JPH0455248B2 - - 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 directional gyro that indicates the direction around the vertical axis. It is equipped with a turn indicator that indicates the turning angular velocity and bank angle of the aircraft, supplementing the pilot's senses and enabling safe flight under any conditions.

Next, the above conventional device will be explained with reference to FIGS. 3 to 6.

FIG. 3 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 11.
Here, a deviation signal between the magnetic azimuth output and the azimuth of the spin axes 100, 100a, that is, the gyro azimuth, is generated, and this deviation signal is sent to the amplifier 114A.
Feedback is provided to the slave torquer 115 provided on the horizontal shaft 103a through the servo axis 103a, and the gyro orientation is made to match the magnetic orientation. This loop is called an orientation constraint loop. When the aircraft is moving violently, the azimuth direction is output, and errors in the azimuth angle due to gyro drift are restrained by the magnetic direction from the flux valve 114 to maintain accuracy. The aircraft direction is
Compass card 109 attached to vertical axis 106a
Read by.

FIG. 4 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 134 is not visible) fixedly installed at corresponding positions on the external gimbal 135. external gimbal 135
has roll shafts 136, 136a at positions orthogonal to the pitch shafts 134, 134a, and these roll shafts 136, 136a are attached to roll bases 138, 138a mounted in the tail direction of the aircraft. It is rotatably fitted into the 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. 5 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. 6 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 axis ZZ' which is perpendicular to both XX', a torque proportional to the turning angular velocity Ω is generated due to the gyro effect. It occurs around the output shaft axis YY'. 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. 5) attached to the output shaft 173a. Output.

[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.

Therefore, the present invention provides a novel attitude detection device that solves the problems of conventional devices.

[Means for solving problems]

The attitude detection device according to the present invention has three main
three gyroscopes 1, 2, 3 and three acceleration deflectors 4, 5, 6, one magnetic sensor 7, which are attached to the navigation vehicle so that their input axes coincide with the respective axes;
a signal conversion unit 8 that receives the outputs of the three gyroscopes, three accelerometers, and one magnetic sensor;
It consists of a calculation section 9 which receives the output of this signal conversion section as input, and a signal output section 10 which receives the output of this calculation section and outputs an attitude signal to the outside.

[Effect]

The outputs of three gyroscopes 1, 2, 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 horizontal component calculation unit 54 uses the outputs of the three accelerations 4, 5, 6 attached to the aircraft body so that the input axes of the gyros 1, 2, 3 and their acquisition axes are in equilibrium, and the CTM signal. ,
Calculate horizontal components α and β of gravitational acceleration. CTM
If is correct, horizontal components α and β are zero, but
If there is an error in the CTM, the horizontal components α and β will be finite values, so the upright torque calculation unit 56
It is converted into a torque king signal that makes the CTM a correct value and sent to the CTM calculation unit 53.

On the other hand, the azimuth angle signal in the CTM calculation unit 53 and the magnetic azimuth signal from the magnetic azimuth sensor 7 are compared and calculated in the azimuth constraint torque calculation unit 57 to create a torque signal around the azimuth axis. Feedback is provided to the CTM calculation unit 53 so that there is no difference between the two.

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 7 are connected to three main orthogonal axes of a navigation object (not shown), namely, X, Y,
Install each input shaft so that it matches the Z-axis (see arrow). Output signals from these are input to the calculation section 9 via the signal conversion section 8. Arithmetic unit 9
After calculating the coordinate transformation matrix CTM, gyro drift correction calculation, acceleration correction calculation, etc., the roll angle, pitch angle, azimuth angle, X, Y, Z angular velocity, bank necessary for control and operation of the navigation object are Signals such as angle, slip angle, etc. are 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, Y from the signal converter 8 shown in FIG.
and acceleration signals from Z accelerometers 4, 5 and 6
XA, YA and ZA are the signals from the CTM calculation unit 53.
Together with the CTM signal CS, it is input to the horizontal component calculation unit 54, where horizontal components α and β of acceleration in both east-west and north-south directions are calculated. 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 56.
After performing the standing torque calculation, the CTM is inputted to the CTM calculation unit 53, and the CTM is torqued so that the horizontal components α and β become zero.

The azimuth signal AS from the CTM calculation unit 53 and the magnetic azimuth signal MAS from the signal conversion unit 8 are supplied to the azimuth constraint torque calculation unit 57, where
Comparison calculation A restraint torque calculation is performed, and the output torque signal TS is fed back to the CTM calculation section 53, and the CTM is mainly torqued around the azimuth axis, whereby the CTM orientation is restrained by the magnetic orientation.

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 cut-off signal COS generated from acceleration signals XA, YA, ZA, etc. is supplied to the acceleration correction calculation section 55,
The input to the upright torque calculating section 58 may be cut off when acceleration is applied.

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.
1 and 52 to a CTM calculation unit 53 which calculates the coordinate transformation matrix CTM, and causes the CTM to be calculated. A horizontal component calculation unit 54 uses the outputs of three X, Y, Z accelerometers 4, 5, 6 installed on the aircraft body so that the input axis of the gyroscope and the acquisition axis thereof are parallel, and the CTM signal CS. , calculate the horizontal components α and β of the gravitational acceleration. If 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 torque signal so that the CTM becomes the correct value. Convert and
It is sent to the CTM calculation unit 53 and rotated so that it faces 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 7.
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 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 signals are input to the gyro bias calculation section 58 together with the CTM signal CS, and the calculated bias correction signal BC is applied to each gyro 1, 2,
3 bias correctors 50, 51, 52,
Correct so that gyroscope 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 maintenance such as bearings or slip rings, so the attitude detection device of the present invention has a long life and The device requires almost no maintenance.

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.

[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 perspective view of a conventional directional dial, and FIG. FIG. 5 is a front view of a display section of a conventional turning indicator, and FIG. 6 is a perspective view of a rate gyro constituting the turning indicator. In the figure, 1, 2, 3 are gyro, 4, 5, 6
is an accelerometer, 7 is a magnetic direction sensor, 8 is a signal conversion section, 9 is a calculation section, 10 is a signal output section, 50,5
1 and 52 are bias correctors, 53 is a CTM calculation unit,
Reference numeral 54 denotes a horizontal component calculation section, 56 a standing torque calculation section, 57 a direction restraint torque calculation section, and 58 a gyro bias calculation section.

Claims (1)

[Claims] 1. Three gyroscopes and three accelerometers are attached to the vehicle so that their input axes coincide with the three main axes of the vehicle, and one magnetic orientation sensor is installed. Attach it to the above navigation vehicle, and
a signal converter that receives the outputs of three gyroscopes, three accelerometers, and one magnetic orientation sensor; a calculation section that receives the output of the signal conversion section; and an external In the attitude detection device, the arithmetic unit includes a CTM (coordinate transformation matrix) arithmetic unit that inputs the gyroscope signals from the three gyroscopes, and a CTM of the CTM arithmetic unit. a horizontal component calculation unit that calculates the horizontal component of gravity from the signal and the acceleration signals from the three accelerometers; and a standing torque calculation unit that receives the output of the horizontal component calculation unit and calculates the standing torque signal of the CTM. and an azimuth constraint torque that compares the azimuth signal from the CTM calculation unit with the magnetic azimuth signal from the magnetic azimuth sensor and calculates a azimuth constraint torque signal for constraining the CTM to the magnetic azimuth signal around the azimuth axis. What is claimed is: 1. A posture detection device comprising: a calculation unit, and configured to feed back the standing torque signal and the azimuth restraint torque signal to the CTM calculation unit.