JP2000159199A - Guiding control device for spacecraft - Google Patents

Abstract

PROBLEM TO BE SOLVED: To come close to a target, to separate from the target, and to turn around the target with one-several time of maneuver by computing the orbital controlled variable so that a spacecraft returns to the real position after one orbital period of the target on the basis of the relative position and speed of the spacecraft at the time of orbital control maneuver. SOLUTION: A maneuver time determining unit 4 determines the appropriate time for maneuver so as to come close to a target, separate from the target and turn around the target. A navigation sensor 10 decides the relative position and the relative speed of the spacecraft to the target per each time. A controlled variable computing unit 5 computes the orbital controlled variable so that the spacecraft returns to the real position after one orbital period of the target. On the basis of this output, driving signal is sent from a controlled variable distributing unit 6 to an actuator 7. With this structure, the spacecraft performs the orbital movement 9 that it returns to the rear position again after one orbital period of the target.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a guidance control device for a spacecraft, and more particularly to a guidance control system for guiding a spacecraft toward or around a target.

[0002]

2. Description of the Related Art In order to perform various services such as checking the state of a target orbiting around the earth, refueling and repairing the target, a comparison is made between the spacecrafts to be serviced from a considerable distance. It is often necessary to approach close distances. Further, in order to perform an appearance inspection on the entire target, it is effective for the spacecraft to fly around the target.

Conventionally, methods and apparatuses for approaching / separating / circulating around such a target have been proposed. FIG. 8 is a diagram showing a method of approaching a spacecraft disclosed in Japanese Patent Application Laid-Open No. 2-200600. In FIG. 8, the origin O of the coordinate system O-XYZ is fixed to the center of mass of the target. The X axis is the direction of the orbital velocity of the target, the Y axis is the direction opposite to the angular momentum of the target, and the Z axis is the direction of the center of the earth as viewed from the target. 1 is a target (hereinafter referred to as a target),
Reference numeral 2 denotes a spacecraft, and reference numeral 3 denotes a line indicating a command direction. FIG.
At the time shown in (a), the initial position of the spacecraft 2 is shown. The angle of the target 1 with respect to the horizontal plane as viewed from the spacecraft 2 is defined as an azimuth angle φ, and the spacecraft 2 performs trajectory control so as to maintain the azimuth angle φ at the commanded azimuth. Here, FIG.
If the orbit control of the spacecraft 2 is not performed in the positional relationship shown in
Since the spacecraft 2 has a higher orbital altitude than the target 1, the spacecraft 2 appears to move in the −X-axis direction according to the orbital motion of the target 1. Therefore, in order to maintain the azimuth angle φ in the command azimuth, it is sufficient to perform trajectory control such that the azimuth angle φ also moves in the + Z-axis direction. At the point in time in FIG. 8B, the spacecraft 2 is closer to the target 1 along the command direction by the above-described trajectory control. At the time point of FIG. 8B, in order to further maintain the azimuth angle φ at the command azimuth, the trajectory control for moving in the + Z-axis direction again is performed. As described above, by repeatedly performing the trajectory control along the line 3, the spacecraft 2 approaches the target 1 along the line 3 as shown in FIG. As the magnitude of the relative altitude difference z from the target 1 decreases, the velocity in the −X-axis direction generated by the orbital motion decreases. Therefore, in the case of a long distance, a relatively quick approach is performed, and the approaching distance is obtained. In, approach at a reduced relative approach speed.

[0004]

In the conventional guidance control system for a spacecraft as described above, the trajectory control is performed so as to approach or separate from the target along a designated line. Orbit control maneuvers must be performed discretely at shorter time intervals than in the case of, and the consumption of propellant tends to be excessive. In addition, when the spacecraft orbits the target, the circumference of the commanded azimuth line is approximated by a polygonal line, and the number of designated lines increases, which tends to complicate operation. Also, the consumption of propellant tends to be excessive due to the increase in the number of lines.

