JPH0569760B2 - - Google Patents

Description

[Detailed description of the invention]

[Technical Field of the Invention] The present invention relates to an attitude control device for an artificial satellite during steady operation. [Prior Art] Conventionally, as this type of control device, there are a bias momentum method as shown in FIG. 1 and a zero momentum method as shown in FIG. 1 is an earth sensor, 2 is an earth sensor electronic circuit, 3 is a yaw sensor, 4 is a yaw sensor electronic circuit, 5 is an attitude control device, 8, 10, and 12 are roll, pitch, and yaw wheels, and 7, 9, and 11 are respectively , roll, pitch, and yaw wheel drive circuits. In the bias momentum method shown in FIG. 1, the roll angle and pitch angle of the satellite's attitude angle are detected by the earth sensor 1, but a sensor for detecting the yaw angle is not required. A large bias momentum is stored in the pitch wheel 10, and the pitch angle is controlled by changing the rotational speed of the pitch wheel 10. The roll angle is controlled by rotating the yaw wheel 12 and by the reaction of the combined momentum of the yaw wheel momentum and the bias momentum of the pitch wheel. Yaw angle control is performed passively by controlling the roll angle using gyro stiffness due to pitch wheel bias momentum and roll/yaw conversion due to satellite orbital motion coupling. In the zero momentum method shown in FIG. 2, three components of the attitude angle of the satellite are always required, of which the roll angle and pitch angle are detected by the earth sensor 1, and the yaw angle is detected by the yaw sensor 3. The roll angle is controlled by changing the rotation speed of the roll wheel 8, and the pitch angle and yaw angle are controlled by changing the rotation speeds of the pitch wheel 10 and yaw wheel 12, respectively. It can be done. In both types, the attitude control device 5 provides control commands to give appropriate rotational speeds to each wheel. In the bias momentum method, the yaw angle is controlled passively according to the law of conservation of angular motion, and in order to improve accuracy, a large pitch wheel 10 is required, which has the drawback of increasing the weight of the entire attitude control method. have The zero momentum method requires a yaw sensor 3 that constantly detects the yaw angle, but at present there is no highly accurate, lightweight, and reliable yaw sensor 3. [Summary of the Invention] This invention is intended to improve this problem, and by estimating the disturbance momentum and compensating for it in advance, it is possible to create a lightweight, highly reliable artificial satellite with high control accuracy. This paper proposes an attitude control device. [Embodiment of the Invention] FIG. 3 is an overall configuration diagram of an embodiment of an attitude control device for an artificial satellite according to the present invention. As is clear from FIG. 3, this embodiment includes an earth sensor 1 and its electronic circuit 2 that detect the roll angle and pitch angle of the attitude angle of the artificial satellite, a sun sensor 3 that detects the yaw angle, and The roll wheel 8, pitch wheel 9, In order to change the rotational speed of the yaw wheel 10, each wheel drive circuit 7, 8,
9 is configured to control the attitude control device 5 using the processing power of the on-board computer 6. FIG. 4 shows the attitude control device 5 in FIG. 3,
It is a processing function block diagram of the on-board computer 6. The wheel control device 15 sends a control command to the wheel drive circuit 13 to give an appropriate rotational speed to the wheel 14.
give to Here, the wheel drive circuit 13 represents the roll wheel drive circuit 7, pitch wheel drive circuit 9, and yaw wheel drive circuit 11 shown in FIG. 3, and the wheel 14 represents the roll wheel drive circuit 7, pitch wheel drive circuit 9, and yaw wheel drive circuit 11 shown in FIG. It represents the yaw wheel 12. A wheel momentum detection device 16 detects the wheel momentum, and when it exceeds a certain value, an unloading control device 17 releases the unnecessary momentum. The amount of momentum released during unloading is estimated by the unloading momentum estimation device 18. At the same time, once the unloading momentum release amount that controls the wheel control device 15 to suppress the transient response during unloading is obtained, the disturbance momentum at each time is estimated by the disturbance momentum estimation device. By driving the wheels to compensate for this disturbance momentum, the attitude angle error of the satellite can be reduced. In FIG. 4, the wheel control device 15, the wheel momentum detection device 16, and the unloading control device 17 constitute the attitude control device 5 in FIG.
8. The disturbance momentum estimating device 19 constitutes the onboard computer 6 in FIG. Next, the operation of the above embodiment will be explained. In FIGS. 3 and 4, pitch wheel 10
rotates at a nearly constant rotational speed, so a constant bias momentum is accumulated. As a result, the satellite attempts to maintain a constant attitude due to the law of conservation of angular momentum. Therefore, if there is an error in the pitch angle, it is detected by the earth sensor 1, and sent to the pitch wheel drive circuit 9 by the attitude control device 5 through the earth sensor electronic circuit 2.
As a result, the rotational speed of the pitch wheel 10 changes, and the pitch angle error is controlled to be zero due to the reaction of the angular momentum. If there is an error in the roll angle, earth sensor 1
The yaw wheel drive circuit 11 is controlled by the attitude control device 5 through the earth sensor electronics 2 so that the yaw wheel 12
The rotational speed of the roller changes, and the reaction of the angular momentum controls the roll angle error to zero. If there is an error in the yaw angle and the yaw angle is detected by the sun sensor 3, the sun sensor electronic circuit 4
Through this, the roll wheel drive circuit 7 is controlled by the attitude control device 5, and as a result, the rotational speed of the roll wheel 8 changes, and the yaw angle error is controlled to be zero due to the reaction of the angular momentum. be done. However, when using the sun sensor 3 as a yaw sensor, the yaw angle cannot always be detected. Figure 6 shows the visible range 3 of the solar sensor.
