JP4191588B2 - Satellite tracking antenna controller - Google Patents

Description

  The present invention relates to an antenna control device that is mounted on an aircraft and tracks a satellite.

An antenna mounted on a mobile body such as an aircraft and communicating with a communication satellite uses an antenna that always tracks the communication satellite even if there is a change in the position or nose direction of the aircraft. In order to increase the sensitivity of the communication equipment, it is necessary to narrow the antenna beam more sharply (narrower). In this case, it is natural that the tracking direction accuracy must be increased. In other words, improving the tracking accuracy is an essential technique for improving the sensitivity of communication equipment.
Regarding the improvement in tracking accuracy of such a satellite tracking antenna control device, in a conventional satellite tracking control device, for example, as shown in Patent Document 1 (FIG. 1 in the publication), a mobile fixed coordinate system and an antenna There is one which can perform satellite tracking control of a highly accurate antenna by calculating and changing the amount of axis deviation with respect to the gimbal coordinate system.

Further, for example, as shown in Patent Document 2 (FIG. 1 in the publication), attitude azimuth estimating means for estimating the absolute attitude and absolute azimuth of the moving body, and satellite absolute direction calculating means for calculating the azimuth angle and elevation angle of the satellite. Satellite relative direction calculating means for calculating the azimuth and elevation angle of the satellite relative to the moving body, satellite relative direction searching means for searching for the satellite relative direction using the reception level of the signal transmitted from the satellite, and satellite relative direction calculation. The antenna or antenna beam is driven in a direction in which the satellite relative direction obtained from the satellite relative direction search means is superimposed on the satellite relative direction obtained from the means to direct the satellite, and the antenna or antenna beam for the moving object at that time By providing an antenna driving device that inputs the azimuth angle and the elevation angle to the posture azimuth estimation means 1, the antenna can be controlled with high accuracy over a long period of time. It has to be able to track.
Further, for example, as shown in Patent Document 3 (FIG. 1 in the publication), a triaxial rate gyro is placed on the base of an antenna that is a device to be controlled, so that the antenna is matched with the reference speed signal even if the hull is shaken. There is something that can stabilize the space.
JP 2003-37424 A (FIG. 1 in the publication) JP 2002-158525 A (FIG. 1 in the publication) JP 50-61589 A (FIG. 1 in the publication)

When such a satellite tracking control device is used and an antenna for tracking a satellite is mounted on a large mobile body (for example, an aircraft having a length of 50 m), there are the following problems.
(1) IRU (Inertia Reference Unit) that is included in an inertial navigation system (INS) mounted on an aircraft and that measures position (latitude, longitude, altitude) and posture (roll, pitch, heading angle) with high accuracy The location is typically mounted on the front of the aircraft (the location where the pilot is maneuvering), but the antenna that tracks the satellite is far away from this location for a variety of reasons (eg, about 20 meters). It is often mounted on the upper part of the fuselage. For this reason, there is a relative angle error between the antenna position and the IRU, and there is a problem in that the airframe is subject to relative displacement of several Hz such as vibration and bending, and the tracking accuracy of the antenna is lowered. .
(2) The position (latitude, longitude, altitude) and attitude (roll, pitch, heading angle) signals output from the IRU (inertia reference unit) mounted on the aircraft are random that cannot be ignored for antenna angle control. Time change and random update rate change (change of data update time interval) may be large, and if the attitude of the antenna is controlled with this time delay included, the airplane will make a turning motion continuously. However, there is a problem that the tracking angle error becomes large.

(3) If an antenna is installed without modifying the aircraft body as much as possible, it is necessary to reduce the height of the antenna as much as possible in order to reduce the effect on the aerodynamic characteristics of the aircraft (not to increase the aerodynamic resistance). is there. It is desirable that the degree of freedom of the antenna has three rotational degrees of freedom like the movement of the aircraft. However, if three degrees of freedom are provided, it is difficult to reduce the height of the antenna in terms of structure. The antenna height can be reduced if the antenna freedom is composed of two degrees of azimuth and elevation, but the singularity (keyhole where the azimuth is not fixed when the elevation is increased due to the geometrical relationship) ) Occurred, and satellite tracking of the antenna in this range became impossible.
(4) When a disturbance such as airframe vibration occurs in the vicinity of the singular point (keyhole) described above (for example, an elevation angle of 80 degrees or more), the azimuth angle of the antenna that tracks the satellite greatly fluctuates, and the tracking accuracy decreases significantly. There was a problem.
(5) Three axis signals of roll axis (X), pitch axis (Y), and yaw axis (true azimuth axis, Z), which are attitude angles output from IRU, output a value of ± 180 degrees. Therefore, since the output value becomes a discontinuous value when passing from +180 degrees to -180 degrees (or vice versa), when this signal is filtered in the control device, normal control cannot be performed unless this is taken into consideration. For example, there was a problem that tracking was lost.

A satellite tracking antenna control device according to the present invention includes an inertial reference unit that is mounted on an aircraft and measures its position and orientation, an antenna frame mounted on the aircraft that is remote from the inertial reference unit, and an antenna frame on the antenna frame. An antenna that is installed in a direction and is driven to communicate with a communication satellite in an arbitrary direction, an angular velocity sensor that is installed in the vicinity of the antenna mount and measures at least a two-dimensional angular velocity of the antenna mount, and a body posture measured by the inertial reference unit A posture estimation filter for estimating a current posture of the antenna mount using a signal and an angular velocity signal of the angular velocity sensor;
This posture estimation filter internally includes a coordinate converter and an integrator for calculating the angular velocity sensor output to obtain an estimated posture value of the antenna mount, and an airframe posture that is output from the inertial sensor unit and taken into the posture estimation filter. An IRU time delay corrector for delaying an input signal by a time corresponding to a signal time delay, and an attitude estimated value of the antenna mount is input to the IRU delay time corrector, and its output is output from the inertia reference unit. Including a signal filter for estimating the drift amount compared with the airframe attitude signal, and using the obtained drift estimation value to correct the output of the angular velocity sensor or its coordinate conversion value to obtain the attitude estimation value of the antenna mount, The navigation is controlled by controlling the antenna based on the estimated attitude of the antenna mount output by the attitude estimation filter. It is corrected for both the measurement delay of the inertial reference unit and distortion of the machine of the aircraft.

It is possible to correct a certain time delay of the signal of the inertial reference unit (IRU), and to improve tracking accuracy as viewed in time average. In addition, since high-frequency disturbances such as airframe vibration can be measured up to a high frequency with an angular velocity sensor installed in the vicinity of the antenna base, a quick response can be achieved and highly accurate antenna tracking control can be realized.
In addition, since a keyhole filter for accurately tracking the satellite is provided even in a high elevation angle antenna orientation, it is possible to obtain high control accuracy that has not been achieved in the past.