SUMMARY OF THE INVENTION The present invention has been made to solve the above-described problems, and has a guidance and control device for a spacecraft capable of approaching / separating from or revolving around a target with a small number of one or several maneuvers. The purpose is to obtain.

[0006]

According to a first aspect of the present invention, there is provided a spacecraft guidance control apparatus for approaching / separating / circulating a target orbiting around the earth. In the control device, a navigation sensor unit for estimating a relative position and a relative speed of the spacecraft with respect to the target, a maneuver time determining unit for determining an orbit control maneuver time for controlling the orbit of the spacecraft, and the orbit control maneuver time A control amount calculation unit that calculates an orbit control amount such that the spacecraft returns to the position at the orbit control maneuver time after one orbit cycle of the target based on the relative position and relative velocity of the spacecraft in the orbit; An actuator for control, and a control amount distribution unit that distributes the trajectory control amount to the actuator and sends a drive signal.

According to a second aspect of the present invention, there is provided a spacecraft guidance control apparatus for approaching / separating / circulating a target which orbits around the earth. A navigation sensor unit for estimating a relative position and a relative speed of the spacecraft with respect to a maneuver time determining unit for determining an orbit control maneuver time for controlling the orbit of the spacecraft, and a relative position of the spacecraft at the orbit control maneuver time. An orbit such that the spacecraft passes through a position symmetrical to the position at the orbit control maneuver time with respect to the target or a virtually determined reference point after a half-orbit cycle of the target based on the position and the relative velocity. A control amount calculating unit for calculating a control amount, an actuator for trajectory control, and a drive signal by distributing the trajectory control amount to the actuator. Is obtained by a that the control amount distribution unit.

According to a third aspect of the present invention, there is provided a spacecraft guidance control device according to the first or second configuration, wherein the maneuver time determination unit sets the sun or sun with respect to a target or a virtually determined reference point. The time when the direction becomes almost the same as that of the spacecraft is determined as the orbit control maneuver time.

In the guidance control apparatus for a spacecraft according to a fourth configuration of the present invention, in the first or second configuration, the maneuver time determination unit determines a point where the target is farthest from the earth or a point closest to the earth. The time near the passing is determined as the orbit control maneuver time.

According to a fifth aspect of the present invention, in the guidance control apparatus for a spacecraft according to any one of the first to fourth aspects, the maneuver time determining unit determines the number of orbit cycles from the orbit control maneuver time. Is determined as the corrected maneuver time, and the control amount calculation unit propagates the relative position and relative velocity of the spacecraft at the corrected maneuver time based on the orbital motion, and half the time from the orbit control maneuver time. Predict the relative position after the orbital cycle or one orbital cycle,
The trajectory control amount for correcting the relative speed of the spacecraft at the corrected maneuver time is calculated so that the predicted value matches the trajectory determined at the trajectory control maneuver time.

[0011]

DESCRIPTION OF THE PREFERRED EMBODIMENTS Embodiment 1 FIG. 1 is a block diagram showing a configuration of a guidance control system for a spacecraft according to Embodiment 1 of the present invention. In FIG. 1, 4 is a maneuver time determination unit, 5 is a control amount calculation unit, 6 is a control amount distribution unit, 7
Is an actuator, 8 is a guidance control device of the spacecraft, 9 is a trajectory relative to a target of the spacecraft, and 10 is a navigation sensor of the spacecraft.

Next, the operation of the first embodiment will be described. The maneuver time determination unit 4 determines a time appropriate for the maneuver for approaching / separating / lapping the target 1. The navigation sensor 10 determines and outputs a relative position and a relative speed of the spacecraft with respect to the target at each time. At the maneuver time determined by the maneuver time determination unit 4, the control amount calculating unit 5 determines that the spacecraft returns to the current position (the position at the maneuver time. A trajectory control amount that returns to position) is calculated. A drive signal is sent from the control amount distribution unit 6 to the actuator 7 based on the output of the control amount calculation unit 5. The spacecraft makes an orbital motion 9 such that the spacecraft returns to the current position again after one orbit period of the target by the drive signal. According to the orbital motion 9, the relative position and relative speed of the spacecraft with respect to the target change.

