A method for inertial measurement device hardware-in-the-loop simulation

Abstract

This invention relates to a hardware-in-the-loop (HIL) simulation method for an inertial measurement unit (IMU), belonging to the field of simulation testing. The method avoids singularities in the turntable's motion by altering its pitch angle travel, thus solving the problem that the turntable cannot provide yaw and roll degrees of freedom within singular angle ranges, preventing the simulation of aircraft attitude angle changes. This method maintains consistent simulation accuracy with traditional methods, does not affect simulation results, exhibits good similarity to the actual field environment, and provides high simulation precision.

Description

Technical Field

[0001] This invention belongs to the field of simulation testing, specifically relating to a semi-physical simulation method for an inertial measurement device. Background Technology

[0002] The structural feature of a vertical three-axis turntable is that the outer ring frame axis is vertically upward, corresponding to the yaw angle motion of the aircraft. It has a large yaw rotation range and is suitable for simulating aircraft with a large yaw angle range. The vertical three-axis turntable has three independently moving gyro ring frames. The inner ring frame is supported on the middle ring frame along its X-axis and can rotate freely around the X-axis. The middle ring frame is supported on the outer ring frame, with its support axis being the Z-axis, which is perpendicular to the X-axis. When the middle ring frame rotates around its own Z-axis, it also drives the inner ring frame to rotate. The outer ring frame is supported on the base of the turntable via its rotation axis Y. When the outer ring frame rotates around its Y-axis, it simultaneously drives the middle and inner ring frames to rotate. Thus, the motion of the inner ring frame relative to the base of the turntable is equivalent to the attitude angle motion of the aircraft in the air around the axial coordinate system. By controlling the rotational motion of the three ring frames, the attitude angle motion of the aircraft in space can be completely reproduced from each ring frame. The coordinate system of the vertical three-axis turntable is as follows: Figure 1 As shown.

[0003] The outer ring axis (yaw axis) and inner ring axis (roll axis) of the vertical three-axis turntable are arranged in series. This type of turntable structure will produce a singularity when the pitch axis approaches or reaches ±90°. Generally, the singularity will appear between 80° and 90° or between -90° and -80° (that is, a singularity may occur when the pitch exceeds 80° or is less than -80°). This angular range is called the singular angle range. When the pitch angle moves within the singular angle range, due to the occurrence of the singularity, the turntable cannot provide the motion in both the yaw and roll directions. Therefore, it cannot simulate the attitude angle motion of the aircraft within the above-mentioned angle range, making it impossible to conduct the hardware-in-the-loop simulation test of the inertial measurement unit successfully. Summary of the Invention

[0004] (a) Technical problems to be solved

[0005] The technical problem to be solved by this invention is: how to provide a hardware-in-the-loop simulation method for an inertial measurement device to solve the problem that the turntable cannot provide the motion of the two degrees of freedom in the yaw and roll directions within the singular angle range, thus making it impossible to simulate the change of the aircraft's attitude angle.

[0006] (II) Technical Solution

[0007] To address the aforementioned technical problems, this invention provides a semi-physical simulation method for an inertial measurement device, comprising the following steps:

[0008] Step 1: The simulation computer generates the aircraft attitude motion signals and outputs them to the communication network. The attitude motion signals include: pitch angle θ, yaw angle ψ, roll angle γ, pitch angular velocity wz, yaw angular velocity wy, and roll angular velocity wx.

[0009] Step 2: The three-axis turntable is loaded with an inertial measurement unit (IMU). The IMU receives the pitch angle θ, yaw angle ψ, and roll angle γ from the simulation computer via the communication network, and uses these three angles as the pitch angle θ to control the attitude motion of the three-axis turntable. zt Yaw angle ψ zt and roll angle γ zt Three commands control the movement of the three-axis turntable to simulate the attitude changes of an aircraft during flight;

[0010] Step 3: Denote the pitch angle variation range of the aircraft during actual flight as α°~β°. When the three-axis turntable exhibits a singularity near the range of α°, let the maximum angle at which the three-axis turntable does not exhibit a singular value be α′°. Then, the difference between α′° and α is Δα=|α-α′|. Keep the pitch angle travel of the three-axis turntable constant, so that the pitch angle of the three-axis turntable moves between α′°~(β-Δα)°. Adjust the initial value of the pitch angle of the three-axis turntable to α′°, and adjust the initial values ​​of the yaw angle and roll angle of the three-axis turntable to ψ0° and γ0° respectively, where α′0 is the changed initial value of the pitch angle of the three-axis turntable, α′0=α-Δα, and ψ0 and γ0 are the initial values ​​of the yaw angle and roll angle respectively.

