Aircraft initial attitude checking method, device and equipment and readable storage medium

Abstract

The invention discloses an aircraft initial attitude checking method, device and equipment and a readable storage medium, and the method comprises the steps: carrying out the continuous updating of the attitude of an aircraft based on a first pitch angle, a first roll angle and a first course angle of the aircraft in the erecting process of the aircraft, calculating a second pitch angle, a second roll angle and a second course angle when the aircraft is in a vertical state after erecting is completed; when the aircraft is in the vertical state, calculating a third pitch angle and a third course angle of the aircraft; taking a course angle output by the inertial navigation system after executing the automatic alignment instruction again as a third roll angle of the aircraft; and respectively comparing the second pitch angle, the second roll angle and the second course angle of the aircraft with the third pitch angle, the third roll angle and the third course angle, and checking the precision of the initial attitude of the aircraft. According to the invention, the precision checking of the initial attitude of the aircraft does not depend on external equipment, and the cost and the risk of the external equipment are reduced.

Description

Technical Field

[0001] The present application relates to the field of aircraft orientation technology, and in particular to a method, device, equipment and readable storage medium for checking the initial attitude of an aircraft. Background Art

[0002] Before launch, aircraft require initial attitude calibration, which is crucial for precise navigation and control. Currently, aircraft typically use alignment and sighting methods to calibrate their initial attitude. Alignment methods include transfer alignment, where the aircraft's inertial navigation system (called the sub-INS) uses information from the aircraft's already aligned inertial navigation system (called the main INS) to perform initial attitude alignment. External sighting uses optical sighting equipment to manually measure the aircraft's azimuth angle, thereby calibrating the aircraft's initial attitude.

[0003] However, the use of the transfer alignment method for initial posture calibration requires two sets of equipment to aim at the same time, which increases costs and risks. The use of the external aiming method for initial posture calibration requires the addition of additional optical aiming equipment, which is complicated to operate and the equipment is easily affected by the external environment, which also increases costs and risks.

[0004] In summary, the current initial attitude calibration of aircraft before launch has problems such as reliance on external equipment, high cost and high risk. Summary of the Invention

[0005] The present application provides a method, device, equipment and readable storage medium for initial attitude calibration of an aircraft, aiming to solve the technical problems that the initial attitude calibration of an aircraft before launch currently has reliance on external equipment, high costs and high risks.

[0006] In a first aspect, an embodiment of the present application provides a method for calibrating an initial attitude of an aircraft, the method comprising:

[0007] When the aircraft is in a horizontal state, a first pitch angle and a first roll angle of the aircraft are calculated based on accelerations output by accelerometers of axes corresponding to the pitch angle and roll angle of an inertial navigation system of the aircraft;

[0008] The heading angle output by the inertial navigation system after executing the automatic alignment command is used as the first heading angle of the aircraft;

[0009] During the erection process of the aircraft, the attitude of the aircraft is continuously updated based on the first pitch angle, the first roll angle, and the first heading angle of the aircraft using the quaternion method and the differential solution method. The second pitch angle, the second roll angle, and the second heading angle of the aircraft when the aircraft is in the vertical state after the erection is completed are calculated based on the updated attitude of the aircraft;

[0010] When the aircraft is in a vertical state, a third pitch angle and a third heading angle of the aircraft are calculated based on the acceleration output again by the accelerometers of the axes corresponding to the pitch angle and roll angle of the aircraft's inertial navigation system;

[0011] The heading angle output by the inertial navigation system after executing the automatic alignment command again is used as the third roll angle of the aircraft;

[0012] The accuracy of the initial attitude of the aircraft is checked by comparing the second pitch angle, the second roll angle and the second heading angle of the aircraft with the third pitch angle, the third roll angle and the third heading angle respectively.

[0013] Optionally, calculating the first pitch angle and the first roll angle of the aircraft according to the accelerations output by accelerometers of axes corresponding to the pitch angle and the roll angle of the aircraft inertial navigation system includes:

[0014] The smoothing method is used to calculate the first pitch angle and the first roll angle of the aircraft using Formula 1 based on the acceleration output by the accelerometer of the axis corresponding to the pitch angle and the roll angle of the aircraft inertial navigation system. The Formula 1 is:

[0015]

[0016] Among them, θ0 is the first pitch angle, γ0 is the first roll angle, Ax i is the i-th acceleration output by the accelerometer of the pitch angle axis of the inertial navigation system during the smoothing calculation cycle, Az i is the i-th acceleration output by the accelerometer of the axis corresponding to the roll angle of the inertial navigation system during the smoothing calculation cycle, g0 is the local acceleration of gravity, and T is the total number of accelerations output by the accelerometer during the smoothing calculation cycle.

[0017] Optionally, the continuously updating the attitude of the aircraft based on the first pitch angle, the first roll angle, and the first heading angle of the aircraft by using the quaternion method and the differential solution method includes:

[0018] The initial attitude quaternion of the aircraft is established according to the first pitch angle, the first roll angle and the first heading angle of the aircraft through Formula 2, which is:

[0019]

[0020] An attitude update differential equation is established based on the initial attitude quaternion. The attitude of the aircraft is continuously updated based on the attitude update differential equation and the initial attitude quaternion, and the updated attitude quaternion is calculated. The attitude update differential equation is:

[0021]

[0022] Where Q is the initial attitude quaternion, q0, q1, q2 and q3 are the four components of the quaternion Q, i, j and k are the rotation components around the three orthogonal axes of the three-dimensional space, θ0 is the first pitch angle, γ0 is the first roll angle, Ψ0 is the first heading angle, is the updated attitude quaternion, which represents the derivative of each component of the quaternion with respect to time, is the quaternion matrix before updating, in and are the angular velocities of the x-axis, y-axis, and z-axis output by the gyroscope at the update time t, respectively. The superscript b represents the body coordinate system of the inertial navigation system.

[0023] Optionally, the calculating, based on the updated attitude of the aircraft, a second pitch angle, a second roll angle, and a second heading angle of the aircraft in a vertical state after the aircraft has been erected includes:

[0024] According to the updated attitude quaternion in the attitude update differential equation, the second pitch angle, second roll angle, and second heading angle of the aircraft in the vertical state after the vertical erection are calculated by formula 3. The formula 3 is:

[0025]

[0026] Among them, t is the second heading angle, θ t is the second pitch angle, γ t is the second roll angle.