For the purpose of explanation, the operation of the control amount calculating section 5 will be described when the target trajectory is a substantially circular trajectory.
It is widely known that the relative motion of a spacecraft with respect to a target in a circular orbit with an orbital angular velocity ω is expressed by the following equation when disturbance such as aerodynamic resistance can be ignored.

X (t _s ) = Φpp (t _s , t ₀ ) * x (t ₀ ) + Φpv (t _s , t ₀ ) * v (t ₀ ) (1)

In the equation (1), "x (t ₀ )", "x
(T _s ) ”is a relative position vector with respect to the spacecraft target at time t = t ₀ and time t = t _s , respectively, and“ v
(T ₀ ) ”is a relative velocity vector with respect to the target of the spacecraft at time t = t ₀ . Matrix "Φpp (t _s, t _0)",
“Φpv (t _s , t ₀ )” is a sub-matrix of the state transition matrix from time t = t ₀ to time t = t _s , and its component is shown in equation (2).

[0016]

(Equation 1)

From equation (2), the target velocity v such that it returns to the current position again after one orbit period of the target from the current time t _0.
d (t ₀ ) is given by equation (3).

[0018]

(Equation 2)

In equation (3), “z ₀ ” represents the z component of the position vector “x (t ₀ )”. The symbol "*" indicates that any value is acceptable. Therefore, the trajectory control amount is given by equation (4). In equation (4), “u ₀ ” is the velocity vector “v
(T ₀ ) ”.

[0020]

(Equation 3)

FIG. 2 is a diagram showing an operation example of the control amount calculating section 5 according to the first embodiment of the present invention. In the figure, 1
Is a target, 2 is a spacecraft, and the origin O of the coordinate system O-XYZ is fixed to the center of mass of the target 1. X
The axis is the direction of the orbital velocity of the target, the Y axis is the direction opposite to the angular momentum of the target, and the Z axis is the direction of the center of the earth as viewed from the target. 2 (a) is a diagram showing an operation example of the control amount calculating section 5 when passing through the point z ₀ on spacecraft 2 + Z axis in the maneuver time. Before Maneuver time,
The spacecraft 2 maintains the altitude difference z ₀ with respect to the target 1 and +
It is assumed that the object is moving in the X direction at the relative speed of Expression (5).

[0022]

(Equation 4)

The trajectory control amount at this time is obtained from the equation (4) using the equation (6).
Is calculated. By performing this orbit control once at the above-mentioned maneuver time, the spacecraft 2 shifts to an elliptical orbit around the target 1 thereafter.

[0024]

(Equation 5)

In this example, the spacecraft 2 is the target 1
There is an advantage that the target 1 is illuminated by the albedo light of the earth and is easily detected by the optical sensor, since the time when the light passes immediately below is the maneuver time.

FIG. 2B is a diagram showing an operation example of the control amount calculating unit 5 when the spacecraft 2 passes the point z ₀ on the -Z axis at the time of the maneuver time. Figure 2 while keeping the altitude difference z ₀
In contrast to (a), the spacecraft 2 that has moved in the −X direction performs the orbit control once with the orbit control amount given by Expression (6) at the maneuver time, so that the spacecraft 2 moves the target 1 The orbit moves to an elliptical orbit. In this example, since the time at which the spacecraft 2 passes directly above the target 1 is the maneuver time, the target 1 is detected by the optical sensor with the earth as a background.

