A quasi-dynamic micro air vehicle attitude measurement method and system

Abstract

The present invention relates to the field of aircraft detection technology, and discloses a quasi-dynamic micro aircraft attitude measurement method and system thereof, wherein the attitude angle of a flapping-wing aircraft is measured on an experimental bench, and a coordinate system suitable for the flapping-wing aircraft is established, including a body coordinate system. b X b Y b Z b , inertial coordinate system O g X g Y g Z g , wing coordinate system, velocity coordinate system O v X v Y v Z v , the relationship between the inertial coordinate system of the aircraft and the body coordinate system is obtained through the coordinate change, and the attitude angle (pitch angle) of the flapping-wing aircraft is determined 、 Yaw angle 、 Roll angle), establish the attitude dynamics equation of the flapping-wing aircraft, and determine its final attitude angle by the rotation angular velocity during flight. Compared with the prior art, the present invention measures the attitude angle of the micro flapping-wing aircraft on the experimental bench, establishes a coordinate system suitable for the flapping-wing aircraft based on the existing coordinate system to analyze its flight attitude angle, and has low cost, small size and simple testing.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft detection, and in particular to a quasi-dynamic micro-aircraft attitude measurement method and system thereof. Background Art

[0002] Due to the small size, strong concealment, and high flexibility of flapping-wing aircraft, they have been widely used in military and civilian fields. However, the flight stability of the currently developed micro flapping-wing aircraft is generally poor. In order to achieve stable flight of the micro flapping-wing aircraft, its attitude must be analyzed and studied. Due to its small size and complex flight environment, the flight attitude of the micro flapping-wing aircraft is difficult to measure. The traditional practice is to install a gyroscope on the micro flapping-wing aircraft for outdoor flight to obtain the attitude angle of the aircraft. The test flight experiment has high requirements for the experimental site, requiring a large open space, and the site needs to be arranged in advance to prevent the aircraft from breaking or injuring others.

[0003] In order to solve the problem that the attitude angle of a micro flapping-wing aircraft is difficult to measure, the present invention provides a quasi-dynamic measurement system and method for the attitude of a micro flapping-wing aircraft, which measures the attitude angle of a micro flapping-wing aircraft on a test bench and has the advantages of low cost, small size, and simple testing. Summary of the invention

[0004] Purpose of the invention: In view of the problems existing in the prior art, the present invention provides a quasi-dynamic micro-aircraft attitude measurement method and system, which can measure the attitude angle of a micro-flapping-wing aircraft on a test bench, and has the advantages of low cost, small size, and simple testing.

[0005] Technical solution: The present invention provides a quasi-dynamic micro-aircraft attitude measurement method, comprising the following steps:

[0006] Step 1: Install the gyroscope at the center of gravity of the flapping-wing aircraft, install the flapping-wing aircraft on the mounting plate, fix the mounting plate to the roll axis of the carbon fiber three-axis bracket, and install the yaw axis of the carbon fiber three-axis bracket on the experimental bench;

[0007] Step 2: The flapping-wing aircraft receives the servo control signal and starts maneuvering. The attitude of the flapping-wing aircraft begins to change, and the gyroscope begins to measure the angular velocity of the flapping-wing aircraft in the three-axis direction.

[0008] Step 3: Establish a coordinate system suitable for flapping-wing aircraft. First, establish the body coordinate system O b X b Y b Z b , inertial coordinate system O g X g Y g Z g , wing coordinate system, velocity coordinate system Ov X v Y v Z v , coordinate changes are performed to obtain the relationship between the inertial coordinate system of the aircraft and the body coordinate system. The attitude angle of the flapping-wing aircraft is determined by the relationship between the inertial coordinate system and the body coordinate system, that is, three Euler angles, namely, the pitch angle θ, the yaw angle ψ, and the roll angle γ;

[0009] Step 4: Establish the attitude dynamics equation of the flapping-wing aircraft, which is the relationship between the body coordinate system and the inertial coordinate system of the flapping-wing aircraft during the entire flight process, for the rotation of the aircraft itself. From this, its attitude angle is determined by the angular velocity of rotation during flight.

[0010] Furthermore, at the initial moment, the flapping-wing aircraft is in a horizontal state, the gyroscope X-axis is installed horizontally and consistent with the forward direction of the aircraft, the gyroscope Y-axis points upward, and the gyroscope Z-axis points to the right side of the flapping-wing aircraft.