Embodiment 1 FIG.
FIG. 1 is a diagram for explaining the satellite tracking antenna control apparatus according to the first embodiment of the present invention, particularly for explaining the arrangement of sensors and antennas, and for illustrating the definitions of the angles and the like of each part to be described later. . An inertial reference unit (IRU) 2 that measures the position (latitude, longitude, altitude) and attitude angle (roll, pitch, heading angle) of the aircraft 1 with high accuracy is provided at the front of the aircraft 1 (sometimes referred to as the aircraft). is set up. The IRU 2 is the ground surface 8 shown in the figure (the local horizontal plane coordinate system 8 serving as a reference for the angle of the aircraft is a theoretical ground surface obtained by regarding the earth as an ellipsoid, not an actual mountain or sea surface. For example, the earth's ellipsoid defined by the geodetic reference system 1980 (GRS80 ellipsoid) is used as a reference.
The aircraft 1 obtains the position and attitude angle when flying over the earth based on these IRU signals, and realizes flight on the target route. An antenna base 3 of an antenna 5 that transmits and receives radio waves to and from a communication satellite is structurally connected and fixed to the aircraft body 1 above the aircraft body in the center of the aircraft 1.
An antenna support mechanism (including a gimbal mechanism) 4 that supports the antenna 5 is provided on the top of the antenna base 3 so that the antenna 5 can rotate with respect to the antenna base 3 with two degrees of freedom or three degrees of freedom. I support it. An angular velocity sensor 6 for measuring the rotational angular velocity of the antenna base is fixed to the antenna base 3. This angular velocity sensor 6 can measure a rotational angular velocity with three degrees of freedom. The angular velocity sensor 6 is not limited to a sensor that can measure only the angular velocity, and may be a simple inertia reference device that combines, for example, an angular velocity sensor and an acceleration sensor, an orientation sensor (magnetic compass), or a position sensor such as GPS.

A relative angle sensor 7 for measuring the relative angle between the antenna base 3 and the antenna 5 is attached to the antenna support mechanism 4. The relative angle sensor 7 can use an encoder, a resolver, and the like. It is possible to measure relative angular velocities by differentiating them manually or by software. Although not shown here, two relative angle sensors 7 are installed when the degree of freedom of the antenna drive mechanism is two degrees of freedom, and three relative angle sensors 7 are installed when the degree of freedom is three degrees of freedom.
In general, large aerodynamic disturbances such as air pockets occur in large aircraft, which causes large bending vibrations or torsional vibrations in the airframe. The frequency causes bending torsional vibration of several Hz, and the relative angle between the antenna base 3 and the inertial reference unit (IRU) 2 also varies about 1 degree. In contrast, the output signal of the inertial reference unit (IRU) 2 that measures the position and attitude angle of the aircraft 1 with high accuracy has high accuracy, but the sampling rate and frequency response characteristics of the output signal are required by the antenna 5. It is not sufficiently short or high compared to the angular accuracy. Furthermore, in the case of a large aircraft, the antenna 5 and the inertial reference unit (IRU) 2 may be installed at positions far away from each other (sometimes 50 m or more). It is impossible to accurately measure the position and orientation information at the position.

  The accuracy of each sensor around the antenna shown in FIG. 1 satisfies the required accuracy of the antenna. The response speed is faster than the bending torsional frequency of the aircraft. By providing the sensor 6 capable of measuring the angular velocity or angle provided on the antenna base 3 at a high sampling rate and up to a high frequency response, the inertial reference unit (IRU) 2 can be used even if disturbance such as airframe vibration occurs in the airframe 1. By combining the above signal and the signal from the sensor 6 as described later, the angle of the location of the antenna base 3 is obtained with high accuracy. Furthermore, since the relative angle between the antenna base 3 and the antenna 5 can be measured with high accuracy and high frequency response by the relative angle sensor 7 provided in the antenna support mechanism, the antenna 5 can always be tracked in the direction of the communication satellite with high accuracy. it can.

The definition of each angle (here, only the elevation angle here) in FIG. 1 will be described.
θ: Directional angle (elevation angle) of the antenna 5 with respect to the theoretical ground surface 8.
Δθ: pointing error from the target of the antenna elevation angle θ.
θ _B : The angle of the antenna base 3 with respect to the ground surface 8.
θ _G : Gimbal angle with respect to the antenna base 3.
θ _T : Target pointing angle (satellite elevation angle).
θ _A : The angle of the aircraft 1 with respect to the ground surface 8.

Next, means for processing the signal of the sensor described above will be described.
FIG. 2 is a block diagram showing the entire control system of the satellite tracking antenna control apparatus.
Attitude reference signals (~ Φ, ~ Θ, ~ Ψ) 10 indicating the roll, pitch, and heading angle of the fuselage 1 are output from an inertial reference unit (IRU) 2 installed at the front of the aircraft.
Each signal includes a random time delay as described above. These posture angle signals (˜Φ, ˜Θ, ˜Ψ) 10 are input to the posture estimation filter 12. An output voltage of an angular velocity sensor (also referred to as a rate gyro) 6 installed on the antenna base 3 at a position away from the IRU of the airframe 1 is converted from a voltage to an angular velocity, and also converted by an attachment angle corrector 9 to be converted into an angular velocity of the rate gyro. Signals (ω _BX , ω _BY , ω _BZ ) 11 are output. Here, ω _BX , ω _BY , and ω _BZ are rate gyro angular velocities around the X, Y, and Z axes of the airframe 1, respectively.
For example, if V _G is the gyro output (V), V _off is the offset power supply voltage (V), ω _G is the angular velocity (deg / s), and SF is the sensitivity (V / (deg / s)), the rate Since the gyro has three axes of the X (roll) axis, the Y (pitch) axis, and the Z (yaw) axis, the conversion from the gyro output voltage V _G to the angular velocity ω _G is expressed by the following equation.

The conversion from the coordinate system of the surface to which the rate gyro 6 is attached (the coordinate system of the antenna base 3) to the coordinate system of the airframe 1 (when the antenna base is installed at an angle) can be calculated by the following equation, for example. That is, it is obtained by rotating the body coordinate system by Φ _G around the X axis, then rotating by Θ _G around the Y axis, and further rotating by Ψ _G around the Z axis.

A combination of the above equations (1) and (2) is an example of the conversion from the voltage to the angular velocity and the mounting angle corrector 9 shown in FIG.
FIG. 3 is a block diagram showing the internal configuration of the posture estimation filter 12 of FIG. The IRU2 angle signals (~ Φ, ~ Θ, ~ Ψ) 10 indicate a roll angle, a pitch angle, and a heading angle, respectively, but each signal includes a random time delay. These IRU2 angle signals (~ Φ, ~ Θ, ~ Ψ) 10 calculate the difference from the outputs of the IRU delay time correctors 27, 28, and 29, respectively, and then the roll axis signal filter 20 and the pitch axis signal filter. 21, input to the yaw axis signal filter 22. In the block diagrams of the filters 20, 21, and 22, ω _Fp is a proportional gain (rad / s), ω _Fi is an integral gain, and s is a Laplace operator. Also, e ^−tXs , e ^−tYs , and e ^−tZs shown in the block diagrams of the time delay correctors 27, 28, and 29 are ^signals of the estimated attitude angles (^ Φ, ^ Θ, ^ Ψ) 13 output to the antenna unit. Is delayed by times t _X , t _Y , and t _Z , respectively.

