CN115200580A - Quasi-dynamic micro aircraft attitude measurement method and system - Google Patents

Quasi-dynamic micro aircraft attitude measurement method and system Download PDF

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CN115200580A
CN115200580A CN202210860655.0A CN202210860655A CN115200580A CN 115200580 A CN115200580 A CN 115200580A CN 202210860655 A CN202210860655 A CN 202210860655A CN 115200580 A CN115200580 A CN 115200580A
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coordinate system
aircraft
wing
axis
flapping
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张锐
刘开一
郑兴
胡薇
胡宇琦
赵瑞
李宇辰
韩秀文
王艳
陆曹荣
李舒凯
周聪慧
吴天昊
燕永智
祁昌庆
韩正儒
张慧贤
陆夏
杜双飞
徐礼超
夏晶晶
戴建国
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Huaiyin Institute of Technology
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    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G01C21/10—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration
    • G01C21/12—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning
    • G01C21/16—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation
    • G01C21/18—Stabilised platforms, e.g. by gyroscope
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G01C21/20—Instruments for performing navigational calculations
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C25/00—Manufacturing, calibrating, cleaning, or repairing instruments or devices referred to in the other groups of this subclass
    • G01C25/005—Manufacturing, calibrating, cleaning, or repairing instruments or devices referred to in the other groups of this subclass initial alignment, calibration or starting-up of inertial devices
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation

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Abstract

The invention relates to the technical field of aircraft detection, and discloses a method and a system for measuring attitude of a quasi-dynamic miniature aircraft 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 Obtaining the relation between the inertial coordinate system of the aircraft and the coordinate system of the aircraft body through coordinate change, and determining the attitude angle (pitch angle) of the flapping wing aircraft、Yaw angle、Roll angle), establishing an attitude dynamics equation of the flapping wing aircraft, and determining the final attitude angle of the flapping wing aircraft according to the rotating angular speed in flight. Compared with the prior art, the utility modelThe attitude angle of the miniature flapping wing aircraft is measured on the experiment bench, and the flight attitude angle of the miniature flapping wing aircraft is analyzed by establishing a coordinate system suitable for the flapping wing aircraft according to the existing coordinate system, so that the miniature flapping wing aircraft attitude angle measuring device is low in cost, small in size and simple in test.

Description

Quasi-dynamic micro aircraft attitude measurement method and system
Technical Field
The invention relates to the technical field of aircraft detection, in particular to a method and a system for measuring the attitude of a quasi-dynamic micro aircraft.
Background
Due to the characteristics of small specific volume, strong concealment, high flexibility and the like of the flapping wing aircraft, the flapping wing aircraft has been widely applied to the fields of military affairs and civil use. However, the flight stability of the micro flapping wing air vehicle developed at present is generally poor. In order to realize the stable flight of the micro flapping wing flying machine, the attitude of the micro flapping wing flying machine must be analyzed and researched. Due to the small volume and the complex flying environment, the flying attitude of the miniature flapping wing aircraft is difficult to measure. Traditionally, a gyroscope is mounted on a miniature flapping-wing aircraft to carry out outdoor flight, and the attitude angle of the aircraft is obtained. The test flight experiment has higher requirements on the experimental site, needs a larger open site, and needs to arrange the site in advance to prevent the aircraft from breaking or injuring others.
In order to solve the problem that the attitude angle of the miniature flapping-wing aircraft is difficult to measure, the invention provides the system and the method for measuring the attitude of the miniature flapping-wing aircraft in a quasi-dynamic manner.
Disclosure of Invention
The purpose of the invention is as follows: aiming at the problems in the prior art, the invention provides a method and a system for measuring the attitude of a quasi-dynamic micro-aircraft, which are used for measuring the attitude angle of a micro-flapping-wing aircraft on an experiment bench and have the advantages of low cost, small volume, simple test and the like.
