JPS6150896A - Control system of attitude of artificial satellite - Google Patents

Abstract

(57) [Summary] This bulletin contains application data before electronic filing, so abstract data is not recorded.

Description

[Detailed Description of the Invention] [Technical Field of the Invention] This invention relates to the attitude of an artificial satellite during steady operation:
This relates to the ritll method.

[Prior art]

Conventionally, this type of control system includes a bias momentum system as shown in FIG. 1 and a zero momentum system as shown in FIG. III l'i earth sensor,
(2) is the earth sensor electronic circuit, (3) is the yaw sensor, (
4) is a yaw sensor electronic circuit, (5) is an attitude control device, (
8), Ql, If;
, ”]-Wheel, (7), 191, t
ill are roll, pitch, and yaw wheel drive circuits, respectively.

In the bias momentum method shown in Figure 1, ;,? j −
The roll angle and the pitch angle of Gin in the r range are detected by Earth-Binsa (1), but a sensor for detecting the yaw angle is not necessarily required. Pitch wheel tlGK has accumulated a large bias momentum, and the pitch angle is controlled by changing the rotational speed of pitch wheel 01. The roll angle is controlled by rotating the yaw wheel t13 and by the reaction of the combined momentum of the yaw wheel momentum and the bias momentum of the pitch wheel.

The yaw angle is controlled passively by controlling the roll angle using gyro stiffness due to the bias momentum of the pitch wheel and roll/yaw conversion due to the orbital motion coupling of the satellite.

In the zero momentum method shown in Figure 2, three components of the attitude angle of one 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 O-wheel wheel (8), and the pitch angle and yaw angle are controlled by changing the rotation speed of the pitch wheel l and yaw wheel (I2), respectively. This is done by making changes.

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 1fill.
1iil is a passive system based on the law of conservation of angular motion, and has the disadvantage that a large pitch wheel is required to improve accuracy, which increases the weight of the entire attitude control system.

The zero momentum method requires a yaw sensor (3) that constantly detects the yaw angle;
A lightweight and reliable yaw sensor (3) does not exist.

[Summary of the invention]

This invention is intended to improve such problems,
By estimating the disturbance momentum and compensating for it in advance, we propose a method for controlling the attitude side ν(1) of a satellite that has a high control angle, is light and heavy, and has high reliability.

[Embodiments of the invention]

FIG. 3 is an overall configuration diagram of an embodiment of the attitude control method for an artificial satellite according to the present invention. This example is explained in 1. As is clear from Figure 3, among the attitude angles of one artificial satellite, the earth sensor il+ and its electronic circuit (2), which produce the roll angle and pitch angle, and the sun sensor (3), which detects the yaw angle, and its The roll wheel (
8), drive circuit for each wheel to change the rotation speed of pitch wheel (9), yaw wheel a value +71, tel, +91, on-board computer (6)
The VC is configured to control the attitude control device (5) using the processing capacity of .

Figure 81!4 shows the attitude control device (5) in Figure 3.

FIG. 3 is a processing function configuration diagram of the onboard aircraft (6). The wheel side leg device a9 gives a command 1b1 to give the wheel I an appropriate rotational speed to the wheel drive circuit.

Here, the wheel drive circuit 0 includes the roll wheel Lp BJ circuit (7) and the pitch wheel drive circuit (9) shown in FIG.
), representing the yaw wheel and drive circuit aυ, wheel 1
4) represents the roll wheel (8), pitch wheel aI, and yaw wheel 13 in FIG. The wheel momentum is detected by the wheel momentum detection device QB, and when it exceeds a certain value, the unloading control device 01 releases unnecessary momentum. The amount of momentum released during unloading is estimated by the unloading momentum estimating mechanism α barrel. The wheel control device [15 is controlled so as to suppress the transient response during simultaneous trapping and unloading. Once the unloading momentum release amount is obtained, the disturbance momentum at each time is estimated by the disturbance momentum estimation mechanism a. By driving the wheels to compensate for this disturbance momentum, the attitude angle errors of the two artificial satellites can be reduced.