The estimated drift values ^ dx, ^ dy, and ^ dz, which are the outputs of the filters 20, 21, and 22, convert the three-axis output signals of the angular velocity sensor 6 provided on the antenna base 3 from voltage to angular velocity, and are further attached. Differences from the angular velocity signals (ω _Bx , ω _By , ω _Bz ) 11 of the rate gyro converted by the angle corrector 9 are calculated, and angular velocity estimated values ( ^{^} ω _Bx , ^{^} ω _By , ^{^} ω _Bz ) 16 are calculated. Further, these estimated angular velocity values 16 are input to a coordinate converter 23 that converts the airframe angular velocity into Euler angular time derivative. Further, the coordinate converter 23 receives the estimated roll angle value ^{^} Φ and the estimated pitch angle value ^{^} E. From the coordinate converter 23, Euler angle time differential estimation values (d ^{^} Φ / dt, d ^{^} Θ / dt, d ^{^} Ψ / dt) are output and integrated by integrators 24, 25, and 26. , The estimated attitude angle (^ Φ, ^ Θ, ^ Ψ) 13 of the antenna unit is output, and the estimated airframe angular velocity ( ^{^} ω _BX , ^{^} ω _BY , ^{^} ω _BZ ) 16 of the antenna unit is also obtained as an output. It is done.

FIG. 4 shows an angular velocity signal (ω _Bx , ω _By , ω _Bz ) 11 and IRU 2 obtained by converting the angular velocity measured by the rate gyro 6 attached to the airframe 1 from voltage to angular velocity and further converted by the attached angle corrector 9. The relationship between the Euler angles (Φ, Θ, Ψ) and the Euler angle time derivatives (dΦ / dt, dΘ / dt, dΨ / dt) is shown. Euler angle time differentiation is also called posture rate. This relationship can be described by the following equation. As shown in FIG. 3, ^{^} Φ and ^{^} Θ of the estimated value 13 are used as Φ and Θ in the following equations.

Equation (3) becomes a singular point at Θ = 90 deg, and no solution can be obtained. Therefore, in the vicinity of Θ = 90 deg, it is discriminated by a mathematical expression (not shown) conditional expression (error processing or only a range close to 90 deg). Special processing).
Here, each variable is defined with a more correct name as follows.
Φ: Roll angle when converting from LLN (Local-Level-north) coordinate system to aircraft coordinate system (3-2-1) Θ: When converting from LLN coordinate system to aircraft coordinate system (3-2-1) Pitch angle Ψ: Heading angle (true orientation) when converting from LLN coordinate system to airframe coordinate system (3-2-1)
The above Φ Θ Ψ is the IRU output.
Also, Euler angle time derivative (posture rate) is
d ^{^} Φ / dt: time derivative of the roll angle,
d ^{^} Θ / dt: time derivative of the above pitch angle,
d ^{^} Ψ / dt: time derivative of the heading angle (true orientation),
Also, the angular velocity in the aircraft coordinate system is
ω _Bx : Roll angular velocity in the aircraft coordinate system,
ω _By : pitch angular velocity in the aircraft coordinate system,
ω _Bz : Yaw angular velocity in the aircraft coordinate system,
ω _Bx, ω _{By and} ω _Bz are rate gyro outputs.
The conversion from Euler angle time differentiation (= posture rate) to airframe angular velocity is expressed by the following equation.

FIG. 5 shows the frequencies of the attitude estimation filter 12 in FIG. 3 when the frequency of the IRU attitude angle sensor 10 and the rate gyro signal 11 of the antenna unit are plotted on the horizontal axis and the gain is plotted on the vertical axis. It is the figure which showed the relationship of the zone | band. The attitude angle signal 10 of the IRU 2 is assigned from 0 Hz to the frequency band of the proportional gain ω _Fp (rad / s) 42 set in the roll, pitch and yaw axis signal filters 20, 21, and 22 (the illustrated line 40) From the frequency to ωg43 (assumed to be 10 Hz in the figure), the rate gyro signal 11 is in charge (illustrated line 41).
By adopting such a configuration, even if an airframe vibration occurs and an angular fluctuation relatively occurs between the antenna base 3 and the IRU 2, the antenna 5 can be controlled so as to cancel the angular fluctuation. The antenna 5 can be tracked in the direction of the satellite with high accuracy.
Further, an average value of the delay time of the IRU 2 is obtained in advance (a method described later may be used, or a value known as the IRU design specification may be used), and this is set in the delay time correctors 27, 28, and 29. As a result, the influence of the random delay time can be reduced, and the time-average tracking accuracy can be maintained with high accuracy.

The posture estimation filter 12 ′ shown in FIG. 6 is a configuration example when coordinate transformation 23 ′ is added to make the posture estimation filter 12 of FIG. 3 more accurate. In the posture estimation filter 12 in FIG. 3, a part of the coordinate conversion is omitted in order to speed up the calculation.
For example, the difference value between the IRU angle signal (˜Φ) 10 and the output of the IRU delay time corrector 27 is input to the roll axis signal filter 20, and an estimated drift value ^ dx is obtained as an output. A difference between dx and the angular velocity signal (ω _Bx ) 11 of the rate gyro is input to the coordinate converter 23. At this time, since the coordinate system of the drift estimated value ^ dx and the angular velocity signal (ω _Bx ) 11 of the rate gyro does not completely coincide with each other, for example, if the aircraft keeps being tilted by about 80 deg, an error will occur. A phenomenon occurs in which the estimated attitude angle ( ^{^} Φ, ^{^} Θ, ^{^} Ψ) 13 of the antenna portion diverges. However, since the large inclination angle is not kept in any of the three axes of roll, pitch, and yaw in normal flight, the posture estimation filter 12 partially omits the coordinate conversion shown in FIG. The output value does not diverge.

On the other hand, the posture estimation filter 12 ′ having the configuration shown in FIG. 6 is obtained by adding some coordinate transformations that are omitted in the posture estimation filter 12 of FIG. The angular velocity signals (ω _Bx , ω _By , ω _Bz ) 11 of the rate gyro are converted from the airframe angular velocity to the Euler angular time derivative by the coordinate converter 23. These (dΦ / dt, dΘ / dt, dΨ / dt) are used for the difference calculation with the drift estimation values ^ dx, ^ dy, ^ dz. Then, Euler angle time differential estimated values (d ^{^} Φ / dt, d ^{^} Θ / dt, d ^{^} Ψ /) estimated by the feedback loops of the integrators 24, 25, 26 and the IRU delay time correctors 27, 28, 29 are used. taking out a signal dt), the estimated angular velocity value after conversion to the aircraft angular velocity from the input Euler angles time derivative to a coordinate converter ^{_{23 '(^ ω Bx, ^}} ω by, the output of the ^{^} omega _Bz) 16 is obtained. Similarly, the value obtained by integrating the Euler angle time derivative estimated values (d ^{^} Φ / dt, d ^{^} Θ / dt, d ^{^} Ψ / dt) is the estimated attitude angle ( ^{^} Φ, ^{^} Θ, ^{^} Ψ) of the antenna unit 13 Is output as
For example, the posture estimation filter 12 ′ configured as described above does not diverge even if the roll axis continues to hold 89 deg. As a result, the tracking control of the antenna can be performed more stably and with high accuracy. If the installed computer has sufficient computing power, it is effective to use the posture estimation filter 12 ′ having the configuration shown in FIG. 6 instead of the configuration shown in FIG.
3 and the posture estimation filter 12 ′ of FIG. 6 perform some flipping or the like, and when the roll angle Θ reaches 90 degrees (when it is very close to 90 degrees in an actual apparatus), the calculation of the coordinate converter 23 is indefinite. Therefore, in this case, another process is required, which will be described later.