As described above, the guidance control apparatus for a spacecraft according to the first embodiment of the present invention determines the navigation sensor for estimating the relative position and relative speed of the spacecraft with respect to the target, and determines the orbit control maneuver time. A maneuver time determination unit,
A control amount calculating unit that calculates a trajectory control amount, an actuator for trajectory control, and a control amount distribution unit that sends a drive signal by allocating the trajectory control amount to the actuator,
Based on the output of the navigation sensor unit, the control amount calculation unit outputs an orbit control amount such that the spacecraft returns to the current position again after one orbit period of the target by one orbit control by the orbit control, and the control amount calculation unit Based on the output, a drive signal is sent from the control amount distribution unit to the actuator to perform trajectory control. After performing one orbit control, the spacecraft returns to the current position in the orbital cycle after performing one orbital control under ideal conditions with no disturbance or observation error, because the orbital control returns to the current position again after one orbital cycle It keeps moving on such an orbit. Such a circular orbit is not always a circular orbit as conventionally performed, but when monitoring or inspecting the target 1, the spacecraft may fly around the target,
The shape of the orbital orbit may not be particularly limited. Therefore,
An orbit control amount for maintaining the orbit is not required. Actually, control for maintaining the orbit is necessary due to disturbances and observation errors, but in such a case, the control amount is expected to be small.

The attitude control during the orbit is considered as follows. In order to keep observing the target 1, it is desirable to perform control (target pointing) in which the attitude with respect to the target 1 is kept constant. However, if the sensor that observes the target 1 has a wide field of view even in the earth-oriented or inertial-oriented direction, it is expected that the target 1 enters the sensor field of view at an appropriate timing while the spacecraft 2 orbits the target 1. You. Since the orbital period of the target 1 orbiting the earth and the orbital period of the spacecraft 2 orbiting the earth are very close, the orbital period of the target 1 may be replaced by the orbital period of the spacecraft 2 in the present invention. Needless to say, there is nothing.

Embodiment 2 FIG. 3 is a block diagram showing a configuration of a guidance control system for a spacecraft according to Embodiment 2 of the present invention. The difference from FIG. 1 lies in the operation of the control amount calculation unit 5. That is, in the control amount calculation unit 5 of FIG. 3, at the maneuver time determined by the maneuver time determination unit 4,
One orbit control based on the output of the navigation sensor 10 provides an orbit control amount such that the spacecraft passes through a position symmetrical to the current position with respect to the target or a virtually determined reference point after a half-orbit period of the target. calculate.

For the sake of explanation, the operation of the control amount calculation unit 5 will be described in the case where the trajectory of the target is a substantially circular trajectory. formula
From (2), the target velocity v _d (t ₀ ) that passes through a position symmetrical to the current position with respect to the target or a virtually determined reference point after a half-orbit cycle of the target from the current time is expressed by the following equation.
Given by (7).

[0031]

(Equation 6)

In equation (7), the position vector of the reference point is as shown in equation (8). Since the target is located on the origin from the definition of the coordinate system, when passing through a position symmetrical with respect to the target, both “x _off ” and “z _off ” in equation (8) are set to 0. Good. (X _off , 0, z _off ) (8)

The trajectory control amount is given by equation (9). formula
In (9), “w ₀ ” is the z of the velocity vector “v (t ₀ )”.
Represents a component.

[0034]

(Equation 7)

[0035] FIG. 4 is a diagram showing an operation example of the control amount calculating section 5 according to the second embodiment of the present invention, at a point x ₀ on the spacecraft 2 + x-axis are the relative stop until the maneuver time The operation in the case of FIG. 4A shows an example in which trajectory control is performed such that the target 1 passes through a position symmetrical to the current position. From equation (9), the trajectory control amount is given by equation (10). This orbit control is performed at the maneuver time as 1
The spacecraft 2 shifts to an elliptical orbit that orbits the target 1 thereafter.

[0036]

(Equation 8)

FIG. 4B and FIG. 4C show examples in which the trajectory control is performed such that the trajectory passes through a position symmetrical to the current position with respect to the reference point (x _off , 0, 0). From equation (9), the trajectory control amount is given by equation (11).