In FIG. 4, the wheel control device, the wheel momentum detection device αe, and the unloading control device aη are as follows:
The attitude control device (5) shown in FIG. 3 is configured, and the unloading momentum estimating mechanism and the disturbance momentum estimating mechanism αj are configured as the on-board computer (6) shown in FIG.

Next, the operation of the above embodiment will be explained.

In FIGS. 3 and 4, the pitch wheel fin is.

Since it rotates at a constant rotational speed, a constant bias momentum is applied. As a result, the angular motion [
According to the law of circumstantial conservation, one artificial satellite tries to maintain a constant attitude state. Therefore, if there is an error in the pitch angle,
1+1) is detected by the ball sensor (1), and through the earth sensor electronics (2), the pitch wheel drive circuit (9) is controlled by the attitude control device (5), so that the rotational speed of the pitch wheel a1 is The pitch angle error is controlled to zero by the reaction of the angular momentum.

If there is an error in the roll angle, it is detected by the earth sensor [1), and the yaw wheel drive circuit αB is controlled by the attitude control device (5) through the earth sensor electronic circuit (2), so that the yaw wheel drive circuit αB is The rotational speed changes and by the reaction of its angular momentum.

The roll angle error is controlled to be zero.

If there is an error in the yaw angle and the yaw angle is detected by the solar sensor (3), the roll wheel drive circuit ()) is controlled by the attitude control device (5) through the solar sensor electronic circuit (4). , resulting in a roll wheel (8)
The rotational speed of changes, and due to the reaction of its angular momentum,
The yaw angle error is controlled to be zero. However, when the sun sensor (3) is used as a yaw sensor, the yaw angle cannot be detected accurately. Figure 6 shows the visible range Gυ of the solar sensor.

o2. It shows the invisible area (to) and (b), and the geostationary orbit Qυ of one artificial satellite is drawn from the direction of the north pole.

Due to the positional relationship between the Earth (2) and the Sun (solstice), the invisible area of the Tai 14 sensor is 03. In (b), the yaw angle cannot be detected.

The disturbance torque around the roll axis of the artificial satellite is accumulated as momentum, and is combined with the bias momentum of the pitch wheel to rotate one artificial satellite around the yaw axis.
This causes a yaw angle error.

Here, the yaw angle error is expressed as linear.

hB However, ψ is the yaw angle error, the satellite momentum due to H door disturbance, hxI'i the momentum of the roll wheel, h
B is the bias momentum of the pitch wheel.

If the sun sensor detects the yaw angle, set hx to 1(X
The yaw angle error can be made zero by matching K, but if the sun sensor cannot detect the yaw angle, the outsourced satellite momentum estimate H that is sufficiently close to HX
Control the roll wheel momentum hx equal to x. At this time, the yaw angle error is.

Ha s It is expressed as

The disturbance momentum estimation mechanism 09 is stored in the satellite. Here, Hx is the estimated disturbance momentum around the roll axis, and Hz is the estimated disturbance momentum around the yaw axis.

H2 can be approximately expressed as a Fourier series whose fundamental harmonic is the orbital period. That is.

However, not only a constant term and a vibration term but also a two-shot number term are required. Also, ω is the orbital angle m(L, (a, 1°Q
Xl(a
), (b), (b21) (1-0,1,2...
).

zi xl 1c , 1 , lr: , ) , Id , 1
, Id, 1 (J-1,2゜XJ
ZJ XJ ZJ・
) are Fourier coefficients.

When there is one solar sensor in the visible range, the hole momentum hx r hz is equal to the 21 star momentum hx and HzK, so there are several thieves.

By measuring h2, the Fourier coefficients can be estimated.

According to equations +31 and +41, a linear 1-knee-1 system of first order exists between the measured wheel morph/tam and the estimated Fourier coefficients.

H-CX+t...(5) However,
Hii jllll ”L wheel momentum vector, X is the estimated Fourier coefficient vector, C is the coefficient between wheel momentum and Fourier thread count 7+
Lf, +matrix, and ε are noise.

