JPH0535971B2 - - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION [Field of Industrial Application] The present invention relates to an attitude detection device for a navigation object such as an aircraft or a car using a gyro, an accelerometer, and a magnetic orientation sensor.

[Conventional technology]

Conventionally, aircraft have a DG (directional gyro) that indicates the direction, a magnetic compass (or flux valve compass), a VG (horizontal gyro) that indicates the attitude angle, and a vertical axis. It is equipped with a turn indicator that indicates the surrounding turning angular velocity and bank angle, supplementing the pilot's senses and enabling safe flight under any conditions.

Next, the above conventional device will be explained with reference to FIGS. 4 to 9.

FIG. 4 is a perspective view of an example of a directional gyro currently used in aircraft. As shown in the figure, this gyro has spin shafts 100, 10
A gyro rotor 101 whose 0a is substantially horizontal and rotates at high speed is rotatably supported by its spin shafts 100, 100a by a horizontal ring 102 which is a gyro case. The horizontal ring 102 has a horizontal axis 10 at a position perpendicular to the spin axes 100, 100a.
3,103a, these horizontal axes 103,10
3a are rotatably fitted into horizontal shaft bearings 105, 105a (105a not shown) fixedly installed at corresponding positions on the vertical ring 104. vertical ring 104
are vertical shafts 106, 106a that protrude vertically at positions orthogonal to the horizontal shaft bearings 105, 105a.
These vertical shafts 106, 106a are fixed to vertical shaft bearings 107, 107a (107a) fixed at corresponding positions of a base 107B fixed to the aircraft.
(not shown). An upright torquer 108 and a compass card 109 are attached to the upper vertical shaft 106a. Lower vertical axis 10
6 includes a receiving synchronizer 110 and a transmitting synchronizer 1.
11 is installed.

The horizontal ring 102 has spin shafts 100, 100a.
An electrolyte level 112 is mounted which detects the inclination of the liquid relative to the horizontal plane. The output of the electrolyte level 112 is fed back to the upright torquer 108 via an amplifier 113 and the output of the gyro rotor 101 is
The spin axes 100, 100a of the spindles 100 and 100a are always held horizontally. This loop is called a stand-up loop. The magnetic azimuth output of the flux valve 114, which magnetically detects the azimuth angle of the aircraft body, is transmitted to the receiving sink 1.
Here, a deviation signal is generated between the magnetic azimuth output and the azimuth of the spin axes 100, 100a, that is, the gyro azimuth, and this deviation signal is sent to the amplifier 114.
A is fed back to the slave torquer 115 provided on the horizontal shaft 103 to match the gyro direction with the magnetic direction. This loop is called an orientation constraint loop. When the aircraft is in intense motion, the gyro azimuth is output, and errors in azimuth due to gyro drift are restrained by the magnetic azimuth from the flux valve 114 to maintain accuracy. The aircraft direction is determined by the compass card 1 attached to the vertical axis 106a.
Read by 09.

FIG. 5 is an example of a horizontal instrument currently in use for detecting the inclination angle (roll angle, pitch angle) of an aircraft body. In this example, internal gimbal 132
has a built-in gyro rotor 130 that rotates at high speed while holding a spin shaft 131 substantially vertically. The internal gimbal 132 is connected to the spin axis 131
Pitch axes 133, 133 at horizontal positions orthogonal to
The pitch shafts 133, 133a are rotatably fitted into pitch shaft bearings 134, 134a (the pitch shaft bearings 134 are not visible) fixedly installed at corresponding positions on the external gimbal 135. external gimbal 13
5 has roll shafts 136, 136a at positions orthogonal to the pitch shafts 134, 134a, and these roll shafts 136, 136a are provided on roll bases 138, 138a attached in the tail direction of the aircraft. It is rotatably fitted into the roll shaft bearings 137, 137a. The internal gimbal 132 has a roll electrolyte level 139 that detects the tilt of the spin axis 131 relative to the horizontal plane about the roll axes 136, 136a, and a pitch electrolyte level 142 that detects the tilt about the pitch axes 133, 133a.

The output of the roll electrolyte level 139 is sent via a roll amplifier 140 to a roll torquer 141 attached to the pitch shaft 133.
Feedback is performed so that the output of 39 becomes zero. This loop is called the roll erector system. On the other hand, the output of the pitch electrolyte level 142 is fed back via a pitch amplifier 143 to a pitch torquer 144 attached to the roll shaft 136, so that the tilt of the spin shaft 131 around the pitch axes 133, 133a is maintained at zero. This loop is called the pitch orthostatic system. The roll angle of the aircraft is roll axis 1
From the roll angle transmitter 145 attached to 36a,
Further, the pitch angles are each output from a pitch angle transmitter 146 attached to the pitch shaft 133a.