[0027] Optionally, the performing accuracy check on the initial attitude of the aircraft by comparing the second pitch angle, the second roll angle, and the second heading angle of the aircraft with the third pitch angle, the third roll angle, and the third heading angle, respectively, includes:

[0028] If the absolute value of the difference between the second pitch angle and the third pitch angle of the aircraft is less than the first preset difference, the absolute value of the difference between the second roll angle and the third roll angle is less than the second preset difference, and the absolute value of the difference between the second heading angle and the third heading angle is less than the third preset difference, then the result of the initial attitude accuracy check of the aircraft is determined to be passed;

[0029] If the absolute value of the difference between the second pitch angle and the third pitch angle of the aircraft is not less than the first preset difference, or the absolute value of the difference between the second roll angle and the third roll angle is not less than the second preset difference, or the absolute value of the difference between the second heading angle and the third heading angle is not less than the third preset difference, then it is determined that the initial attitude accuracy calibration result of the aircraft is failed.

[0030] In a second aspect, an embodiment of the present application provides an aircraft initial attitude calibration device, the aircraft initial attitude calibration device comprising:

[0031] a first calculation module, configured to calculate a first pitch angle and a first roll angle of the aircraft based on accelerations output by accelerometers of axes corresponding to the pitch angle and roll angle of an inertial navigation system of the aircraft when the aircraft is in a horizontal state;

[0032] A first acquisition module is configured to use the heading angle output by the inertial navigation system after executing the automatic alignment instruction as the first heading angle of the aircraft;

[0033] a continuous updating module for continuously updating the attitude of the aircraft based on the first pitch angle, the first roll angle, and the first heading angle of the aircraft during the erection process of the aircraft using a quaternion method and a differential solution method, and calculating, based on the updated attitude of the aircraft, a second pitch angle, a second roll angle, and a second heading angle of the aircraft when the aircraft is in a vertical state after the erection is completed;

[0034] a second calculation module, configured to calculate, when the aircraft is in a vertical state, a third pitch angle and a third heading angle of the aircraft based on accelerations output again by the accelerometers of the axes corresponding to the pitch angle and roll angle of the aircraft's inertial navigation system;

[0035] A second acquisition module is used to use the heading angle output by the inertial navigation system after executing the automatic alignment instruction again as the third roll angle of the aircraft;

[0036] The calibration module is used to perform accuracy calibration on the initial attitude of the aircraft by comparing the second pitch angle, the second roll angle and the second heading angle of the aircraft with the third pitch angle, the third roll angle and the third heading angle respectively.

[0037] Optionally, the first calculation module is configured to:

[0038] The smoothing method is used to calculate the first pitch angle and the first roll angle of the aircraft using Formula 1 based on the acceleration output by the accelerometer of the axis corresponding to the pitch angle and the roll angle of the aircraft inertial navigation system. The Formula 1 is:

[0039]

[0040] Among them, θ0 is the first pitch angle, γ0 is the first roll angle, Ax i is the i-th acceleration output by the accelerometer of the pitch angle axis of the inertial navigation system during the smoothing calculation cycle, Az i is the i-th acceleration output by the accelerometer of the axis corresponding to the roll angle of the inertial navigation system during the smoothing calculation cycle, g0 is the local acceleration of gravity, and T is the total number of accelerations output by the accelerometer during the smoothing calculation cycle.

[0041] Optionally, the continuous updating module is used to:

[0042] The initial attitude quaternion of the aircraft is established according to the first pitch angle, the first roll angle and the first heading angle of the aircraft through Formula 2, which is:

[0043]

[0044] An attitude update differential equation is established based on the initial attitude quaternion. The attitude of the aircraft is continuously updated based on the attitude update differential equation and the initial attitude quaternion, and the updated attitude quaternion is calculated. The attitude update differential equation is:

[0045]

[0046] Where Q is the initial attitude quaternion, q0, q1, q2 and q3 are the four components of the quaternion Q, i, j and k are the rotation components around the three orthogonal axes of the three-dimensional space, θ0 is the first pitch angle, γ0 is the first roll angle, Ψ0 is the first heading angle, is the updated attitude quaternion, which represents the derivative of each component of the quaternion with respect to time, is the quaternion matrix before updating, in and are the angular velocities of the x-axis, y-axis, and z-axis output by the gyroscope at the update time t, respectively. The superscript b represents the body coordinate system of the inertial navigation system.

[0047] In the third aspect, an embodiment of the present application provides an aircraft initial attitude calibration device, which includes a processor, a memory, and an aircraft initial attitude calibration program stored in the memory and executable by the processor, wherein when the aircraft initial attitude calibration program is executed by the processor, the steps of the aircraft initial attitude calibration method described above are implemented.

[0048] In a fourth aspect, an embodiment of the present application provides a readable storage medium, on which is stored an aircraft initial attitude calibration program, wherein when the aircraft initial attitude calibration program is executed by a processor, the steps of the aircraft initial attitude calibration method as described above are implemented.

[0049] The beneficial effects of the technical solutions provided in the embodiments of the present application include:

[0050] In the embodiment of the present application, when the aircraft is in a horizontal state, the first pitch angle and the first roll angle of the aircraft are calculated according to the acceleration output by the accelerometer of the axis corresponding to the pitch angle and the roll angle of the aircraft's inertial navigation system; the heading angle output after the inertial navigation system executes the automatic alignment instruction is used as the first heading angle of the aircraft; during the vertical take-off process of the aircraft, the quaternion method and the differential solution method are used to continuously update the attitude of the aircraft based on the first pitch angle, the first roll angle and the first heading angle of the aircraft, and the vertical position of the aircraft after the vertical take-off is completed is calculated based on the updated attitude of the aircraft. the second pitch angle, second roll angle and second heading angle in the vertical state; when the aircraft is in the vertical state, the third pitch angle and third heading angle of the aircraft are calculated according to the acceleration output again by the accelerometer of the axis corresponding to the pitch angle and roll angle of the aircraft's inertial navigation system; the heading angle output after the inertial navigation system executes the automatic alignment command again is used as the third roll angle of the aircraft; the initial attitude of the aircraft is accurately calibrated by comparing the second pitch angle, second roll angle and second heading angle of the aircraft with the third pitch angle, third roll angle and third heading angle respectively. According to the embodiments of the present application, during the erection process of an aircraft, the quaternion method and differential solution method are used to continuously update the attitude of the aircraft based on the first pitch angle, first roll angle, and first heading angle of the aircraft in the horizontal state. The second pitch angle, second roll angle, and second heading angle of the aircraft in the vertical state after the erection is completed are calculated. The first pitch angle, first roll angle, and first heading angle of the aircraft are obtained based on the acceleration output by the inertial navigation system and the automatic alignment function when the aircraft is in the horizontal state. When the aircraft is in the vertical state, the third pitch angle, third roll angle, and third heading angle of the aircraft are again obtained based on the acceleration output by the inertial navigation system and the automatic alignment function. The second pitch angle, second roll angle, and second heading angle of the aircraft are compared with the third pitch angle, third roll angle, and third heading angle respectively to perform an accuracy check on the initial attitude of the aircraft. During the entire accuracy check process of the initial attitude of the aircraft, no external equipment is required, and only the aircraft's own inertial navigation system is required. This reduces the cost and risk of relying on external equipment, and has the advantages of accurate and simple calibration. BRIEF DESCRIPTION OF THE DRAWINGS

[0051] Figure 1 This is a flow chart of an embodiment of a method for calibrating an aircraft's initial attitude;

[0052] Figure 2 For this application Figure 1 Detailed flow chart of step S60;

[0053] Figure 3 This is a schematic diagram of the functional modules of an embodiment of the aircraft initial attitude calibration device of the present application;

[0054] Figure 4 This is a schematic diagram of the hardware structure of the aircraft initial attitude calibration equipment involved in the embodiment of the present application. DETAILED DESCRIPTION

[0055] In order to enable those skilled in the art to better understand the present invention, the following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of this application.

[0056] In order to make the objectives, technical solutions and advantages of this application clearer, the implementation methods of this application will be further described in detail below with reference to the accompanying drawings.

[0057] In a first aspect, an embodiment of the present application provides a method for calibrating an aircraft's initial attitude.

[0058] In one embodiment, referring to Figure 1 , Figure 1 This is a flow chart of an embodiment of the method for checking the initial attitude of an aircraft of the present application. Figure 1 As shown, the method for checking the initial attitude of the aircraft includes:

[0059] Step S10 , when the aircraft is in a horizontal state, a first pitch angle and a first roll angle of the aircraft are calculated based on the accelerations output by the accelerometers of the axes corresponding to the pitch angle and the roll angle of the aircraft's inertial navigation system.

[0060] In this embodiment, the aircraft's initial attitude needs to be accurately calibrated before launch to achieve precise navigation and control. The inertial navigation system (INS) consists of an inertial measurement unit (IMU) and a data processing and computing module. The IMU includes a gyroscope and an accelerometer. The gyroscope measures the aircraft's angular velocity, i.e., the aircraft's rotation rate around three spatial axes, while the accelerometer measures the aircraft's linear acceleration. Before launch, the aircraft is horizontally positioned on a launch platform and kept stationary. The INS is installed inside the aircraft. When the aircraft is horizontal, the X- and Z-axes of the INS are horizontal, and the Y-axis is vertically pointing upward. Rotation about the X-axis represents the roll angle, rotation about the Y-axis represents the heading angle, and rotation about the Z-axis represents the pitch angle. The aircraft's attitude angles remain consistent with those of the INS. Therefore, based on the accelerations output by the accelerometers of the axes corresponding to the pitch angle and roll angle of the aircraft's inertial navigation system, namely, the z-axis and x-axis accelerometers, the pitch angle and roll angle of the inertial navigation system are calculated to be the first pitch angle and first roll angle of the aircraft. The first pitch angle and first roll angle are the first pitch angle and first roll angle of the aircraft when it is in a horizontal state before launch.

[0061] Step S20: Using the heading angle output by the inertial navigation system after executing the automatic alignment instruction as the first heading angle of the aircraft.

[0062] In this embodiment, an auto-alignment command is issued to the inertial navigation system. The inertial navigation system executes the command and then uses an internal algorithm to calculate and output a heading angle. Because the attitude angles of the aircraft in a horizontal state are consistent with those of the inertial navigation system, the heading angle calculated and output by the inertial navigation system is used as the aircraft's first heading angle. The first heading angle is the first heading angle of the aircraft in a horizontal state before launch. Auto-alignment is a built-in function of the inertial navigation system. The inertial navigation system can use a variety of internal algorithms to calculate the heading angle for auto-alignment, and these are not limited here.

[0063] It should be noted that the reason why the heading angle output by the inertial navigation system after executing the automatic alignment command is used as the first heading angle of the aircraft is that when the aircraft's transmitter is in a horizontal state, the accelerometer can sense the gravity component when it is stationary or at a constant speed. The first pitch angle and first roll angle of the aircraft can be calculated by decomposing the acceleration output by the accelerometer of the axis corresponding to the pitch angle and roll angle, but the heading angle cannot be directly obtained through the gravity component because it only reflects the inclination of the carrier relative to the vertical line of the ground and does not contain direction information. Therefore, it is necessary to use the heading angle output by the inertial navigation system after executing the automatic alignment command as the first heading angle of the aircraft.

[0064] Step S30: During the vertical takeoff process of the aircraft, the attitude of the aircraft is continuously updated based on the first pitch angle, the first roll angle and the first heading angle of the aircraft using the quaternion method and the differential solution method. The second pitch angle, the second roll angle and the second heading angle of the aircraft when the aircraft is in a vertical state after the vertical takeoff is completed are calculated based on the updated attitude of the aircraft.

[0065] In this embodiment, the aircraft's erection process refers to the critical stage in which the aircraft adjusts from a stationary or horizontally parked position on the ground to a vertical or predetermined takeoff attitude, completing preparations for launch. Based on the first pitch angle, first roll angle, and first heading angle acquired in steps S10 and S20 while the aircraft was in a horizontal position before launch, the aircraft's attitude is continuously updated during the erection process using quaternion methods and differential solution methods until the aircraft is in a vertical position after erection is complete. The second pitch angle, second roll angle, and second heading angle are then calculated for the aircraft in the vertical position.