[0011] Step 4: Input the initial pitch angle θ0° into the inertial measurement unit under test via the simulation information interface. After the test begins, the simulation computer calculates the aircraft's attitude motion signals in real time, namely pitch angle θ, yaw angle ψ, and roll angle γ, and simultaneously generates three-axis turntable control commands:

[0012] θ zt =θ-Δα (1)

[0013] ψ zt =ψ (2)

[0014] γ zt =γ (3)

[0015] In formulas (1), (2), and (3), θ zt ψ zt γ zt In order to obtain the aircraft attitude motion signals through the inertial measurement device, the pitch angle, yaw angle and roll angle commands are used to control the motion of the three-axis turntable. The three-axis turntable adjusts its motion attitude according to the commands.

[0016] Step 5: The data acquisition device collects and records the command output results of the inertial measurement unit and transmits them to the data analysis computer. The data analysis computer compares the command output results of the inertial measurement unit with the actual response information of the three-axis turntable, including the turntable pitch angle, turntable yaw angle, turntable roll angle, turntable pitch angular velocity, turntable yaw angular velocity, and turntable roll angular velocity, to determine the accuracy of the inertial measurement unit.

[0017] In step 3, the travel of the pitch angle of the three-axis turntable is denoted as η. Then η = |β-α|. The three-axis turntable moves a total of η° in the pitch angle of the simulated aircraft attitude angle change.

[0018] The pitch angle θ obtained by the inertial measurement device in step 4 gd It is calculated using the following formula:

[0019] θ gd =θ gd0 +∫wz gd *t (4)

[0020] In formula (4), θ gd0 Let wz be the initial value of the pitch angle of the inertial measurement unit. gd θ is the real-time pitch angular velocity sensed by the angular velocity sensor in the inertial measurement device. From equation (4), it can be seen that θ gd It is independent of the commands controlling the turntable, that is, independent of formulas (1), (2), and (3), so that θ gd0 =θ0, then the inertial measurement device calculates θ gd The simulation results were consistent with the actual results obtained when the variation was made within the range of α° to β°.

[0021] In step 3, when a singularity occurs near the range of β°, let the maximum angle at which the three-axis turntable does not exhibit a singular value be β′°. Then, the difference between β′° and β° is Δβ=|β-β′|. Keeping the travel η of the pitch angle of the three-axis turntable constant, the pitch angle of the three-axis turntable moves between (α+Δβ)° and β′°, thus avoiding the range where the three-axis turntable exhibits a singular value.

[0022] When the three-axis turntable exhibits an anomaly near β°, the initial values ​​of its pitch, yaw, and roll angles are adjusted to α0′°, ψ0°, and γ0°, respectively. Here, α0′ is the initial pitch angle value after changing the pitch angle travel of the three-axis turntable, α′0 = α + Δβ, and ψ0 and γ0 are the initial values ​​of the yaw and roll angles, respectively. The actual initial pitch angle value θ0° is loaded into the measured inertial measurement device via the simulation information interface. After the test begins, the simulation computer runs the aircraft dynamics and kinematics model, calculating the aircraft's attitude motion signals in real time, namely pitch angle θ, yaw angle ψ, and roll angle γ, and simultaneously generating turntable control commands.

[0023] θ zt =θ+Δβ (5)

[0024] ψ zt =ψ (6)

[0025] γ zt =γ (7)

[0026] In formulas (5) to (7), θ zt ψ zt γ zt These are the pitch, yaw, and roll control commands for controlling the turntable's movement.

