JPH0733095A - Spacecraft control device and thruster control method - Google Patents

Abstract

PURPOSE:To provide a spacecraft control device capable of controlling the position and attitude of a spacecraft independently without using its center of gravity, and a thruster control method enabling efficient thruster distribution with a small amount of calculation. CONSTITUTION:A spacecraft control device is provided with a noninterference computing part 8 for eliminating interference on the dynamics of the operation quantity of position control and attitude control obtained from deviation between the target position value 3 and the output of a position detector 2 and from deviation between the target attitude value 6 and the output of an attitude detector 5, and a feedforward computing part 9 for compensating the acceleration component of the target value and inertial force generated on the dynamics from each target value and the output of each detector. Offset-permutation jet logic utilizing pattern matching is applied in a thruster selecting rule so as to obtain the jet distribution of a thruster small in the total jet quantity.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a spacecraft control device for controlling a spacecraft such as an artificial satellite which needs to control six degrees of freedom in position and attitude, and a thruster control method thereof.

[0002]

2. Description of the Related Art FIG. 23 shows, for example, a material of the 8th guidance control symposium sponsored by the Society of Instrument and Control Engineers (1991 1
Block diagram showing a conventional spacecraft controller shown in "Research on Space Free-Flight Telerobots-Guidance Control and Manipulation by Ground Experiment Model-" (Yoshitsugu Toda et al.), Pages 41-49 (issued on May 25). Is. In the figure, 1
Is a spacecraft body of a spacecraft such as an artificial satellite, 2 is a position detector that detects the position of the spacecraft body 1, 3 is a position target value generator that generates a position target value, and 4 is position control calculation It is a position control calculation unit for performing. Reference numeral 5 is a posture detector that detects the posture of the spacecraft body 1, 6 is a posture target value generation unit that generates a target posture value, and 7 is a posture control calculation unit that performs posture control calculation. Reference numeral 14 is a center-of-mass position calculation unit that determines the center-of-mass position of the spacecraft body 1, and reference numeral 15 is an inertial force calculation unit that determines an inertial force on the dynamics of the spacecraft body 1.

Next, the operation will be described. Spacecraft body 1
The position detector 2 and the posture detector 5 respectively detect the position and the posture of the. For the position detector 2, for example, an inertial sensor such as an acceleration sensor or a camera type sensor (hereinafter referred to as a proximity sensor) that measures a distance to a target is used. A gyro or earth sensor is used as the attitude detector 5. Since the position detector 2 does not always detect the position of the center of mass of the spacecraft body 1, the center of mass of the spacecraft body 1 is calculated by the center-of-mass position calculation unit 14 from the outputs of the position detector 2 and the attitude detector 5. Calculate the position. The position of the center of mass is important because the position and posture do not interfere with each other in dynamics when the position of center of mass is used, and thus the position control and the posture control can be independently performed. The position target value generation unit 3 outputs the target value of the center of mass position, and the position control calculation unit 4 outputs the position target value generation unit 3
The operation amount of the position control is obtained from the deviation between the output of (1) and the output of the mass center position calculation unit 14 (actual mass center position). The inertial force calculation unit 15 obtains an inertial force generated by using polar coordinates for expressing the center of mass position, and adds the inertial force to the output of the position control calculation unit 4 to compensate the inertial force. The sum of the outputs of the position control calculation unit 4 and the inertial force calculation unit 15 is input to the spacecraft body 1 and position control is performed by a position control actuator such as a thruster. On the other hand, the posture is controlled similarly to the position control. First, the attitude detector 5 detects the attitude of the spacecraft body 1, and the attitude target value generator 6 generates a target attitude value. The deviation between the two is input to the attitude control calculation unit 7 to obtain the operation amount of the attitude control. This manipulated variable is input to the spacecraft body 1 independently of the manipulated variable of the position control, and the attitude control is performed by an attitude control actuator such as a thruster.

Further, FIG. 24 shows, for example, Japanese Patent Laid-Open No. 62-592.
It is a block diagram which shows the thruster control method of the conventional spacecraft shown by the 00 publication, and FIG. 25 is a perspective view which shows the thruster arrangement | positioning. 24 and 25, 1
Is the spacecraft body, 18a is the first thruster, and 18b
Is a second thruster, 18c is a third thruster, and 18d is a fourth thruster. Reference numeral 16 is a thruster selection unit by an arithmetic circuit that performs thruster distribution according to the thruster selection rule, and 17 is a thruster modulator including a drive circuit. Reference numeral 20 is a thruster valve which is opened and closed by the thruster modulator 17, 21 is a propellant pipe connected to the thruster valve 20, and 22 is a propellant tank.

Next, the operation will be described. The attitude disturbance signal A about each of the three axes of pitch, yaw, and roll occurring in the spacecraft is input to the thruster selection unit 16 to synthesize attitude disturbance vectors, and the thrusters 18a to 18d corresponding to the synthesized vectors are selected. At the same time, the control amounts of the thrusters 18a to 18d are calculated. FIG. 26 shows the injection state of the thrusters 18a to 18d corresponding to each control axis. In the table, the circle marks indicate the thruster jets, and by simultaneously jetting the thruster circles, only the thrust forces for the axes to be controlled are generated and the thrust forces for the other two axes are canceled. As described above, the thruster control signal of the thruster selector 16 is input to the thruster modulator 17 to input the thruster valve 20.
The propellant flowing from the propellant tank 22 through the propellant pipe 21 is supplied to the thrusters 18a to 18d in accordance with the control amount and is injected.