[0038]

(Equation 9)

FIG. 4B shows a case where the x component of the reference point is larger than half of the current position (x _off <x _{0 in} consideration of the sign).
/ 2), the spacecraft 2 transitions to a circular orbit such that the spacecraft 2 approaches the target 1 and returns without approaching. FIG. 4C shows a case where the x component of the reference point is smaller than half of the current position (x _{of f when} considering the sign _).
In> corresponds to the case where the x _0/2), spacecraft 2 moves to orbit circling the target 1. As described above, in the second embodiment, various inspection trajectories, such as a trajectory that is inserted into the target orbit from the x-axis or a trajectory that approaches the target and returns to the original position by selecting the reference point, by the same calculation procedure. There is an effect that it can be realized. Also, the amount of orbit control for maintaining such an orbit becomes unnecessary, and by performing this orbit control once, the spacecraft continues to move on the same orbit thereafter.

Further, as shown in FIG.
Under the condition that the component x _off is set to be larger than half of the current position of the spacecraft, the position of the reference point x _off is made closer to the target 1 for each half-orbit cycle of the target. If the spacecraft outputs a trajectory control amount such that the spacecraft passes a position symmetrical to the position of the spacecraft with respect to the reference point after the half orbit period of The number of maneuvers required, and the consumption of propellant is small.

As described above, the guidance control device for a spacecraft according to the second embodiment of the present invention
By one orbit control, an orbit control amount is output such that the spacecraft passes through a position symmetrical to the current position with respect to the target or a virtually determined reference point after a half-orbit cycle of the target, and the control amount is calculated. A drive signal is sent from the control amount distribution unit to the actuator based on the output of the unit, and the trajectory control is performed. Various orbital control is performed by selecting the reference point because the orbit control that the spacecraft passes through a position symmetrical to the current position with respect to the target or the virtual reference point after the target half orbit cycle Is possible with a small number of maneuvers.

Embodiment 3 FIG. 5 is a diagram for explaining the operation of the maneuver time determination unit according to the third embodiment of the present invention. The operation example of the maneuver time determination unit 4 in the guidance control device according to the first or second embodiment is described. FIG. For the purpose of explanation, the trajectory of the target 1 is a substantially circular trajectory.

Since the orbital period is from several hours to one day, the direction of the sun 11 viewed from the earth hardly changes during that time. In addition, since the y-axis is defined as the direction opposite to the orbital angular momentum, the y-axis is oriented in substantially the same direction during one orbit period. Therefore, the angle formed by the sun 11 with respect to the y-axis can be considered constant during one orbit period.

On the other hand, since the z-axis points in the direction of the center of the earth viewed from the target 1 and the x-axis points in the direction of the orbital velocity of the target 1, it rotates 360 degrees around the y-axis during one orbital period. Conversely, when the x-axis and the z-axis are fixed, the sun 11 appears to rotate while maintaining a certain angle with respect to the y-axis.
This is the cone 12 representing the movement in the sun direction in FIG.

The spacecraft 2 can orbit around the target 1 by orbit control. Now, it is assumed that the spacecraft 2 keeps its orbit at a point 13 on the x-axis. Assuming that a point at which a straight line extending along the y-axis from the point 13 intersects the cone 12 is a point 14, the time when the sun 11 comes in a direction from the target 1 toward the point 14 is defined as the orbit control maneuver time by the maneuver time determination unit 4. judge. That is, the time at which the sun direction is substantially the same as the direction of the spacecraft with respect to the target or the virtually determined reference point is determined as the orbit control maneuver time. In the control amount calculation unit 5, at the above-mentioned maneuver time,
The control amount is calculated by Expression (10), and the control amount distribution unit 6 sends a drive signal to the actuator 7.

When the spacecraft 2 moves to the orbit around the target 1 at such timing, the angle α between the direction of the target 1 viewed from the spacecraft 2 and the direction of the sun 11 viewed from the spacecraft 2 becomes FIG. 6 shows an example of a numerical simulation of how it changes during orbit. In FIG. 6, time 15 is the orbit control maneuver time. When the angle α is close to 180 degrees, the target 1 is in front of light when viewed from the spacecraft 2, and when the angle α is close to 0 degrees, the target 1 is in backlight when viewed from the spacecraft 2. By shifting to the orbit at the timing described above, α is always about 140 degrees or more,
That is, the target 1 is kept in direct light, and a situation is created in which the target 1 can be easily observed by the optical camera over one cycle of the orbit.