[0027] (III) Beneficial Effects

[0028] Compared with existing technologies, the present invention has the following advantages: the method maintains the same simulation accuracy as traditional simulation methods, does not affect the simulation effect, has a good similarity to the actual field environment, has high simulation accuracy, and is simple, effective, practical and versatile. Attached Figure Description

[0029] Figure 1 This is a schematic diagram of the coordinate system of the vertical three-axis rotary table of the present invention;

[0030] Figure 2 This is a block diagram illustrating the principle of the present invention. Detailed Implementation

[0031] To make the objectives, contents, and advantages of the present invention clearer, the specific embodiments of the present invention will be described in further detail below with reference to the accompanying drawings and examples.

[0032] To address the aforementioned technical problems, this embodiment provides a semi-physical simulation method for an inertial measurement device, such as... Figures 1-2 As shown, it includes the following steps:

[0033] Step 1: The simulation computer generates the aircraft attitude motion signals and outputs them to the communication network. The attitude motion signals include: pitch angle θ, yaw angle ψ, roll angle γ, pitch angular velocity wz, yaw angular velocity wy, and roll angular velocity wx.

[0034] Step 2: The three-axis turntable is loaded with an inertial measurement unit (IMU). The IMU receives the pitch angle θ, yaw angle ψ, and roll angle γ from the simulation computer via the communication network, and uses these three angles as the pitch angle θ to control the attitude motion of the three-axis turntable. zt Yaw angle ψ zt and roll angle γ zt Three commands control the movement of the three-axis turntable to simulate the attitude changes of an aircraft during flight;

[0035] Step 3: Denote the pitch angle variation range of the aircraft during actual flight as α°~β°. When the three-axis turntable exhibits a singularity near the range of α°, let the maximum angle at which the three-axis turntable does not exhibit a singular value be α′°. Then, the difference between α′° and α is Δα=|α-α′|. Keep the pitch angle travel of the three-axis turntable constant, so that the pitch angle of the three-axis turntable moves between α′°~(β-Δα)°. Adjust the initial value of the pitch angle of the three-axis turntable to α′°, and adjust the initial values ​​of the yaw angle and roll angle of the three-axis turntable to ψ0° and γ0° respectively, where α0′ is the changed initial value of the pitch angle of the three-axis turntable, α′0=α-Δα, and ψ0 and γ0 are the initial values ​​of the yaw angle and roll angle respectively.

[0036] Step 4: Input the initial pitch angle θ0° into the inertial measurement unit under test via the simulation information interface. After the test begins, the simulation computer calculates the aircraft's attitude motion signals in real time, namely pitch angle θ, yaw angle ψ, and roll angle γ, and simultaneously generates three-axis turntable control commands:

[0037] θ zt =θ-Δα (1)

[0038] ψ zt =ψ (2)

[0039] γ zt =γ (3)

[0040] In formulas (1), (2), and (3), θ zt ψ zt γ zt In order to obtain the aircraft attitude motion signals through the inertial measurement device, the pitch angle, yaw angle and roll angle commands are used to control the motion of the three-axis turntable. The three-axis turntable adjusts its motion attitude according to the commands.

[0041] Step 5: The data acquisition device collects and records the command output results of the inertial measurement unit and transmits them to the data analysis computer. The data analysis computer compares the command output results of the inertial measurement unit with the actual response information of the three-axis turntable, including the turntable pitch angle, turntable yaw angle, turntable roll angle, turntable pitch angular velocity, turntable yaw angular velocity, and turntable roll angular velocity, to determine the accuracy of the inertial measurement unit.

[0042] In step 3, the travel of the pitch angle of the three-axis turntable is denoted as η. Then η = |β-α|. The three-axis turntable moves a total of η° in the pitch angle of the simulated aircraft attitude angle change.

[0043] The pitch angle θ obtained by the inertial measurement device in step 4 gd It is calculated using the following formula:

[0044] θ gd =θgd0 +∫wz gd *t (4)

[0045] In formula (4), θ gd0 Let wz be the initial value of the pitch angle of the inertial measurement unit. gd θ is the real-time pitch angular velocity sensed by the angular velocity sensor in the inertial measurement device. From equation (4), it can be seen that θ gd It is independent of the commands controlling the turntable, that is, independent of formulas (1), (2), and (3), so that θ gd0 =θ0, then the inertial measurement device calculates θ gd The simulation results were consistent with the actual results obtained when the variation was made within the range of α° to β°.

