JPH0478519B2 - - Google Patents

Description

[Detailed description of the invention] The present invention relates to an attitude control method for an artificial satellite.

To improve the accuracy of satellite attitude control,
It is considered necessary to develop a satellite attitude control method based on attitude determination using stellar sensors and target attitude state determination using orbit information.

By the way, this type of attitude control method that has been developed so far rotates the two viewing directions of the earth sensor and calculates the relative deviation (φ, θ) between the earth center direction and the satellite axis direction, as shown in Figure 1. More specifically, the roll deviation φ is detected from the difference in earth scanning width between the sensor fields of view 1 and 2, and the pitch deviation θ is detected from the deviation of the reference pulse from the center of the scanning pulse. It was something to control.
However, as is well known, in the conventional method described above, the accuracy of detecting deviations by the earth sensor is poor due to radiation fluctuations in the earth's atmosphere, and therefore it is difficult to control the attitude of the artificial satellite with high precision. Furthermore, since the conventional attitude control method is a control method in which only the attitude deviation is subject to control compensation, it is difficult to control the attitude change rate so that it coincides with the target value at the same time. There were drawbacks such as:

This invention relates to an attitude control method for an artificial satellite that will be necessary to realize a high-precision three-axis attitude control satellite that is expected to be developed in the future. The present invention attempts to provide an attitude control method for an artificial satellite configured to control the attitude state to a desired state based on this and a target attitude state calculated using orbit information from ground station commands.

An embodiment of the present invention will be described in detail below with reference to the drawings.

FIG. 2 is a diagram showing the general concept of attitude control using wheels. In this figure, the artificial satellite 1 is equipped with wheels 2 as torque generators in the satellite axis coordinates X _B , Y _B , and Z _B directions, and when the voltage or current applied to the wheels 2 is increased or decreased, the wheels 2 The rotation speed of increases or decreases. The artificial satellite 1 can be controlled around the X _B , Y _B , and Z _B axes by using the reaction caused by the electromagnetic force generated at this time.

FIG. 3 is a diagram showing the concept of the attitude control system according to the present invention. In the figure, 3 is a star sensor, 4 is an inertial sensor, 5 is an attitude determination device, 6 is a target attitude calculation device, 7 is an attitude control device, 8 is a computer, and 9 is an actuator control circuit.

In such a configuration, the attitude determining device 5 receives data from the stellar sensor 3 and data from the inertial sensor 4 as input, and calculates and outputs the attitude of the artificial satellite. The target attitude calculation device 6 reads the orbit information periodically transmitted as a command signal from the ground station by closing the switch SW, and reads this and the operation information of the secondary propulsion system, such as the jet, which is input from the attitude control device 7. Using this, the target attitude state, for example, the unit vector in the direction of the earth's center with respect to the satellite and its rate of change are calculated and output. The attitude control device 7 receives as input the current attitude state given by the attitude determination device 5, the future target attitude state given by the target attitude calculation device 6, and the wheel rotation angular velocity which is the output from the actuator control circuit 9. , calculates and outputs the operation amount for attitude control.
Using the output signal of the control computer configured as described above as input, the actuator control circuit 9 generates an analog voltage corresponding to the input signal and applies this to the wheel to control the rotational angular velocity of the wheel. This system realizes posture control.

Next, an example of a more specific configuration of each device of the above invention will be described.

FIG. 4 is a diagram showing the configuration of the attitude control system according to the present invention.

In the figure, the attitude determination device 5 includes a stellar vector calculation device 10, a database 11, a subcatalog editing device 12, a star identification processing device 13, an attitude displacement calculation device 14, and an attitude calculation device 15. Further, the attitude control device 7 includes a predicted attitude state calculation device 16, a control parameter calculation device 17, and a control variable calculation device 18. In such a configuration, the attitude determining device 5 determines the stellar coordinates (y _i ,
z _i ), [where i is the number of the observed star] as input,
The stellar vector calculation device 10 calculates a unit vector S _i (t ₀ ) in the stellar direction with respect to the satellite axis coordinate.
The subcatalog editing device 12 inputs the stellar catalog in the database 11 and edits the subcatalog S _j ^a
Edit (j is the catalog star number). At this time, posture information is required, and in the first processing, for example, a predicted value of the posture set in advance in the database 11 is used. The star identification processing device 13 determines the catalog star S _i ^a _j corresponding to the above S _i (t).
The attitude displacement calculation device 14 calculates the angular velocity ω _l (t) around the satellite axis measured by the inertial sensor 4 (where l=1,
2, 3) as input, calculate the change in attitude of the artificial satellite [〓C _B ] _T from the fixed star observation time t ₀ to the present time (for example, let this be t ₀ +T). The posture calculation device 15 inputs the above-mentioned (S _i (t ₀ ), S _ij ^a ) and [C _B ] _T , calculates and outputs the posture [ _B C _I ] _t at the current moment t=(t ₀ +T). Also, for the second time (i.e. t=t ₀ +
2T) For posture calculations after that, the above posture calculation device 15 is used.
The configuration is such that the posture [ _B C _I ] _t calculated in is fed back to the sub-catalog editing device 12 to provide posture information.