The technical scheme is as follows: the invention provides a quasi-dynamic micro aircraft attitude measurement method, which comprises the following steps:
step 1: installing a gyroscope at the gravity center of the flapping wing aircraft, installing the flapping wing aircraft on a mounting plate, fixing the mounting plate with a rolling shaft of a carbon fiber triaxial bracket, and installing a yaw shaft of the carbon fiber triaxial bracket on an experiment bench;
and 2, step: the flapping wing air vehicle receives a steering engine control signal, starts maneuvering flight, starts changing the attitude of the flapping wing air vehicle, and starts measuring the angular speed of the flapping wing air vehicle in the three-axis direction by the gyroscope;
and step 3: establishing a coordinate system suitable for the flapping wing aircraft, and firstly establishing a 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 Changing coordinates to obtain a relation between an inertial coordinate system of the flapping-wing aircraft and a body coordinate system, wherein the attitude angle of the flapping-wing aircraft is determined by the relation between the inertial coordinate system and the body coordinate system, namely three euler angles which are a pitch angle theta, a yaw angle psi and a roll angle gamma respectively;
and 4, step 4: an attitude dynamics equation of the flapping-wing aircraft is established, and the attitude angle of the flapping-wing aircraft is determined according to the rotation of the aircraft, namely the relation between a body coordinate system and an inertia coordinate system of the flapping-wing aircraft in the whole flight process.
Further, at the initial moment, the flapping wing air vehicle is in a horizontal state, the X axis of the gyroscope is horizontally arranged and is consistent with the advancing direction of the air vehicle, the Y axis of the gyroscope points upwards, and the Z axis of the gyroscope points to the right side of the flapping wing air vehicle.
Furthermore, the wing coordinate system in the step 3 comprises four wings, and a corresponding wing coordinate system is established for each wing, namely a left front wing LF, a left back wing LH, a right front wing RF and a right back wing RH;
origin O of the coordinate system of the left rear wing LH At the wing root of the left hind wing, O LH Z LH The axis always coincides with the turning axis of the wing and points to the direction of the wing tip, O LH X LH In the left rear fin and with O LH Z LH Perpendicular, pointing in the direction of the leading edge, O LH Y LH Determined by the right hand coordinate system;
origin O of coordinate system of right front wing RF At the root of the right anterior wing, O RF Z RF The axis is positioned in the wing, is parallel to the overturning axis of the wing and points to the direction of the wing tip, and O RF X RF Also inside the wing, perpendicular to O RF Z RF Axial and pointing in the direction of the leading edge, O RF Y RF Determined by the right hand coordinate system;
the other two wing coordinate systems are similar to the wing systems in the respective ipsilateral directions, except that the origin of coordinates is located at the respective wing root.
Further, the origin Ob of the body coordinate system is positioned on the center of mass of the ornithopter, O b X b The axis being parallel to the axis of the fuselage of the aircraft and being defined as pointing forward of the fuselage as positive, O b Y b The axis lying in the plane of symmetry of the fuselage and being parallel to O b X b The axis is vertical and the prescribed direction is positive, O b Z b Determined by the right hand coordinate system;
origin O of the inertial frame g Is the center of mass, O, of the flapping wing aircraft during takeoff g X g In a horizontal plane and defining the direction of advance at take-off as positive, O g Y g Pointing vertically upwards, O g Z g 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 coinciding with the speed direction of dragonfly flight, O a Y a The axis is in the dragonfly symmetry plane and perpendicular to O a X a Axis, provided that pointing upwards is positive, and vice versa negative, O a Z a The axes are then determined by the right hand coordinate system.
Further, the specific step of obtaining the relationship between the inertial coordinate system of the aircraft and the body coordinate system in step 3 is:
step 3.1: firstly, the origin and each axis of the inertial coordinate system are respectively superposed with the origin and each axis of the body coordinate system, and the inertial coordinate system is arranged along O g Y g Axial rotation by an angle phi, O g X g Shaft and O g Z g The shafts respectively turn to O g X 1 Shaft and O g Z 1 Axes forming a transitional coordinate system O g X 1 Y g Z 1 ；
Step 3.2: then the transition coordinate system O g X 1 Y g Z 1 Along O g Z 1 Rotation of the shaft by angle theta, O g X 1 Shaft and O g Y g The shafts respectively turn to O g X b Shaft and O g Y 2 Axes forming another transitional coordinate system O g X b Y 2 Z 1 ；
Step 3.3: finally, the transition coordinate system O is processed g X b Y 2 Z 1 Along O g X b Rotation of the shaft by an angle of gamma, O g Y 2 Shaft and O g Z 1 The shafts are respectively turned to O g Y b Shaft and O g Z b Axis, finally obtaining a coordinate system O b X b Y b Z b The posture of (2).