[0006]

Since the conventional spacecraft control device is configured as described above, the position control amount (the amount to be controlled) must be controlled in order to control the position and the attitude independently.
As the center of mass of the spacecraft must be used as, it is difficult to freely select the control amount of position control,
In addition, there is a problem that the control error is apt to increase because the error of the detector is easily influenced in obtaining the center of mass position. Further, although inertial force compensation is performed, there is no feedforward compensation for the target value, so there is a problem that the tracking characteristics for the target value deteriorate when the target value changes rapidly.

When attention is paid to the thruster control method, the injection logic of the thrusters 18a to 18d is configured as compared to the case of controlling only the pitch axis, the yaw axis, or the roll axis only, and the case of simultaneously controlling a plurality of axes. Controlled by adding the logic for each axis. However, even if it is the optimum injection logic in the sense that the total injection amount decreases when controlling only individual axis components, when controlling multiple axis components at the same time, the logic of each axis component There is a problem in that the addition is not always optimal, and as a result, the total injection amount increases and valuable propellant is wasted. This problem could not be avoided even if the arrangement of the thrusters 18a to 18d was simple. As the simplest thruster arrangement, the force in only one axis, that is, the pure force,
It is conceivable to arrange the thrusters 18a to 18d so as to generate torque only around one axis, that is, pure torque. In order to generate a pure torque by the thruster, since couples are used, two thrusters 18a to 18d are indispensable. And the couple force is to offset the translational force of the two thrusters and to extract only the rotational force. Thrusters 18a-18
Considering the injection amount of d, it is not a good idea to cancel the translational force. That is, when it is required to generate both the force and the torque, the ratio of the injection amounts of the two thrusters forming the pure torque is changed rather than the realization by adding the pure force and the pure torque. It is considered that the total injection amount will be smaller if it is realized. As is clear from this example, both the force and the torque are generally applied to the thrusters 18a to 18d.
, The thrusters 18a-1
There is a problem that propellant is wasted if the logic addition for each axis is used regardless of the arrangement of 8d. On the other hand, if a general solution such as linear programming is used, an optimal solution can be obtained in the sense that the total injection amount is minimized, but it is far beyond the capacity of the current onboard computer. However, there is a problem that it is practically impossible to solve the 6-axis components by using linear programming in real time. Further, as the equipment mounted on the outer surface of the satellite such as the solar cell paddle, antenna, and arm increases in size and types, there is an increasing demand for preventing the jet gases of the thrusters 18a to 18d from hitting these equipments. The placement of the thrusters 18a to 18d is greatly restricted, and if the redundancy is also taken into consideration, 10 to 10
Twenty thrusters are normally arranged in various positions and orientations. As a result, it becomes difficult to intuitively grasp the optimum combination of the thrusters 18a to 18d in each axis component, and there is a problem that the designer has to make a trial and error request.

The first aspect of the present invention is made to solve the above problems, and position control and attitude control can be executed independently, and a target value for position control can be freely selected. It is an object of the present invention to provide a control device for a spacecraft that can improve the control error as a result.

It is another object of the present invention to provide a spacecraft control device capable of improving feed-forward compensation with respect to a target value and improving tracking characteristics with respect to the target value.

The invention described in claim 3 is the same as claim 1
A spacecraft controller capable of independently controlling the relative position and relative attitude between the spacecraft and the target by using camera-type proximity sensors for the position detector and the attitude detector in the invention described in Aim to get.

Further, according to the invention described in claim 4, even when the noise of the proximity sensor poses a problem, the influence of the noise is reduced by the estimator, and the relative space between the spacecraft and the target is reduced to 6%. It is an object of the present invention to obtain a spacecraft control device capable of controlling the degree of freedom with high precision.

Further, the invention according to claim 5 of the spacecraft is capable of accurately executing control such that the spacecraft faces the target and approaches the target while keeping the attitude of the inertial space constant. The purpose is to obtain a control device.

According to the invention of claim 6, the spacecraft orbits the target while facing the target on a surface having an inertial space without using an acceleration sensor or the like for detecting the position. An object of the present invention is to obtain a control device for a spacecraft that can execute such control with high accuracy.

Further, according to the invention described in claim 7, the spacecraft orbits around the target while facing the target on the surface having the inertial space without using an acceleration sensor or the like for detecting the position. It is also an object of the present invention to obtain a spacecraft control device capable of accurately executing control such that its orbiting surface rotates in inertial space.

Further, according to the invention described in claim 8, the useless portion in the injection pattern obtained by the look-up table is detected by pattern matching, and is removed or replaced by the cancellation / replacement injection logic to perform fuel injection. It is an object to obtain an efficient thruster control method with a smaller total amount.

A ninth aspect of the invention is to facilitate pattern matching in the invention according to the seventh aspect by assigning a code to the component of the reference pattern and performing scaling.