The attitude determination device 5 configured in this manner is configured to determine the star observation time (k-1)T obtained based on star identification.
Since the attitude at the time is corrected by the attitude change from (k-1)T to T time obtained by the attitude displacement calculation device 14, the attitude at the current moment kT (however, t ₀ = 0) is determined with high accuracy. This is to realize a system that

Next, details of each device constituting the attitude determining device 5 will be explained using FIGS. 5, 6, and 7.

FIG. 5 is a diagram showing the relationship between stellar sensor coordinates (X _s , Y _s , Z _s ) and satellite axis coordinates (X _B , Y _B , Z _B ). In this example, the number of installed star sensors 1 is 2.
It is said that

In the figure, (α ₁ , δ ₁ ) and (α ₁ , δ ₂ ) are the respective stellar sensors, and (α ₂ , δ ₂ ) are the (azimuth angle, elevation angle) of the optical axis direction X _s of the stellar sensor 1, respectively. In such a relationship, the coordinate transformation matrices [ _B C _S ] ₁ and [ _B C _S ] ₂ that provide the relationship between the stellar sensor coordinates and the satellite axis coordinates are given by the following equations.

[ _B C _S ] ₁ = cosα ₁ sinα ₁ sinδ ₁ cosδ ₁ cosδ ₁ −sinα ₁ cosα ₁ 0 −cosα ₁ −sinα 1 cosδ ₁ sinδ _{1 sinδ 1 (} 1) [ _B C _S ] ₂ = −sinα ₂ cosα ₂ sinδ ₂ cosδ ₂ cosδ ₂ cosα ₂ −sinα ₂ sinα ₂ −cosα ₂ cosδ ₂ sinδ ₂ sinδ ₂ (2) Figure 6 is a diagram showing the observation concept of the unit vector S _i (t) in the stellar direction with respect to the satellite axis coordinate. be.

In the figure, (y _i , z _i ) are stellar coordinates within the field of view of the stellar sensor, and are values observed by the stellar sensor 1. (α _l ,
δ _l ) is the (azimuth angle, elevation angle) of the star sensor number l (l=1, 2) in the optical axis direction.

In this relationship, the stellar vector calculation device 2 calculates a unit vector S _i (t) in the stellar direction with respect to the satellite axis coordinate using the following equation.

S _i (t)=[ _B C _S ] _l cosyi cosz _i siny _i cosz _i sinz _i l=1,2 i=1,2,3 (3) However, in this embodiment, multiple Out of the observed values, a total of 3, i.e. S ₁
(t) S ₂ (t) and S ₃ (t) are selected and used in the following processing.

On the other hand, the sub-catalog editing device 12 is provided with the data from the database 11 and the attitude calculation device 15 [ _B C _I ]
_t (However, only for the first processing, for example, [ _B C _I ] ₀ set in advance in the database) as input,
Edit subcatalog S _j ^a using the following method.

First, the unit vector i _Sl in the optical axis direction of the stellar sensor
Calculate (t) using the following formula.

i _Sl (t)=[ _S C _B ] _l [ _B C _I ] _t [1,0,0] ^T , l=1,
2
(4) Next, the stellar catalog ^a in database 4
From the following conditional expression, select a star _j ^a that is predicted to exist within the field of view of the star sensor.

S _j ^a ∈ [ ^a | cos ^-1 { ^a・i _sl } | < ε ₁ ] (5) However, ε ₁ is set to a constant value taking into account the size of the field of view of the stellar sensor.