[0011] Furthermore, the wing coordinate system in step 3 includes four wings, and a corresponding wing coordinate system is established for each wing, namely, the left forewing LF, the left hindwing LH, the right forewing RF, and the right hindwing RH;

[0012] The origin of the left rear wing coordinate system is O LH Located at the root of the left hind wing, O LH Z LH The axis always coincides with the wing's flip axis and points to the wing tip. LH X LH In the left hind wing plane, and connected with O LH Z LH Vertical, pointing to the leading edge, O LH Y LH It is determined by the right-hand coordinate system;

[0013] The origin of the right front wing coordinate system is O RF Located at the root of the right front wing, O RF Z RF The axis is inside the wing, parallel to the wing's flip axis, and points toward the wing tip. RF X RF Also inside the wing, perpendicular to O RF Z RF Axis and points to the leading edge, O RF Y RF It is determined by the right-hand coordinate system;

[0014] The other two wing coordinate systems are similar to the wing systems on the same side, except that the coordinate origins are located at the roots of their respective wings.

[0015] Furthermore, the origin O of the body coordinate system bAt the center of mass of the flapping-wing aircraft, O b X b The axis is parallel to the fuselage axis of the aircraft, and it is stipulated that the direction pointing to the front of the fuselage is positive. b Y b The axis is located in the symmetry plane of the fuselage and is aligned with the O b X b The axis is vertical, and the direction is defined as upward. b Z b It is determined by the right-hand coordinate system;

[0016] The origin O of the inertial coordinate system g is the center of mass of the flapping-wing aircraft when it takes off, O g X g In the horizontal plane, the direction of travel during takeoff is defined as positive. g Y g Pointing vertically upward, O g Z g The direction of is determined by the right-hand coordinate system;

[0017] The velocity coordinate system origin O a Fixed at the center of mass of the dragonfly, O a X a The axis coincides with the direction of the dragonfly's flight speed, O a Y a The axis is in the dragonfly symmetry plane and is perpendicular to O a X a Axis, pointing upward is positive, and pointing the other way is negative. a Z a The axes are determined by the right-hand coordinate system.

[0018] Furthermore, the specific steps of obtaining the relationship between the aircraft inertial coordinate system and the body coordinate system in step 3 are:

[0019] Step 3.1: First, make the origin and axes of the inertial coordinate system coincide with the origin and axes of the body coordinate system, and move the inertial coordinate system along O g Y g Axis rotation angle ψ, O g X g Axis and O g Z g The axis is rotated to O g X1 axis and O g Z1 axis, forming a transitional coordinate system O g X1Y g Z1;

[0020] Step 3.2: Then, transform the transition coordinate system O g X1Y g Z1 along O gZ1 axis rotation angle θ, O g X1 axis and O g Y g The axis is rotated to O g X b Axis and O g Y2 axis, forming another transitional coordinate system O g X b Y2Z1;

[0021] Step 3.3: Finally, the transition coordinate system O g X b Y2Z1 along O g X b Axis rotation angle γ, O g Y2 axis and O g The Z1 axis is rotated to O g Y b Axis and O g Z b axis, and finally obtain the coordinate system O b X b Y b Z b posture.

[0022] Furthermore, the matrix expression of the relationship between the inertial coordinate system of the vehicle and the body coordinate system is:

[0023]

[0024] Where L(γ,θ,ψ) is the transformation matrix from the inertial coordinate system to the body coordinate system.

[0025] Furthermore, the specific steps of establishing the attitude dynamics equation of the flapping-wing aircraft in step 4 are as follows:

[0026] Step 4.1: Set the angular velocity and angular momentum of the flapping-wing aircraft during flight to be and And select the origin O of the body coordinate system b As a base point, we have:

[0027]

[0028] Where I is the moment of inertia of the flapping-wing aircraft in the body coordinate system, H X , H Y , H Z is the momentum of the flapping-wing aircraft model The components in each direction in the body coordinate system, ω X ,ω Y ,ω Z is the component of the angular velocity of the flapping-wing aircraft model in each axis in the body coordinate system, IX ,I Y ,I Z is the moment of inertia of the flapping-wing aircraft model about each axis of the body coordinate system, I XY ,I XZ ,I YZ ,I YX ,I ZX ,I ZY is the product of inertia of the flapping-wing aircraft model with respect to each axis of the body coordinate system;