[0046] In step 3, when a singularity occurs near the range of β°, let the maximum angle at which the three-axis turntable does not exhibit a singular value be β′°. Then, the difference between β′° and β° is Δβ=|β-β′|. Keeping the travel η of the pitch angle of the three-axis turntable constant, the pitch angle of the three-axis turntable moves between (α+Δβ)° and β′°, thus avoiding the range where the three-axis turntable exhibits a singular value.

[0047] When the three-axis turntable exhibits an anomaly near β°, the initial values ​​of its pitch, yaw, and roll angles are adjusted to α0′°, ψ0°, and γ0°, respectively. Here, α′0 is the initial pitch angle value after changing the pitch angle travel of the three-axis turntable, α′0 = α + Δβ, and ψ0 and γ0 are the initial values ​​of the yaw and roll angles, respectively. The actual initial pitch angle value θ0° is loaded into the measured inertial measurement device via the simulation information interface. After the test begins, the simulation computer runs the aircraft dynamics and kinematics model, calculating the aircraft's attitude motion signals in real time, namely pitch angle θ, yaw angle ψ, and roll angle γ, and simultaneously generating turntable control commands.

[0048] θ zt =θ+Δβ (5)

[0049] ψ zt =ψ (6)

[0050] γ zt =γ (7)

[0051] In formulas (5) to (7), θ zt ψ zt γ zt These are the pitch, yaw, and roll control commands for controlling the turntable's movement.

[0052] Example 2

[0053] In a hardware-in-the-loop simulation test of an inertial measurement unit (IPU) for a certain type of missile, the component involved in the test was the inertial navigation system. The simulation equipment used included a vertical three-axis turntable, a simulation computer, a simulation information interface, a data recording device, and a data analysis computer. The actual range of pitch angle movement during the missile's flight was 90° to -50°. The maximum angle at which the turntable did not exhibit singular values ​​was 80°. The initial values ​​of the actual pitch angle, yaw angle, and roll angle were 90°, 0°, and 0°, respectively. The specific implementation steps of the test are as follows:

[0054] (1) The inertial measurement device to be measured is mounted on a vertical three-axis turntable.

[0055] (2) Adjust the turntable to the ready-to-launch state, that is, adjust the initial values ​​of the turntable's attitude angles to the positions of pitch angle 80°, yaw angle 0°, and roll angle 0°.

[0056] (3) The initial value of the actual pitch angle of 90° is loaded into the inertial measurement device under test through the simulation information interface.

[0057] (4) After the test begins, the simulation computer runs the aircraft dynamics and kinematics model to calculate the aircraft's attitude motion signals, pitch angle θ, yaw angle ψ, and roll angle γ, in real time; at the same time, it generates turntable control commands:

[0058] θ zt =θ-(90-80) (1)

[0059] ψ zt =ψ (2)

[0060] γ zt =γ (3)

[0061] Using θ zt ψ zt γ zt The pitch, yaw, and roll angles of the turntable are controlled to move the turntable.

[0062] (5) The data acquisition device collects and records the output results of the inertial measurement device and transmits them to the data analysis computer. The data analysis computer compares the output results of the inertial measurement device with the actual response of the turntable, such as the turntable pitch angle, turntable yaw angle, turntable roll angle, turntable pitch angular velocity, turntable yaw angular velocity, and turntable roll angular velocity, and analyzes and evaluates whether the performance indicators of the inertial measurement device meet the design technical requirements.

[0063] The above description is only a preferred embodiment of the present invention. It should be noted that for those skilled in the art, several improvements and modifications can be made without departing from the technical principles of the present invention, and these improvements and modifications should also be considered within the scope of protection of the present invention.