The star identification processing device 13 is given by the above formula (3)
The following identification process is performed using S _i (t), i=1, 2, 3 and the stellar catalog value S _j ^a given by equation (5).

a ₁ = S ₁ (t)・S ₂ (t) a ₂ = S ₂ (t)・S ₃ (t) (6) Calculate a ₃ = S ₃ (t)・S ₁ (t).

Next, take one star S ₁ ^a from the subcatalog and use the preset constant ε ₂ for the remaining j−1 stars in the subcatalog to calculate S ₁ ^a , S _j−1 ^a −a It is tested whether there is even one combination that satisfies _i > cosε ₂ , i=1, 2, 3 (7). If even one pair exists ^, that star _1a is left as an identification candidate. Also, if no pair exists, that star is removed from the subcatalog. Repeat this operation for all stars in the subcatalog. If ε ₂ is set small enough, this will result in a set of catalog stars _{S 1j a} ^, S _2j ^a , S _2j a , S _3j ^{a ,} S _3j corresponding to a 1 , a 2 , _{a 3} in the subcatalog. ^a , S _1j ^a remains.
Therefore, the catalog value of S ₁ is S _1j ^a , the catalog value of S ₂ is S _2j ^a , and the catalog star of S ₃ is S _3j ^a , and the identification is completed.

Figure 7 shows the star observation time t ₀ by star sensor 1.
Assuming that t ₀ =0, the relative relationship between the attitude of the artificial satellite at time (n-1)τ and time nτ is shown. However, τ
is the sampling time of stellar observation. In the figure, (i _Bo-1 , j _Bo-1 , k _Bo-1 ) is a unit vector in the satellite axis direction at time (n-1)τ, (i _Bo , j _Bo , k _Bo )
is the unit vector in the satellite axis direction at time nτ, (ΔΨ _o , Δφ _o , Δθ _o ) are the Euler angles, and (ω _1o-1 ,
ω _2o-1 , ω _3o-1 ) and (ω _1o , ω _2o , ω _3o ) are rotational angular velocities around the satellite axis, respectively.

In such a relationship, the posture displacement calculation device 7
calculates the change in attitude [ ^ΔC B] from time t ₀ to current time t=nτ using the following formula.

[〓C _B ]=Δρ ₁ ² −Δρ ₂ ² −Δρ ₃ ² +Δρ ₄ ² 2 (−Δρ ₄
Δρ ₃ +Δρ ₁ Δρ ₂ ) 2(Δρ ₄ Δρ ₂ +Δρ ₁ Δρ ₂ ) 2(Δρ ₄ Δρ ₂ +Δρ ₁ Δρ ₂ ) 2(Δρ ₄ Δρ ₃ +Δρ ₁ Δρ ₂ )−Δρ ₁ ² +Δρ ₂ ² +Δρ
₃ ² +Δρ ₄ ² 2 (−Δρ ₄ Δρ ₁ +Δρ ₂ Δρ ₃ ) 2 (−Δρ ₄ Δρ ₁ +Δρ ₂ Δρ ₃ ) 2 (−Δρ ₄ Δρ ₂ +Δρ ₁ Δρ ₃ ) 2 (Δρ ₄ Δρ ₁ +Δρ
₂ Δρ ₃ ) −Δρ ₁ ² −Δρ ₂ ² +Δρ ₃ ² +Δρ ₄ ² (8) However _, Δρ ₁ Δρ ₂ Δρ ₃ Δρ _4o = Δq _{4 −Δq 3 Δq 2} Δq 1 Δq 3 Δq ₄ Δq ₁ Δq ₂ −Δq ₂ Δq ₁ Δq ₄ Δq ₃ −Δq ₁ −Δq ₂ −Δq 3 Δq ₄ Δρ ₁ Δρ ₂ Δρ _{3 Δρ 4o-1} (9) Δq ₁ = sinΔθ _o ／2cosΔφn／2sinΔΨn／2−co
sΔθn／2sinΔφn／2cosΔΨn／2 Δq ₂ =−sinΔθn／2cosΔφn／2cosΔΨn／2−
cosΔθn／2sinΔφn／2sinΔΨn／2 Δq ₃ = sinΔθn／2cosΔφn／2cosΔΨn／2−co
sΔθn／2cosΔφn／2sinΔΨn／2 Δq ₄ =cosΔθn／2cosΔφn／2cosΔΨn／2−si
nΔθn/2sinΔφn/2sinΔΨn/2 The unknown quantity here is the Euler angle (Δφ,
Δθ, ΔΨ) _o , but the initial value Δφ ₀ at time t=0
= Δθ ₀ = ΔΨ ₀ =, and the rotational angular velocity around the satellite axis measured by the inertial sensor (2) ω _uo (u=1, 2, 3)
is read with a period τ and calculated using the following formula.