FIG. 6 shows the display section of a turn indicator currently used in aircraft. Using the base line 151 and the pointer 152, the turning angular velocity of the aircraft is displayed by the gyro shown in FIG. The lower half of the display section is a bank angle display section 154, which outputs and displays the bank angle based on the position of a ball 155 enclosed within a circular ring having curvature.

FIG. 7 shows a portion of a rate gyro for detecting the turning angular velocity of the turning meter. Gyroscope rotor 1
A gyro case 171 containing the gyro rotor 170 has output shafts 173, 173a at positions orthogonal to the axis XX' of the spin shaft 172 of the gyro rotor 170,
These output shafts 173, 173a are fixed to output shaft bearings 175, 1 fixed to a base 174 fixed to the aircraft body.
It is rotatably fitted into 75a. Gyro case 17
1 and the base 174, a restoring spring 176 and a damping pot 177 are provided.

Output shaft axis YY′ and axis of spin shaft 172
When a turning angle Ω is applied around the input shaft ZZ' which is orthogonal to both XX', a torque proportional to the turning angular velocity Ω is generated around the output shaft axis YY' due to the gyro effect. This torque rotates the inside of the gyro case 171 around the output shaft axis YY', and a torque is generated by the restoring spring 176 in accordance with the angle change, creating a balanced state. In other words, the input angular velocity Ω has been converted into a rotation angle around the output shaft axis YY', and this angle of change is displayed by the pointer 178 (corresponding to the pointer 152 in FIG. 6) attached to the output shaft 173a. Output.

FIG. 8 is a partial sectional view showing an example of the magnetic direction sensor 7 called a flux valve compass. In this example, the magnetic direction detection unit 7 is connected from the center of the base 7-1 via the universal joint 7-2.
-3 is suspended in a pendulum shape. A bowl 7-4 is attached to the bottom surface of the base 7-1, and a magnetic direction detection section 7-4 is attached to the bottom surface of the base 7-1.
Construct a container for 3. This container is filled with damping oil 7-5 to damp the pendulum motion of the magnetic direction detecting section 7-3. Due to the action of the universal joint 7-2, the magnetic azimuth detecting section 7-3 is always held horizontally regardless of the inclination of the navigation object, and the occurrence of azimuth errors due to inclination can be prevented.

FIG. 9 is a perspective view of the magnetic direction detecting section 7-3. There is an exciter coil 70 wound in the center.
is excited with an AC power supply. Made from high permeability material 3
Three pick-up coils 72-1, 7 are wound around spokes 71-1, 71-2, 71-3 fixed to the exciter coil 70 at equal angular intervals.
At 2-2 and 72-3, AC voltages corresponding to the azimuth of the navigation object are generated with a phase of 120°.

[Problem that the invention seeks to solve]

However, the above-mentioned conventional devices are mechanically complex, require skill and time for assembly and adjustment, are expensive, and have wear parts such as ball bearings and sliding electrical circuits, so they require periodic It required regular maintenance and inspection, and it also had problems such as being susceptible to vibration and shock.

In addition, the magnetic orientation sensor supports its magnetic orientation detection portion with a universal joint, and requires oil for damping, making it large, heavy, and expensive. Therefore, if the magnetic direction detection unit is attached directly to the navigation object, it can be made smaller and
Although the weight can be reduced, the vertical component of the earth's magnetism is detected due to the inclination of the navigation vehicle, which has the disadvantage of increasing azimuth errors.

Therefore, the present invention provides a new attitude detection device that solves the problems of conventional devices.
Three gyroscopes 1, 2, 3, three accelerometers 4, 5, 6, and one magnetic sensor 7 attached to the navigation vehicle so that their input axes coincide with the respective axes.
A, the above 3 gyroscopes, 3 accelerometers and 1
A signal conversion section 8 receives the outputs of the magnetic sensors, a calculation section 9 receives the outputs of the signal conversion sections, and a signal output section 10 receives the outputs of the calculation sections and outputs attitude signals to the outside. The magnetic azimuth detection section 7-3 of the magnetic azimuth sensor is directly attached to the navigation object, and the tilt error calculation section 60 is used to correct errors caused by the tilt of the magnetic azimuth sensor.
The azimuth error of the magnetic azimuth sensor is calculated and corrected from the azimuth, roll angle, and pitch angle of the vehicle obtained from the CTM calculation unit and the horizontal and vertical components of the geomagnetic field provided in the tilt error correction calculation unit. do.