Returning again to the overall block diagram of FIG. An estimated attitude angle (^ Φ, ^ Θ, ^ Ψ) 13 that is an output of the attitude estimation filter 12 is input to the coordinate converter 15. Further, the coordinate converter 15 receives signals such as the latitude, longitude and altitude of the aircraft which are the output of the IRU 2 and signals such as the latitude and longitude of the satellite from an aircraft command device (not shown). The The coordinate converter 15 calculates the elevation angle command value ^ θ _EL and the azimuth command value ^ θ _AZ of the antenna based on the above input. An example of the calculation method is shown below according to the procedure.
(Procedure 1)
Calculation of the radius of curvature R _A of the bow line at the latitude of the aircraft

Where R _{e is the} equator radius, e ² is the eccentricity of the ellipsoid, and ψ _A is the latitude of the aircraft.
(Procedure 2)
Calculation of radius of curvature R _S of the bow line at the satellite's latitude

Where ψ _S is the latitude of the satellite.
(Procedure 3)
Calculation of aircraft position in the fixed earth coordinate system,
A relational expression between Cartesian coordinates and geodetic coordinates in the earth fixed coordinate system.

here,
( ^E x _A , ^E y _A , ^E z _A ): Coordinate position of the center of mass of the aircraft in the fixed earth coordinate system
(ψ _A , λ _A , h _A ): Geodetic coordinate position of the center of mass of the aircraft in the fixed earth coordinate system ψ _A : Aircraft latitude, λ _A : Aircraft longitude, h _A : Aircraft altitude (step 4)
Calculation of satellite position in the fixed earth coordinate system,
Relational expression between the earth fixed coordinate system and the geodetic coordinate system.

here,
( ^E x _S , ^E y _S , ^E z _S ): Coordinate position of the center of mass of the satellite in the fixed earth coordinate system,
(ψ _S , λ _S , h _S ): Geodetic coordinate position of the center of mass of the satellite in the earth fixed coordinate system,
ψ _{S is} the latitude of the satellite, λ _{S is} the longitude of the satellite, and h _S is the altitude of the satellite.
(Procedure 5)
Calculation of the relative position of the aircraft and satellite in the Earth fixed coordinate system (satellite pointing direction as seen from the aircraft)

(Procedure 6)
Calculation of the relative position of the aircraft and satellite in the LLN coordinate system (satellite pointing direction as seen from the aircraft)
Conversion from the earth fixed coordinate system to the LLN coordinate system is obtained by rotating the earth fixed coordinate system by λ _A around the Z axis and then rotating by −π / 2−ψ _A around the Y axis.

(Procedure 7)
Calculation of the relative position of the aircraft and satellite in the aircraft coordinate system (satellite pointing direction as seen from the aircraft)
Rotate LLN coordinate system around Z axis, rotate Θ around Y axis, and then rotate around X axis Obtained by rotating.
Incidentally, X _B-axis in body coordinate system: roll angle [Phi, Y _B axis around: pitch angle theta, Z _B axis: a heading angle [psi.

(Procedure 8)
Calculation of the relative position of the aircraft and satellite in the antenna coordinate system (satellite pointing direction as seen from the aircraft)
Further, the mounting surface correction is performed, and further, the mounting surface correction of the mounting correction angles ΔΨ, ΔΘ, ΔΦ between the antenna base 3 and the inertial reference unit (IRU) 2 is performed.

(Procedure 9)
Conversion of aircraft roll, pitch, and heading rotation to AZ and EL angles in the antenna coordinate system (Satellite pointing angle as seen from AZ-EL antenna mounted on aircraft)
Note: AZ axis defines a q _AZ angle reverse rotation as forward direction ^B Z-axis heading the ^(A X-axis) as 0 deg. The angle range is -180 to +180 deg. The EL axis angle is 0 deg on the ^A X- ^A Y plane, and the positive direction is toward the ^A Z axis. The angle range is 0 to 90 deg.

^A θ _AZ and ^A θ _EL shown in the above equations (13) and (14) correspond to the azimuth command value (^ θ _AZ ) 120 and the elevation command value (^ θ _EL ) 122 in FIG.
Based on the input of the azimuth angle command value (^ θ _AZ ) 120 and the elevation angle command value (^ θ _EL ) 122 obtained by the above calculation, the coordinate converter 17 estimates the angular velocity ( ^{^} ω _Bx , ^{^} ω _By , ^{^} ω _Bz ) 16 is calculated as an antenna elevation angular velocity command value (^ ω _EL ) 127 and an azimuth angular velocity command value (^ θ _AZ ) 124. The calculation formula is as follows.

Elevation angle obtained by the coordinate converter 15 as described above, the azimuth angle command value ^(A θ _AZ) 120 and ^(A θ _EL) 122 and determined by the coordinate converter 17 angular velocity _{(^ ω AZ,) 124,} (^ ω EL) 127 can be used as the space stable input of the elevation gimbal servo controller 18 and the azimuth gimbal servo controller 19. It should be noted that in the equation (15), when ^ θ _EL = 90 deg, a solution for tan ^ θ _EL cannot be obtained, so another processing is required.
In this way, by configuring the posture estimation filter 12, it is possible to correct a certain time delay of the signal of the inertial reference unit (IRU) 2, and it is possible to improve tracking accuracy as viewed in time average. In addition, high frequency disturbances such as airframe vibrations can be measured up to a high frequency with the angular velocity sensor 6 installed in the vicinity of the antenna base 3, so that a quick response can be achieved and highly accurate antenna tracking control can be realized.
Although the angular velocity sensor 6 is provided on the antenna base 3 in FIG. 1, the antenna tracking control can be similarly performed with high accuracy by providing each velocity sensor 6 on the antenna 5.

Embodiment 2. FIG.
FIG. 7 is a conceptual diagram showing the intensity distribution of radio waves received by the satellite tracking antenna according to the second embodiment of the present invention. The degree of freedom of the antenna is two degrees of freedom, but here, the distribution of the radio wave intensity is shown when the uniaxial antenna is slightly swung by Δθ in the satellite pointing direction. The distribution of received intensity at the present time is indicated by a quadratic curve with the angle X2 as the apex. When the + Δθ angle is changed, the quadratic curve with the X3 as the apex, and when the −Δθ angle is changed, the X2 is the apex 2 The next curve is shown. The reception intensity of the radio wave at each antenna angle is Y2, Y3, Y1. By measuring the angles X1, X2, and X3 and the radio wave strengths Y1, Y2, and Y3, the true satellite direction in which the radio wave strength is strongest can be calculated based on these angles. As an example, the following calculation is performed to determine the angle error Err from the current X2 angle to the true satellite direction.