Δφ _o = 1/6 (d ₁ +2d ₂ +2d ₃ +d ₄ ) Δθ _o = 1/6 (η ₁ +2η ₂ +2η ₃ +η ₄ ) (11) ΔΨ _o = 1/6 (f ₁ +2f ₂ +2f ₃ +f ₄ ) However, d ₁ = τω _1o η ₁ = τω _2o f ₁ = τω _3o d ₂ = τ｛ω _1o cos(ω _2o τ/2) + ω _3o sin(ω _2o τ/
2)} η ₂ = τ{ω _1o tan (ω _1o τ/2) sin (ω _2o τ/2) +
ω _2o −ω _3o tan (ω _1o τ/2) cos (ω _2o τ/2)} f ₂ = τ{−ω _1o sin (ω _2o τ/2)/cos (ω _1o τ/
2) + ω _3o cos (ω _2o τ/2)/cos (ω _1o τ/2)} d ₃ = τ{ω _1o cos (η ₂ /2) + ω _3o sin (η ₂ /2)} η ₃ = τ{ω _1o tan (d ₂ /2) sin (η ₂ /2) + ω _2o − ω _3o tan (d ₂ /2) cos (η ₂ /2)} f ₃ = τ{−ω _1o sin (η ₂ /2) /cos (d ₂ /2) + ω _3o cos (η ₂ /2) / cos (d ₂ /2)} d ₄ = τ{ω _1o cosη ₃ +ω _3o sinη ₃ } η ₄ = τ{ω _1o tand ₃ sinη ₃ +ω _2o −ω _3o tand ₃ cosη ₃ } f ₄ =τ{−ω _1o sinη ₃ /cosd ₃ +ω _3o cosη ₃ /cosd ₃ } In addition, [Δρ ₁ , Δρ ₂ , Δρ ₃ , Δρ ₄ ] ₀ ^T = [0,0,0,1
] ₀ ^T
(12).

The attitude calculation device 15 is the above-mentioned star identification processing device 13.
That is, (S ₁ , S ₂ , S ₃ ) and (S _1j ^a , S _2j ^a ,
S _3j ^a ), first, the attitude of the satellite at time t=0, that is, the satellite axis coordinates (X _B , Y _B , Z _B ) and the inertial space coordinates (X _I , Y _I , _ZI )
The relationship with [ _B C _I ] _t=0 is calculated using the following formula.

[ _B C _I ] _t=0 = S ₁ S ₂ S ₃ [S _1j ^a , S _2j ^a , S _3j ^a ] (13) Next, using equations (13) and (8), calculate the current posture. [ _B C _I ] _t=o 〓 is calculated using the following formula.

[ _B C _I ] _t=o = [〓C _B ] [ _B C _I ] _t=0 (14) The attitude calculation device 15 calculates the above equation (14) until the attitude initial value is updated again after T time. ) is used as the initial value, the attitude is calculated and output using the following formula.

Here, if the update time t=nτ of the attitude initial value is set to t=0 again, the attitude of the satellite [i _B , j _B , k _B ]
_tk is [i _B , j _B , k _B ] _tk = [ _B C _I ] _tk [i _I , j _I , k _I ] (15) [ _B C _I ] _tk = [〓C _B ] [ _B C _I ] It is given by _tK-1 (16).

However, for [ _B C _I ] ₀ , the calculation result of equation (14) is used.
In addition, [〓C _B ] reads inertial sensor data ω _l (l = 1, 2, 3) every time τ = t _k - _{t k-1} , and formula (11),
Calculate Euler angles Δφ, Δθ, ΔΨ using (12), substitute them into equation (16), and obtain Euler parameters (Δρ ₁ , Δρ ₂ , Δρ ₃ , Δρ ₄ ) from equations (9) and (10). is calculated and substituted into the right-hand side of equation (8) to sequentially calculate [〓C _B ].