The present invention as set forth in claim 10 is intended to secure functional redundancy against a thruster failure by expanding the replacement injection pattern in the invention as set forth in claim 7.

Further, the invention according to claim 11 solves the on-board linear programming after reducing the required calculation amount, and obtains a combination of thrusters that minimizes the total fuel injection amount. It is an object of the present invention to obtain a thruster control method capable of controlling the thruster.

The present invention according to claim 12 aims to ensure functional redundancy by software against a thruster failure.

Further, the invention as set forth in claim 13 aims to reduce the total weight of the thrusters to be deployed while ensuring the redundancy of the thrusters.

[0021]

A control device for a spacecraft according to a first aspect of the present invention constitutes a position control system based on a deviation between a position target value and an output of a position detector, and similarly, a position target value and an attitude The attitude control system is constructed from the deviation of the detector output, and the interference on the dynamics of the manipulated variables of both is removed by the decoupling calculation.

Further, the control device for a spacecraft according to a second aspect of the present invention further calculates the acceleration of the target value and the inertial force generated on the dynamics by feedforward calculation using the target value and the detector output. It is intended to compensate.

Further, the control device for a spacecraft according to a third aspect of the present invention uses, in the position detector and the attitude detector, a proximity sensor that detects a relative position / relative attitude between the spacecraft and the target, The relative position and relative attitude between the spacecraft and the target are independently controlled.

Further, the control device for a spacecraft according to a fourth aspect of the present invention uses, in the position detector and the attitude detector, a proximity sensor that detects a relative position / relative attitude between the spacecraft and the target, Also, the angular velocity of the spacecraft is detected by the gyro, and the relative attitude and the angular velocity of the target are estimated by the output of the proximity sensor and the gyro, and the results are used for position / attitude control calculation and feedforward calculation, and The relative position and relative attitude with respect to the target are controlled.

According to a fifth aspect of the present invention, there is provided a control device for a spacecraft, wherein a distance between a spacecraft and a target (hereinafter referred to as a range) and an angle at which the spacecraft looks at the target (hereinafter, referred to as a range) in a position detector. , LOS angle), and a attitude detector that detects the attitude of the spacecraft with respect to the inertial space. The range between the spacecraft and the target and the LOS angle and the attitude of the spacecraft in the inertial space are used. And are controlled independently.

According to a sixth aspect of the present invention, there is provided a spacecraft control device, wherein a position detector includes a sensor for detecting a range and a LOS angle between the spacecraft and a target, and the attitude detector includes a spacecraft. By using a sensor that detects the attitude with respect to the inertial space, and by giving a attitude that rotates at a constant angular velocity with respect to the inertial space as a target attitude value, and independently controlling the position and attitude, the spacecraft It is controlled so as to fly around the target while facing the target on a certain surface.

According to a seventh aspect of the present invention, there is provided a spacecraft control device which uses a sensor for detecting a range and a LOS angle between a spacecraft and a target as a position detector and a spacecraft as an attitude detector. By using a sensor that detects the attitude with respect to the inertial space of the spacecraft, and by providing the attitude target value such that the attitude angular velocity fluctuates at a certain ratio with respect to the inertial space, and by controlling the position and attitude independently, It is controlled so as to fly around the target while facing the target on a surface having an inertial space, and the orbiting surface rotates in the inertial space.

Further, in the thruster control method according to the invention described in claim 8, a pattern in which waste occurs among the injection patterns obtained by the look-up table method is matched with the prepared offset injection pattern or replacement injection pattern. The useless pattern is detected by applying the offsetting / replacement injection logic, and as a result, the total injection amount is reduced.

According to the ninth aspect of the thruster control method of the present invention, the reference pattern component is given the same sign in the case of the offset injection pattern and a different sign in the case of the replacement injection pattern, and The scaling is performed for each pattern.

In the thruster control method according to the tenth aspect of the present invention, the pattern for replacing one thruster with another thruster in the replacement injection pattern is a replacement injection pattern for failure.

Further, in the thruster control method according to the invention described in claim 11, the 6-axis component is not simultaneously solved, but the 3-axis component +3 so that the linear programming can be executed onboard.
It is designed to solve by dividing into axial components and reducing the amount of calculation.

According to the twelfth aspect of the invention, the thruster control method provides a look-up table corresponding to a thruster failure in addition to the normal look-up table.

The thruster control method according to a thirteenth aspect of the present invention uses the degree of difficulty of replacing each thruster with another remaining thruster as an index, and sets the redundant number of thrusters to be large or small. .

[0034]

In the spacecraft control device according to the present invention, the position is the center of mass of the spacecraft by constructing the position control system from the deviation between the target position value and the output of the position detector. Eliminate the need, and likewise for posture,
By constructing the attitude control system from the deviation between the attitude target value and the attitude detector output, if the position control system does not control the mass center position of the spacecraft, the operation amount of the position control and the operation amount of the attitude control are dynamics. Although there is interference, the position control system and the attitude control system ignore the interference and are configured independently, and the interference component is removed by decoupling calculation, which is not the case when the center of mass position is targeted. Also, the position control system and the attitude control system can be independently configured, and the degree of freedom in selecting the control amount of the position control system is increased and the influence of the detector error on the control amount can be reduced.

