CN106595657B - Axial symmetry aircraft attitude measurement device and its measurement method - Google Patents
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- G—PHYSICS
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Abstract
Axial symmetry aircraft attitude measurement device, including solar irradiation can be made to be mapped to transparent outer cover, frustum cone structure body, photovoltaic cell arrays, IMU Inertial Measurement Unit, gyroscope, accelerometer, real-time clock and master controller in the photocell display of laid inside, the frustum cone structure body includes four rotary tables at axial symmetry and frustum cone structure body, is provided with a photovoltaic cell arrays above each rotary table；The photovoltaic cell arrays are connected with master controller, for obtaining the solar vector at a certain moment；Real-time clock is connected with master controller, the IMU Inertial Measurement Unit is connected with master controller, the gyroscope and accelerometer are embedded in IMU Inertial Measurement Unit the angular speed and acceleration that can measure aircraft in real time, the present invention improves the precision of row device attitude measurement, promotes the development of aircraft attitude measurement technology；There is very big facilitation in terms of design and processing and fabricating for Aircraft structural design, aircraft material application field and sensor.
Description
Technical Field
The invention belongs to the technical field of aircraft attitude measurement, and particularly relates to an attitude measurement device of an axisymmetric aircraft.
Background
Flight attitude is an important index for describing an aircraft, and flight attitude measurement is also an essential link in an aircraft control system. The commonly used method for measuring the attitude of the aircraft is to perform data fusion on measured data by using a gyroscope, an accelerometer and a magnetometer to obtain the attitude of the aircraft. However, in general, since the housing or other parts of the aircraft are made of ferromagnetic materials, the gyroscope, the accelerometer, and the magnetometer are seriously interfered, so that the accuracy of the attitude measurement data is reduced, and the flight state of the aircraft cannot be accurately measured.
The problem that the prior art has shortcomings is that an attitude measurement device of an aircraft with high measurement accuracy is needed to be provided.
Disclosure of Invention
The invention provides an attitude measurement device of an axisymmetric aircraft, which utilizes a method of determining a sun position vector by a photocell array to replace a geomagnetic vector to complete attitude measurement of the aircraft.
The invention is realized by the following technical scheme: the attitude measuring device of the axisymmetric aircraft comprises a transparent shell, a circular truncated cone structure body, a photocell array, an IMU (inertial measurement unit), a gyroscope, an accelerometer, a real-time clock and a main controller, wherein the transparent shell can enable sunlight to irradiate the photocell array laid inside the transparent shell; the photovoltaic cell array is connected with the main controller and is used for obtaining a sun vector at a certain moment; the real-time clock is connected with the main controller, the IMU inertia measurement unit is connected with the main controller, and the gyroscope and the accelerometer are embedded in the IMU inertia measurement unit and can measure the angular velocity and the acceleration of the aircraft in real time.
As a preferable technical scheme, the transparent shell is made of polycarbonate engineering plastic materials.
As a preferred technical solution, the photocell array is arranged on the circular truncated cone according to an angle range of 20 ° to 90 °, and 60 ° is preferred.
As a preferable technical solution, the photocell array can determine the position vector of the sun according to the time given by the real-time clock, the revolution and rotation of the earth, the sun orbit information, and the angle of the photocell array arranged on the circular table.