[Effect]

The outputs of three gyroscopes 12 and 3 installed with their input axes aligned with the three main axes of the navigation object are sent to a coordinate transformation matrix (CTM) calculation unit 53 via bias correctors 50, 51 and 52. Input,
Calculate CTM. The outputs of three accelerometers 4, 5, 6 installed on the aircraft so that the input axes of the gyroscopes 1, 2, 3 and their acquisition axes are parallel,
A horizontal component calculating section 54 calculates horizontal components α and β of the gravitational acceleration from the above-mentioned CTM signal.
If the CTM is correct, the horizontal components α and β are zero, but if there is an error in the CTM, the horizontal components α and β become finite values, so the upright torque calculation unit 56 calculates the correct value for the CTM. It is converted into east-west and north-south direction torqueing signals ET1 and ET2 and sent to the CTM calculation unit 53.

On the other hand, the azimuth signal AS in the CTM calculation unit 53,
Magnetic direction signal MAS from magnetic direction sensor 7A
The magnetic azimuth signal MASI whose inclination error has been corrected through the inclination error correction calculation unit 60 is compared with the magnetic azimuth signal MASI in the azimuth constraint torque calculation unit 57 to create a torque signal around the azimuth axis, which eliminates the difference between the two. Feedback is provided to the CTM calculation unit 53 as follows.

On the other hand, if there is a drift in the gyro, the outputs of the upright torque calculation section 56 and the azimuth restraint torque calculation section 57 will not be zero, but will have finite values corresponding to the gyro drift. These output signals are inputted together with the CTM signal to the gyro bias calculating section 58, and the calculated bias correction signals are inputted to the bias correctors 50, 51, 52 of each gyro so that the gyro drift becomes zero. Fix it.

Roll angle, pitch angle, azimuth angle, and bias corrector 50, 51, 5 of the navigation object from the CTM calculation unit 53
The angular velocity of the navigation object from 2 and the sideslip signal from the Y accelerometer 5 are output, respectively. When the navigation object is making acceleration movements such as turning or increasing/decreasing speed,
Cut off the input to the standing torque calculation unit 56,
Remove acceleration effects. When a speed signal is obtained from another sensor, this signal and the CTM signal are supplied to the acceleration correction calculation unit 55 to remove the influence of acceleration.

〔Example〕

An embodiment of the present invention will be described below with reference to FIGS. 1 and 2.

FIG. 1 is a block diagram showing the entirety of an example of the attitude detection device of the present invention. In the example shown in the figure, an X gyro 1, a Y gyro 2, which are non-rotating gyros such as a vibrating gyro or a gas rate gyro,
Z gyroscope 3, X accelerometer 4, Y accelerometer 5,
The Z accelerometer 6 and the magnetic orientation sensor 7A without a gimbal are installed so that their respective input axes coincide with the three main orthogonal axes of the navigation object (not shown), that is, the X, Y, and Z axes ( (see arrow). The output signals from these are sent to the calculation unit 9 via the signal conversion unit 8.
Enter. The calculation unit 9 calculates the coordinate transformation matrix CTM, gyro drift correction calculation,
After performing acceleration correction calculations, etc., control of the navigation object,
Roll angle, pitch angle, azimuth angle, X, necessary for operation
Signals such as Y, Z angular velocity, bank angle, slip angle, etc.
The signal is sent out via the signal output section 10.

FIG. 2 is a block diagram showing an example of the configuration of the calculation section 9 of FIG. 1. The gyro signals XG, YG, and ZG from the X, Y, and Z gyros 1, 2, and 3 from the signal converter 8 shown in FIG. and the signal output section 1 shown in FIG. 1 as the Z rate.
0, and is also input to a CTM (coordinate transformation matrix) calculation unit 53 to calculate CTM.

On the other hand, X from the signal converter 8 shown in FIG.
Acceleration signals XA, YA, and ZA from the Y, Z, and Z accelerometers 4, 5, and 6 are input to the horizontal component calculation unit 54 together with the CTM signal CS from the CTM calculation unit 53, where the Calculate horizontal components α and β of acceleration. The horizontal components α and β are input to the acceleration correction calculation unit 55 together with the speed signal SS from the accelerometer of the aircraft, and there, after removing the component of the motion acceleration of the navigation object, they are input to the upright torque calculation unit 5.
6, and after calculating the standing torque, CTM
The signal is input to the calculation unit 53, and the CTM is torqued so that the horizontal components α and β become zero.