Next, the attitude control device (7) will be explained.
The attitude control device (7) inputs the wheel rotation angular velocity δ(0) which is the output from the actuator control circuit (9), and the predicted attitude state calculation device (16) calculates the end time t of each control interval [0, t _f ]. Satellite attitude θ^(t _f ) at _f
and the attitude change speed ω^(t _f ) are calculated and output.
The target attitude state generator 6 outputs the target attitude θ° (t _f ) to be reached at the terminal time t _f and the attitude change rate ω° (t _f ). The control parameter calculation device 17 calculates the above ω^(t _f ), θ^(t _f ) and ω゜(t _f ), θ゜(
t _f ) as input, intermediate parameters λ ₁ and λ ₂ for determining control variables are calculated and output so that the sum of squares of wheel applied voltages is minimized. Control variable calculation device 18
calculates and outputs a control variable V _j (j=1, 2, . . . , m) for wheel operation using the above λ ₁ and λ ₂ as input. Using the output signal of the attitude control device 7 configured as described above as input, the actuator control circuit 9 controls the rotational angular velocity of the wheel in the same manner as the conventional method, thereby realizing attitude control of the artificial satellite.

Hereinafter, details of each device constituting the attitude control device will be explained using FIG. 8.

Assume that the wheel is a DC plasticless motor shown in FIG. 8a. In the figure, V is voltage, R is resistance, L is coil, i is current, 19 is rotor, 20
is a magnet. Now, each control interval [0, t _f ] is defined as m
The voltage V changes in steps such that the magnitude is constant in each small section as shown in FIG. 8b.
Consider control by (t). The predicted attitude state calculation device 16 calculates the wheel speed δ(0) at the initial time 0.
As input, the constant parameters a _j (j=1, 2,..., m), b that give the predicted values θ^(t _f ), ω^(t _f ) of the posture state at the terminal time t _f using the following equations: _j (j=1, 2,...,
m) and Ω ₁ , Ω ₂ .

ω^(t _f )=a ₁ V ₁ +a ₂ V ₂ +…+a _n V _n +Ω ₁ (17) θ^(t _f )=b ₁ V ₁ +b ₂ V ₂ +…b _n V _n +Ω ₂ ( 18) However Ω ₁ = −μe ^-c 1 ^t mδ(0) + f(t _f ) (20) Ω ₂ =-μ1/C ₁ [1-e ^-c 1 ^t f] δ(0)+F(t _f )(2
2) C ₁ = K ₁ K ₂ /RI _f , C ₂ = K ₂ /RI _f , μ = I _f /I _y (23) I _f : Moment of inertia of wheel K ₁ : Back electromotive force K ₂ : Torque constant I _y : Moment of inertia about the satellite axis Y _B (t _f ): Integral value of disturbance torque F (t _f )=∫ ₀ ^tf f (τ) dτ (24) μ=I _f /I _y (25) The target attitude state generator 6 incorporates a control target generation function unique to each artificial satellite, and its output is the attitude θ゜(t _f ) of the target to be reached at the end of each control interval [0, t _f ]. and the attitude change speed ω゜(t _f ).

The control parameter calculation device 17 uses equation (19) ~
Using the values calculated in (25) and ω° (t _f ) and θ° (t _f ) as input, the control intermediate parameters λ ₁ and λ ₂ are calculated using the following equations.

λ ₁ = −2 {(ω _y ⁰ (t _f ) − Ω ₁ ) (b ₁ ² + b ₂ ² + …+b _n ² ) −(θ _y ⁰ (t _f )−Ω ₂ ) (a ₁ b ₁ + a ₂ b ₂ +… +a _n b _n )}/ {(a ₁ ² +a ₂ ² +…+a _n ² )(b ₁ ² +b ₂ ² +…+b _n ² ) −(a ₁ b ₁ +a ₂ b ₂ +a _n b _n ) ² } (26) λ ₂ = 2 {(ω _y ⁰ (t _f )−Ω ₁ ) (a ₁ b ₁ +a ₂ b ₂ +… +a _n b _n )−(θ _y ⁰ (t _f )−Ω ₂ )a ₁ ² +a ₂ ² +…+a _n
² )}/ {(a ₁ ² +a ₂ ² +…a _n ² )(b ₁ ² +b ₂ ² +b ₂ ² +…+b _n ² )−(a ₁ b ₁ +a ₂ b ₂ +…+a _n b _n ) ² }(27) The control variable calculation device 18 uses the above equation (26), (27)
V _j is calculated using the following formula using the value of the formula as input.