A small displacement ± Δθ is given to the antenna directing direction, and based on the intensity change of the received radio wave from the communication satellite at this time, an error in the antenna directing direction is obtained as described above. It is called an antenna directivity direction control means based on strength.
In this way, by measuring the satellite pointing direction with higher accuracy as required using the intensity distribution of the received radio wave, the tracking control of the antenna can be performed with higher accuracy based on this direction. . While the antenna is shaken by a small angle Δθ, the radio wave intensity decreases and the amount of communication decreases.Therefore, the antenna intensity distribution measurement is performed at a certain time interval or when the satellite direction is lost due to disturbance, etc. During other times, the tracking of the antenna can be performed with high accuracy over a long period of time by tracking the satellite with the output of the attitude estimation filter described above.

Embodiment 3 FIG.
FIG. 8 is an overhead view when the tilt angle sensor 9 is provided in the base portion of the satellite tracking antenna according to the third embodiment of the present invention. An angular velocity sensor 6 is provided on the antenna base 3, and the relative three-axis angular relationship between the angular velocity sensor 6 and the inertial reference unit 2 installed in the front part of the aircraft 1 greatly affects the tracking accuracy. The relationship is shown in the above equation (12). Among the mounting surface corrections of the mounting correction angles ΔΨ, ΔΘ, ΔΦ between the antenna base 6 and the inertial reference unit (IRU) 2, the roll angle ΔΨ and the pitch angle ΔΘ are obtained by providing an inclination angle sensor 9 on the antenna base 3. It can be measured with high accuracy. For example, when a servo accelerometer capable of measuring from a direct current component is used for the tilt angle sensor 9, the tilt angle can be measured with high accuracy based on gravitational acceleration. In addition, since the heading angle ΔΦ cannot be measured by an inclination angle sensor using gravity, a magnetic sensor or the like can be used when the accuracy is low. It is necessary to measure the relative positional relationship between the inertial reference unit (IRU) 2 and the antenna base 3 with high accuracy using a typical position measurement sensor, and to install the antenna base 3 in the aircraft 1.
As described above, by providing the antenna base 3 with the tilt angle sensor, it is possible to facilitate attachment to an aircraft of at least two axes (roll, pitch axis) and to track the antenna with high accuracy.

Embodiment 4 FIG.
FIG. 9 is a time history response waveform showing an example of the angle output signal of the inertial reference unit (IRU) 2 and the output angle signal of the attitude estimation filter in the satellite tracking antenna control apparatus according to the fourth embodiment of the present invention. The example 80 of the IRU output heading angle signal is discontinuously switched to -180 deg when the aircraft 1 continues to make a turn when the turn becomes greater than +180 deg. Similarly, when it becomes −180 deg or less, the value is discontinuously switched to +180 deg. The IRU angle signals (~ Φ, ~ Θ, ~ Ψ) 10 are angle signals, and therefore output a value of ± 180 deg. When a signal having this discontinuous ± 180 deg switching is input to the attitude estimation filter 12, for example, the attitude tracking filter output heading angle signal 81 in FIG. When the heading angle signal (˜Ψ) has a value such as −180 deg, a step input response to the low-pass filter is generated, and low-frequency vibration is generated. Further, it outputs a vibration waveform that should not occur when the IRU heading angle signal (˜Ψ) is turned back by ± 180 deg.
FIG. 10 shows a block diagram in which initialization processing and ± 180 deg folding processing are added to the attitude estimation filter 12 in the satellite tracking antenna control apparatus according to the present embodiment. In this figure, for example, a case where the heading angle signal (˜Ψ) of the IRU causes a turn of ± 180 deg is shown. In general passenger aircraft, the heading angle (~ Ψ) rotates continuously by 180 deg or more by turning, but the roll (~ Φ) and pitch (~ Θ) are ± 90 deg or less unless special flight such as somersault is performed, There are no discontinuities.

  For example, the heading angle signal (˜Ψ) 10 of the IRU is detected by the ± 180 deg folding detector 90 to determine whether or not the folding is ± 180 deg. For example, by monitoring the difference value of the heading angle signal (˜Ψ), if a change in a value corresponding to 360 deg is suddenly detected, a turn of ± 180 deg can be detected. ± 180 deg initialization processing of the yaw axis signal filter 14 ′, ± 180 deg initialization processing of the integrator 21 ′, and ± 180 deg initialization processing of the IRU delay time corrector 24 ′ based on the detection signal of the ± 180 deg folding detector 90 I do. Specifically, when these filters are constituted by digital filters, it can be dealt with by increasing / decreasing the value of ± 360 deg with respect to the value of each buffer storing the previous value at the time when folding of ± 180 deg occurs. . Further, when the satellite tracking antenna control device is activated or when an initialization operation is performed, a roll axis signal filter 12 ′, a pitch axis signal filter 13 ′, and a yaw axis signal for the initialization signal 91 are used. The filter 14 ′, integrators 19 ′, 20 ′, 21 ′, and IRU time delay correctors 22 ′, 23 ′, 24 ′ are initialized. More specifically, for example, when these filters, integrators, and delay time correctors are configured by digital filters, the values of the buffers that store the previous values are set to match the input values. Equivalent to.

In FIG. 10, ± 180 deg folding processing is performed only on the IRU heading angle signal (˜ψ), but similarly, ± 180 deg processing can be performed for the roll angle and pitch angle.
By adding the initialization process and the ± 180 deg process as described above, as shown in FIG. 9B, an example of the attitude estimation filter output heading angle signal (with a ± 180 deg folding process) 82, the folding is also ± 180 deg even in the initial state. Even if this occurs, a correct output angle signal can be obtained, and highly accurate antenna tracking control can be performed.

Embodiment 5 FIG.
FIG. 11 is a block diagram showing the concept of a technique for estimating the delay time of the inertial reference unit (IRU) in the satellite tracking antenna control apparatus according to the fifth embodiment of the present invention.
When the aircraft 1 repeats the turning motion 100 with respect to the three axes of roll, pitch, and heading, the aircraft vibration 101 of the IRU unit (front of the aircraft) is added to these motions, and the inertial reference unit (IRU) 102 The angle is measured and the IRU angle 103 is output. However, this signal contains a random time delay. On the other hand, an angular velocity signal obtained by adding the body vibration 104 of the antenna unit (the center of the body) is measured by an angular velocity sensor (rate gyro) 105. This signal is converted by a coordinate transformation 106 based on the positional relationship between the rate gyro 105 and the antenna base 3 and the relative angular relationship between the inertial reference unit (IRU) 2 and the antenna base 3 to obtain a rate gyro angular velocity 107. These two types of signals IRU angle 103 and rate gyro angular velocity 107 are input to the IRU delay time estimator 108, and a delay time estimated value (τ) 109 is output.