Further, in the spacecraft control device according to the second aspect of the present invention, the feedforward calculation unit compensates for the change in the target value and the influence of the inertial force generated on the dynamics. Improve control system characteristics such as tracking characteristics.

Further, the control device for a spacecraft according to a third aspect of the present invention detects the relative position and relative attitude between the spacecraft and the target by the proximity sensor, and detects the deviation between the target value and the detector output. By constructing a control system independent of position and attitude to eliminate interference on dynamics by decoupling calculation, and by compensating for the change in target value and the influence of inertial force generated on dynamics with the feedforward calculation unit, the universe It is possible to precisely control the relative 6 degrees of freedom between the machine and the target.

Further, the control device for a spacecraft according to a fourth aspect of the present invention detects the relative position and relative attitude between the spacecraft and the target by the proximity sensor, and detects the deviation between the target value and the detector output. A control system is configured independently of position and attitude to eliminate interference on dynamics by decoupling calculation, and the angular velocity of the spacecraft is detected by a gyro. To control the relative 6 degrees of freedom between the spacecraft and the target more accurately by estimating the attitude and angular velocity of the target, and using the output of the estimator that performs the estimation for control calculation and feedforward calculation. The estimator also reduces the effect of noise contained in the proximity sensor whose output is also the relative orientation.

Further, in the spacecraft control device according to a fifth aspect of the invention, a sensor for detecting the range and the LOS angle between the spacecraft and the target is used for the position detector, and the spacecraft is used for the attitude detector. By using a sensor that detects the attitude of the spacecraft with respect to the inertial space, the LOS angle of the spacecraft with respect to the target is output from the position detector.
The angle can be controlled independently, and for example, the spacecraft can be controlled to face the target and approach the target while keeping the attitude with respect to the inertial space constant.

Further, the control device for a spacecraft according to a sixth aspect of the present invention provides a posture for rotating the spacecraft at a constant angular velocity with respect to the inertial space as a target value of the posture of the spacecraft. By controlling the position control so that the range and LOS angle with respect to the target will be constant, the spacecraft will always face the target while maintaining a constant distance from the target and rotating its attitude with respect to the inertial space. It is possible to control, and as a result, the spacecraft orbits around the target so that the orbiting surface is constant in inertial space.

Further, the control device for a spacecraft according to a seventh aspect of the present invention provides a posture in which the posture angular velocity fluctuates at a certain ratio with respect to the inertial space as a target value of the posture of the spacecraft. By controlling the position control of the spacecraft so that the range and LOS angle with respect to the target are constant, the spacecraft always faces the target, the distance from the target is constant, and the attitude angular velocity vector rotates with respect to the inertial space. It is possible to control
As a result, the spacecraft flies around the target such that its orbiting surface rotates in inertial space, and orbits the target in its entirety.

The thruster selection rule in the invention described in claim 8 is a relatively small calculation by reducing the useless portion in the thruster distribution obtained by the table lookup method by the offsetting / replacement injection logic. A method is provided for achieving a quantity-efficient thruster distribution.

According to the thruster selection rule in the invention described in claim 9, a plurality of canceling injection patterns or replacement injection patterns are sequentially matched with the injection pattern obtained by the table lookup method, and the reference pattern code is obtained. It is possible to obtain the goodness of fit by the same algorithm by obtaining the goodness of fit according to the difference between the cancellation and the substitution included. Since each pattern is appropriately scaled, the degree of conformity directly represents the decrease in the total injection amount.

According to the thruster selection rule in the tenth aspect of the present invention, the replacement injection logic is preferentially applied to the thruster that has failed at the time of thruster failure, and after the replacement, the offset / replacement injection that does not include the failed thruster. Applying logic also reduces the total thruster firing.

According to the thruster selection rule in the eleventh aspect of the present invention, the 6-axis component is divided into the 3-axis component and the 3-axis component to reduce the calculation amount, thereby enabling the on-board execution of the linear programming method. To do.

According to the thruster selection rule of the invention described in claim 12, in addition to the normal look-up table, a look-up table at the time of failure is provided, so that the function redundancy of the thruster by software against the failure of the thruster is provided. Secure.

According to the thirteenth aspect of the present invention, in the case where each thruster cannot be replaced with the other remaining thrusters, the redundant number of thrusters is preferentially assigned, and conversely, a certain thruster is changed to another. When the amount of increase in the total injection amount is small when replaced with the remaining thrusters, the allocation of the redundant number of the thrusters is reduced as necessary, and the redundant configuration of the thrusters is performed according to the need. Reduce the total weight of deployed thrusters while ensuring thruster redundancy.

[0047]

【Example】

Example 1. Embodiment 1 of the present invention will be described below with reference to the drawings. FIG. 1 is a block diagram showing an embodiment of the invention described in claims 1 and 2. In the figure, 1 is a spacecraft body, 2 is a position detector, 3 is a position target value generator, 4 is a position control calculator, 5 is a posture detector, 6 is a posture target value generator, and 7 is a posture control calculator. 23, which are the same as or equivalent to those of the related art denoted by the same reference numerals in FIG. 23, detailed description thereof will be omitted. Further, 8 is a decoupling calculation unit that removes the dynamic interference between the position control operation amount output from the position control calculation unit 4 and the posture control operation amount output from the posture control calculation unit 7, and 9 indicates the position. A feed for obtaining the compensation amount for the acceleration of the target value based on the outputs of the target value generating unit 3 and the posture target value generating unit 6 and the inertial force compensation amount based on the outputs of the position detector 2 and the posture detector 5. It is a forward calculation unit. The spacecraft body 1 is composed of the thruster selector 16, the thruster modulator 17, and the thruster 18 shown in FIG. 24, and the spacecraft dynamics 19.