Further, the matrix expression of the relation between the traveling device inertial coordinate system and the machine body coordinate system is as follows:
Figure BDA0003756256320000031
where L (γ, θ, ψ) is a transformation matrix from the inertial coordinate system to the body coordinate system.
Further, the step 4 of establishing an attitude dynamics equation of the flapping wing aircraft comprises the following specific steps:
step 4.1: the angular velocity and the momentum moment of the flapping wing aircraft in the flying process are respectively set as
Figure BDA0003756256320000032
And
Figure BDA0003756256320000033
and selecting a coordinate system of the bodyOrigin O b As base points, there are:
Figure BDA0003756256320000034
Figure BDA0003756256320000035
wherein I is the rotational inertia of the flapping wing aircraft in a body coordinate system, H X 、H Y 、H Z Moment of momentum for a model of a flapping wing aircraft
Figure BDA0003756256320000036
Component, omega, of each direction in the machine body coordinate system X 、ω Y 、ω Z Component of angular velocity of the flapping wing aircraft model in the body coordinate system, I X 、I Y 、I Z For the moment of inertia, I, of the flapping wing aircraft model to each axis of the aircraft body coordinate system XY 、I XZ 、I YZ 、I YX 、I ZX 、I ZY Inertia products of each axis of a coordinate system of the flapping wing aircraft model to the aircraft body are obtained;
step 4.2: setting the flapping wing aircraft model as a symmetrical model, simplifying the formula in the step 4.1 when the inertia product is zero, and then performing vector cross product on the simplified formula;
step 4.3: the total moment received by the flapping wing aircraft is respectively set as M in each axial component in a body coordinate system Xb 、M Xb 、M Xb Then:
Figure BDA0003756256320000037
step 4.4: the angular velocity of rotation is equal to the vector sum of the angular velocities of the flapping-wing aircraft model rotation about the respective axes:
Figure BDA0003756256320000041
step 4.5: the transformation relation from the inertial coordinate system to the body coordinate system is as follows:
Figure BDA0003756256320000042
Figure BDA0003756256320000043
the rotation angular speed (omega) of the known model is obtained by connecting the rotation angular speed and the attitude angle of the flapping wing aircraft during flight X ，ω Y ，ω Z And (e, ψ, γ) are obtained.
The invention also discloses a quasi-dynamic micro aircraft attitude measurement system which comprises the flapping wing aircraft, a carbon fiber three-axis support and a gyroscope, wherein the gyroscope is installed at the gravity center of the flapping wing aircraft, the flapping wing aircraft is installed on a mounting plate, the carbon fiber three-axis support comprises a yaw axis, a pitch axis and a roll axis, the mounting plate is fixed with the roll axis of the carbon fiber three-axis support, the yaw axis of the carbon fiber three-axis support 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; also included is a coordinate data processing module configured to perform the steps of the quasi-dynamic micro aerial vehicle attitude measurement method described above.
Has the advantages that:
the attitude angle of the miniature flapping wing aircraft is measured on the experiment bench, and the flight attitude angle of the miniature flapping wing aircraft is analyzed by establishing a coordinate system suitable for the flapping wing aircraft according to the existing coordinate system, so that the invention has the advantages of low cost, small volume and simple test.
Drawings
FIG. 1 is a schematic diagram of a quasi-dynamic micro-aircraft attitude measurement device according to the present invention;
FIG. 2 is a flow chart of the operation of the quasi-dynamic micro aircraft attitude measurement system of the present invention;
FIG. 3 is a schematic view of coordinate systems of the present invention;
FIG. 4 is a schematic diagram of coordinate system transformation according to the present invention.