V _j = -1/2 (λ ₁ a _j + λ ₂ b _j ), j = 1, 2,..., m (28) Note that this equation (28) satisfies equations (17) and (18) Conditional expression that minimizes V ₁ ² +V ₂ ² +…+V _n ² (29) out of V _j , that is, F= _n 〓 ^j=1 V _j ² +λ ₁ ( _n 〓 ^j=1 a _j V _j +Ω ₁ −ω゜(t _f )) +λ ₂ ( _n 〓 ^j=1 b _j V _j +Ω ₂ −θ゜(t _f )) (30), and the condition for the minimum is ∂F/∂V _j = It is derived from 0,j=1,2,...,m (31) ∂F/∂λ ₁ =0,i=1,2.

Equations (17) to (31) above show the calculation formula for the control operation amount for one wheel, but the calculation formula for the control operation amount for wheels mounted in two other axes is also the same as above. Needless to say, they are given in exactly the same way.

Finally, the ninth section describes an example in which the target attitude calculation device 6 is a geostationary three-axis satellite pointing to the center of the earth.
This will be explained using figures.

In the figure, 21 is an orbit calculation device, 22 is an earth direction unit vector calculation device, and 23 is a target control amount calculation device. In such a configuration, the orbit calculation device 21 calculates predicted values of the satellite position vector and velocity vector at any time by inputting orbit information periodically given by ground commands and the output signal of the control variable calculation device 18. Calculate. The earth direction unit vector calculation device 22 receives the output signal from the orbit calculation device 21 and calculates a unit vector in the direction of the earth center as seen from the satellite. The target control amount calculation device 23 inputs the output signals from the earth direction unit vector calculation device 22 and the attitude calculation device 15, and calculates ω^(t _f ), used in equations (17) and (18),
Calculate and output θ^(t _f ).

The details of each device constituting the target posture calculation device 6 will be explained below using FIG. 10.

FIG. 10 is a diagram showing the mathematical concept for satellite position calculation in the case of a geostationary satellite. Orbital information provided by ground commands includes ascending node right ascension Ω, orbital inclination i,
Assume period T and ascending node elongation f ₀ at time t ₀ . The position (RA, DI) of the satellite at an arbitrary time t is given as follows, with time t ₀ being 0 o'clock.

sinDI=sin i sinωt where ω=2π/T (32) RA=Ω+α where (33) cosα=cosωt/cosDI The unit vector e from the satellite center to the earth center is calculated using equations (32) and (33) It is given by the following formula using RA and DI.

e=cos(-DI)cos(RAS+180°) cos(-DI)sin(RAS+180°) sin(-DI) (34) Here, for example, let's talk about the field that controls the machine axis _YB direction toward the center of the earth. Using the attitude at the current time t = t _k given by equation (15) and the unit vector e at time t _f = t _k + τ given by equation (34),
The target attitude is given by the following equation.

The control targets around the X _B axis θ゜_X (t _f ), ω゜_X (t _f ) are The control target θ゜_Z (t _f ), ω _Z (t _f ) around the Z _B axis is is given by Note that for artificial satellites flying at medium or low altitudes, orbit calculations using special perturbation methods or the like are required, but even if such calculation methods are used, the effects of the present invention cannot be prevented.

As is clear from the above description, the attitude control method according to the present invention uses a stellar sensor and an inertial sensor to determine the attitude of an artificial satellite with respect to inertial space coordinates and measure the rotational angular velocity around the axis.
Using this and the target attitude state calculated using orbit information from ground station commands as input, it is possible to control the attitude of the satellite and the rate of change of attitude so that they simultaneously match the target state. without using an earth sensor,
Earth-centered three-axis attitude control can be realized. It has the following advantages.