[0029] Step 4.2: Set the flapping-wing aircraft model to be a symmetrical model with a zero product of inertia, simplify the formula in step 4.1, and then perform a vector cross product on it;

[0030] Step 4.3: The components of the total moment on the flapping-wing aircraft in each axis in the body coordinate system are set as M Xb 、M Xb 、M Xb ,but:

[0031]

[0032] Step 4.4: The angular velocity of rotation is equal to the vector sum of the angular velocities of the flapping-wing aircraft model around each axis:

[0033]

[0034] Step 4.5: The transformation relationship from the inertial coordinate system to the body coordinate system is:

[0035]

[0036] The angular velocity and attitude angle of the flapping-wing aircraft during flight are linked together. The angular velocity (ω X ,ω Y ,ω Z ,) can be used to obtain its attitude angle (θ, ψ, γ).

[0037] The present invention also discloses a quasi-dynamic micro-aircraft attitude measurement system, comprising a flapping-wing aircraft, a carbon fiber three-axis bracket, and a gyroscope. The gyroscope is installed at the center of gravity of the flapping-wing aircraft, and the flapping-wing aircraft is installed on a mounting plate. The carbon fiber three-axis bracket comprises a yaw axis, a pitch axis, and a roll axis. The mounting plate is fixed to the roll axis of the carbon fiber three-axis bracket. The yaw axis of the carbon fiber three-axis bracket is installed on an experimental bench. The pitch axis is rotatably connected to the yaw axis, and the roll axis is installed on the pitch axis. The system also includes a coordinate data processing module, which is configured to execute the steps of the quasi-dynamic micro-aircraft attitude measurement method as described above.

[0038] Beneficial effects:

[0039] The invention measures the attitude angle of the micro flapping-wing aircraft on an experimental bench, establishes a coordinate system suitable for the flapping-wing aircraft according to the existing coordinate system to analyze its flight attitude angle, and has low cost, small volume and simple testing. BRIEF DESCRIPTION OF THE DRAWINGS

[0040] Figure 1 This is a structural diagram of the quasi-dynamic micro-aircraft attitude measurement device of the present invention;

[0041] Figure 2 This is a working diagram of the quasi-dynamic micro-aircraft attitude measurement system of the present invention;

[0042] Figure 3 Schematic diagram of each coordinate system of the present invention;

[0043] Figure 4 Schematic diagram of coordinate system transformation of the present invention.

[0044] Among them, 1-experimental bench, 2-mounting plate, 3-carbon fiber three-axis bracket, 301-yaw axis, 302-pitch axis, 303-roll axis. DETAILED DESCRIPTION

[0045] The present invention will be further described below in conjunction with the accompanying drawings. The following embodiments are only used to more clearly illustrate the technical solution of the present invention, and cannot be used to limit the protection scope of the present invention.

[0046] The present invention discloses a quasi-dynamic micro-aircraft attitude measurement method and system. Dragonflies in nature can achieve a variety of maneuverable flight actions such as forward flight, hovering, side flight, and backward flight. Therefore, dragonflies have become a biological model for the research of micro flapping-wing aircraft. The following uses a dragonfly model for illustration, but the method is not limited to dragonflies.

[0047] The quasi-dynamic micro-aircraft attitude measurement system consists of an experimental bench, a carbon fiber three-axis bracket, a mounting plate, a flapping-wing aircraft, a micro-gyroscope installed on the flapping-wing aircraft, a remote controller, a coordinate data processing module, a computer, etc. Figure 1 The measurement system workflow diagram is shown in Figure 2 As shown. The gyroscope is installed at the center of gravity of the flapping-wing aircraft, the flapping-wing aircraft is installed on the mounting plate, the carbon fiber three-axis bracket includes a yaw axis, a pitch axis and a roll axis, the mounting plate is fixed to the roll axis of the carbon fiber three-axis bracket, the yaw axis of the carbon fiber three-axis bracket is installed on the experimental bench, the pitch axis is rotatably connected to the yaw axis, and the roll axis is installed on the pitch axis; and a coordinate data processing module is also included, which is configured to perform the following steps of the quasi-dynamic micro-aircraft attitude measurement method.