1 (As a constant method, according to the least squares method that minimizes the square of noise ε, the Fourier coefficient vector is
CH...(6) and IE are determined. So, the transposed matrix, -1 is the inverse 2
It means ri.

When the Fourier system nX is estimated, +31, +41
Estimated values of disturbance momentum Hx, Hz are obtained from the equation.

Once the disturbance momentum estimation coefficient f(,,H,) is obtained by the disturbance momentum estimator t; daαj, these are constantly inserted into the wheel control device a9 and the feed 7
Perform off-food control. If the disturbance momentum can be estimated with high accuracy, the load on the wheel control device a'1) (1!9) can be reduced. Also, by the operation of the disturbance momentum estimation device 09, the yaw in the solar sensor visible range can be reduced. Angular control can be performed with high precision.

The wheel 041 can absorb the periodic disturbance torque and change the attitude state of the two satellites to the target attitude state KfM, but in response to the permanent disturbance torque, the wheel speed increases and becomes the controller nhi. Sometimes. Therefore, it is necessary to constantly monitor the momentum of the wheel α with the wheel momentum detection device Rue and, if it exceeds a certain value, release the unnecessary angular momentum accumulated in the wheel 6a using the unloading control device αD. At this time, the reaction causes a harmful attitude 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 It is effective to reduce the wheel momentum command.

The unloading momentum estimator (4q8) performs this function.

In addition, the disturbance moment homme estimation mechanism a! In I.

The disturbance momentum cannot be correctly estimated unless the amount of momentum change due to unloading is considered.

That is, the unloading momentum estimating mechanism tll 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 one unloading momentum machine. FIG.
i Gradually decrease. At this time, the wheel momentum) is.

It is equal to the accumulation momentum (A) of 1ρI.

Further, the unloading momentum estimated value (E) holds the cumulative unloading momentum up to the previous time.

Suppose that unloading is now carried out at time 2 (tl). The unloading thruster torque (0) generates torque only during t. In order to suppress this transient response, the unloading thruster torque (0) is
The unloading momentum equivalent to the momentum generated by the wheel is calculated as the wheel l-lux bias (c). 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 to 12.

The unloading momentum at time t2 is here:
CX is a transformation matrix from the inertial coordinate system to the Wy road coordinate system, and C1 is its inverse Dali. HoQ
uM is the estimated cumulative unloading momentum up to the previous time, tl is the actual unloading moment (on is the actual unloading time), and T is the unloading thruster thrust level. Because there is.

At time t2, the wheel momentum 0) does not match the satellite's bond momentum 0), resulting in a slight transient response.

Since equation (7) is a prediction, it is necessary to actually measure the change in momentum of unloading If and use it as an estimated value of unloading momentum.

Assuming that the transient response that occurred at r t2 has subsided at time t3, the wheel momentum () is measured N times from time t3 to time t4, and the average thereof is taken as the wheel momentum after unloading.

That is, the unloading moment at time t4 is △ 4 p m constant 1) 1 & HUM can be expressed as a linear equation.

Here, hl is the wheel momentum i1 at time t1
)+1 constant one shift, ht is the wheel momentum measurement value from t to t4, and N is the number of measurements.

As described above, the unloading momentum estimation mechanism (
1υ, when the estimated momentum value HUM by unloading is obtained, the disturbance momentum estimation mechanism u calculates the measured wheel momentum value H (i.e., the cage, h
Correct 2).

~U~ha -H+ HUM (9) Also, it is necessary to subtract the estimated disturbance momentum estimation) momentum change due to direct H.

haUhaha H-HHuM’・1)1 Therefore, if unloading is performed.

Posture accuracy can be improved by inserting the estimated disturbance value corrected by equation 1 into the wheel control device 4.

In addition, in order to reduce the load on the onboard computer (6) that executes the disturbance momentum estimation mechanism (1), in the geostationary satellite orbit, based on the latest information from the past 24 hours, WL Road - Shu 8
It is also possible to estimate the disturbance momentum during one round of the second order by off-line processing of the 14 i++++ constant data in Hoy/Hoy. Figure 6 is.