[Brief explanation of the drawing]

Fig. 1 is a diagram showing the concept of earth-oriented three-axis attitude control using the conventional method, Fig. 2 is a diagram showing the concept of satellite axis coordinates and wheel mounting, and Fig. 3 is a diagram showing the concept of the attitude control method according to the present invention. 4 is a conceptual diagram of the configuration of an attitude control system showing a specific embodiment of the present invention, FIG. 5 is a diagram showing the relationship between star sensor coordinates and satellite axis coordinates, and FIG. 6 is a diagram showing the relationship between star sensor coordinates and satellite axis coordinates. A diagram showing the concept of observation, Figure 7 is a diagram showing the relationship between the previous stage posture and the current posture in sequential posture determination,
Figure 8a is a configuration diagram of a DC brushless motor, Figure 8b is a conceptual diagram of the wheel control voltage given by the control variable calculation device, Figure 9 is a diagram showing the configuration concept of the control target calculation device, and Figure 10 is a satellite diagram. It is a diagram showing an example of a mathematical model for position calculation, in which 1 is an artificial satellite, 2 is a wheel, 3 is a stellar sensor, 4 is an inertial sensor, 5 is an attitude determination device, 6 is a target attitude calculation device, and 7 is an attitude. Control device, 8 is a computer, 9 is an actuator control circuit, 10 is a stellar vector calculation device, 11 is a database, 12 is a subcatalog editing device, 13 is a star identification processing device, 14 is an attitude displacement calculation device, 15 is an attitude calculation device , 16 is a predicted attitude state calculation device, 17 is a control parameter calculation device, 18 is a control variable calculation device, 19 is a rotor, 2
0 is a magnet, 21 is an orbit calculation device, 22 is an earth direction unit vector calculation device, and 23 is a target control amount calculation device. Note that the same or corresponding parts in the figures are indicated by the same reference numerals.

Claims (1)

[Claims] 1. Attitude determining means for determining the attitude state of the satellite by inputting outputs from a stellar sensor and an inertial sensor, and an attitude determining means for inputting wheel rotation angle information from an actuator control circuit for attitude control. An attitude control means for calculating the amount of operation, orbit information transmitted as a command signal from the ground station, an output signal from the attitude control means, and an output signal from the attitude determination means to arrive at a certain time in the future. A computer equipped with a target attitude calculation means for calculating the desired attitude state of the artificial satellite is mounted on the artificial satellite, and a control operation signal outputted from the computer is used to calculate the attitude and attitude change rate of the artificial satellite to the target. An attitude control method for an artificial satellite, which is characterized by controlling the attitude and the speed of attitude change. 2. The attitude determining means is a stellar vector calculation device that calculates a unit vector in the stellar direction with respect to the satellite axis coordinates by inputting the output signal from the stellar sensor, and a predicted attitude value used at a stage before obtaining the stellar catalog and attitude determination results. a sub-catalog editing device that reads the information from the database and edits a catalog of stars predicted to be within the field of view of the star sensor, and inputs the output signals of the above-mentioned stellar vector calculation device and sub-catalog editing device to correspond to the observed stars. A fixed processing device that identifies catalog stars, and an attitude displacement calculation device that calculates the change in the satellite's attitude from the time when the star was observed by the star sensor to the current moment by inputting the output signal from the inertial sensor that measures the angular velocity around the satellite's axis. and an attitude calculation device that calculates the current attitude of the artificial satellite by inputting the output signals of the identification processing device and the attitude displacement calculation device, and the attitude control means is configured to calculate the attitude of the satellite at each control period [0, t _f ] The rotational angular velocity of the wheel at the initial time 0 is input, and the predicted attitude state calculation device calculates the predicted value of the attitude state of the satellite at the end time t _f of the control interval, and periodically as command data from the ground station. The end time of the control section is determined by inputting the transmitted orbit information at an arbitrary time, the output signal from the control variable calculation device, and the output signal from the attitude calculation device.
By inputting the output signals of the target attitude calculation device that calculates the target value of the attitude state that you want to reach at t _f , the predicted attitude state calculation device, and the target attitude calculation device, the sum of squares of the control operation amount is minimized. A control parameter calculation device that determines intermediate parameters for control so as to satisfy conditions, and a control variable calculation device that inputs the output signal of the control variable calculation device and calculates a manipulated variable for attitude control. An attitude control system for an artificial satellite according to claim 1, characterized in that: 3 Feedback the output signal of the attitude calculation device to the sub-catalog editing device, the output signal of the actuator control circuit to the predicted attitude state calculation device, and the output signal of the control variable calculation device to the target attitude calculation device. An attitude control system for an artificial satellite according to claim 1 or 2, characterized in that the system is configured to perform feedback.