The device comprises an experiment bench 1, a placement plate 2, a carbon fiber three-axis support 3, a yaw axis 301, a pitch axis 302 and a roll axis 303.
Detailed Description
The invention is further described below with reference to the accompanying drawings. The following examples are only for illustrating the technical solutions of the present invention more clearly, and the protection scope of the present invention is not limited thereby.
The invention discloses a quasi-dynamic micro aircraft attitude measurement method and a system thereof, wherein a dragonfly in nature can realize various maneuvering flight actions such as forward flight, hovering, lateral flight, reverse flight and the like, so the dragonfly becomes a biological sample for the research of a micro flapping wing aircraft. The following description will be made using a dragonfly model, but the method is not limited to a dragonfly.
The quasi-dynamic micro aircraft attitude measurement system consists of an experiment bench, a carbon fiber three-axis support, a mounting plate, a flapping wing aircraft, a micro gyroscope arranged on the flapping wing aircraft, a remote controller, a coordinate data processing module, a computer and the like, and is shown in figure 1. The measurement system work flow diagram is shown in fig. 2. The method comprises the following steps of installing a gyroscope at the gravity center of a flapping wing aircraft, installing the flapping wing aircraft on a mounting plate, wherein a carbon fiber triaxial support comprises a yaw axis, a pitch axis and a roll axis; also included is a coordinate data processing module configured to perform the steps of the quasi-dynamic micro-aircraft attitude measurement method as follows.
Before the test is started, at the initial moment, the flapping wing air vehicle is in a horizontal state, the X axis of the gyroscope is horizontally arranged and is consistent with the advancing direction of the flapping wing air vehicle, the Y axis of the gyroscope points upwards, and the Z axis of the gyroscope points to the right side of the flapping wing air vehicle. The flapping wing air vehicle and the gyroscope are powered by a high-performance polymer lithium battery, and the battery is fixed on the air vehicle. During testing, the flapping wing aircraft receives a steering engine control signal sent by the remote controller, the flapping wing aircraft starts to maneuver, the attitude of the aircraft starts to change, the micro gyroscope starts to measure the angular velocity of the flapping wing aircraft in the three-axis direction, the data is received by the receiver and then transmitted to the coordinate data processing module, and the attitude angle of the aircraft relative to the ground is obtained through coordinate transformation and calculation, so that the flight stability of the aircraft is further evaluated.
At present, the existing coordinate systems are all directed to fixed-wing aircraft, and there is no coordinate system specially directed to flapping-wing aircraft, so that a coordinate system suitable for the flapping-wing aircraft needs to be established according to the existing coordinate system to analyze the flight attitude angle of the flapping-wing aircraft. The flapping wing aircraft coordinate system includes an inertial coordinate system, a body coordinate system, and a wing coordinate system, each coordinate system being shown in FIG. 3.
The flapping wing aircraft is obviously different from the traditional aircraft in that the wings are in a moving state all the time, a wing coordinate system must be established in order to analyze the flapping of the wings and the aerodynamic force applied to the wings, and a dragonfly with four wings needs to establish a corresponding wing coordinate system for each wing. LF, LH, RF and RH in the figure represent the left anterior wing, the left posterior wing, the right anterior wing and the right posterior wing respectively, only the coordinate systems of the left posterior wing and the right anterior wing are shown in the figure, and the origin O of the left posterior wing system LH At the wing root of the left hind wing, O LH Z LH The axis always coincides with the turning axis of the wing and points to the direction of the wing tip, O LH X LH In the left back wing and with O LH Z LH Perpendicular, pointing in the direction of the leading edge, O LH Y LH Determined by the right hand coordinate system; origin O of right front wing system RE At the wing root of the right front wing, O RF Z RF The axis is positioned in the wing, is also parallel to the overturning axis of the wing and points to the direction of the wing tip, O RF X RF Also inside the wing, perpendicular to O RF Z RF Axial and pointing in the direction of the leading edge, O RF Y RF It is determined from the right-hand coordinate system and the other two wing coordinate systems are similar to the wing systems in the respective ipsilateral directions, except that the origin of coordinates is located at the respective wing root.