[0048] Before the test begins, at the initial moment, the flapping-wing aircraft is in a horizontal state, the gyroscope X-axis is installed horizontally and in line with the forward direction of the flapping-wing aircraft, the gyroscope Y-axis points upward, and the gyroscope Z-axis points to the right side of the flapping-wing aircraft. The flapping-wing aircraft and gyroscope are powered by high-performance polymer lithium batteries, which are fixed on the aircraft. During the test, the flapping-wing aircraft receives the servo control signal sent by the remote control, the flapping-wing aircraft begins to maneuver, the aircraft attitude begins to change, and the micro gyroscope begins to measure the angular velocity of the flapping-wing aircraft in the three-axis direction. The data is then received by the receiver and transmitted to the coordinate data processing module. After coordinate transformation and calculation, the attitude angle of the aircraft relative to the ground is obtained, thereby further evaluating the flight stability of the aircraft.

[0049] At present, the existing coordinate systems are all for fixed-wing aircraft, and there is no coordinate system specifically for flapping-wing aircraft. Therefore, it is necessary to establish a coordinate system suitable for flapping-wing aircraft based on the existing coordinate system to analyze its flight attitude angle. The flapping-wing aircraft coordinate system includes the inertial coordinate system, the body coordinate system and the wing coordinate system. Figure 3 shown.

[0050] The obvious difference between flapping-wing aircraft and traditional aircraft is that the wings are always in motion. In order to analyze the flapping of the wings and the aerodynamic forces on the wings, a wing coordinate system must be established. Since dragonflies have four wings, a corresponding wing coordinate system needs to be established for each wing. LF, LH, RF, and RH in the figure represent the left forewing, left hindwing, right forewing, and right hindwing, respectively. Only the coordinate systems of the left hindwing and the right forewing are given in the figure. The origin of the left hindwing system is O LH Located at the root of the left hind wing, O LH Z LH The axis always coincides with the wing's flip axis and points to the wing tip. LH X LH In the left hind wing plane, and connected with O LH Z LH Vertical, pointing to the leading edge, O LH Y LH It is determined by the right-hand coordinate system; the origin O of the right front wing system RF Located at the root of the right front wing, O RF Z RF The axis is inside the wing, parallel to the wing's flip axis, and points toward the wing tip. RF X RF Also inside the wing, perpendicular to O RF Z RF Axis and points to the leading edge, O RF Y RF It is determined by the right-hand coordinate system. The other two wing coordinate systems are similar to the wing systems on the same side, except that the coordinate origin is located at the root of each wing.

[0051] First, establish the body coordinate system O b X b Y b Z b , inertial coordinate system O g X g Y g Z g , wing coordinate system, velocity coordinate system O v X v Y v Z v .

[0052] The body coordinate system is a coordinate system introduced to study the spatial attitude of flapping-wing aircraft. It is fixed to the flapping-wing aircraft and moves with the flapping-wing aircraft. The origin O b At the center of mass of the flapping-wing aircraft, O b X b The axis is parallel to the fuselage axis of the flapping-wing aircraft, and it is stipulated that the direction pointing to the front of the fuselage is positive. b Y b The axis is located in the symmetry plane of the fuselage and is aligned with the O b X b The axis is vertical, and the direction is defined as upward. b Z b It is determined by the right-hand coordinate system.

[0053] The inertial coordinate system (or ground coordinate system) is fixed to the earth. The origin of the inertial coordinate system is O g is the center of mass of the dragonfly when it takes off, O g X g In the horizontal plane, the direction of travel during takeoff is defined as positive. g Y g Pointing vertically upward, O g Z g The orientation is determined by the right-hand coordinate system.

[0054] Velocity coordinate system (or airflow coordinate system) origin O a Fixed at the center of mass of the dragonfly, O a X a The axis coincides with the direction of the dragonfly's flight speed, O a Y a The axis is in the dragonfly symmetry plane and is perpendicular to O a X a Axis, pointing upward is positive, and pointing the other way is negative. a Z a The axes are determined by the right-hand coordinate system.