[0066] Step S40 , when the aircraft is in a vertical state, a third pitch angle and a third heading angle of the aircraft are calculated based on the accelerations output again by the accelerometers of the axes corresponding to the pitch angle and roll angle of the aircraft inertial navigation system.

[0067] In this embodiment, the same method as step S10 is used, with the difference that when the aircraft is in a vertical state, the y-axis and z-axis of the inertial navigation system are horizontal, the x-axis is vertically upward, rotation around the x-axis is the heading angle, rotation around the y-axis is the roll angle, and rotation around the z-axis is the pitch angle. The attitude angles of the aircraft in the vertical state and the horizontal state are consistent, that is, the pitch angle of the inertial navigation system in the vertical state corresponds to the pitch angle of the aircraft, the roll angle of the inertial navigation system corresponds to the heading angle of the aircraft, and the heading angle of the inertial navigation system corresponds to the roll angle of the aircraft. Therefore, the pitch angle and roll angle of the inertial navigation system calculated here based on the acceleration output again by the accelerometers of the axes corresponding to the pitch angle and roll angle of the aircraft's inertial navigation system, that is, the z-axis and y-axis accelerometers, are the third pitch angle and third heading angle when the aircraft is in the vertical state.

[0068] Step S50: The heading angle output by the inertial navigation system after executing the automatic alignment instruction again is used as the third roll angle of the aircraft.

[0069] In this embodiment, the same method as step S20 is used to again send an automatic alignment command to the inertial navigation system. The inertial navigation system executes the automatic alignment command and then uses an internal algorithm to calculate and output a heading angle. The difference is that the heading angle output by the inertial navigation system in the vertical state corresponds to the third roll angle of the aircraft.

[0070] Step S60 , performing accuracy check on the initial attitude of the aircraft by comparing the second pitch angle, the second roll angle, and the second heading angle of the aircraft with the third pitch angle, the third roll angle, and the third heading angle respectively.

[0071] In this embodiment, the aforementioned steps indicate that the second pitch angle, second roll angle, and second heading angle of the aircraft, as well as the third pitch angle, third roll angle, and third heading angle, are all attitude angles of the aircraft in its current vertical state. The difference is that the second pitch angle, second roll angle, and second heading angle of the aircraft are calculated by continuously updating the aircraft's attitude when the aircraft is in its horizontal state based on the first pitch angle, first roll angle, and first heading angle, while the third pitch angle, third roll angle, and third heading angle of the aircraft are again calculated based on the acceleration output by the inertial navigation system and the auto-alignment function. Therefore, by comparing the second pitch angle, second roll angle, and second heading angle calculated using the two methods with the third pitch angle, third roll angle, and third heading angle, a self-calibration of the aircraft's initial attitude can be achieved. Specifically, the entire initial attitude accuracy calibration process for the aircraft does not require reliance on external equipment, but only utilizes the aircraft's own inertial navigation system, reducing the cost and risk of reliance on external equipment while offering the advantages of accurate and simple calibration. Among them, respectively comparing means comparing the second pitch angle and the third pitch angle of the aircraft, comparing the second roll angle and the third roll angle, and comparing the second heading angle and the third heading angle, such as checking whether the difference between each attitude angle is too large through comparison, and whether it meets the accuracy verification requirements of the initial attitude of the aircraft.

[0072] In this embodiment, the aircraft is horizontally positioned on a launch platform and remains stationary before launch. An inertial navigation system is installed within the aircraft. The pitch and roll angles of the inertial navigation system, i.e., the first pitch and roll angles of the aircraft in the horizontal state, are calculated based on the accelerations output by the accelerometers of the axes corresponding to the pitch and roll angles of the inertial navigation system. An automatic alignment command is then issued to the inertial navigation system. The inertial navigation system executes the automatic alignment command and then uses an internal algorithm to calculate and output a heading angle. Since the attitude angles of the aircraft in the horizontal state are consistent with the attitude angles of the inertial navigation system, the heading angle calculated and output by the inertial navigation system is used as the first heading angle of the aircraft in the horizontal state. Based on the first pitch, roll, and heading angles of the aircraft in the horizontal state before launch, the attitude of the aircraft is continuously updated during the erection process using a quaternion method and a differential solution method until the aircraft is in the vertical state after erection is completed. A second pitch, roll, and heading angle of the aircraft in the vertical state are then calculated. When the aircraft is in the vertical state, the second pitch, roll, and heading angles are calculated. In the vertical state, the pitch and roll angles of the inertial navigation system are calculated based on the accelerations output again by the accelerometers of the axes corresponding to the pitch and roll angles of the aircraft's inertial navigation system in the vertical state, i.e., the third pitch angle and third heading angle of the aircraft in the vertical state. An automatic alignment command is again issued to the inertial navigation system. The inertial navigation system executes the automatic alignment command and then uses an internal algorithm to calculate and output a heading angle. The heading angle again output by the inertial navigation system corresponds to the third heading angle of the aircraft in the vertical state. Furthermore, by comparing the second pitch angle, second roll angle, and second heading angle calculated using the two methods with the third pitch angle, third roll angle, and third heading angle, respectively, the accuracy of the aircraft's initial attitude can be self-calibrated. That is, the entire process of calibrating the aircraft's initial attitude does not require reliance on external equipment. This eliminates the need for external equipment, reduces the cost of external equipment, the complexity of manual operations, and the risk of error. Only the aircraft's own inertial navigation system is required, and this offers the advantages of accurate and simple calibration, significantly shortening the time required for initial attitude accuracy calibration before launch.