Firstly, a coordinate system O of the body is established b X b Y b Z b Inertial coordinate system O g X g Y g Z g Wing and wingCoordinate system, velocity coordinate system O v X v Y v Z v 。
The body coordinate system is a coordinate system introduced for researching the space attitude of the flapping wing aircraft, is fixedly connected with the flapping wing aircraft and moves together with the flapping wing aircraft, and has an origin O b At the center of mass of the ornithopter, O b X b The axis being parallel to the axis of the fuselage of the ornithopter, the front of the fuselage being defined as positive, O b Y b The axis lying in the plane of symmetry of the fuselage and being parallel to O b X b The axis is vertical and the prescribed direction is positive, O b Z b Is determined by the right hand coordinate system.
The inertial coordinate system (or ground coordinate system) is fixed to earth, and the origin O of the inertial coordinate system g Is the center of mass of dragonfly in takeoff, O g X g In a horizontal plane and defining a positive heading direction at takeoff, O g Y g Pointing vertically upwards, O g Z g Is determined by the right-hand coordinate system.
Velocity coordinate system (or airflow coordinate system) origin O a Fixed at the center of mass of the dragonfly, O a X a Axis coinciding with the dragonfly flight speed, O a Y a The axis is in the dragonfly symmetry plane and perpendicular to O a X a Axis, defined as positive when pointing upwards, negative when pointing downwards, O a Z a The axes are then determined by the right hand coordinate system.
The attitude angle of the ornithopter is determined by the relationship between the inertial coordinate system and the body coordinate system, which is known as the euler angles, and the three euler angles are defined as follows:
pitch angle θ: longitudinal axis of the body coordinate system (i.e. O) b X b Axle) and horizontal ground (i.e. O) g X g Z g Faces) and specifies the angle between O b X b Theta is positive when the axis points above the horizontal plane, and negative otherwise.
Yaw angle ψ: o of the coordinate system of the body b X b Projection of axes in the horizontal plane and O of the inertial frame g X g Angle between axes, viewed from the rear of the bodyWhen, stipulate O b X b Axial direction O g X g Y g Psi is positive on the left side of the plane, and vice versa negative.
Roll angle γ: o of the coordinate system of the body b Y b Axis and O comprising a coordinate system of the body b X b The angle between the vertical planes of the shafts, viewed from the rear of the machine body, defines O b Y b Gamma is positive when the axis is to the right of the plumb face, and negative when the axis is negative.
After the coordinate system is established, coordinate change is needed to obtain the relation between the inertial coordinate system of the aircraft and the coordinate system of the body. The inertial coordinate system and the body coordinate system can be obtained by the variation of fig. 4.
First, the origin and each axis of the inertial coordinate system are respectively superposed on the origin and each axis of the body coordinate system, and the inertial coordinate system is positioned along the line O g Y g Axial rotation by an angle phi, O g X g Shaft and O g Z g The shafts respectively turn to O g X 1 Shaft and O g Z 1 Axes forming a transitional coordinate system O g X 1 Y g Z 1 And the relation between the inertial coordinate system and the transition coordinate system after the first rotation adopts the expression of a matrix as follows:
Figure BDA0003756256320000071
then, the transition coordinate system O is set g X 1 Y g Z 1 Along O g Z 1 Rotation of the shaft by an angle theta, O g X 1 Shaft and O g Y g The shafts respectively turn to O g X b Shaft and O g Y 2 Axes forming another transitional coordinate system O g X b Y 2 Z 1 The matrix expression between the two is:
Figure BDA0003756256320000072
finally, the transition coordinate system O is set g X b Y 2 Z 1 Along O g X b Rotation of the shaft by an angle of gamma, O g Y 2 Shaft and O g Z 1 The shafts are respectively turned to O g Y b Shaft and O g Z b Axis of ultimately obtaining the coordinate system O b X b Y b Z b The matrix expression between the two is:
Figure BDA0003756256320000073
the transformation relationship from the inertial coordinate system to the body coordinate system can be obtained by the equations (1), (2) and (3):
Figure BDA0003756256320000074
where L (γ, θ, ψ) — the transformation matrix of the inertial coordinate system to the body coordinate system.