The measuring method of the attitude measuring device of the axisymmetric aircraft comprises the following steps:
step 1: calculating a sun position vector and a gravity acceleration vector under a navigation coordinate system;
the gravity acceleration vector under the navigation coordinate system is as follows: [0, 0, g ]]TThe calculation of the sun position vector under the navigation coordinate system relates to the transformation of the coordinate system, and comprises the following specific calculation steps:
step 1.1: calculating the sun position vector under the geocentric inertial coordinate system:
the sun is arranged to rotate around the earth according to the sun orbit elements, wherein the sun orbit elements comprise a ascension point and right ascension channel omega, an amplitude angle omega of a near-sun point, a true near-sun point angle f,The orbit inclination angle i, the orbit semimajor axis a and the orbit eccentricity e are calculated, and the sun position vector S under the geocentric inertial coordinate system is calculatedIThe specific process is as follows:
wherein, E is a near point angle, x and y are the positions of the sun in a rectangular coordinate system, r is the module of the sun radial direction, and r is a (1-E cos E);
then the sun vector S under the earth' S center inertial coordinate systemIAs shown in equation (2):
wherein,
step 1.2: calculating sun position vector under navigation coordinate system
The transformation of the sun position vector from the geocentric inertial coordinate system to the navigation coordinate system needs to be multiplied by a plurality of rotation matrixes to obtain the sun position vector S expressed in the navigation coordinate systemNAs shown in formula (3):
SN＝RZ(90°)·RY(-L)·RZ(ωet)·SI (3)
in the formula: rZ(ωet) is a rotation matrix of the earth's rotation; rY(-L) is the moment of rotation R of the latitude of the aircraftZ(90 °) array; a rotation matrix rotated 90 degrees around the Z axis;
step 2: an accelerometer 7 embedded in the IMU inertia measurement unit 5 measures gravity acceleration G, and a photocell array 3 measures a solar vector S;
and step 3: : correcting errors of a gyroscope 6 embedded in the IMU inertial measurement unit 5 and obtaining an accurate attitude angle of the aircraft, and specifically comprising the following steps:
step 3.1: gravity acceleration G measured under navigation coordinate systemNAnd sun vector SNRespectively multiplying by attitude rotation matrixes represented by quaternions to obtain the gravity acceleration G under a carrier coordinate systemBAs shown in equation (4), there is:
GB＝Rbn(q)GN (4)
obtaining the sun vector S under the carrier coordinate systemBGBAs shown in equation (5), there is:
SB＝Rbn(q)SN (5)
wherein G isNRepresenting gravitational acceleration in a navigational coordinate system, GBGravitational acceleration, S, under a carrier system obtained after coordinate transformationNRepresenting the sun vector, S, in a navigational coordinate systemBThe sun vector R under the carrier system obtained after coordinate transformationbn(q) is a transformation matrix from the navigation coordinate system to the carrier coordinate system, as shown in equation (6):
wherein q is0，q1，q2，q3Is a quaternion;
step 3.2: gravitational acceleration G transformed from navigation coordinate system to carrier coordinate systemBAnd sun vector SBAnd (3) cross-multiplying the gravity acceleration G measured by the accelerometer 7 embedded in the IMU inertia measurement unit 5 and the sun vector S measured by the photocell array 3 to obtain a correction error delta, as shown in a formula (7):
Δ＝G×GB+S×SB (7)
wherein Δ represents a correction error;
step 3.3: the correction error Δ calculated in step 3.2 is used as a PI correction error of the gyroscope 6 embedded in the IMU inertial measurement unit 5 to correct the angular velocity measured by the gyroscope 6, as shown in formula (8):
ωg＝ω0+KPΔ+KI∫Δ (8)
wherein, ω isgIndicating the corrected angular velocity, ω0Representing the initial angular velocity, K, of the gyroscope 6PDenotes the proportional amplification factor, KIRepresents an integral amplification factor;
step 3.4: updating quaternion by using quaternion differential equation, and obtaining the corrected attitude angle of the aircraft by using a conversion formula of quaternion and attitude angle, wherein the formulas are shown in formulas (8) and (9):
wherein q is (q)0，q1，q2，q3)TAnd represents a vector of quaternions,represents the derivative of q with respect to time;
wherein omegagThe matrix formed by the measured accelerations for the gyroscope 6 is shown in equation (10):
where Φ, θ, Ψ represent the roll angle, pitch angle, and yaw angle of the aircraft, respectively.
Through the steps, the accurate attitude of the axisymmetric aircraft can be measured.
The aircraft outer shell is dug out to be equal to the transparent outer shell in size from the outside of the aircraft, and the transparent outer shell is laid on the part. And reinforcing the joint of the aircraft shell and the transparent shell by using a special processing technology in the aircraft interior, so that the transparent shell and the aircraft shell are integrated.
The total number of the photocells laid on the surfaces of the circular truncated cones can be set according to different aircrafts, so that the photocells of the four circular truncated cones can receive solar radiation and convert the energy of the solar radiation into electromotive force. The energy of the solar radiation received by each circular table is different, and the magnitude of the solar radiation energy is related to the magnitude of the generated electromotive force.
The photocell array can determine the position vector of the sun according to the given time of a real-time clock, the revolution and rotation of the earth, the sun orbit element information, the laying angle of the photocell array and the difference of the solar radiation energy received by the four round tables.
The main controller is connected with the photocell array, the IMU inertial measurement unit and the real-time clock, can process data measured by each sensor in real time, and can fuse the data by using corresponding algorithms to calculate the real attitude of the aircraft.
The photocell array and the real-time clock are connected with the main controller and used for obtaining the position vector information of the sun.