Next, the operation will be described. First, the equation of motion of the spacecraft body 1 is expressed by the following equation (1).

M *** x + d = u _g (1)

Here, M is a 6 × 6 mass matrix of the spacecraft body 1, x is a sixth-order variable consisting of three position variables and three attitude variables, ** is a symbol indicating the second derivative with respect to time, and d Is an inertial force term such as centrifugal force or Coriolis force, u is a control input such as a thruster for a spacecraft, and these x, d and u _g are vector quantities. If the position variable at x is the center of mass position of the spacecraft body 1 with respect to the target, M will be a block diagonal matrix represented by the following equation (2).

[0051]

[Equation 1]

That is, since the position and orientation are made non-interfering in terms of dynamics, the position and orientation can be easily controlled independently. On the other hand, if the position variable at x is not the center of mass of the spacecraft body 1, M does not have such a block diagonal matrix, but the position and orientation interfere. For example, the operation amount of the attitude control affects the control amount of the position control.

In order to deal with this problem, consider a virtual input v instead of the control input u and set it as the following equation (3). This virtual input v is also a vector quantity.

U = M · v + d (3)

By substituting the equation (3) into the equation (1), the following equation (4) is obtained, and the position and orientation are decoupled under the virtual input v.

** x = v (4)

In this way, the position and attitude control systems are independently configured for the decoupling dynamics as in the equation (4), and the actual control input u is set to the obtained virtual control input v. Just ask. The target value for the variable x is now x _h
Assuming (vector amount), the virtual control input v may be given by, for example, the following expression (5).

V = ** x _h + k _d (* x _h − * x) + k _p (x _h −x) (5)

Where * represents the first derivative with respect to time,
k _d is a speed control gain, and k _p is a position control gain.
By substituting the equation (5) into the equation (3), the actual control input u is given by the following equation (6).

[0060] _{u = M {k d (*} x h - * x) + k p (x h -x)} + (M · ** x h + d) (6)

In FIG. 1, the position detector 2 detects the position variable of x, and the position target value generator 3 generates a target value for the position variable. In the position control calculation unit 4, from the deviation between the position target value and the output of the position detector 2,
The three components relating to the position control of the virtual control input v are determined with respect to the dynamics of the quadratic integral system as in the equation (4).
The function of the position control calculation unit 4 is equivalent to the calculation within {} of the expression (6), for example. This also applies to the posture. In FIG. 1, the posture detector 5 detects the posture variable x, and the posture target value generator 6 generates a target value for the posture variable. The attitude control calculation unit 7 determines three components relating to the attitude control of the virtual control input v from the deviation between the attitude target value and the attitude detector 5 with respect to the dynamics of the quadratic integral system as shown in equation (4). The function of the attitude control calculation unit 7 corresponds to the calculation in {} of the expression (6), for example.

The decoupling calculator 8 has a function of decoupling the output of the position control calculator 4 and the output of the attitude control calculator 7 in terms of dynamics. This corresponds to the operation of multiplying the calculation amount in {} by the mass matrix M from the left in the equation (6). Further, the feedforward calculation unit 9 obtains a compensation term associated with the acceleration of the target value from the outputs of the position target value generation unit 3 and the posture target value generation unit 6. This calculation corresponds to the calculation of M · ** x _{h in} the equation (6).
The feedforward calculation unit 9 also obtains the inertial force term on the dynamics from the outputs of the position detector 2 and the attitude detector 5. This calculation corresponds to the calculation of d in equation (6). That is, the feedforward calculation unit 9 obtains the second term on the right side of the equation (6).

As described above, even when the center of mass of the spacecraft body 1 is not used as the position variable, the position and attitude can be controlled independently. The obtained actual control input u _g may be passed through a low-pass filter in order to reduce the control gain in the high frequency range. If the spacecraft body 1 and the target value do not move very fast, the feedforward calculator 9 can be omitted.

Example 2. Next, a second embodiment of the present invention will be described with reference to the drawings. 2 is a block diagram showing an embodiment of the invention described in claim 3, and the portions corresponding to those in FIG. 1 are designated by the same reference numerals and the description thereof will be omitted. In FIG. 2, 10 is a proximity sensor that detects a relative position or a relative attitude with respect to the target, and 11 is a gyro that detects the angular velocity of the spacecraft body 1.

Next, the operation will be described. Spacecraft body 1
Detects the relative position and relative attitude with respect to the target by the proximity sensor 10. The position control calculation unit 4 and the posture control calculation unit 7 perform, for example, PID (proportional / derivative / integral) control on the target values of the position and the posture obtained from the position target value generation unit 3 and the posture target value generation unit 6, respectively. . Each of the computing units 4, 7, 8 and 9 needs a time differential value of the relative position and the relative posture, which can be approximately obtained by pseudo differentiation or the like.

