixedly installed on the spin-integrated aircraft (1), and an inclinometer (4) is placed on the horizontal support plate in the left and right directions. The relevant parameters of the rope traction system (3) are adjusted according to the angle displayed by the inclinometer (4) until the deviation between the roll angle of the spin-integrated aircraft (1) and the zero-degree angle is less than the first threshold.

3. The method for testing the attitude of a suspended rotating aircraft in a closed wind tunnel test as described in claim 2, characterized in that, in When calibrating the pitch angle of the spin-type aircraft (1), a horizontal support plate is fixedly installed on the spin-type aircraft (1), and an inclinometer (4) is placed on the horizontal support plate along the front-back direction. The relevant parameters of the rope traction system (3) are adjusted according to the angle displayed by the inclinometer (4) until the deviation between the pitch angle of the spin-type aircraft (1) and the zero-degree angle is less than the first threshold.

4. The method for testing the attitude of a suspended rotating aircraft in a closed wind tunnel test as described in claim 3, characterized in that, in When calibrating the yaw angle of the spin-type aircraft (1), a first measurement point and a second measurement point at the same altitude but different positions are determined on the spin-type aircraft (1), and the distance between the first measurement point and the second measurement point is measured. Then, the distance between the first measurement point and the second measurement point relative to the same side wall of the closed wind tunnel (2) is measured by using an altitude vernier caliper (5). The relevant parameters of the rope traction system (3) are adjusted based on the yaw angle calculated from the above distance data until the deviation between the yaw angle and zero degree angle of the spin-type aircraft (1) is less than the first threshold.

5. The method for testing the attitude of a suspended rotating aircraft in a closed wind tunnel test as described in claim 4, characterized in that, When calibrating the center of mass position of the spin-type aircraft (1), the reference position is determined according to the size of the closed wind tunnel (2). Then, the distance between the center of mass of the spin-type aircraft (1) and the side wall and bottom wall of the closed wind tunnel (2) is measured by using a height vernier caliper (5). The relevant parameters of the rope traction system (3) are adjusted according to the distance until the deviation of the center of mass position of the spin-type aircraft (1) from the reference position in the front-back direction and the left-right direction is less than the second threshold.

6. A method for testing the attitude of a suspended spinning aircraft in a closed wind tunnel as described in any one of claims 1-5, characterized in that, In the measurement step, obtaining the attitude change data of the rotating aircraft (1) based on the attitude change calculation of the calibration plate (6) includes: Calibrate the camera intrinsic parameters and obtain the camera coordinate system; Calibrate the camera's extrinsic parameters and obtain the transformation relationship between the world coordinate system and the camera coordinate system; The transformation relationship between the projectile coordinate system and the camera coordinate system is obtained by taking an image of the calibration plate (6) with the camera. The projectile coordinate system is transformed to the world coordinate system, and the attitude change data of the rotating body aircraft (1) is obtained by using the rotation matrix method.

7. The method for testing the attitude of a suspended rotating aircraft in a closed wind tunnel test as described in claim 6, characterized in that, In the measurement step, obtaining the attitude change data of the rotating aircraft (1) based on the attitude change calculation of the calibration plate (6) includes: Calibrate camera intrinsic parameters and obtain camera coordinate system O c -x c y c z c ; Calibrate camera extrinsic parameters and obtain world coordinate system O. w -x w y w z w and camera coordinate system O c -x c y c z c The transformation relationship is expressed by Equation 1: The projectile coordinate system O is obtained by taking an image of the calibration plate (6) with a camera. a -x a y a z a With camera coordinate system O c -x c y c z c The transformation relationship is expressed by Equation 2: Transform the projectile coordinate system to the world coordinate system, as shown in Equation 3: The attitude change data of the rotating body aircraft (1) is obtained by solving the rotation matrix method, including: The relationship between the rotation matrix and the angle is expressed as Equation 4: By comparing the corresponding position values ​​of the matrices in Equations 3 and 4 using the method of undetermined coefficients, and combining this with Equation 3, the attitude change data of the rotating aircraft (1) can be obtained, i.e., Equation 5: Pose = [t x t y t z arctan(r5 / r8) arctan r2 arctan(-r1 / r0)] T (Equation 5); Among them, T aw =[t x ,t y ,t z ] T The displacements of the rotating aircraft (1) in the front-to-back direction (x-axis), the up-down direction (y-axis), and the left-to-right direction (z-axis) are given. Let φ be the roll angle, θ be the yaw angle, and ψ be the pitch angle of the rotating aircraft (1); [r0r3r6] T For O w -x w Axis at O c -x c y c z c The direction vector of the coordinate system; [r1r4r7] T For O w -y w Axis at O c -x c y c z c The direction vector of the coordinate system; [r2r5r8] T For O w -z w Axis at O c -x c y c z c The direction vector of the coordinate system; c H w This is the pose matrix of the camera coordinate system relative to the world coordinate system; a H w Let be the pose matrix of the projectile coordinate system relative to the world coordinate system.

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