[0073] Furthermore, in one embodiment, calculating the first pitch angle and the first roll angle of the aircraft based on the accelerations output by the accelerometers of the axes corresponding to the pitch angle and the roll angle of the aircraft inertial navigation system includes:

[0074] The smoothing method is used to calculate the first pitch angle and the first roll angle of the aircraft using Formula 1 based on the acceleration output by the accelerometer of the axis corresponding to the pitch angle and the roll angle of the aircraft inertial navigation system. The Formula 1 is:

[0075]

[0076] Among them, θ0 is the first pitch angle, γ0 is the first roll angle, Ax i is the i-th acceleration output by the accelerometer of the pitch angle axis of the inertial navigation system during the smoothing calculation cycle, Az i is the i-th acceleration output by the accelerometer of the axis corresponding to the roll angle of the inertial navigation system during the smoothing calculation cycle, g0 is the local acceleration of gravity, and T is the total number of accelerations output by the accelerometer during the smoothing calculation cycle.

[0077] In this embodiment, a smoothing method is used to obtain multiple acceleration values ​​output by the accelerometers of the axes corresponding to the pitch angle and roll angle of the aircraft inertial navigation system during a smoothing calculation period. Formula 1 is used to perform smoothing calculation based on the multiple acceleration values ​​output by the accelerometers of the axes corresponding to the pitch angle and roll angle, which can make the calculated first pitch angle and first roll angle of the aircraft more accurate.

[0078] Furthermore, in one embodiment, the continuously updating the attitude of the aircraft based on the first pitch angle, the first roll angle, and the first heading angle of the aircraft using the quaternion method and the differential solution method includes:

[0079] The initial attitude quaternion of the aircraft is established according to the first pitch angle, the first roll angle and the first heading angle of the aircraft through Formula 2, which is:

[0080]

[0081] An attitude update differential equation is established based on the initial attitude quaternion. The attitude of the aircraft is continuously updated based on the attitude update differential equation and the initial attitude quaternion, and the updated attitude quaternion is calculated. The attitude update differential equation is:

[0082]

[0083] Where Q is the initial attitude quaternion, q0, q1, q2 and q3 are the four components of the quaternion Q, i, j and k are the rotation components around the three orthogonal axes of the three-dimensional space, θ0 is the first pitch angle, γ0 is the first roll angle, Ψ0 is the first heading angle, is the updated attitude quaternion, which represents the derivative of each component of the quaternion with respect to time, is the quaternion matrix before updating, in and are the angular velocities of the x-axis, y-axis, and z-axis output by the gyroscope at the update time t, respectively. The superscript b represents the body coordinate system of the inertial navigation system.

[0084] In this embodiment, quaternions are an extension of complex numbers. The quaternion method describes three-dimensional rotations through four-dimensional hypercomplex numbers, avoiding the universal lock problem of Euler angles while having the advantages of high computational efficiency and no singularity. Describes the change of posture over time, where ω x 、ω y and ω z is the angular velocity measured by the gyroscope, Represents quaternion multiplication, Q is the quaternion of the current attitude, The quaternion of the attitude at the next moment can be expanded to obtain the above-mentioned attitude update differential equation. During the verticalization process of the aircraft, the above-mentioned attitude update differential equation is used to continuously update the attitude of the aircraft. That is, in each current cycle, the attitude quaternion of the current cycle is calculated based on the attitude quaternion of the previous cycle. In this way, the attitude quaternion of the aircraft in the vertical state after the verticalization is completed can be obtained.

[0085] Furthermore, in one embodiment, the calculating of the second pitch angle, the second roll angle, and the second heading angle of the aircraft in the vertical state after the aircraft is erected based on the updated attitude of the aircraft includes:

[0086] According to the updated attitude quaternion in the attitude update differential equation, the second pitch angle, second roll angle, and second heading angle of the aircraft in the vertical state after the vertical erection are calculated by formula 3. The formula 3 is:

[0087]

[0088] Among them, t is the second heading angle, θ t is the second pitch angle, γ t is the second roll angle.

[0089] In this embodiment, during the erection process of the aircraft, the above-mentioned attitude update differential equation is used to continuously update the attitude of the aircraft. The latest attitude quaternion after the aircraft is erected is the attitude quaternion of the aircraft in the vertical state. Through Formula 3, the attitude quaternion of the aircraft in the vertical state can be used to calculate the second pitch angle, second roll angle and second heading angle of the aircraft in the vertical state after the erection is completed.

[0090] Furthermore, in one embodiment, referring to Figure 2 , Figure 2 For this application Figure 1 The detailed flow chart of step S60 is as follows: Figure 2 As shown, step S60 includes:

[0091] Step S601: If the absolute value of the difference between the second pitch angle and the third pitch angle of the aircraft is less than the first preset difference, the absolute value of the difference between the second roll angle and the third roll angle is less than the second preset difference, and the absolute value of the difference between the second heading angle and the third heading angle is less than the third preset difference, then the result of the initial attitude accuracy calibration of the aircraft is determined to be passed;

[0092] In step S602, if the absolute value of the difference between the second pitch angle and the third pitch angle of the aircraft is not less than the first preset difference, or the absolute value of the difference between the second roll angle and the third roll angle is not less than the second preset difference, or the absolute value of the difference between the second heading angle and the third heading angle is not less than the third preset difference, then it is determined that the initial attitude accuracy calibration result of the aircraft is failed.

[0093] In this embodiment, the absolute values ​​of the differences between the attitude angles calculated using the two methods are compared with the corresponding preset differences. If the absolute values ​​of the differences between the two methods for each attitude angle are less than the corresponding preset differences, it means that the initial attitude angles of the aircraft meet the accuracy verification requirements. If the absolute values ​​of the differences between the two methods for a certain attitude angle are not less than the corresponding preset differences, it means that the attitude angle of the initial attitude of the aircraft does not meet the accuracy verification requirements. The initial attitude of the aircraft as a whole cannot meet the accuracy verification requirements. The reason may be that there is a problem with the equipment of the inertial navigation system or a problem with the data acquisition calculation, and further problem troubleshooting is required. In the case that the initial attitude accuracy verification result of the aircraft fails, in order to ensure the accurate navigation and control of the aircraft, the aircraft can be launched after the problem is eliminated or the accuracy verification is passed.

[0094] In a second aspect, an embodiment of the present application also provides an aircraft initial attitude calibration device.