[0055] The attitude angle of a flapping-wing aircraft is determined by the relationship between the inertial coordinate system and the body coordinate system, which is commonly known as the Euler angle. The three Euler angles are defined as follows:

[0056] Pitch angle θ: the vertical axis of the body coordinate system (i.e. O b X b axis) and the horizontal ground (i.e. O g X g Z g The angle between the two surfaces is specified as O b X b When the axis points upwards, θ is positive; otherwise, it is negative.

[0057] Yaw angle ψ: O of the body coordinate system b X b The projection of the axis in the horizontal plane and the O of the inertial coordinate system g X g The angle between the axes, when viewed from the rear of the fuselage, is specified as O b X b Axis points to O g X g Y g When it is on the left side of the plane, ψ is positive, otherwise it is negative.

[0058] Roll angle γ: O of the body coordinate system b Y b Axis and O containing the body coordinate system b X b The angle between the vertical planes of the axes, when viewed from the rear of the fuselage, is specified as O b Y b When the axis is on the right side of the plumb line, γ is positive, otherwise it is negative.

[0059] After the coordinate system is established, the coordinates need to be changed to obtain the relationship between the aircraft's inertial coordinate system and the aircraft's body coordinate system. Figure 4 The changes are obtained.

[0060] First, the origin and each axis of the inertial coordinate system are aligned with the origin and each axis of the body coordinate system, and the inertial coordinate system is aligned along O g Y g Axis rotation angle ψ, O g X g Axis and O g Z g The axis is rotated to O g X1 axis and O g Z1 axis, forming a transitional coordinate system O g X1Y g Z1, the relationship between the inertial coordinate system and the transition coordinate system after the first rotation is expressed by the matrix:

[0061]

[0062] Next, the transition coordinate system O g X1Y g Z1 along O g Z1 axis rotation angle θ, O g X1 axis and O g Y g The axis is rotated to O g X b Axis and O g Y2 axis, forming another transitional coordinate system O g X b Y2Z1, the matrix expression between the two is:

[0063]

[0064] Finally, the transition coordinate system O g X b Y2Z1 along O g X b Axis rotation angle γ, O g Y2 axis and O g The Z1 axis is rotated to O g Y b Axis and O g Z b axis, and finally obtain the coordinate system O b X b Y b Z b The matrix expression between the two is:

[0065]

[0066] From equations (1), (2) and (3), we can get the transformation relationship from the inertial coordinate system to the body coordinate system:

[0067]

[0068] Where L(γ,θ,ψ) is the transformation matrix from the inertial coordinate system to the body coordinate system.

[0069] To achieve stable flight of flapping-wing aircraft, it is necessary to analyze its attitude and establish the attitude dynamics equation of flapping-wing aircraft. The dynamics equation here mainly focuses on the rotation of the aircraft itself, that is, the relationship between the body coordinate system and the ground coordinate system of the flapping-wing aircraft during the entire flight. The angular velocity and angular momentum of the flapping-wing aircraft during flight are set to and And select the origin O of the body coordinate system b (that is, the center of mass of the dragonfly) as the base point, then:

[0070]

[0071] Where I is the moment of inertia of the flapping-wing aircraft in the body coordinate system, H X , H Y , H Z is the component of the momentum H of the dragonfly model in each direction in the body coordinate system, ω X ,ω Y ,ω Z is the component of the angular velocity of the dragonfly model in each axis in the body coordinate system, I X ,I Y ,I Z is the moment of inertia of the dragonfly model about each axis of the body coordinate system, I XY ,I XZ ,I YZ ,I YX ,I ZX ,I ZY is the inertia product of the dragonfly model to each axis of the body coordinate system. In the present invention, the dragonfly is a symmetrical model, so the inertia product is zero, and then formula (6) is transformed into:

[0072]

[0073] Then, from the vector cross product formula, we get:

[0074]

[0075] The total moment of the dragonfly in each axis in the body coordinate system is set as M Xb 、M Xb 、M Xb , Substituting equations (7) and (8) into equation (5), we get:

[0076]

[0077] Here, the present invention describes the posture of the dragonfly flapping-wing aircraft from the perspective of Euler angles. The angular velocity of the dragonfly model should be equal to the vector sum of the angular velocities of the dragonfly model rotating around each axis:

[0078]

[0079] The transformation relationship from the inertial coordinate system to the body coordinate system is:

[0080]

[0081] Transforming formula (11) yields:

[0082]

[0083] The above equation is the attitude kinematic equation of the dragonfly model during flight, which links the angular velocity and attitude angle of the dragonfly during flight. The angular velocity of the dragonfly model (ω X ,ω Y ,ω Z ,) can be used to obtain its attitude angle (θ, ψ, γ).