Azimuth signal AS from CTM calculation unit 53 and magnetic azimuth signal MAS1 from tilt error correction calculation unit 60
is supplied to the azimuth constraint torque calculation section 57, where a comparative computation constraint torque calculation is performed, and the output torque TS is fed back to the CTM calculation section 53, and the CTM is mainly torqued around the azimuth axis. Constrain the CTM orientation to the magnetic orientation. The slope error correction calculation unit 60 is
The azimuth angle of the navigation object obtained from the CTM calculation unit 53,
The azimuth error of the magnetic azimuth sensor shown in the following formula (1) is calculated from the roll angle, pitch angle, and the values of the horizontal component and vertical component of the earth's magnetism provided in the tilt error correction calculation unit, and the error caused by the tilt of the magnetic azimuth sensor is calculated. Correct any errors that occur.

Δφ＝H _h cosφ{sinφ(cosθ−cosφ)＋cosφ・sinθ
sinψ}/H _h (sin ² φcosθ＋cos ² φcosφ＋cosφsinφs
inθsinψ) −H _v (sinφSinθ−cosφcosθsinψ)/
…(1) Here, φ: Azimuth angle of the vehicle θ: Pitch angle of the vehicle φ: Roll angle of the vehicle H _h : Horizontal component of geomagnetism H _v : Vertical component of geomagnetism Δφ: Direction error of magnetic azimuth sensor The tilt error correction calculation unit converts the orientation error Δφ obtained by the above equation (1) into the output signal MAS of the magnetic orientation sensor.
and correct the error.

Note that in equation (1), when the roll angle and pitch angle of the vessel are small, the following equation (2) may be used to simplify the calculation.

Δφ＝H _h θ・cos ² φ−H _v (φsinφ−φcosφ)
/H _h (1+θφcosφsinφ)...(2) Furthermore, when the roll angle and pitch angle of the navigation vehicle are small, the following equation (3) may be used instead of equation (2).

Δφ＝θφcos ² φ−H _v /H _h (θsinφ−φcosφ) …(3) In this formula (3), H _v /H _h = tan (inclination angle), and in this case, the horizontal and vertical components of the geomagnetism Instead of providing the value of the angle of inclination in the inclination error correction calculation section, a value of the inclination angle may be provided.

The outputs of the orientation restraint torque calculation section 57 and the standing torque calculation section 56 are the CTM signal of the CTM calculation section 53.
Together with CS, it is input to the gyro bias calculating section 58, where each bias correction signal of 5
1,52.

In addition, when the speed signal SS with the desired accuracy cannot be obtained, the magnetic direction signal MAS or X, Y, Z
A cutoff signal COS generated from acceleration signals XA, YA, ZA, etc. may be supplied to the acceleration correction calculation unit 55, and the input to the standing torque calculation unit 56 may be cut off when acceleration is applied.

FIG. 3 shows a magnetic orientation sensor 7A according to the present invention.
FIG. In the figure, 7A-1 is a box, and inside thereof there is a magnetic direction detecting section 7-3 similar to that shown in FIGS. 8 and 9. 7A-2 is box body 7
A-1 cover, 7A-3 is the lead wire, box body 7
It supplies electric power to the magnetic direction detecting section 7-3 disposed inside A-1 and the cover 7A-2, and extracts a magnetic direction signal therefrom.

The output signals XG, YG, and ZG of the three X, Y, and Z gyroscopes 1, 2, and 3, which are installed with their input axes aligned with the three main axes of the navigation object, are sent to bias correctors 50 and 5.
Coordinate transformation matrix (CTM) via 1,52
is input to the CTM calculation unit 53 which calculates CTM. Three X, Y,
Outputs of Z accelerometers 4, 5, 6 and the above CTM signal
A horizontal component calculation unit 54 calculates horizontal components α and β of the gravitational acceleration from the CS. If CTM is correct, horizontal components α and β are zero, but CTM
If there is an error in , the horizontal components α and β will have finite values.
The torque signal is converted into a torque signal that makes the CTM a correct value, and sent to the CTM calculation unit 53, which rotates it so that it points in the correct direction.

On the other hand, the azimuth signal AS from the CTM calculation unit 53
and the magnetic direction signal from the magnetic direction sensor 7A.
MAS in the azimuth constraint torque calculation unit 57,
Perform comparison calculations, etc., create a torque signal around the azimuth axis, and use this to eliminate the difference between the two.
Feedback is provided to the CTM calculation unit 53.