1, 32, and invisible areas 33, 34, the geostationary orbit 21 of the artificial satellite is drawn from the direction of the North Pole. Due to the positional relationship between the earth 20 and the sun 30, the yaw angle cannot be detected in the invisible areas 33 and 34 of the sun sensor. Disturbing torque about the satellite's roll axis accumulates as momentum and is combined with the bias momentum of the pitch wheel to rotate the satellite about the yaw axis, resulting in a yaw angle error. Here, the yaw angle error is expressed as follows. ψ = H _{X −} h _X / h _B ... (1) where ψ is the yaw _angle error, _H If the sun sensor detects the yaw angle, set _h
The yaw angle error can be made zero by matching H _X , but if the sun sensor cannot detect the _yaw angle, the roll _wheel control the momentum _hX . At this time, the yaw angle error is expressed as ψ= _HX −H^ _X / _hB ...(2). The disturbance momentum estimating device 19 is for determining the estimated disturbance momentum value H^ _XH ^ _Z accumulated in the satellite. Here, H^ _X is the estimated value of the disturbance momentum around the roll axis, and H^ _Z is the estimated value of the disturbance momentum around the yaw axis. H^ _X and H^ _Z can be approximately expressed as a Fourier series with the orbital period as the fundamental harmonic. ^That _{is , H ^ _ _ _ _} ^_ _{_ _ _ _ _ z0} + _o 〓 ^k=1 (a _zk coskω _pt +b _zk sinω _pt ) + _o 〓 ^k=1 (c _zkt coskω _pt +d _zkt sinkω _pt ) ...(4) However, not only the constant term and the vibration term, but also the divergence section is required. Also, ω _p is the orbital angular velocity, {a _xi },
{a _zi }, {b _xi }, {b _zi }, (i=0, 1, 2...)
，
{c _xj }, {c _zj }, {d _xj }, {d _zj }, (j=1, 2,
…)
is the Fourier coefficient. When the sun sensor is in the visible range, the wheel momentum h _X and h _Z are the satellite momentum, respectively.
Since H _X and H _Z are equal, the Fourier coefficient can be estimated by measuring several h _X and h _Z. According to equations (3) and (4), the following linear relationship exists between the measured wheel momentum and the estimated Fourier coefficient. H~=Cx^+ε...(5) However, H is the measured wheel momentum vector, x^ is the estimated Fourier coefficient vector,
C is a coefficient matrix between wheel momentum and Fourier coefficients, and ε is noise. As an estimation method, according to the least squares method that minimizes the square of the noise ε, the Fourier coefficient vector is estimated as X^=(C ^T C) ^-1 C ^T H~ (6). However, T means a transposed matrix, and -1 means an inverse matrix. When _the Fourier _coefficient Once the _estimated disturbance momentum values H _^
Performs feed forward control. If the disturbance momentum is accurately estimated, the load on the wheel control device 15 can be reduced. In addition, the operation of the disturbance momentum estimating device 19 allows accurate yaw angle control in the sun sensor invisible region. The wheel 14 absorbs periodic disturbance torque,
Although the attitude of the satellite can be maintained at the target attitude, the wheel speed may increase and become uncontrollable in response to permanent disturbance torque.
Therefore, the wheel momentum detection device 16 constantly monitors the momentum of the wheel 14, and if it exceeds a certain value, the unloading control device 1
7, it is necessary to release the unnecessary angular momentum stored in the wheel 14. At this time, the reaction causes an attitude error that is harmful to the satellite, but in order to reduce this attitude error, the unloaded momentum is added to the wheel control system as a feedforward signal, and the effective wheel control system is It is effective to reduce the momentum command. The unloading momentum estimation device 18 performs this function. Further, in the disturbance momentum estimating device 19, the disturbance momentum cannot be correctly estimated unless the amount of change in momentum due to unloading is considered. That is, the unloading momentum estimating device 18 estimates the momentum released to the outside world by unloading in order to fulfill the above two functions. FIG. 5 is an explanatory diagram of the operation of the unloading momentum estimation device 18. The accumulated momentum of the satellite in Figure 5A gradually increases or decreases depending on the disturbance torque. At this time, the wheel momentum MU is approximately equal to the satellite's accumulated momentum MU. Furthermore, the estimated unloading momentum value E holds the cumulative unloading momentum up to the previous time. Assume that unloading is now performed at time _t1 . Unloading thruster torque cross
Generates torque only during t _po . In order to suppress this transient response, unloading momentum equivalent to the momentum generated by the unloading thruster torque roller is added as a wheel torque bias. Since the wheel torque bias is small compared to the unloading thruster torque, it is necessary to continue applying the wheel torque bias from t ₁ to t ₂ . The unloading momentum estimate H^ _UM ² at time t ₂ can be expressed as: H^ _UM ² =C ⁰ _I [H^ ⁰ _UM +C ^I ₀ ∫ ^t1 _t1 ^+t _po T _C dt]……(7
) Here, C ⁰ _I is the transformation matrix from the inertial coordinate system to the orbital coordinate system, and C ^I ₀ is its inverse matrix. H ⁰ _UM is the estimated cumulative unloading momentum up to the previous time, t ₁ is the unloading time, t _po is the unloading thruster on time, and T _C is the unloading thruster thrust level. Since the unloading momentum is a predicted value, at time _t2 , the wheel momentum does not match the satellite's accumulated momentum, resulting in a slight transient response. Since equation (7) is a predicted value, it is necessary to actually measure the amount of change in momentum before and after unloading and use it as an estimated unloading momentum value. Assuming that the transient response that occurred at _t2 has subsided at time _t3 , the wheel momentum is measured N times from time _t3 to _t4 , and the average thereof is taken as the wheel momentum after unloading. That is, the unloading momentum estimated value H^ _UM ⁴ at time t ₄ can be expressed as follows.