[0095] In one embodiment, referring to Figure 3 , Figure 3 This is a functional module diagram of an embodiment of the aircraft initial attitude calibration device of the present application, as shown in FIG. Figure 3 As shown, the aircraft initial attitude calibration device includes:

[0096] A first calculation module 10 is configured to calculate a first pitch angle and a first roll angle of the aircraft based on accelerations output by accelerometers of axes corresponding to the pitch angle and roll angle of an inertial navigation system of the aircraft when the aircraft is in a horizontal state;

[0097] A first acquisition module 20 is configured to use the heading angle output by the inertial navigation system after executing the automatic alignment instruction as the first heading angle of the aircraft;

[0098] a continuous updating module 30 for continuously updating the attitude of the aircraft based on the first pitch angle, the first roll angle, and the first heading angle of the aircraft during the erection process using a quaternion method and a differential solution method, and calculating, based on the updated attitude of the aircraft, a second pitch angle, a second roll angle, and a second heading angle of the aircraft when the aircraft is in an upright state after the erection is completed;

[0099] A second calculation module 40 is configured to calculate a third pitch angle and a third heading angle of the aircraft based on the accelerations output again by the accelerometers of the axes corresponding to the pitch angle and the roll angle of the aircraft's inertial navigation system when the aircraft is in a vertical state;

[0100] A second acquisition module 50 is configured to use the heading angle output by the inertial navigation system after executing the automatic alignment instruction again as a third roll angle of the aircraft;

[0101] The calibration module 60 is used to perform accuracy calibration on the initial attitude of the aircraft by comparing the second pitch angle, the second roll angle and the second heading angle of the aircraft with the third pitch angle, the third roll angle and the third heading angle respectively.

[0102] Furthermore, in one embodiment, the first calculation module 10 is configured to:

[0103] The smoothing method is used to calculate the first pitch angle and the first roll angle of the aircraft using Formula 1 based on the acceleration output by the accelerometer of the axis corresponding to the pitch angle and the roll angle of the aircraft inertial navigation system. The Formula 1 is:

[0104]

[0105] Among them, θ0 is the first pitch angle, γ0 is the first roll angle, Ax i is the i-th acceleration output by the accelerometer of the pitch angle axis of the inertial navigation system during the smoothing calculation cycle, Az i is the i-th acceleration output by the accelerometer of the axis corresponding to the roll angle of the inertial navigation system during the smoothing calculation cycle, g0 is the local acceleration of gravity, and T is the total number of accelerations output by the accelerometer during the smoothing calculation cycle.

[0106] Furthermore, in one embodiment, the continuous updating module 30 is configured to:

[0107] The initial attitude quaternion of the aircraft is established according to the first pitch angle, the first roll angle and the first heading angle of the aircraft through Formula 2, which is:

[0108]

[0109] An attitude update differential equation is established based on the initial attitude quaternion. The attitude of the aircraft is continuously updated based on the attitude update differential equation and the initial attitude quaternion, and the updated attitude quaternion is calculated. The attitude update differential equation is:

[0110]

[0111] Where Q is the initial attitude quaternion, q0, q1, q2 and q3 are the four components of the quaternion Q, i, j and k are the rotation components around the three orthogonal axes of the three-dimensional space, θ0 is the first pitch angle, γ0 is the first roll angle, Ψ0 is the first heading angle, is the updated attitude quaternion, which represents the derivative of each component of the quaternion with respect to time, is the quaternion matrix before updating, in and are the angular velocities of the x-axis, y-axis, and z-axis output by the gyroscope at the update time t, respectively. The superscript b represents the body coordinate system of the inertial navigation system.

[0112] Furthermore, in one embodiment, the continuous updating module 30 is configured to:

[0113] According to the updated attitude quaternion in the attitude update differential equation, the second pitch angle, second roll angle, and second heading angle of the aircraft in the vertical state after the vertical erection are calculated by formula 3. The formula 3 is:

[0114]

[0115] Among them, t is the second heading angle, θ t is the second pitch angle, γ t is the second roll angle.

[0116] Furthermore, in one embodiment, the verification module 60 is configured to:

[0117] If the absolute value of the difference between the second pitch angle and the third pitch angle of the aircraft is less than the first preset difference, the absolute value of the difference between the second roll angle and the third roll angle is less than the second preset difference, and the absolute value of the difference between the second heading angle and the third heading angle is less than the third preset difference, then the result of the initial attitude accuracy check of the aircraft is determined to be passed;

[0118] If the absolute value of the difference between the second pitch angle and the third pitch angle of the aircraft is not less than the first preset difference, or the absolute value of the difference between the second roll angle and the third roll angle is not less than the second preset difference, or the absolute value of the difference between the second heading angle and the third heading angle is not less than the third preset difference, then it is determined that the initial attitude accuracy calibration result of the aircraft is failed.

[0119] Among them, the functional implementation of each module in the above-mentioned aircraft initial attitude calibration device corresponds to the various steps in the above-mentioned aircraft initial attitude calibration method embodiment, and its functions and implementation processes will not be repeated here one by one.

[0120] In a third aspect, an embodiment of the present application provides an aircraft initial attitude calibration device.

[0121] Reference Figure 4 , Figure 4 Schematic diagram of the hardware structure of the aircraft initial attitude calibration device involved in the embodiment of the present application. In the embodiment of the present application, the aircraft initial attitude calibration device may include a processor, a memory, a communication interface and a communication bus.

[0122] The communication bus may be of any type and is used to interconnect the processor, memory, and communication interface.

[0123] Communication interfaces include input / output (I / O) interfaces, physical interfaces, and logical interfaces. These interfaces interconnect components within the aircraft initial attitude calibration equipment, as well as interfaces that connect the aircraft initial attitude calibration equipment to other devices (such as other computing devices or user devices). Physical interfaces can include Ethernet, fiber optic, and ATM interfaces; user devices can include displays and keyboards.

[0124] The memory can be various types of storage media, such as random access memory (RAM), read-only memory (ROM), non-volatile RAM (NVRAM), flash memory, optical storage, hard disk, programmable ROM (PROM), erasable PROM (EPROM), electrically erasable PROM (EEPROM), etc.

[0125] The processor may be a general-purpose processor that can call an aircraft initial attitude calibration program stored in a memory and execute the aircraft initial attitude calibration method provided in the embodiments of the present application. For example, the general-purpose processor may be a central processing unit (CPU). The method executed when the aircraft initial attitude calibration program is called can be referenced to the various embodiments of the aircraft initial attitude calibration method of the present application and will not be further described here.