The relative position of the proximity sensor 10 with respect to the target is not the relative position of the center of mass of the spacecraft body 1 but the attachment point of the proximity sensor 10. Therefore, the dynamics of the spacecraft body 1 expressed by the relative position and the relative attitude interfere with each other. This interference can be removed by inputting the outputs of the position control calculation unit 4 and the attitude control calculation unit 6 to the decoupling calculation unit 8 as described above.

Further, the angular velocity of the spacecraft body 1 is set to the gyro 11
Detected and input to the feed-forward calculation unit 9. This value is used to calculate the inertial force on the dynamics. Further, the output of the proximity sensor 10 and the outputs of the position target value generation unit 3 and the attitude target value generation unit 6 are input to the feedforward calculation unit 9 to perform calculation for compensating the acceleration amount of the target value and the inertial force on the dynamics. To do. By adding this result to the output of the decoupling calculator 8, a control input to the spacecraft body 1 can be obtained.

As a result, the spacecraft body 1 can accurately control the relative position and relative attitude with respect to the target. For example, the spacecraft body 1 accurately follows the target.
Such control is necessary when the spacecraft body 1 docks to the target by rendezvous. In addition, in order to reduce the control gain in the high frequency range, the obtained control input to the spacecraft body 1 may be passed through a low-pass filter. Further, when the movement of the target is not so fast, the feedforward calculation unit 9 can be omitted.

Example 3. Next, a third embodiment of the present invention will be described with reference to the drawings. FIG. 3 is a block diagram showing an embodiment of the invention described in claim 4, and the portions corresponding to those in FIG. 2 are designated by the same reference numerals and the description thereof will be omitted. In FIG. 3, reference numeral 12 is an estimator that estimates the relative attitude with respect to the target and the angular velocity of the target from the outputs of the proximity sensor 10 and the gyro 11.

Next, the operation will be described. The estimator 12 estimates the relative attitude with respect to the target and the angular velocity of the target from the relative attitude with respect to the target (including noise) output from the proximity sensor 10 and the angular velocity of the spacecraft body 1 output from the gyro 11. But
This is based on the following principle. That is, if the target is moving freely, the angular velocity of the target can be regarded as a substantially constant value. 1 which is the output of the proximity sensor intended for freedom relative orientation of the orientation for now easy q _1, the angular velocity of the target q _2, the gyro output and u _g. When the angular velocity of the target is added to the time derivative of the relative attitude, the angular velocity of the spacecraft body 1 is obtained, so the following equation (7) is established.

* Q ₁ + q ₂ = u _g (7)

On the other hand, since the angular velocity of the target is almost constant with respect to time, the following equation (8) approximately holds.

* Q ₂ = 0 (8)

These two expressions can be collectively expressed as the following (9).
It becomes an expression.

[0075]

[Equation 2]

Here, the relative attitude q ₁ is determined by the proximity sensor 1
Since the output of 0 is known, the following equation (10) is obtained as the output equation for the above equation (9).

[0077]

[Equation 3]

However, the noise term included in the output is omitted here. Since the expressions (9) and (10) are observable, the observer and the Kalman filter can be configured based on the expression (10). Now, q ₁ and q ₂ are collectively referred to as q, and the following equation (11) is used. In this case, q
Is a vector quantity, and A, B, and C are matrices.

* Q = A · q + B · u, y = C · q (11)

At this time, the observer is expressed as follows.

* Z = (A−E · C) z + E · y + B · u _g (12) Here, z is a quadratic vector and is an estimated value of q.
E is a matrix of observer gains, and AEC is selected so as to be a stable matrix. Since q is composed of the relative attitude q ₁ with respect to the target and the target angular velocity q ₂ as described above, the relative attitude and the target angular velocity can be estimated by z in the equation (12). This is the function of the estimator 12.

The relative attitude with respect to the target estimated in this way reduces the influence of noise contained in the output of the proximity sensor 10. Further, since the angular velocity of the target can be estimated, the time differential value of the relative attitude can be estimated from this estimated value and the output of the gyro 11.

The output of the estimator 12 and the output of the gyro 11 are input to the attitude control calculation unit 7 and the feedforward calculation unit 9. The attitude control calculation unit 7 performs feedback control calculation of attitude control from the attitude target value output from the attitude target value generation unit 6 and these amounts. Regarding position control, as in the case of the second embodiment, the position control calculation unit 4 performs feedback control calculation from the deviation between the output of the position target value generation unit 3 and the relative position output of the proximity sensor 10. The output of the position control calculation unit 4 and the output of the attitude control calculation unit 7 are input to the decoupling calculation unit 8 to remove interference on the dynamics.

Further, the feedforward calculation unit 9 performs a calculation for compensating the acceleration component of the target value and the inertial force on the dynamics, and by adding the result to the output of the decoupling calculation unit 8, the spacecraft body 1 Get control input for. In this case, compared to the second embodiment, when the noise of the proximity sensor 10 becomes a problem, the estimation accuracy of the relative attitude with respect to the target and the target angular velocity is improved, and thus the final control accuracy can be improved. it can. In order to reduce the control gain in the high frequency range, the obtained control input to the spacecraft body 1 may be passed through a low pass filter. Further, when the movement of the target is not so fast, the feedforward calculation unit 9 can be omitted.