〔Effect of the invention〕

As described above, according to the present invention, it is possible to perform attitude control of an artificial satellite with high control accuracy, light weight, and high reliability.

[Brief explanation of the drawing]

FIG. 1 is an overall configuration diagram of a bias momentum system, which is an example of a conventional attitude control device, and FIG. 2 is an overall configuration diagram of a zero momentum system, which is an example of a conventional attitude control device.
FIG. 3 is an overall configuration diagram of an embodiment of an attitude control device for an artificial satellite according to the present invention, FIG. 4 is a diagram of the processing function configuration of the attitude control device and onboard computer in FIG. 3, and FIG. FIG. 6 is an explanatory diagram of the operation of the dinging momentum estimating mechanism, and is an explanatory diagram of the invisible area of the sun sensor and the wheel momentum measurement orbit position. In the figure, 1 is the earth sensor, 2 is the same electronic circuit, 3
is the sun sensor or yaw sensor, 4 is the electronic circuit, 5 is the attitude control device, 6 is the onboard computer, 8 is the roll wheel, 7 is the same drive circuit, 10 is the pitch wheel, 9 is the same drive circuit, 12 is the yaw wheel ,
11 is the drive circuit, 15 is a wheel control device, 1
6 is a wheel momentum detection device, 17 is an unloading control device, 18 is an unloading momentum estimation device, and 19 is a disturbance momentum estimation device. It should be noted that the same or corresponding parts in the figures are indicated by the same reference numerals.

Claims (1)

[Claims] 1. An earth sensor, a sun sensor, and corresponding electronic circuits for detecting the attitude angle of an artificial satellite;
Using the output signals of the earth sensor and the sun sensor as input, the roll wheel, pitch wheel, and yaw wheel are rotated so that the attitude angle and attitude angle change rate of the artificial satellite reach the target attitude angle and attitude angle change rate. A wheel control device that controls a wheel drive circuit for changing speed, a wheel momentum detection device that detects the momentum of the wheel, and when the wheel momentum detected by the wheel momentum detection device exceeds a certain value. , an unloading control device for releasing unnecessary momentum built up in the wheel;
When the unloading momentum estimating device estimates the unloading momentum release amount and supplies the unloaded momentum to the wheel control device, and the unloading momentum estimating device obtains the unloading momentum release amount, An attitude control device for an artificial satellite, comprising: a disturbance momentum estimating device that supplies a disturbance momentum estimated value corrected by a momentum change amount estimated value due to unloading to the wheel control device. 2. An attitude control device for an artificial satellite according to claim 1, characterized in that the disturbance momentum estimating device is assigned to a computer in a ground station, and the artificial satellite is equipped with a transmitter/receiver for real-time processing. .