Embodiment 4 FIG. 7 is a diagram for explaining the operation of the maneuver time determination unit according to the fourth embodiment of the present invention. The operation of the maneuver time determination unit 4 in the guidance control device for a spacecraft according to the first or second embodiment will be described. It is a figure showing another example of operation. Unlike FIGS. 5 and 6,
The trajectory of the target 1 is an elliptical trajectory.

In FIG. 7, the target 1 (or the spacecraft 2) determines that the maneuver time determination unit 4 is near the apogee 16.
The passing time is determined as the orbit control maneuver time. When the control amount calculation unit 5 outputs the orbit control amount with the target speed of the spacecraft 2 set to 0 so that the spacecraft 2 and the target 1 are relatively stationary at the apogee 16, the control position calculation unit 5 returns to the current position after one orbit cycle from this time. It is expected to return. At the same time
Since the trajectory of the target 1 is an elliptical trajectory, assuming that the eccentricity is e, the relative distance of the spacecraft 2 from the target 1 near the perigee 17 is (1 + e) / (1) of the relative distance near the apogee 16. -E) Magnified according to orbital motion by a factor of two. This magnification is equal to the ratio of the distance a * (1 + e) from earth 18 to apogee 16 to the distance a * (1-e) from earth 18 to perigee 17. Here, the symbol “a” is the major radius of the trajectory of the target 1.

In particular, when the apogee 16 of the target 1 is close to the altitude of the geosynchronous orbit, the movements of the target 1 and the spacecraft 2 seen from the ground station appear almost stationary near the apogee 16 like a geostationary satellite. Therefore, apogee 16
It is relatively easy to operate the orbit control of the spacecraft 2 from the ground station in the vicinity. If the relative distance near the apogee 16 is appropriately maintained by orbit control, the relative distance increases in accordance with the orbital movement near the perigee 17 as described above. It is considered that the danger that the spacecraft 2 collides with the target 1 can be avoided.

In the fourth embodiment, the maneuver time determination unit 4 determines the time at which the target 1 (or the spacecraft 2) passes near the apogee 16 as the orbit control maneuver time, and performs the orbit control. However, the maneuver time determination unit 4 may determine the time at which the target 1 (or the spacecraft 2) passes near the perigee 17 as the orbit control maneuver time to perform the orbit control.

In this case, the relative distance between the target 1 and the spacecraft 2 increases in accordance with the orbital movement near the perigee 17 as described above, so that the relative distance becomes the longest. On the other hand, at the apogee 16, the relative distance is the shortest. Therefore, if the orbit control is performed from the ground station near the perigee 17, there is an effect that the orbit control can be performed with high accuracy.

Embodiment 5 FIG. In the first to fourth embodiments, the trajectory control is performed so as to move to the target point after a half orbit cycle or one orbit cycle. However, the maneuver time determination unit 4 may set the maneuver time at intervals smaller than one orbit cycle or half orbit cycle. In an ideal case with no error, the spacecraft 2 moves on the orbit determined at the time of orbit control at the maneuver time set at fine intervals, so there is no need to perform the maneuver again. However, there is a case where the orbit deviates from the orbit determined by the orbit control due to a disturbance, an observation error, or the like. In this case, control for maintaining the orbit determined above is necessary. Therefore, a maneuver time set at an interval smaller than one orbit cycle or half-orbit cycle from the first maneuver time is referred to as a modified maneuver time. At each corrected maneuver time, the navigation sensor 10 measures the relative position and relative speed of the spacecraft at the corrected maneuver time, and the control amount calculator 5 calculates the above-mentioned measured values and the orbital motion 9.
The relative position after a half orbit cycle or one orbit cycle is predicted from the orbit control maneuver time based on the mathematical model of the orbit, and the corrected maneuver time is adjusted so that the predicted value matches the target orbit when performing the orbit control maneuver The trajectory control amount for correcting the relative speed at is calculated. As a result, an effect of reducing the influence of noise such as disturbance and observation error can be expected.