Example 4. Next, a fourth embodiment of the present invention will be described with reference to the drawings. FIG. 4 is a block diagram showing an embodiment of the invention described in claim 5, and the portions corresponding to those in FIG. 2 are designated by the same reference numerals and the description thereof will be omitted. In FIG. 4, 13 is an attitude estimator that estimates the attitude of the spacecraft body 1 with respect to the inertial space from the output of the gyro 11.

Next, the operation will be described. Proximity sensor 1
At 0, the relative distance (range) to the target and the angle at which the target is seen (called the LOS angle and having two components) are detected. A position control system is constructed by regarding these quantities as position control quantities. That is, the target position value generator 3 generates target values for these control amounts, and the position control calculator 4 calculates feedback control from the deviation from the target values.

On the other hand, the attitude control uses the gyro 11 to control the attitude with respect to the inertial space. Such a control is also possible because the attitude with respect to the inertial space and the position control amount are independent. Therefore, the output of the gyro 11 (spacecraft body 1
Input angular velocity) to the posture estimator 13 to estimate the posture with respect to the inertial space. This estimation is in principle as follows. Now, let U be the 3 × 3 direction cosine matrix with respect to the inertial space of the spacecraft body 1, and let ω be the angular velocity of the spacecraft body 1. U
Since there is a relationship of the following expression (13) between the time derivative * U of ω and ω, the posture with respect to the inertial space can be assumed based on this expression (13). Note that ω is a vector amount, and “×” in the equation (13) means that vector product calculation is performed for each column.

* U = ω × U (13)

When the attitude and the angular velocity of the spacecraft body 1 are obtained, the attitude control calculation unit 7 calculates the feedback control from the deviation between them and the output of the attitude target value generation unit 6.

The output of the position control calculation unit 4 and the output of the attitude control calculation unit 7 are input to the decoupling calculation unit 8 and interference on dynamics is removed. Further, the feedforward calculation unit 9 performs a calculation for compensating for the acceleration component of the target value and the inertial force on the dynamics, and by adding the result to the output of the decoupling calculation unit 8, the control input to the spacecraft body 1 is performed. To get

As a result, the spacecraft body 1 can accurately control the range with respect to the target, the LOS angle, and the attitude with respect to the inertial space. Such control is performed when the spacecraft body 1 approaches the target from a relatively distant location,
It is necessary when flying around the target.

Although the proximity sensor 10 detects the range and the LOS angle with respect to the target in the fourth embodiment, another sensor (for example, a rendezvous radar) may detect these amounts. Although the case where the attitude with respect to the inertial space is estimated from the output of the gyro 11 is shown, the attitude with respect to the inertial space may be directly detected by the earth sensor or the sun sensor. Further, in order to reduce the control gain in the high frequency range, the obtained control input to the spacecraft body 1 may be passed through a low pass filter. Further, when the movement of the target is not so fast, the feedforward calculation unit 9 can be omitted.

Example 5. Next, a fifth embodiment of the present invention will be described with reference to the drawings. FIG. 5 shows an embodiment of the invention described in claim 6, and in the above-mentioned Embodiment 4, the attitude target value when the spacecraft body 1 orbits around the target on the surface having the inertial space. 9 is a flowchart when the generation unit 6 is realized by software inside a computer.

Next, the operation will be described. First, in step ST1, the value of the target angular velocity of the spacecraft body 1 (this is ω _d ) is set. This target angular velocity ω _d is a vector quantity and may be set to have a constant value in the inertial space. Then target attitude of the spacecraft main body 1 in step ST2 sets the initial value of (referred to as U _d). This target posture U
_d is a matrix similar to the attitude U of the spacecraft body 1, and the attitude in the inertial space of the spacecraft body 1 at that time is used as its initial value. Next, in step ST3, the target attitude U _d of the spacecraft body 1 is updated. For this updating, the relational expression shown in the following Expression (14) similar to Expression (13) is used.

* U _d = ω _d × U _d (14)

Next, in step ST4, the target attitude U _d of the spacecraft body 1 is output. Although U _d has nine components, the attitude control calculation unit 7 calculates the attitude control by obtaining three independent component deviations from the target attitude U _d of the spacecraft body 1 and the attitude U of the spacecraft body 1. It will be. Next, in step ST5, the time is updated and the process returns to step ST3.

If the target angular velocity and the target attitude of the spacecraft main body 1 are output by the attitude target value generator 6 in this manner, the attitude control system causes the spacecraft main body 1 to have a constant angular speed with respect to the inertial space. It is controlled to rotate with. on the other hand,
The position target value generator 3 generates a target value so that the range and the LOS angle with respect to the target of the spacecraft body 1 are constant, and feedback control is performed by the position control system.

As a result, control can be realized in which the spacecraft body 1 orbits the target on a surface having an inertial space while keeping the distance and the direction to the target constant. At that time, a sensor (for example, an acceleration sensor and its integrator) that detects the position of the spacecraft body 1 in the inertial space is not required.