Claims

1. A hardware-in-the-loop simulation method for an inertial measurement device, characterized in that, It includes the following steps: Step 1: The simulation computer generates the aircraft's attitude motion signals and outputs them to the communication network. The attitude motion signals include: pitch angle. Yaw angle Roll angle Pitch angular velocity wz yaw rate wy Roll angular velocity wx ; Step 2: The three-axis turntable is loaded with an inertial measurement unit (IMU). The IMU receives the pitch angle from the simulation computer via the communication network. Yaw angle Roll angle These three angles are used as pitch angles to control the attitude motion of the three-axis turntable. Yaw angle and roll angle Three commands control the movement of the three-axis turntable to simulate the attitude changes of an aircraft during flight; Step 3: Denote the range of pitch angle changes of the aircraft during actual flight as α°~β°. When a singularity occurs near α° on the three-axis turntable, let the maximum angle at which the three-axis turntable does not exhibit a singular value be α′°. Then, α′° and... The difference between To maintain a constant travel in the pitch angle of the three-axis turntable, the pitch angle is allowed to move between α′° and (β-Δα)°. The initial pitch angle of the three-axis turntable is adjusted to α0′°, and the initial yaw and roll angles are adjusted to ψ0° and γ0°, respectively. The initial value of the pitch angle of the modified three-axis turntable. , , These are the initial values ​​for the yaw angle and roll angle, respectively. Step 4: Input the initial pitch angle value θ0° into the inertial measurement unit under test via the simulation information interface. After the test begins, the simulation computer calculates the aircraft's attitude motion signal, i.e., the pitch angle, in real time. Yaw angle Roll angle Simultaneously, three-axis rotary table control commands are generated: （1） （2） （3） Among them, in formulas (1), (2), and (3) , , In order to obtain the aircraft attitude motion signals through the inertial measurement device, the pitch angle, yaw angle and roll angle commands are used to control the motion of the three-axis turntable. The three-axis turntable adjusts its motion attitude according to the commands. Step 5: The data acquisition device collects and records the command output results of the inertial measurement unit and transmits them to the data analysis computer. The data analysis computer compares the command output results of the inertial measurement unit with the actual response information of the three-axis turntable, including the turntable pitch angle, turntable yaw angle, turntable roll angle, turntable pitch angular velocity, turntable yaw angular velocity, and turntable roll angular velocity, to determine the accuracy of the inertial measurement unit. In step 3, the travel of the three-axis rotary table's pitch angle is recorded as... ,but = The three-axis turntable moved η° in the pitch angle of the simulated aircraft attitude angle change. The pitch angle obtained by the inertial measurement device in step 4 It is calculated using the following formula: （4） In formula (4), The initial value of the pitch angle of the inertial measurement device. Let be the real-time pitch angular velocity sensed by the angular velocity sensor in the inertial measurement device. As can be seen from equation (4), It is unrelated to the commands controlling the turntable, that is, unrelated to formulas (1), (2), and (3), making The inertial measurement device then calculates... The simulation results are consistent with the actual results obtained when the variation is made within the range of α° to β°. In step 3, when a singularity occurs in the range near β° of the three-axis turntable, let β′° be the maximum angle at which the three-axis turntable does not exhibit singular values. Then, the difference between β′° and β° is... Maintain the travel of the pitch angle of the three-axis turntable. By keeping the pitch angle of the three-axis turntable constant, the pitch angle of the turntable can be moved between (α+Δβ)° and β′°, thus avoiding the range of singular values ​​that may occur in the three-axis turntable. When a singularity occurs in the range near β° of the three-axis turntable, the initial values ​​of the pitch angle, yaw angle, and roll angle of the three-axis turntable are adjusted to positions α0′°, ψ0°, and γ0°, respectively. To change the initial pitch angle value after the pitch angle travel of the three-axis rotary table, , , The initial values ​​for yaw and roll angles are given. The actual initial value for pitch angle θ0° is then loaded into the inertial measurement unit (IMU) under test via the simulation information interface. After the experiment begins, the simulation computer runs the aircraft dynamics and kinematics model, and calculates the aircraft's attitude motion signal, i.e., pitch angle, in real time. Yaw angle Roll angle Simultaneously, turntable control commands are generated: （5） （6） （7） Among them, in formulas (5) to (7), , , These are the pitch, yaw, and roll control commands for controlling the turntable's movement.

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