[0084] The above embodiments are only for illustrating the technical concept and features of the present invention, and their purpose is to enable people familiar with the technology to understand the content of the present invention and implement it accordingly, and they cannot be used to limit the protection scope of the present invention. Any equivalent transformation or modification made according to the spirit of the present invention should be included in the protection scope of the present invention.

Claims

1. A quasi-dynamic micro air vehicle attitude measurement method, It is characterized in that The steps include: Step 1: Install the gyroscope at the center of gravity of the flapping-wing aircraft, install the flapping-wing aircraft on the mounting plate, fix the mounting plate to the roll axis of the carbon fiber three-axis bracket, and install the yaw axis of the carbon fiber three-axis bracket on the experimental bench; Step 2: The flapping-wing aircraft receives the servo control signal and starts maneuvering. The attitude of the flapping-wing aircraft begins to change, and the gyroscope begins to measure the angular velocity of the flapping-wing aircraft in the three-axis direction. Step 3: Establish a coordinate system suitable for flapping-wing aircraft. First, establish the body coordinate system O b X b Y b Z b , inertial coordinate system O g X g Y g Z g , wing coordinate system, velocity coordinate system O v X v Y v Z v , coordinate changes are performed to obtain the relationship between the inertial coordinate system of the aircraft and the body coordinate system. The attitude angle of the flapping-wing aircraft is determined by the relationship between the inertial coordinate system and the body coordinate system, that is, three Euler angles, namely, the pitch angle θ, the yaw angle ψ, and the roll angle γ; Step 4: Establish the attitude dynamics equation of the flapping-wing aircraft, which is the relationship between the body coordinate system and the inertial coordinate system of the flapping-wing aircraft during the entire flight process, and thus determine its attitude angle through the angular velocity of rotation during flight; The specific steps to establish the attitude dynamics equation of a flapping-wing aircraft are: Step 4.1: Set the angular velocity and angular momentum of the flapping-wing aircraft during flight to be and And select the origin O of the body coordinate system b As a base point, we have: Where I is the moment of inertia of the flapping-wing aircraft in the body coordinate system, H X , H Y , H Z is the momentum of the flapping-wing aircraft model The components in each direction in the body coordinate system, ω X ,ω Y ,ω Z is the component of the angular velocity of the flapping-wing aircraft model in each axis in the body coordinate system, I X ,I Y ,I Z is the moment of inertia of the flapping-wing aircraft model about each axis of the body coordinate system, I XY ,I XZ ,I YZ ,I YX ,I ZX ,I ZY is the product of inertia of the flapping-wing aircraft model with respect to each axis of the body coordinate system; Step 4.2: Set the flapping-wing aircraft model to be a symmetrical model with a zero product of inertia, simplify the formula in step 4.1, and then perform a vector cross product on it; Step 4.3: The total torque on the flapping-wing aircraft is set as M in each axis in the body coordinate system. Xb 、M Xb 、M Xb ,but: Step 4.4: The angular velocity of rotation is equal to the vector sum of the angular velocities of the flapping-wing aircraft model around each axis: Step 4.5: The transformation relationship from the inertial coordinate system to the body coordinate system is: The angular velocity and attitude angle of the flapping-wing aircraft during flight are linked together. The angular velocity (ω X ,ω Y ,ω Z ,) can be used to obtain its attitude angle (θ, ψ, γ).

2. The quasi-dynamic micro-aircraft attitude measurement method according to claim 1, It is characterized in that At the initial moment, the flapping-wing aircraft is in a horizontal state, the gyroscope X-axis is installed horizontally and consistent with the forward direction of the aircraft, the gyroscope Y-axis points upward, and the gyroscope Z-axis points to the right side of the flapping-wing aircraft.