This figure shows the wheel momentum measurement trajectory position in the above case. Expressing the orbit position in local time, I23 is 12 o'clock and Q3 is 15 o'clock. Below, a41. (c), (to).

(5), @, (to) songs mean local time of 18:00, 21:00, 24:00, 3:00, 6:00, and 9:00. Of these, Q
Since the solar sensor cannot be used for the orbital position of 2 and the country,
Only h, can be measured. Another kan, Q heart.

(c), Qη, (to), @ orbit position r hx
r h z can be measured.

As mentioned above, if the foot trajectory position is determined, the coefficient matrix C in equation (5) is known, and (6
) matrix operations can be performed in advance on the ground with high precision.

Furthermore, in the case of a stationary 1) 9j star, in order to reduce the processing load on the onboard counter, the processing of the unloading momentum estimation mechanism (IQ) and the disturbance momentum estimation mechanism α9 are performed using a ground division i installed at the ground station. In this case, the D groove diagram will be as shown in the lower part of Fig. 4.The wheel momentum signal that has passed through the unloading control device 1η is transmitted by the transmitter Gυ. , is received by the ground antenna G. After being converted into an electrical signal by the telemeter receiving device ■, the ground meter n1 of the ground station
Processing is performed in real time by machine (2). When the wheel command signal is output, after passing through the command transmitting device (to) once, the commands for a certain period of time 1d are collectively transmitted to the artificial satellite by the transmitting antenna (g). On the i♀1 star side, the reception is received at wholesale η.

Estimating the disturbance momentum of the 17L aircraft (6) once installed ``1
3. Wheel as a continuous command after being stored in the structure:
'414% Inserted into device.

O and V using a large j↑ calculator as a ground word 1n machine (b)
This enables high-speed processing in real time, and the on-board meter 'J1
- The processing load of the machine can be reduced by most of 1%.

〔Effect of the invention〕

As described above, according to the present invention, highly reliable attitude control of an artificial satellite can be performed with high and light JBLL and FB1 degrees.

[Brief explanation of drawings]

FIG. 1 is an overall configuration diagram of a bias momentum method, which is an example of a conventional attitude control method, and FIG. 2 is an overall configuration diagram of a zero momentum method. The third factor is an overall configuration diagram of an embodiment of the attitude control method 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. FIG. 6 is an explanatory diagram of the loading momentum estimator (1)'s dynamic P\ and FIG. 6 is an explanatory diagram of the invisible area of the sun sensor and the wheel momentum measurement trajectory position. In the figure, (1) is the earth sensor, (2) is the electronic circuit, (
3) is a sun sensor or yaw sensor, and (4) is the same electronic circuit. (5) is the attitude control bag fX<, fGl is the onboard computer,
(8) is the roll wheel, (7) is the same drive circuit, al
is the pitch wheel. (9) is a four-way motion circuit, i12 is a yaw wheel, (lυ
is the same drive circuit, a9 is the wheel control device, QE9 is the wheel momentum detection device, t1η is the unloading control device, and 0 is the unloading momentum estimation mechanism. +19 +i Disturbance momentum estimation machine groove. Note that the same or corresponding parts in the two figures are designated by the same reference numerals.

Claims (2)

[Claims] (1) An earth sensor, a sun sensor, and corresponding electronic circuits that detect the attitude angle of the artificial satellite, and the output signals of the earth sensor and the sun sensor as input, and the attitude angle and attitude angle change rate of the artificial satellite. and an attitude control device that controls a wheel drive circuit to change the rotational speed of the roll wheel, pitch wheel, and yaw wheel so as to reach a target attitude angle and attitude angle change rate. An attitude control method for an artificial satellite, characterized in that the attitude is controlled using a disturbance momentum estimating mechanism corrected by unloading that releases unnecessary momentum. (2) An artificial satellite according to claim (1), characterized in that a disturbance momentum estimation mechanism is assigned to a computer in a ground station, and the artificial satellite is equipped with a transceiver for real-time processing. attitude control method.