[0126] Those skilled in the art will understand that Figure 4 The hardware structure shown in the figure does not constitute a limitation to the present application and may include more or fewer components than shown in the figure, or a combination of certain components, or a different arrangement of components.

[0127] In a fourth aspect, an embodiment of the present application also provides a readable storage medium.

[0128] The readable storage medium of the present application stores an aircraft initial attitude calibration program, wherein when the aircraft initial attitude calibration program is executed by the processor, the steps of the aircraft initial attitude calibration method as described above are implemented.

[0129] Among them, the method implemented when the aircraft initial attitude calibration program is executed can refer to the various embodiments of the aircraft initial attitude calibration method of the present application, and will not be repeated here.

[0130] It should be noted that the serial numbers of the above-mentioned embodiments of the present application are for description only and do not represent the advantages or disadvantages of the embodiments.

[0131] The terms "including" and "having" and any variations thereof in the specification and claims of this application and the above-mentioned drawings are intended to cover non-exclusive inclusions. For example, a process, method, system, product or device that includes a series of steps or units is not limited to the listed steps or units, but optionally includes steps or units that are not listed, or optionally includes other steps or units inherent to these processes, methods, products or devices. The terms "first", "second" and "third" are used to distinguish different objects, etc., and do not represent a sequence, nor do they limit the "first", "second" and "third" to different types.

[0132] In the description of the embodiments of this application, the words "exemplary," "for example," or "for example" are used to indicate examples, illustrations, or descriptions. Any embodiment or design described as "exemplary," "for example," or "for example" in the embodiments of this application should not be construed as being preferred or advantageous over other embodiments or designs. Rather, the use of words such as "exemplary," "for example," or "for example" is intended to present the relevant concepts in a concrete manner.

[0133] In the description of the embodiments of the present application, unless otherwise specified, “ / ” means or, for example, A / B can mean A or B; “and / or” in the text is merely a description of the association relationship of associated objects, indicating that three relationships may exist, for example, A and / or B can mean: A exists alone, A and B exist at the same time, and B exists alone. In addition, in the description of the embodiments of the present application, “multiple” refers to two or more than two.

[0134] In some processes described in the embodiments of the present application, multiple operations or steps are included that appear in a specific order. However, it should be understood that these operations or steps may not be performed in the order in which they appear in the embodiments of the present application or may be performed in parallel. The sequence numbers of the operations are only used to distinguish between different operations, and the sequence numbers themselves do not represent any order of execution. In addition, these processes may include more or fewer operations, and these operations or steps may be performed in sequence or in parallel, and these operations or steps may be combined.

[0135] Through the description of the above implementation methods, those skilled in the art can clearly understand that the above-mentioned embodiment methods can be implemented by means of software plus the necessary general hardware platform, of course, it can also be implemented by hardware, but in many cases the former is a better implementation method. Based on this understanding, the technical solution of the present application, or the part that contributes to the prior art, can be embodied in the form of a software product, which is stored in a storage medium (such as ROM / RAM, magnetic disk, optical disk) as described above, and includes a number of instructions for enabling a terminal device to execute the methods described in each embodiment of the present application.

[0136] The above are only preferred embodiments of the present application and do not limit the patent scope of the present application. Any equivalent structure or equivalent process transformation made using the contents of the present application specification and drawings, or directly or indirectly applied in other related technical fields, are also included in the patent protection scope of the present application.

Claims

1. A method for checking the initial attitude of an aircraft, characterized in that: The method for checking the initial attitude of an aircraft comprises: When the aircraft is in a horizontal state, a first pitch angle and a first roll angle of the aircraft are calculated based on accelerations output by accelerometers of axes corresponding to the pitch angle and roll angle of an inertial navigation system of the aircraft; The heading angle output by the inertial navigation system after executing the automatic alignment command is used as the first heading angle of the aircraft; During the erection process of the aircraft, the attitude of the aircraft is continuously updated based on the first pitch angle, the first roll angle, and the first heading angle of the aircraft using the quaternion method and the differential solution method. The second pitch angle, the second roll angle, and the second heading angle of the aircraft when the aircraft is in the vertical state after the erection is completed are calculated based on the updated attitude of the aircraft; When the aircraft is in a vertical state, a third pitch angle and a third heading angle of the aircraft are calculated based on the acceleration output again by the accelerometers of the axes corresponding to the pitch angle and roll angle of the aircraft's inertial navigation system; The heading angle output by the inertial navigation system after executing the automatic alignment command again is used as the third roll angle of the aircraft; The accuracy of the initial attitude of the aircraft is checked by comparing the second pitch angle, the second roll angle and the second heading angle of the aircraft with the third pitch angle, the third roll angle and the third heading angle respectively.

2. The method for checking the initial attitude of an aircraft according to claim 1, wherein: The calculating the first pitch angle and the first roll angle of the aircraft according to the acceleration output by the accelerometer of the axis corresponding to the pitch angle and the roll angle of the aircraft inertial navigation system includes: The smoothing method is used to calculate the first pitch angle and the first roll angle of the aircraft using Formula 1 based on the acceleration output by the accelerometer of the axis corresponding to the pitch angle and the roll angle of the aircraft inertial navigation system. The Formula 1 is: Among them, θ0 is the first pitch angle, γ0 is the first roll angle, Ax i is the i-th acceleration output by the accelerometer of the pitch angle axis of the inertial navigation system during the smoothing calculation cycle, Az i is the i-th acceleration output by the accelerometer of the axis corresponding to the roll angle of the inertial navigation system during the smoothing calculation cycle, g0 is the local acceleration of gravity, and T is the total number of accelerations output by the accelerometer during the smoothing calculation cycle.