For the purpose of explanation, the operation of the control amount calculation unit 5 will be described for the case where the trajectory of the target is a substantially circular trajectory. formula
From (2), the target value “v” of the relative speed of the spacecraft at the modified maneuver time t _c that coincides with the orbit determined at the orbit control maneuver time after a half orbit cycle or one orbit cycle after the orbit control maneuver time. _d (t _c ) ”is the formula
Given by (12). In equation (12), “Φpv
(T _s , t _c ) ”is a sub-matrix of the state transition matrix from the modified maneuver time“ t = t _c ”to the time“ t = t _s ”, and“ x
(T _c ) ”is a relative position vector of the spacecraft 2 with respect to the target 1 at time t = t _c .

V _d (t _c ) = {Φ pv (t _s , t _c )} ⁻¹ * [Φ pp (t _s , t ₀ ) * x (t ₀ ) +... (12) Φ pv (t _s , t ₀ ) * {v (t ₀ ) + δ _v (t ₀ )} − x (t _c )]

From equation (12), the trajectory control amount is given by equation (13). “V (t _c )” is a relative velocity vector of the spacecraft 2 with respect to the target 1 at time t = t _c . δ _v (t _c ) = v _d (t _c ) −v (t _c ) (13)

[0056]

As described above, according to the first configuration of the present invention, in the guidance control apparatus for a spacecraft that approaches / separates / circulates a target that orbits around the earth. A navigation sensor unit for estimating a relative position and a relative speed of the spacecraft with respect to a target; a maneuver time determining unit for determining an orbit control maneuver time for controlling the orbit of the spacecraft; A control amount calculation unit that calculates an orbit control amount such that the spacecraft returns to the position at the orbit control maneuver time after one orbit period of the target based on the relative position and the relative speed, and an actuator for orbit control And a control amount distribution unit that distributes the trajectory control amount to the actuator and sends a drive signal, so that a small amount of propellant circulates around the target. There is an effect that allows the trajectory retention for.

According to the second configuration of the present invention, a target approaching orbiting around the earth is approached /
In a guidance control device for a spacecraft performing separation / circulation, a navigation sensor unit for estimating a relative position and a relative speed of the spacecraft with respect to the target, and a maneuver time for determining a trajectory control maneuver time for controlling a trajectory of the spacecraft A determination unit, based on a relative position and a relative speed of the spacecraft at the time of the orbit control maneuver time, the spacecraft moves the orbit relative to the target or a virtually determined reference point after a half-orbit cycle of the target; A control amount calculation unit that calculates a trajectory control amount that passes through a position symmetric to the position at the control maneuver time, and an actuator for trajectory control,
Since the control amount distribution unit for transmitting the drive signal by distributing the trajectory control amount to the actuator is provided, there is an effect that various trajectory controls can be performed with a small number of maneuvers by selecting a reference point.

According to the third configuration of the present invention, in the first or second configuration, the time at which the sun direction becomes substantially the same as that of the spacecraft with respect to the target or the virtually determined reference point is set. Since the orbit control maneuver time is determined, there is an effect that the time during which the lighting condition is good can be extended.

Further, according to the fourth configuration of the present invention, in the first or second configuration, the time at which the target passes the point farthest from the earth or near the point closest to the earth is determined as the orbit control maneuver time. I decided to
It is possible to improve the safety or the accuracy of the orbit control.

Further, according to the fifth configuration of the present invention, in any one of the first to fourth configurations, the calculation of the orbit control amount requires the target of the spacecraft after one orbit cycle or half orbit cycle of the target. While using the position, a time that is about a fraction of the orbital period from the orbit control maneuver time is set as the corrected maneuver time, and the control amount calculation unit further calculates the relative position and relative velocity of the spacecraft at the corrected maneuver time in orbital motion. The relative maneuver time is predicted after a half orbit cycle or one orbit cycle from the orbit control maneuver time, and the corrected maneuver time is adjusted so that the predicted value coincides with the orbit determined at the orbit control maneuver time. The orbit control amount that corrects the relative speed of the spacecraft is calculated, so the effect of mitigating the effects of noise such as disturbances and observation errors while maintaining the effects described above is maintained. A.