On the other hand, if there is a gyro drift, the outputs ET1 and ET2 of the upright torque calculation section 56 and the output AT of the azimuth restraint torque calculation section 57 do not become zero, but have finite values corresponding to the gyro drift. These signals are sent to the gyro bias calculation unit 58 as a CTM signal.
Bias correction signal BC input and calculated with CS
, the bias corrector 5 of each gyro 1, 2, 3
Input 0, 51, 52 and correct it so that the gyro drift is zero.

Roll angle, pitch angle, azimuth angle, and bias corrector 50, 51, 5 of the navigation object from the CTM calculation unit 53
Angular velocity of the navigation object from 2, X rate, Y rate,
Side slip signal from Z rate and Y accelerometer 5
Output each LS.

When the navigation object is making an acceleration motion such as turning, increasing or decelerating, the input to the upright torque calculating section 56 is cut off to eliminate the influence of acceleration. Although not shown, when a speed signal SS or the like is obtained from another sensor, this and the CTM signal CS are supplied to the acceleration correction calculation unit 55 to remove the influence of acceleration.

〔Effect of the invention〕

In the present invention, the conventional rate gyroscope, horizontal gyroscope, and directional gyroscope are replaced with three gyroscopes and three accelerometers, so complex elements such as gimbals, transmitters, torquers, and erecting devices of conventional devices are replaced. is unnecessary, and a simple and low-cost attitude detection device can be obtained.

Furthermore, by using a non-rotating type gyroscope such as a vibrating gyroscope or a gas rate gyroscope, there are no parts that require bearings or slip rings, so the attitude detection device of the present invention has a long life and requires almost no maintenance. It becomes an unnecessary device.

Furthermore, since there are no parts such as bearings, synchro transmitters, torquers, etc., no skill is required to assemble the device.

In addition, by providing a gyro bias calculating section and correcting the gyro bias within the system, good system performance can be ensured even if an inexpensive vibrating gyro or gallate gyro is used.

In addition, a magnetic orientation sensor without a universal joint is used, and a tilt error correction calculation unit is provided to correct orientation errors caused by tilt, making the device compact, lightweight, and highly accurate.
It also reduces the price.

[Brief explanation of the drawing]

FIG. 1 is a system configuration diagram of the attitude detection device of the present invention, FIG. 2 is a block diagram of the calculation unit 9 in FIG. 1, FIG. 3 is a sectional view of the magnetic orientation sensor of the present invention, and FIG. 4 is a conventional A perspective view of a direction gyro, FIG. 5 is a perspective view of a conventional horizon, FIG. 6 is a front view of the display section of a conventional turning indicator, and FIG. 7 is a perspective view of a rate gyro that constitutes the turning indicator. , 8th
The figure is a partial cross-sectional view of a conventional magnetic direction sensor.
The figure is a perspective view of the magnetic direction detection section. In the figure, 1, 2, 3 are gyro, 4, 5, 6
is an accelerometer, 7A is a magnetic direction sensor, 8 is a signal conversion section, 9 is a calculation section, 10 is a signal output section, 50,
51 and 52 are bias correctors, 53 is a CTM calculation unit, 54 is a horizontal component calculation unit, 56 is an upright torque calculation unit, 57 is an azimuth restraint torque calculation unit, 58 is a gyro bias calculation unit, and 60 is a tilt error correction calculation unit. The parts are shown respectively.

Claims (1)

[Scope of Claims] 1. Three gyroscopes and three accelerometers, and one magnetic orientation sensor, which are attached to the navigation vehicle so that their input axes coincide with the three main axes of the navigation vehicle, respectively; Posture including a calculation section that receives the outputs of the three gyroscopes, three accelerometers, and one magnetic orientation sensor as input, and a signal output section that receives the output of the calculation section and outputs a posture signal to the outside. In the detection device, the magnetic azimuth detection section of the magnetic azimuth sensor is directly attached to the navigation object, and the azimuth error caused by the inclination of the magnetic azimuth sensor is corrected by a coordinate transformation matrix calculation section that calculates a coordinate transformation matrix in the calculation section. A tilt error correction calculation unit is provided to calculate the azimuth, roll angle, and pitch angle of the vehicle obtained from the coordinate transformation matrix calculation unit and the horizontal and vertical components of the geomagnetic field provided in the tilt error correction calculation unit. An attitude detection device configured to calculate and correct an azimuth error caused by the inclination of the magnetic azimuth sensor based on the value.