The IMU inertia measurement unit is connected with the main controller, can measure the angular velocity and the acceleration of the aircraft in real time, and sends the measured data to the main controller.
The main controller is connected with the photocell array, the IMU inertia measurement module and the real-time clock, can process data measured by each sensor in real time, and can fuse the data by using corresponding algorithms to calculate the real attitude of the aircraft.
Compared with the prior art, the invention has the beneficial effects that: the precision of the aircraft attitude measurement is greatly improved, and the development of the aircraft attitude measurement technology is further promoted. Meanwhile, the method has great promotion effects on the aspects of aircraft structure design, aircraft material application field and design, processing and manufacturing of the sensor.
Drawings
FIG. 1 is a schematic structural diagram of a measuring device according to an embodiment of the present invention;
FIG. 2 is a block diagram of an aircraft attitude measurement system;
FIG. 3 is a sun orbit element diagram;
FIG. 4 is a diagram of the relationship between the geocentric inertial coordinate system and the navigation coordinate system;
FIG. 5 is a schematic diagram of a photovoltaic cell array structure according to an embodiment of the present invention;
FIG. 6 is a top view of the photovoltaic cell array of FIG. 5;
fig. 7 is a diagram of the sun position.
Description of reference numerals: the device comprises a transparent shell 1, a structural body 2 used for laying a photocell, a photocell array 3, a real-time clock 4, an inertial measurement unit 5 IMU, a gyroscope 6, an accelerometer 7, a main controller 8 and a circular truncated cone 9.
Detailed Description
The invention will be further explained with reference to the drawings.
Axisymmetric aircraft attitude measurement device: the solar photovoltaic power generation system comprises a transparent shell 1, a circular truncated cone structure body 2, a photovoltaic cell array 3, an IMU inertia measurement unit 5, a gyroscope 6, an accelerometer 7, a real-time clock 4 and a main controller 8, wherein sunlight can irradiate on a photovoltaic cell array laid inside, the circular truncated cone structure body 2 is axisymmetric, the circular truncated cone structure body 2 comprises four circular truncated cones 9, and one photovoltaic cell array 3 is arranged on each circular truncated cone 9; the photocell array 3 is connected with the main controller 8 and is used for obtaining a sun vector at a certain moment; the real-time clock 4 is connected with the master controller 8, the IMU inertia measurement unit 5 is connected with the master controller 8, and the gyroscope 6 and the accelerometer 6 are embedded in the IMU inertia measurement unit 5 and can measure the angular velocity and the acceleration of the aircraft in real time.
As a preferable technical scheme, the transparent shell 1 is made of polycarbonate engineering plastic materials.
As a preferred technical scheme, the installation angle of the photocell array 3 is tau1，τ2The circular truncated cone 9 is provided with an angle range of 20-90 degrees, and 60 degrees is preferably selected.
As a preferred technical solution, the photocell array 3 can determine the position vector of the sun according to the time given by the real-time clock 4, the revolution and rotation of the earth, the sun orbit information, and the angle of the photocell array 3 arranged on the circular table.
The measuring method of the attitude measuring device of the axisymmetric aircraft comprises the following steps:
fig. 1 shows a structure of a measuring device of the present invention, which includes a transparent casing 1, a circular truncated cone structure 2, a photocell array 3, an IMU inertia measuring unit 5, a gyroscope 6, an accelerometer 7, a real-time clock 4 and a master controller 8, which enable sunlight to irradiate on an array of photocells laid inside, a structural block diagram of the measuring device for measuring the real attitude angle of an aircraft is shown in fig. 2, when sunlight irradiates on the photocell array 3 arranged on the circular truncated cone structure 2 through the transparent casing 1 made of polycarbonate engineering plastics, the photocell array 3 obtains a solar vector at this moment in combination with real-time measured by the real-time clock 4, the solar vector measured by the photocell array 3 is input into the master controller 8 together with the acceleration of gravity and the angular velocity of the aircraft measured by the accelerometer 7 and the gyroscope 6 embedded in the IMU inertia measuring unit 5, the main controller 8 corrects the angular velocity measured by the gyroscope through a corresponding algorithm, so that the real attitude angle of the aircraft is obtained.