[Brief description of the drawings]

FIG. 1 is a block diagram showing a configuration of a guidance control device for a spacecraft according to Embodiment 1 of the present invention.

FIG. 2 is a diagram illustrating an operation example of a control amount calculating unit according to the first embodiment of the present invention.

FIG. 3 is a block diagram showing a configuration of a guidance control device for a spacecraft according to Embodiment 2 of the present invention.

FIG. 4 is a diagram illustrating an operation example of a control amount calculating unit according to Embodiment 2 of the present invention;

FIG. 5 is a diagram illustrating an operation of a maneuver time determination unit according to Embodiment 3 of the present invention.

FIG. 6 is a graph showing an operation example of a guidance control device for a spacecraft according to Embodiment 3 of the present invention.

FIG. 7 is a diagram illustrating an operation of a maneuver time determination unit according to Embodiment 4 of the present invention.

FIG. 8 is a diagram showing a conventional approach method of a spacecraft to a target.

[Explanation of symbols]

1 target, 2 spacecraft, 4 maneuver time judgment unit, 5 control amount calculation unit, 6 control amount distribution unit, 7 actuator, 8 guidance control device, 10 navigation sensor.

Claims (5)

[Claims] A navigation sensor for estimating a relative position and a relative speed of the spacecraft with respect to the target in a spacecraft guidance control device which approaches / separates / circulates a target orbiting around the earth. Unit, a maneuver time determining unit that determines an orbit control maneuver time for controlling the orbit of the spacecraft, and based on the relative position and relative speed of the spacecraft at the orbit control maneuver time, after one orbit cycle of the target. A control amount calculation unit that calculates an orbit control amount such that the spacecraft returns to the position at the orbit control maneuver time again, an actuator for orbit control, and a drive signal by distributing the orbit control amount to the actuator. A guidance control device for a spacecraft, comprising a control amount distribution unit for sending. 2. A spacecraft guidance controller for approaching / separating / circulating a target orbiting around the earth in a navigation sensor for estimating a relative position and a relative speed of the spacecraft with respect to the target. Unit, a maneuver time determining unit that determines an orbit control maneuver time for controlling the orbit of the spacecraft, and based on the relative position and relative speed of the spacecraft at the orbit control maneuver time, after a half orbit cycle of the target. A control amount calculating unit that calculates an orbit control amount such that the spacecraft passes through a position symmetrical to the position at the orbit control maneuver time with respect to the target or the virtually determined reference point, and And a control amount distribution unit that distributes the orbit control amount to the actuator and sends a drive signal. Guidance control device. 3. A maneuver time judging unit judges a time at which the sun direction becomes substantially the same as that of the spacecraft with respect to a target or a virtually determined reference point as an orbit control maneuver time. Item 3. The guidance control device for a spacecraft according to item 1 or 2. 4. The universe according to claim 1, wherein the maneuver time determination unit determines a time near the time when the target passes a point farthest from the earth or a point closest to the earth as an orbit control maneuver time. Machine guidance control device. 5. A maneuver time determining unit determines a time that is about a fraction of an orbit cycle away from the orbit control maneuver time as a corrected maneuver time, and the control amount calculating unit determines a relative time of the spacecraft at the corrected maneuver time. Propagating the position and relative velocity based on the orbital motion, predicting the relative position after a half orbital cycle or one orbital cycle from the orbit control maneuver time,
5. The trajectory control amount for correcting the relative speed of the spacecraft at the corrected maneuver time such that the predicted value coincides with the trajectory determined at the trajectory control maneuver time. The guidance and control device for a spacecraft according to Claim 1.