To realize the stable flight of the flapping wing aircraft, the attitude of the flapping wing aircraft needs to be analyzed, and an attitude kinetic equation of the flapping wing aircraft is established, wherein the kinetic equation mainly aims at the rotation of the aircraft, namely the relation between a body coordinate system and a ground coordinate system of the flapping wing aircraft in the whole flight process. The angular velocity and the momentum moment of the flapping wing aircraft in the flying process are respectively set as
Figure BDA0003756256320000075
And
Figure BDA0003756256320000076
and selecting the origin O of the coordinate system of the body b (i.e. the mass center of the dragonfly) as the base point, there are:
Figure BDA0003756256320000077
Figure BDA0003756256320000078
wherein I is the rotational inertia of the flapping wing aircraft in a body coordinate system, H X 、H Y 、H Z Moment of momentum as dragonfly model
Figure BDA0003756256320000081
Component, omega, of each direction in the machine body coordinate system X 、ω Y 、ω Z Is the component of the dragonfly model angular velocity of each axis in the body coordinate system, I X 、I Y 、I Z Is the moment of inertia of each axis of a dragonfly model to a body coordinate system, I XY 、I XZ 、I YZ 、I YX 、I ZX 、I ZY The inertia product of the dragonfly model to each axis of the body coordinate system is obtained. In the invention, the dragonfly is a symmetrical model, so that the inertia product is zero, and the formula (6) is:
Figure BDA0003756256320000082
and then the cross product formula of the vector is obtained:
Figure BDA0003756256320000083
the total moment received by the dragonfly is set as M in each axial component in the body coordinate system Xb 、M Xb 、M Xb Substituting formulae (7) and (8) for formula (5) to obtain:
Figure BDA0003756256320000084
the invention describes the attitude of a dragonfly flapping wing aircraft in the Euler angle. The angular velocity of the dragonfly model rotation should be equal to the vector sum of the angular velocities of the dragonfly model rotation about the respective axes:
Figure BDA0003756256320000085
the transformation relation from the inertial coordinate system to the body coordinate system is as follows:
Figure BDA0003756256320000086
transforming equation (11) to obtain:
Figure BDA0003756256320000091
the above equation is a kinematic equation of attitude of the dragonfly model during flight, which links the rotation angular velocity and attitude angle of the dragonfly model during flight, and the rotation angular velocity (ω) of the dragonfly model is known X ，ω Y ，ω Z And (θ, ψ, γ) to obtain the attitude angle.
The above embodiments are only for illustrating the technical idea and features of the present invention, and the purpose of the embodiments is to enable those skilled in the art to understand the content of the present invention and implement the present invention, and not to limit the protection scope of the present invention by this means. All equivalent changes and modifications made according to the spirit of the present invention should be covered within the protection scope of the present invention.

Claims (8)

1. A quasi-dynamic micro aircraft attitude measurement method is characterized by comprising the following steps:
step 1: installing a gyroscope at the gravity center of the flapping wing aircraft, installing the flapping wing aircraft on a mounting plate, fixing the mounting plate with a rolling shaft of a carbon fiber triaxial bracket, and installing a yaw shaft of the carbon fiber triaxial bracket on an experiment bench;
and 2, step: the flapping wing air vehicle receives a steering engine control signal, starts maneuvering flight, starts changing the attitude of the flapping wing air vehicle, and starts measuring the angular speed of the flapping wing air vehicle in the three-axis direction by the gyroscope;
and step 3: establishing a coordinate system suitable for the flapping wing aircraft, and firstly establishing a 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 Changing coordinates to obtain a relation between an inertial coordinate system of the flapping-wing aircraft and a body coordinate system, wherein the attitude angle of the flapping-wing aircraft is determined by the relation between the inertial coordinate system and the body coordinate system, namely three euler angles which are a pitch angle theta, a yaw angle psi and a roll angle gamma respectively;
and 4, step 4: an attitude dynamics equation of the flapping-wing aircraft is established, and the attitude angle of the flapping-wing aircraft is determined according to the rotation of the aircraft, namely the relation between a body coordinate system and an inertia coordinate system of the flapping-wing aircraft in the whole flight process.