When processing and fusing sensor data, the main controller 8 mainly relates to calculation of solar vectors and gravitational acceleration under a navigation coordinate system and a carrier coordinate system and transformation of the coordinate system to solve attitude angles, and specifically comprises the following steps:
step 1: calculating a sun position vector and a gravity acceleration vector under a navigation coordinate system;
the gravity acceleration vector under the navigation coordinate system is as follows: [0, 0, g ]]TThe calculation of the sun position vector under the navigation coordinate system relates to the transformation of the coordinate system, and comprises the following specific calculation steps:
step 1.1: calculating the sun position vector under the geocentric inertial coordinate system:
the sun is set to rotate around the earth, and according to the sun orbit elements, the sun orbit elements comprise a rising point right ascension omega, a near-sun amplitude omega, a true near point angle f, an orbit inclination angle i, an orbit semimajor axis a and an orbit eccentricity e, a sun position vector S under the geocentric inertial coordinate system is calculatedIAs shown in fig. 3, the specific process is as follows:
wherein, E is a near point angle, x and y are the positions of the sun in a rectangular coordinate system, r is the module of the sun radial direction, and r is a (1-E cos E);
then the sun vector S under the earth' S center inertial coordinate systemIAs shown in equation (2):
wherein,
step 1.2: calculating sun position vector under navigation coordinate system
In connection with FIG. 4, ωeThe rotation angular velocity of the earth is represented, the sun position vector is converted from the geocentric inertia coordinate system to the navigation coordinate system and needs to be multiplied by a plurality of rotation matrixes, and the rotation matrixes are solvedThe sun position vector expressed under the navigation coordinate system is shown as the formula (3):
SN＝RZ(90°)·RY(-L)·RZ(ωet)·SI (3)
in the formula: rZ(ωet) is a rotation matrix of the earth's rotation; rY(-L) is the rotation matrix of the latitude of the aircraft; rZ(90 °) is a rotation matrix rotated 90 degrees around the Z-axis;
step 2: an accelerometer 7 embedded in the IMU inertia measurement unit 5 measures gravity acceleration G, and a photocell array 3 measures a solar vector S; the structural schematic diagram of the arrangement of the photovoltaic cell array 3 is shown in fig. 5, and the top view is shown in fig. 6, wherein the areas indicated by the thick lines are laid photovoltaic cells; the photovoltaic cell array 3 can convert solar energy into electromotive force, the magnitude of the electromotive force is related to the magnitude of solar radiation energy, the photovoltaic cell array can receive different solar energy values through the angle set by the photovoltaic cell array 3 and different positions, and according to the accurate time provided by a real-time clock, the solar position S at the time can be measured as shown in FIG. 7;
and step 3: : correcting errors of a gyroscope 6 embedded in the IMU inertial measurement unit 5 and obtaining an accurate attitude angle of the aircraft, and specifically comprising the following steps:
step 3.1: gravity acceleration G measured under navigation coordinate systemNAnd sun vector SNRespectively multiplying by attitude rotation matrixes represented by quaternions to obtain the gravity acceleration G under a carrier coordinate systemBAs shown in equation (4), there is:
GB＝Rbn(q)GN (4)
obtaining the sun vector S under the carrier coordinate systemB GBAs shown in equation (5), there is:
SB＝Rbn(q)SN (5)
wherein G isNRepresenting gravitational acceleration in a navigational coordinate system, GBGravitational acceleration, S, under a carrier system obtained after coordinate transformationNRepresenting the sun vector, S, in a navigational coordinate systemBThe sun vector R under the carrier system obtained after coordinate transformationbn(q) is a transformation matrix from the navigation coordinate system to the carrier coordinate system, as shown in equation (6):
wherein q is0，q1，q2，q3Is a quaternion;
step 3.2: gravitational acceleration G transformed from navigation coordinate system to carrier coordinate systemBAnd sun vector SBAnd (3) cross-multiplying the gravity acceleration G measured by the accelerometer 7 embedded in the IMU inertia measurement unit 5 and the sun vector S measured by the photocell array 3 to obtain a correction error delta, as shown in a formula (7):
Δ＝G×GB+S×SB (7)
wherein Δ represents a correction error;
step 3.3: the correction error Δ calculated in step 3.2 is used as a PI correction error of the gyroscope 6 embedded in the IMU inertial measurement unit 5 to correct the angular velocity measured by the gyroscope 6, as shown in formula (8):
ωg＝ω0+KPΔ+KI∫Δ (8)
wherein, ω isgIndicating the corrected angular velocity, ω0Representing the initial angular velocity, K, of the gyroscope 6PDenotes the proportional amplification factor, KIRepresents an integral amplification factor;
step 3.4: updating quaternion by using quaternion differential equation, and obtaining the corrected attitude angle of the aircraft by using a conversion formula of quaternion and attitude angle, wherein the formulas are shown in formulas (8) and (9):
wherein q is (q)0，q1，q2，q3)TAnd represents a vector of quaternions,represents the derivative of q with respect to time;
wherein omegagThe matrix formed by the measured accelerations for the gyroscope 6 is shown in equation (10):
where Φ, θ, Ψ represent the roll angle, pitch angle, and yaw angle of the aircraft, respectively.