3. The method for checking the initial attitude of an aircraft according to claim 1, wherein: The method of continuously updating the attitude of the aircraft based on the first pitch angle, the first roll angle, and the first heading angle of the aircraft by using the quaternion method and the differential solution method includes: The initial attitude quaternion of the aircraft is established according to the first pitch angle, the first roll angle and the first heading angle of the aircraft through Formula 2, which is: An attitude update differential equation is established based on the initial attitude quaternion. The attitude of the aircraft is continuously updated based on the attitude update differential equation and the initial attitude quaternion, and the updated attitude quaternion is calculated. The attitude update differential equation is: Where Q is the initial attitude quaternion, q0, q1, q2 and q3 are the four components of the quaternion Q, i, j and k are the rotation components around the three orthogonal axes of the three-dimensional space, θ0 is the first pitch angle, γ0 is the first roll angle, Ψ0 is the first heading angle, is the updated attitude quaternion, which represents the derivative of each component of the quaternion with respect to time, is the quaternion matrix before updating, in and are the angular velocities of the x-axis, y-axis, and z-axis output by the gyroscope at the update time t, respectively. The superscript b represents the body coordinate system of the inertial navigation system.

4. The method for checking the initial attitude of an aircraft according to claim 3, wherein: The calculation of the second pitch angle, the second roll angle, and the second heading angle of the aircraft in the vertical state after the aircraft is erected based on the updated aircraft attitude includes: According to the updated attitude quaternion in the attitude update differential equation, the second pitch angle, second roll angle, and second heading angle of the aircraft in the vertical state after the vertical erection are calculated by formula 3. The formula 3 is: Among them, t is the second heading angle, θ t is the second pitch angle, γ t is the second roll angle.

5. The method for checking the initial attitude of an aircraft according to claim 1, wherein: The accuracy check of the initial attitude of the aircraft by comparing the second pitch angle, the second roll angle, and the second heading angle of the aircraft with the third pitch angle, the third roll angle, and the third heading angle respectively includes: If the absolute value of the difference between the second pitch angle and the third pitch angle of the aircraft is less than the first preset difference, the absolute value of the difference between the second roll angle and the third roll angle is less than the second preset difference, and the absolute value of the difference between the second heading angle and the third heading angle is less than the third preset difference, then the result of the initial attitude accuracy check of the aircraft is determined to be passed; If the absolute value of the difference between the second pitch angle and the third pitch angle of the aircraft is not less than the first preset difference, or the absolute value of the difference between the second roll angle and the third roll angle is not less than the second preset difference, or the absolute value of the difference between the second heading angle and the third heading angle is not less than the third preset difference, then it is determined that the initial attitude accuracy calibration result of the aircraft is failed.

6. An aircraft initial attitude calibration device, characterized in that: The aircraft initial attitude checking device comprises: a first calculation module, configured to calculate a first pitch angle and a first roll angle of the aircraft based on accelerations output by accelerometers of axes corresponding to the pitch angle and roll angle of an inertial navigation system of the aircraft when the aircraft is in a horizontal state; A first acquisition module is configured to use the heading angle output by the inertial navigation system after executing the automatic alignment instruction as the first heading angle of the aircraft; a continuous updating module for continuously updating the attitude of the aircraft based on the first pitch angle, the first roll angle, and the first heading angle of the aircraft during the erection process of the aircraft using a quaternion method and a differential solution method, and calculating, based on the updated attitude of the aircraft, a second pitch angle, a second roll angle, and a second heading angle of the aircraft when the aircraft is in a vertical state after the erection is completed; a second calculation module, configured to calculate, when the aircraft is in a vertical state, a third pitch angle and a third heading angle of the aircraft based on accelerations output again by the accelerometers of the axes corresponding to the pitch angle and roll angle of the aircraft's inertial navigation system; A second acquisition module is used to use the heading angle output by the inertial navigation system after executing the automatic alignment instruction again as the third roll angle of the aircraft; The calibration module is used to perform accuracy calibration on the initial attitude of the aircraft by comparing the second pitch angle, the second roll angle and the second heading angle of the aircraft with the third pitch angle, the third roll angle and the third heading angle respectively.

7. The aircraft initial attitude calibration device according to claim 6, characterized in that: The first calculation module is configured to: The smoothing method is used to calculate the first pitch angle and the first roll angle of the aircraft using Formula 1 based on the acceleration output by the accelerometer of the axis corresponding to the pitch angle and the roll angle of the aircraft inertial navigation system. The Formula 1 is: Among them, θ0 is the first pitch angle, γ0 is the first roll angle, Ax i is the i-th acceleration output by the accelerometer of the pitch angle axis of the inertial navigation system during the smoothing calculation cycle, Az i is the i-th acceleration output by the accelerometer of the axis corresponding to the roll angle of the inertial navigation system during the smoothing calculation cycle, g0 is the local acceleration of gravity, and T is the total number of accelerations output by the accelerometer during the smoothing calculation cycle.

8. The aircraft initial attitude calibration device according to claim 6, characterized in that: The continuous updating module is used to: The initial attitude quaternion of the aircraft is established according to the first pitch angle, the first roll angle and the first heading angle of the aircraft through Formula 2, which is: An attitude update differential equation is established based on the initial attitude quaternion. The attitude of the aircraft is continuously updated based on the attitude update differential equation and the initial attitude quaternion, and the updated attitude quaternion is calculated. The attitude update differential equation is: Where Q is the initial attitude quaternion, q0, q1, q2 and q3 are the four components of the quaternion Q, i, j and k are the rotation components around the three orthogonal axes of the three-dimensional space, θ0 is the first pitch angle, γ0 is the first roll angle, Ψ0 is the first heading angle, is the updated attitude quaternion, which represents the derivative of each component of the quaternion with respect to time, is the quaternion matrix before updating, in and are the angular velocities of the x-axis, y-axis, and z-axis output by the gyroscope at the update time t, respectively. The superscript b represents the body coordinate system of the inertial navigation system.

9. An aircraft initial attitude calibration device, characterized in that: The aircraft initial attitude calibration device includes a processor, a memory, and an aircraft initial attitude calibration program stored in the memory and executable by the processor, wherein when the aircraft initial attitude calibration program is executed by the processor, the steps of the aircraft initial attitude calibration method as described in any one of claims 1 to 5 are implemented.

10. A readable storage medium, characterized in that: The readable storage medium stores an aircraft initial attitude calibration program, wherein when the aircraft initial attitude calibration program is executed by the processor, the steps of the aircraft initial attitude calibration method according to any one of claims 1 to 5 are implemented.

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