More specifically, for example, when the IRU angle 103 is differentiated with respect to time, the IRU angular velocity is obtained. If sinusoidal turning is repeatedly performed as the airframe turning motion, the IRU angle and angular velocity become sinusoidal signals. When this is compared with the signal of the rate gyro angular velocity 107, the difference in time during which the signal crosses zero corresponds to the delay time (τ) 109. In practice, it is possible to estimate the delay time with a more sophisticated algorithm combined with a filter.
By providing the delay time estimator 108 as described above, the delay time value set in the posture estimation filter can be given with higher accuracy, so that more accurate antenna tracking control can be performed.

Embodiment 6 FIG.
FIG. 12 is a bird's-eye view of a satellite tracking antenna having a two-degree-of-freedom configuration according to Embodiment 6 of the present invention. The antenna has two degrees of freedom: an azimuth axis that rotates horizontally and an elevation axis that moves the elevation angle of the antenna up and down. The fixed shaft of the azimuth rotation actuator 110 is fixed to the antenna base 3 (not shown), and the antenna support mechanism and the gimbal 4 are connected to the rotation shaft. The antenna support mechanism and the gimbal 4 are attached to the fixed side of the elevation rotation actuator 111, and the antenna 5 is attached to the rotation side. With such a configuration, the antenna can be directed in an arbitrary direction of the satellite by rotating the two-degree-of-freedom actuators of the azimuth angle actuator 110 and the elevation angle rotation actuator 111.

In order to mount an antenna with as few aircraft changes as possible, it is necessary to reduce the height of the antenna as much as possible so that the aerodynamic characteristics of the aircraft are not changed as much as possible (aerodynamic resistance is not increased). It is desirable that the degree of freedom of the antenna has three rotational degrees of freedom that are the same as the attitude angle motion of the aircraft. However, if there are three degrees of freedom, it is structurally difficult to reduce the antenna height. Therefore, if antenna freedom is composed of two azimuth and elevation angles, the antenna height can be lowered, but the singular point (keyhole) where the azimuth angle cannot be determined when the elevation angle is high is geometric. There is a problem that the satellite tracking of the antenna becomes impossible. An example of a control law 2 DOF antenna is shown in the aforementioned equation (15), the rewritten angular velocity command value ^ omega _AZ azimuth as follows.

For example, assuming that ^ ω _BZ = 0 and the aircraft body continues to turn at an angular velocity of ^ ω _BX around the X axis (around the roll axis), the angular velocity command value for the antenna azimuth is tan ^ It is a function of θ _EL . When the elevation angle of the antenna is ^ θ _EL = 0 deg, 45 deg, 80 deg, 89 deg, 90 deg, the value of tan ^ θ _EL is obtained as follows.

tan (0 deg) = 0
tan (45 deg) = 1
tan (80 deg) = 5.67
tan (89 deg) = 57.29
tan (90 deg) = ∞ (18)
As can be seen, the coefficient increases rapidly as the angle of elevation of the antenna approaches 90 degrees. In particular, in the case of ^ θ _EL 90 deg, it becomes infinite and the angular velocity command value ^ ω _{AZ of the} antenna azimuth cannot be determined. As can be seen from equation (17), when the rotation or vibration (^ ω _BX , ^ ω _BY ) occurs in the roll and pitch directions, the azimuth of the antenna needs to be moved at an angular velocity about 57 times faster when the elevation angle is 89 deg. Occurs.

However, it is difficult to rotate the azimuth at such a high speed because of the limitation of the output torque of the azimuth actuator 110, and if the azimuth actuator 110 is constantly moved greatly by the airframe vibration, the power consumption becomes large and practical. I'm in trouble.
FIG. 13 is a block diagram relating to a keyhole filter that controls the antenna without greatly moving the azimuth angle even when a disturbance such as airframe vibration occurs when the antenna is at a high elevation angle. FIG. 13A is a block diagram showing a state in which a keyhole filter 128 is added to the block diagram of the control system described above.

FIG. 13B is a detailed block diagram showing the internal configuration of the keyhole filter 128. An azimuth command value (^ θAZ) 120 and an elevation command value (^ θEL) 122, which are outputs from the coordinate converter 15, are input to the keyhole filter 128, respectively. Further, the azimuth angular velocity command value (^ ωAZ) 124 that is the output of the coordinate converter 17 is also input to the keyhole filter 128. The azimuth command value (^ θAZ) 120 is input to a low-pass filter 121 that varies the cut-off frequency according to the elevation angle command value (^ θEL) 122, and outputs an azimuth command value (^ θAZ) 123. Similarly, the azimuth angular velocity command value (^ ωAZ) 124 is input to the low-pass filter 125 that varies the cutoff frequency according to the elevation angle command value (^ θEL) 122, and outputs the azimuth angular velocity command value (^ ωAZ) 126. .
These outputs 123 and 126 are respectively input to the azimuth servo motor controller 19 as shown in FIG. 13A, so that the azimuth of the antenna can be controlled to the satellite with high accuracy.
When the elevation angle (^ θEL) 122 of the antenna is large due to the low-pass filter, the cut-off frequency of the low-pass filters 121 and 125 is lowered to increase the normal angular velocity when aircraft vibration occurs at a high frequency. Since the high-frequency vibration is reduced by the low pass filter using the keyhole filter, it is possible to prevent a sudden change in the angular velocity ωAZ of the azimuth angle. This can also prevent a rapid increase in power consumption.

Embodiment 7 FIG.
FIG. 14 is an example of a function indicating the characteristics of the keyhole filter of the satellite tracking antenna control apparatus according to the seventh embodiment of the present invention.
For example, when a cubic function is used for the cut-off frequency curve 130 of the keyhole filter that changes continuously according to the elevation angle θEL shown in FIG.

The cutoff frequency f _c3 [Hz] is set high from the antenna elevation angle θEL = 0 [deg] to θEL = _{θEL_key} [deg]. However, when θEL = _{θEL_key} [deg] or higher, the third order of Equation (19) As shown in the function, the cut-off frequency suddenly decreases as it approaches 90 deg, and is set to f _c1 [Hz] at 90 deg.
FIG. 15 is a Bode diagram showing an example of a low-pass filter that realizes the above-described cutoff frequency. The low-pass filter of FIG. 15 is given by the following equation, for example.

Here, f _{C is a} cutoff frequency. An example of the transfer characteristic when f _C = 0.5 Hz is shown in FIG.
Transform by bilinear sz transform. First, formulas (21) and (22) are rewritten as follows.

Here, T _{s is the} sampling time [sec].
A low-pass filter is configured by serially combining the above transfer functions. This can be described as a block diagram in FIG. In the block diagram, a01, a11, b11, a02, a12, b12, a03, a13, and b13 are varied according to the antenna elevation angle (θEL) 150 based on the relational expressions of Expressions (19) and (22).
As described above, in the “keyhole filter” according to claim 7 described above, the cutoff frequency of the high-frequency cutoff filter is given by a high-order function or the like, and is continuously changed according to the change of the elevation angle. High-accuracy antenna tracking control can be realized even for a continuous change in elevation angle.