2. The method of claim 1, wherein the flapping wing aircraft is initially in a horizontal position with the gyroscope X axis mounted horizontally and aligned with the forward direction of the aircraft, the gyroscope Y axis pointing upward, and the gyroscope Z axis pointing to the right side of the flapping wing aircraft.
3. The quasi-dynamic micro aircraft attitude measurement method according to claim 1, wherein the step 3 wing coordinate system includes four wings, and a corresponding wing coordinate system is established for each wing, which is respectively a left front wing LF, a left back wing LH, a right front wing RF, and a right back wing RH;
origin O of the coordinate system of the left rear wing LH At the wing root of the left hind wing, O LH Z LH The axis always coincides with the overturning axis of the wing and points to the direction of the wing tip O LH X LH In the left back wing and with O LH Z LH Perpendicular, pointing in the direction of the leading edge, O LH Y LH Determined by the right hand coordinate system;
origin O of coordinate system of right front wing RF At the wing root of the right front wing, O RF Z RF The axis is positioned in the wing, is parallel to the overturning axis of the wing and points to the direction of the wing tip, and O RF X RF Is also atInside the wing, perpendicular to O RF Z RF Axial and pointing in the direction of the leading edge, O RF Y RF Determined by the right hand coordinate system;
the other two wing coordinate systems are similar to the wing systems in the respective ipsilateral direction, except that the origin of coordinates is located at the respective wing root.
4. The quasi-dynamic micro aircraft attitude measurement method of claim 1, wherein the body coordinate system origin O b At the center of mass of the ornithopter, O b X b The axis being parallel to the axis of the fuselage of the aircraft and being defined as pointing forward of the fuselage as positive, O b Y b The axis lying in the plane of symmetry of the fuselage and being parallel to O b X b The axis is vertical and the prescribed direction is positive, O b Z b Determined by the right hand coordinate system;
origin O of the inertial frame g Is the center of mass, O, of the flapping wing aircraft during takeoff g X g In a horizontal plane and defining the direction of advance at take-off as positive, O g Y g Pointing vertically upwards, O g Z g Is determined by the right-hand coordinate system;
the velocity coordinate system origin O a Fixed to the dragonfly center of mass, O a X a Axis coinciding with the dragonfly flight speed, O a Y a The axis is in the dragonfly symmetry plane and perpendicular to O a X a Axis, defined as positive when pointing upwards, negative when pointing downwards, O a Z a The axes are then determined by the right hand coordinate system.
5. The quasi-dynamic micro aircraft attitude measurement method according to claim 1, wherein the specific step of obtaining the relationship between the aircraft inertial coordinate system and the airframe coordinate system in step 3 is:
step 3.1: firstly, the origin and each axis of the inertial coordinate system are respectively superposed with the origin and each axis of the body coordinate system, and the inertial coordinate system is arranged along the line O g Y g Axial rotation by an angle psi, O g X g Shaft and O g Z g The shafts are respectively turned to O g X 1 Shaft and O g Z 1 Axes forming a transitional coordinate system O g X 1 Y g Z 1 ；
Step 3.2: then the transition coordinate system O g X 1 Y g Z 1 Along O g Z 1 Rotation of the shaft by angle theta, O g X 1 Shaft and O g Y g The shafts are respectively turned to O g X b Shaft and O g Y 2 Axes forming another transitional coordinate system O g X b Y 2 Z 1 ；
Step 3.3: finally, the transition coordinate system O is set g X b Y 2 Z 1 Along O g X b Rotation of the shaft by an angle of gamma, O g Y 2 Shaft and O g Z 1 The shafts respectively turn to O g Y b Shaft and O g Z b Axis, finally obtaining a coordinate system O b X b Y b Z b The posture of (2).