Through the steps, the accurate attitude of the axisymmetric aircraft can be measured.
The aircraft outer shell is dug out to be equal to the transparent outer shell in size from the outside of the aircraft, and the transparent outer shell is laid on the part. And reinforcing the joint of the aircraft shell and the transparent shell by using a special processing technology in the aircraft interior, so that the transparent shell and the aircraft shell are integrated.
The total number of the photocells laid on the surfaces of the circular truncated cones can be set according to different aircrafts, so that the photocells of the four circular truncated cones can receive solar radiation and convert the energy of the solar radiation into electromotive force. The energy of the solar radiation received by each circular table is different, and the magnitude of the solar radiation energy is related to the magnitude of the generated electromotive force.
The photocell array can determine the position vector of the sun according to the given time of a real-time clock, the revolution and rotation of the earth, the sun orbit element information, the laying angle of the photocell array and the difference of the solar radiation energy received by the four round tables.
The main controller is connected with the photocell array, the IMU inertial measurement unit and the real-time clock, can process data measured by each sensor in real time, and can fuse the data by using corresponding algorithms to calculate the real attitude of the aircraft.
The photocell array and the real-time clock are connected with the main controller and used for obtaining the position vector information of the sun.
The IMU inertia measurement unit is connected with the main controller, can measure the angular velocity and the acceleration of the aircraft in real time, and sends the measured data to the main controller.
The main controller is connected with the photocell array, the IMU inertia measurement module and the real-time clock, can process data measured by each sensor in real time, and can fuse the data by using corresponding algorithms to calculate the real attitude of the aircraft.
The invention greatly improves the precision of the aircraft attitude measurement and further promotes the development of the aircraft attitude measurement technology. Meanwhile, the method has great promotion effects on the aspects of aircraft structure design, aircraft material application field and design, processing and manufacturing of the sensor.
In summary, the preferred embodiments of the present invention are not intended to limit the scope of the invention, and all equivalent changes and modifications made according to the content of the claims of the present invention should fall within the technical scope of the present invention.
Claims (4)
1. An axisymmetric aircraft attitude measurement method adopts an axisymmetric aircraft attitude measurement device, the device comprises a transparent shell (1) which can enable sunlight to irradiate a photocell array laid inside, a circular truncated cone structure body (2), the photocell array (3), an IMU inertia measurement unit (5), a gyroscope (6), an accelerometer (7), a real-time clock (4) and a main controller (8), the circular truncated cone structure body (2) is axisymmetric, the circular truncated cone structure body (2) comprises four circular truncated cones (9), and one photocell array (3) is arranged on each circular truncated cone (9); the photovoltaic cell array (3) is connected with the main controller (8) and is used for obtaining a sun vector at a certain moment; real-time clock (4) link to each other with master controller (8), IMU inertial measurement unit (5) link to each other with master controller (8), but gyroscope (6) and accelerometer (7) are embedded in IMU inertial measurement unit (5) real-time measurement aircraft's angular velocity and acceleration, its characterized in that: the method comprises the following steps:
step 1: calculating a sun position vector and a gravity acceleration vector under a navigation coordinate system; gravitational acceleration G under navigation coordinate systemNAnd the gravity acceleration vector under the navigation coordinate system is as follows: [0, 0, g ]]TThe calculation of the sun position vector under the navigation coordinate system relates to the transformation of the coordinate system, and comprises the following specific calculation steps:
step 1.1: calculating the sun position vector under the geocentric inertial coordinate system
The sun is set to rotate around the earth, and according to the sun orbit elements, the sun orbit elements comprise a rising point right ascension omega, a near-sun amplitude omega, a true near point angle f, an orbit inclination angle i, an orbit semimajor axis a and an orbit eccentricity e, a sun position vector S under the geocentric inertial coordinate system is calculatedIThe specific process is as follows:
wherein E is a near point angle, x and y are the positions of the sun in a rectangular coordinate system, r is a radial module of the sun, and r is a (1-ecosE);
then the sun vector S under the earth' S center inertial coordinate systemIAs shown in equation (2):
wherein
Step 1.2: calculating sun position vector under navigation coordinate system
The transformation of the sun position vector from the geocentric inertial coordinate system to the navigation coordinate system needs to be multiplied by a plurality of rotation matrixes to obtain the sun position vector S expressed in the navigation coordinate systemNAs shown in formula (3):
SN＝RZ(90°)·RY(-L)·RZ(ωet)·SI (3)
in the formula: rZ(ωet) is a rotation matrix of the earth's rotation; rY(-L) is the rotation matrix of the latitude of the aircraft; rZ(90 °) a rotation matrix rotated 90 degrees around the Z-axis;
step 2: an accelerometer (7) embedded in the IMU inertia measurement unit (5) measures gravity acceleration G, and a photocell array (3) measures a solar vector S;
and step 3: correcting errors of a gyroscope (6) embedded in an inertial measurement unit (5) and an Inertial Measurement Unit (IMU) to obtain an accurate attitude angle of the aircraft, and specifically comprising the following steps:
step 3.1: gravity acceleration G measured under navigation coordinate systemNAnd sun vector SNRespectively multiplying by attitude rotation matrixes represented by quaternions to obtain the gravity acceleration G under a carrier coordinate systemBAs shown in equation (4), there is:
GB＝Rbn(q)GN (4)
obtaining the sun vector S under the carrier coordinate systemBAs shown in equation (5), there is:
SB＝Rbn(q)SN (5)
wherein G isNRepresenting gravitational acceleration in a navigational coordinate system, GBGravitational acceleration, S, under a carrier system obtained after coordinate transformationNRepresenting the sun vector, S, in a navigational coordinate systemBThe sun vector R under the carrier system obtained after coordinate transformationbn(q) is a transformation matrix from the navigation coordinate system to the carrier coordinate system, as shown in equation (6):
wherein q is0，q1，q2，q3Is a quaternion;
step 3.2: gravitational acceleration G transformed from navigation coordinate system to carrier coordinate systemBAnd sun vector SBAnd (3) performing cross multiplication on the gravity acceleration G measured by an accelerometer (7) embedded in the IMU inertia measurement unit (5) and the solar vector S measured by the photocell array (3) to obtain a correction error delta, as shown in the formula (7):
△＝G×GB+S×SB (7)
wherein Δ represents a correction error;
step 3.3: and 3.2, taking the correction error delta calculated in the step 2 as a PI correction error of the gyroscope (6) embedded in the IMU inertia measurement unit (5), and correcting the angular velocity measured by the gyroscope (6), as shown in a formula (8):
ωg＝ω0+Kp△+KI∫△ (8)
wherein, ω isgIndicating the corrected angular velocity, ω0Represents the initial angular velocity, K, of the gyroscope (6)pDenotes the proportional amplification factor, KIRepresents an integral amplification factor;
step 3.4: updating quaternion by using quaternion differential equation, and obtaining the corrected attitude angle of the aircraft by using a conversion formula of quaternion and attitude angle, wherein the formulas are shown in formulas (9) and (10):
wherein q is [ q ]0,q1,q2,q3]TRepresents a quaternion vectorRepresents the derivative of q with respect to time;
θ＝-arcsin(2(q1q3+q0q2))
wherein omegagA matrix of measured accelerations for the gyroscope (6), as shown in equation (11):
wherein phi, theta and psi respectively represent the roll angle, the pitch angle and the yaw angle of the aircraft;
through the steps, the accurate attitude of the axisymmetric aircraft can be measured.
2. The axisymmetric aircraft attitude measurement device of claim 1, being characterized in that: the transparent shell (1) is made of polycarbonate engineering plastic materials.
3. The axisymmetric aircraft attitude measurement device of claim 1, being characterized in that: the photocell array (3) is arranged on the circular table (9) according to an angle range of 20-90 degrees, and 60 degrees is preferably selected.
4. The axisymmetric aircraft attitude measurement device of claim 1, being characterized in that: the photocell array (3) can determine the position vector of the sun according to the given time of the real-time clock (4), the revolution and rotation of the earth, the sun orbit information and the angle of the photocell array (3) arranged on the circular truncated cone.
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| CN201610945011.6A CN106595657B (en) | 2016-11-02 | 2016-11-02 | Axial symmetry aircraft attitude measurement device and its measurement method |
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| CN106595657A CN106595657A (en) | 2017-04-26 |
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