Embodiment 8 FIG.
FIG. 17 is an example of an elevation angle time history waveform showing the azimuth stop condition at the high elevation angle of the satellite tracking antenna control apparatus according to the eighth embodiment of the present invention.
The antenna elevation angle command value time trajectory 161 approaches 90 deg as time passes, and the azimuth angle command value (^ θAZ) stops at the fixed start point (P1) 162 of the azimuth angle (θAZ). Servo control fixes the azimuth angle. The elevation angle θEL at the fixed start point (P1) 162 is set to (90−ε1) deg. ε1 is set to a minute angle, for example, about 0.3 deg. Here, the time locus 161 of the elevation angle command value of the antenna further increases and starts to decrease after a certain time, but the azimuth angle (θAZ) remains fixed even after passing through (90−ε1) deg. Yes, driving of the azimuth angle (θAZ) is started only after the fixed end point (P2) 163 of the azimuth angle (θAZ). Here, the elevation angle θEL at the fixed end point (P2) 162 is set to (90−ε2) deg, and there is a relationship of ε1 <ε2, for example, ε2 = 0.5 deg. With this setting, the elevation angle θEL = 90 deg to (90−ε1) deg is in the azimuth (θAZ) fixed zone 164 and the elevation angle θEL = (90−ε1) deg to (90−ε2) deg. Can be referred to as the hysteresis zone 165.

  If there is no hysteresis zone 165, the curve 161 in FIG. 17 is very unstable, such as flapping and azimuth moving ± 180 degrees each time (90−ε1) deg. Thus, in addition to the “keyhole filter” of the above-mentioned claim 7, by providing the hysteresis zone and the fixed zone, even when the elevation angle is closest to the zenith 90 deg (for example, 90 deg to 89.5 deg) ( A stable azimuth tracking motion can be realized even in the case of a singular point with an azimuth angle of 90 deg.

Embodiment 9 FIG.
FIG. 18 is an example of a time history response in the vicinity of the high elevation angle of the azimuth and elevation angles to show the effect of limiting the maximum velocity of the azimuth angle of the satellite tracking antenna according to the ninth embodiment of the present invention.
18 shows a time history response when the maximum elevation angle θEL = 88 deg, FIG. 19 shows a time history response when the maximum elevation angle θEL = 89 deg, and FIG. 20 shows a time history response when the maximum elevation angle θEL = 90 deg. Yes.
A curve 170 in FIG. 18 is an example of the time change of the azimuth angle (command value, actual angle) [maximum elevation angle θ _EL = 88 deg], 171 is an example of the time change of the azimuth angular velocity [maximum elevation angle θ _EL = 88 deg], and 172 is Example of time change of elevation angle (command value, actual angle) [maximum elevation angle θ _EL = 88 deg], 173 is an example of time change of elevation angle velocity [maximum elevation angle θ _EL = 88 deg], and 174 is an example of time change of satellite pointing error. [Maximum elevation angle θ _EL = 88 deg]. In this case, there is no speed limitation on the angular velocity of the azimuth angle of the curve 171.
A curve 180 in FIG. 19 is an example of the time change of the azimuth angle (command value, actual angle) [maximum elevation angle θ _EL = 89 deg], 181 is an example of the time change of the azimuth angular velocity [maximum elevation angle θ _EL = 89 deg], and 182 is Example of time change of elevation angle (command value, actual angle) [maximum elevation angle θ _EL = 89 deg], 183 is an example of time change of elevation angle velocity [maximum elevation angle θ _EL = 89 deg], and 184 is an example of time change of satellite pointing error [Maximum elevation angle θ _EL = 89 deg]. Here, a speed limit of about 40 deg / s is provided for the angular velocity of the azimuth angle of the curve 181 so that the azimuth angle is not rotated at a speed higher than this.

A curve 190 in FIG. 20 is an example of the time change of the azimuth angle (command value, actual angle) [maximum elevation angle θ _EL = 90 deg], 191 is an example of the time change of the azimuth angle velocity [maximum elevation angle θ _EL = 90 deg], and 192 is Example of time variation of elevation angle (command value, actual angle) [Maximum elevation angle θ _EL = 90 deg], 193 is an example of time variation of elevation angle velocity [Maximum elevation angle θ _EL = 90 deg], 194 is an example of time variation of satellite pointing error [Maximum elevation angle θ _EL = 90 deg], 195 is an example [maximum elevation angle θ _EL = 90 deg] of an azimuth angle theoretical value. Again, a speed limit of about 40 deg / s is provided for the angular velocity of the azimuth angle of the curve 191 so that the azimuth angle is not rotated at a higher speed.
In the case of the two-degree-of-freedom antenna configuration similar to that of the above-described sixth embodiment, for example, when the aircraft passes slowly under the equator and the elevation angle becomes 90 deg, the satellite cannot be tracked unless the azimuth is instantaneously rotated 180 deg. . In addition, when aircraft vibrations are applied, the azimuth angle may repeat the rotation of ± 180 degrees. To solve this problem, the motor torque is not increased rapidly even when passing through the singular point of the azimuth angle by adding the control law that limits the maximum speed of the azimuth angle to a constant value as described above. A stable tracking movement can be realized.

Embodiment 10 FIG.
FIG. 21 is a block diagram showing an azimuth angle ± 180 deg process and an initialization process when the antenna is activated in the keyhole filter of the satellite tracking antenna control apparatus according to the tenth embodiment of the present invention.
Whether the signal of the azimuth angle command value 120 returns at ± 180 deg is determined by the ± 180 deg return detector 200. Based on this determination, ± 180 deg processing of the low-pass filter 202 is performed. Further, when the initialization signal 202 is input when the antenna is activated, the low-pass filters 202 and 203 are initialized. As a result, outputs 123 and 126 are obtained. The contents of these processes are equivalent to the ± 180 deg process and initialization process of the attitude estimation filter shown in FIG.

  The “keyhole filter” in the above-described sixth embodiment is applied to two signals of antenna azimuth and azimuth velocity. However, since the azimuth angle rotates 360 degrees continuously, the actual signal becomes a discontinuous signal of ± 180 degrees. (On the other hand, the azimuth velocity does not become discontinuous.) If this discontinuous azimuth is input to the “keyhole filter”, a desired output cannot be obtained at the discontinuity point. Therefore, similarly to the above-described claim 7, by providing the keyhole filter considering the discontinuity of 180 deg as described above, even if the azimuth angle of the antenna continues to rotate continuously by 360 deg, highly accurate tracking is possible. Control can be realized.

  The satellite tracking antenna control device according to the present invention can be used not only for tracking satellites but also for tracking radio signals flying from other celestial bodies, and for tracking radio waves of other aircraft.