6. The quasi-dynamic micro aircraft attitude measurement method of claim 5, wherein the matrix expression of the relationship between the traveling device inertial coordinate system and the aircraft body coordinate system is:
Figure FDA0003756256310000021
where L (γ, θ, ψ) is a conversion matrix of the inertial coordinate system to the body coordinate system.
7. The method for measuring the attitude of the quasi-dynamic micro aircraft according to claim 1, wherein the step 4 of establishing the attitude dynamics equation of the flapping wing aircraft comprises the following specific steps:
step 4.1: the angular velocity and the momentum moment of the flapping wing aircraft in the flying process are respectively set as
Figure FDA0003756256310000031
And
Figure FDA0003756256310000032
and selecting the origin O of the coordinate system of the machine body b As base points, there are:
Figure FDA0003756256310000033
Figure FDA0003756256310000034
wherein I is the rotary inertia of the flapping wing aircraft in the body coordinate system, H X 、H Y 、H Z Moment of momentum for a model of a flapping wing aircraft
Figure FDA0003756256310000035
Component, omega, of each direction in the machine body coordinate system X 、ω Y 、ω Z Component of angular velocity of the flapping wing aircraft model in the coordinate system of the aircraft body, I X 、I Y 、I Z For the moment of inertia, I, of the flapping wing aircraft model to each axis of the aircraft body coordinate system XY 、I XZ 、I YZ 、I YX 、I ZX 、I ZY Inertia products of the flapping wing aircraft model to each axis of the body coordinate system are obtained;
step 4.2: setting the flapping wing aircraft model as a symmetrical model, simplifying the formula in the step 4.1 when the inertia product is zero, and then performing vector cross product on the simplified formula;
step 4.3: the total moment borne by the flapping wing aircraft is respectively set as M in each axial component in a body coordinate system Xb 、M Xb 、M Xb And then:
Figure FDA0003756256310000036
step 4.4: the angular velocity of rotation is equal to the vector sum of the angular velocities of the flapping-wing aircraft model rotation about the respective axes:
Figure FDA0003756256310000037
step 4.5: the transformation relation from the inertial coordinate system to the body coordinate system is as follows:
Figure FDA0003756256310000041
Figure FDA0003756256310000042
the rotation angular speed (omega) of the known model is obtained by connecting the rotation angular speed and the attitude angle of the flapping wing aircraft during flight X ，ω Y ，ω Z And (θ, ψ, γ) to obtain the attitude angle.
8. A quasi-dynamic micro aircraft attitude measurement system is characterized by comprising a flapping wing aircraft, a carbon fiber three-axis support and a gyroscope, wherein the gyroscope is installed at the gravity center of the flapping wing aircraft, the flapping wing aircraft is installed on a mounting plate, the carbon fiber three-axis support comprises a yaw axis, a pitch axis and a roll axis, the mounting plate is fixed with the roll axis of the carbon fiber three-axis support, the yaw axis of the carbon fiber three-axis support is installed on an experiment bench, the pitch axis is rotatably connected to the yaw axis, and the roll axis is installed on the pitch axis; further comprising a coordinate data processing module configured to perform the steps of the quasi-dynamic micro aircraft attitude measurement method according to any one of claims 1 to 7.
CN202210860655.0A 2022-07-20 2022-07-20 Quasi-dynamic micro aircraft attitude measurement method and system Pending CN115200580A (en)

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN116461714A (en) * 2023-05-17 2023-07-21 北京科技大学 Butterfly-simulated ornithopter flight parameter determination method based on dynamic capturing system

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN116461714A (en) * 2023-05-17 2023-07-21 北京科技大学 Butterfly-simulated ornithopter flight parameter determination method based on dynamic capturing system
CN116461714B (en) * 2023-05-17 2024-01-16 北京科技大学 Butterfly-simulated ornithopter flight parameter determination method based on dynamic capturing system

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