3. The quasi-dynamic micro air vehicle attitude measurement method according to claim 1, It is characterized in that The wing coordinate system in step 3 includes four wings, and a corresponding wing coordinate system is established for each wing, namely, the left forewing LF, the left hindwing LH, the right forewing RF, and the right hindwing RH; The origin of the left rear wing coordinate system is O LH Located at the root of the left hind wing, O LH Z LH The axis always coincides with the wing's flip axis and points to the wing tip. LH X LH In the left hind wing plane, and connected with O LH Z LH Vertical, pointing to the leading edge, O LH Y LH It is determined by the right-hand coordinate system; The origin of the right front wing coordinate system is O RF Located at the root of the right front wing, O RF Z RF The axis is inside the wing, parallel to the wing's flip axis, and points toward the wing tip. RF X RF Also inside the wing, perpendicular to O RF Z RF Axis and points to the leading edge, O RF Y RF It is determined by the right-hand coordinate system; The other two wing coordinate systems are similar to the wing systems on the same side, except that the coordinate origins are located at the roots of their respective wings.

4. The quasi-dynamic micro air vehicle attitude measurement method according to claim 1, It is characterized in that The origin O of the body coordinate system b At the center of mass of the flapping-wing aircraft, O b X b The axis is parallel to the fuselage axis of the aircraft, and it is stipulated that the direction pointing to the front of the fuselage is positive. b Y b The axis is located in the symmetry plane of the fuselage and is aligned with the O b X b The axis is vertical, and the direction is defined as upward. b Z b It is determined by the right-hand coordinate system; The origin O of the inertial coordinate system g is the center of mass of the flapping-wing aircraft when it takes off, O g X g In the horizontal plane, the direction of travel during takeoff is defined as positive. g Y g Pointing vertically upward, O g Z g The direction of is determined by the right-hand coordinate system; The velocity coordinate system origin O a Fixed at the center of mass of the dragonfly, O a X a The axis coincides with the direction of the dragonfly's flight speed, O a Y a The axis is in the dragonfly symmetry plane and is perpendicular to O a X a Axis, pointing upward is positive, and pointing the other way is negative. a Z a The axes are determined by the right-hand coordinate system.

5. The quasi-dynamic micro air vehicle attitude measurement method according to claim 1, It is characterized in that The specific steps for obtaining the relationship between the aircraft inertial coordinate system and the body coordinate system in step 3 are: Step 3.1: First, make the origin and axes of the inertial coordinate system coincide with the origin and axes of the body coordinate system, and move the inertial coordinate system along O g Y g Axis rotation angle ψ, O g X g Axis and O g Z g The axis is rotated to O g X 1 Axis and O g Z 1 axis, forming a transitional coordinate system O g X 1 Y g Z 1 ; Step 3.2: Then, transform the transition coordinate system O g X 1 Y g Z 1 Along O g Z 1 Axis rotation angle θ, O g X 1 Axis and O g Y g The axis is rotated to O g X b Axis and O g Y 2 Axis, forming another transitional coordinate system O g X b Y 2 Z 1 ; Step 3.3: Finally, the transition coordinate system O g X b Y 2 Z 1 Along O g X b Axis rotation angle γ, O g Y 2 Axis and O g Z 1 The axis is rotated to O g Y b Axis and O g Z b axis, and finally obtain the coordinate system O b X b Y b Z b posture.

6. The quasi-dynamic micro air vehicle attitude measurement method according to claim 5, It is characterized in that The matrix expression of the relationship between the inertial coordinate system of the traveler and the body coordinate system is: Where L(γ,θ,ψ) is the transformation matrix from the inertial coordinate system to the body coordinate system.

7. A quasi-dynamic micro air vehicle attitude measurement system, It is characterized in that The invention comprises a flapping-wing aircraft, a carbon fiber three-axis bracket, and a gyroscope. The gyroscope is installed at the center of gravity of the flapping-wing aircraft. The flapping-wing aircraft is installed on a mounting plate. The carbon fiber three-axis bracket comprises a yaw axis, a pitch axis, and a roll axis. The mounting plate is fixed to the roll axis of the carbon fiber three-axis bracket. The yaw axis of the carbon fiber three-axis bracket is installed on an experimental bench. The pitch axis is rotatably connected to the yaw axis, and the roll axis is installed on the pitch axis. The invention also comprises a coordinate data processing module. The coordinate data processing module is configured to execute the steps of the quasi-dynamic micro-aircraft attitude measurement method as described in any one of claims 1 to 6.

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