It is a figure which shows the definition about the arrangement | positioning of the sensor of the satellite tracking antenna control apparatus by Embodiment 1 of this invention, an antenna, the angle of each part, etc. It is a block diagram which shows the control system of the antenna control apparatus for satellite tracking of FIG. FIG. 3 is a block diagram illustrating an internal configuration of a posture estimation filter in FIG. 2. The relationship between the angular velocity signal measured by the rate gyro attached to the airframe and the Euler angle and Euler angle time derivative measured by the IRU 2 is shown. FIG. 4 is a diagram illustrating a control frequency band when the frequency is plotted on the horizontal axis and the gain is plotted on the vertical axis in the posture estimation filter of FIG. 3. It is an example of a structure at the time of adding coordinate transformation 23 'to the attitude | position estimation filter of FIG. It is a conceptual diagram which shows the intensity distribution of the electromagnetic wave which the satellite tracking antenna by Embodiment 2 of this invention receives. It is an overhead view at the time of providing a tilt angle sensor in the base part of the satellite tracking antenna by Embodiment 3 of this invention. It is a time history response waveform which shows an example of the angle output signal of an inertial reference unit (IRU), and the output angle signal of an attitude | position estimation filter in the satellite tracking antenna control apparatus by Embodiment 4 of this invention. FIG. 10 is a block diagram in which initialization processing and ± 180 deg folding processing are added to the attitude estimation filter in the satellite tracking antenna control apparatus according to the fifth embodiment of the present invention. It is a block diagram which shows the concept of the method of estimating the delay time of an inertial reference unit (IRU) in the satellite tracking antenna control apparatus by Embodiment 5 of this invention. It is a bird's-eye view of a satellite tracking antenna having a two-degree-of-freedom configuration according to Embodiment 6 of the present invention. It is a block diagram explaining a keyhole filter. It is a detailed block diagram which shows the internal structure of the keyhole filter of Fig.13 (a). It is an example of the function which shows the characteristic of the keyhole filter of the satellite tracking antenna control apparatus by Embodiment 7 of this invention. It is a Bode diagram which shows an example of the low pass filter which implement | achieves a cutoff frequency. It is a block diagram which shows an example which comprised the transfer function shown in FIG. 15 which changes a cutoff frequency according to an antenna elevation angle with the digital filter. It is an example of the time history waveform of the elevation angle which shows the stop condition of the azimuth angle at the time of the high elevation angle of the satellite tracking antenna control apparatus according to the eighth embodiment of the present invention. An example of a time history response in the vicinity of a high elevation angle of the azimuth angle and elevation angle to show the effect of limiting the maximum velocity of the azimuth angle of the satellite tracking antenna according to the ninth embodiment of the present invention. Time history response. It is a time history response in the case of the maximum elevation angle θEL = 89 deg in FIG. This is a time history response when the maximum elevation angle θEL = 90 deg in FIG. It is a block diagram which shows the +180 deg process of the azimuth in the keyhole filter of the satellite control antenna control apparatus by Embodiment 10 of this invention, and the initialization process at the time of antenna starting.

Explanation of symbols

1 Aircraft (airframe), 2 Inertial reference unit (IRU), 3 Antenna base,
4 Antenna indication mechanism (gimbal), 5 Antenna, 6 Angular velocity sensor,
7 Relative angle sensor, 8 Ground surface,
9 Mounting angle corrector (voltage / angle conversion),
10 attitude angle sensor, 11 rate gyro signal, 12 attitude estimation filter,
13 Estimated posture angle, 15 Coordinate converter, 16, 17 Angular velocity estimated value,
20 roll axis signal filter, 21 pitch axis signal filter,
22 yaw axis signal filter, 23 coordinate converter,
24, 25, 26 integrator, 27, 28, 29 IRU delay time corrector,
110 azimuth rotation actuator, 111 elevation rotation actuator,
128 keyhole filter, 165 hysteresis zone,
202, 203 low-pass filter,
Radius of curvature of the shoreline at the latitude of the _RA aircraft,
Radius of curvature of the shoreline at the latitude of the Rs satellite,
Re equator radius,
e ² Eccentricity of the earth (ellipsoid),
[psi _S satellite latitude, lambda _S satellite longitude, altitude h _S satellite,
ψ _A aircraft latitude, λ _A aircraft longitude, h _A aircraft altitude.

Claims (8)

An inertial reference unit mounted on an aircraft that measures its own position and orientation, an antenna mount mounted on the aircraft that is remote from the inertial reference unit, and is installed on the antenna mount and is directional-driven and communicated An antenna that communicates with a satellite, an angular velocity sensor that is installed in the vicinity of the antenna mount, measures at least a two-dimensional angular velocity of the antenna mount, and uses a body posture signal measured by the inertial reference unit and an angular velocity signal of the angular velocity sensor. It has a posture estimation filter that estimates the current posture of the antenna mount,
This posture estimation filter internally includes a coordinate converter and an integrator for calculating the angular velocity sensor output to obtain an estimated posture value of the antenna mount, and an airframe posture that is output from the inertial sensor unit and taken into the posture estimation filter. An IRU time delay corrector for delaying an input signal by a time corresponding to a signal time delay, and an attitude estimated value of the antenna mount is input to the IRU delay time corrector, and its output is output from the inertia reference unit. Includes a signal filter that estimates the drift amount compared to the aircraft attitude signal,
Using the obtained drift estimation value, the output of the angular velocity sensor or its coordinate conversion value is corrected to obtain the estimated posture value of the antenna mount, and the antenna is based on the estimated posture of the antenna mount output by the posture estimation filter. The antenna control device for satellite tracking, wherein both the distortion of the aircraft body and the measurement delay of the inertial reference unit are corrected by controlling the aircraft. In claim 1, instead of correcting the output of the angular velocity sensor using the drift estimation amount, the drift estimation amount is removed from the Euler angular time derivative obtained by coordinate conversion of the output of the angular velocity sensor, Antenna control for satellite tracking , wherein both the distortion of the aircraft body and the measurement delay of the inertial reference unit are corrected by controlling the antenna based on the estimated attitude of the antenna mount obtained by integration apparatus. Roll axis the posture estimation filter is output from the inertial reference unit, the pitch axis, any of the attitude angle signal of the yaw axis is determined whether passed through the ± 180 deg, the type of the passed signal and ± 180 degrees folding detector that notifies the passage and the ± 180 degree folding detector by Ri said pose estimation filter, the integrator, or the initialization signal generator to initialize the IR U delay corrector The antenna control device for satellite tracking according to claim 1, comprising:   The antenna mount has a rotation angle of two degrees of freedom including one degree of freedom for controlling an azimuth angle around a vertical axis and one degree of freedom for controlling an elevation angle around one horizontal axis perpendicular to the vertical axis, 2. The satellite tracking according to claim 1, wherein a signal of the angular velocity sensor installed in the vicinity of the antenna mount is output through a keyhole filter that varies a cutoff frequency according to the value of the elevation angle. Antenna control device. 5. The satellite tracking antenna control device according to claim 4 , wherein the cutoff frequency of the keyhole filter is configured to continuously change in accordance with a change in the value of the elevation angle. 5. The satellite tracking antenna control device according to claim 4 , wherein the azimuth angle is fixed when the elevation angle value is closer to the zenith than a predetermined angle. The predetermined angle has a hysteresis characteristic, and the predetermined angle when the elevation angle approaches the zenith is larger than the predetermined angle when the elevation angle moves away from the zenith. An antenna control device for satellite tracking according to claim 6 . 7. The control method according to claim 6 , wherein when the value of the elevation angle is closer to the zenith than a predetermined angle, the control of the azimuth includes a control law that limits the maximum speed to a predetermined constant value. An antenna control apparatus for satellite tracking as described.

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