Example 6. Next, a sixth embodiment of the present invention will be described with reference to the drawings. FIG. 6 shows an embodiment of the invention described in claim 7. In the embodiment 4, the spacecraft body 1 makes an orbit around the target, and the orbiting surface rotates in the inertial space. 6 is a flowchart when the posture target value generator 6 for performing the above control is realized by software inside the computer.

Next, the operation will be described. First, at step ST6, the rotational angular velocity of the orbiting surface of the spacecraft body 1 (
_r )) is set. The rotational angular velocity ω _r of the orbiting surface is a vector quantity and may be set so as to have a constant value in the inertial space. Next, in step ST7, an initial value of the target angular velocity of the spacecraft body 1 (this is ω _d ) is set. This target angular velocity ω _{d is} also a vector quantity, and the initial value thereof is the angular velocity of the spacecraft body 1 at that time. Then target attitude of the spacecraft main body 1 in step ST8 to set the initial value of (referred to as U _d). This target attitude U _d is a matrix similar to the attitude U of the spacecraft body 1, and the attitude in the inertial space of the spacecraft body 1 at that time is used as its initial value. Next, in step ST9, the target angular velocity ω of the spacecraft body 1
Update _d . The following formula (15) is used for this update.

* Ω _d = ω _r × ω _d (15)

However, in the above equation, * ω _d represents the time derivative of ω _{d in} the inertial space, and x means the vector product of both. Next, in step ST10, the target attitude U _d of the spacecraft body 1 is updated. The formula (14) described in the fifth embodiment is used for this updating. Next, in step ST11, the target angular velocity ω _d of the spacecraft body 1 is output, and in step ST12
Outputs the target attitude U _d of the spacecraft body 1. Although U _d has nine components, the attitude control calculation unit 7 calculates the attitude control by obtaining three independent component deviations from the target attitude U _d of the spacecraft body 1 and the attitude U of the spacecraft body 1. It will be.
Next, in step ST13, the time is updated and step ST9
Return to.

When the attitude target value generator 6 outputs the target angular velocity and the target attitude of the spacecraft body 1 as described above, the spacecraft body 1 is caused to rotate in the inertial space by the action of the attitude control system. Controlled by. On the other hand, the position target value generation unit 3 generates a target value so that the range and the LOS angle with respect to the target of the spacecraft body 1 become constant, and the position control system performs feedback control.

As a result, control can be realized in which the spacecraft body 1 orbits around the target while keeping the distance and direction to the target constant, and the orbiting surface rotates at a constant angular velocity in the inertial space. That is, the body 1 of the spacecraft can orbit the entire target. At that time, a sensor (for example, an acceleration sensor and its integrator) that detects the position of the spacecraft body 1 in the inertial space is not required.

Example 7. Next, a seventh embodiment of the present invention will be described with reference to the drawings. FIG. 7 is a block diagram showing a thruster control method according to an embodiment of the present invention described in claim 8. The corresponding parts are designated by the same reference numerals as those in FIG. 1 and their description is omitted. In the figure, reference numeral 23 denotes the position target value generator 3 shown in FIG.
Position control calculation unit 4, posture target value generation unit 6, posture control calculation unit 7, decoupling calculation unit 8, feedforward calculation unit 9
Is a control unit that executes the control law, and 24 is a navigation sensor such as the position sensor 2 and the attitude detector 5 in FIG. Further, FIG. 8 is an explanatory diagram showing a thruster selection rule when performing thruster distribution in the thruster selection unit 16, and in this case, pattern matching to which a cancellation / replacement injection pattern is applied is used.

Next, the operation will be described. Navigation sensor 2
4 is the position of the spacecraft body 1 using the proximity sensor 10, etc.
Attitude, velocity, and angular velocity are estimated, and the estimation results are output as navigation signals. The control unit 23 compares the navigation signal received from the navigation sensor 24 with the guidance signal to determine the control force / torque required for the position, attitude, velocity and angular velocity of the spacecraft body 1 to have desired values. The control signal b is output as the control signal b to the thruster selection unit 16. The control signal b is a vector quantity. When the thruster selection unit 16 receives the control signal b from the control unit 23, it performs thruster distribution according to the thruster selection rule and outputs the duty ratio vector x to the thruster modulator 17 so as to generate the necessary control force / torque as the resultant force. . Here, since each thruster 18 can generate only a pulse-like force, the thruster modulator 17 adjusts the density of the injection interval of the thruster 18 by opening and closing the thruster valve to generate a thruster amount corresponding to the duty ratio vector x.

Next, the internal operation of the thruster selecting section 16 will be described. First, the control signal b is received from the control unit 23, and the duty ratio vector x is obtained according to the look-up table. This x is "x before rule application"
Is.

Next, the cancellation / substitution injection pattern created in advance at the time of design is applied. When the canceling injection pattern is included in x, as shown by the downward-sloping hatching in FIG. 8, the part that can be matched is deleted (= canceling injection logic). When the replacement injection pattern is included in x, as shown by hatching to the left in FIG.
The matching part is replaced with another pattern having a smaller injection amount (= replacement injection logic). After such a procedure is repeated until there are no more applicable patterns, or after repeating an appropriate number of times, the obtained x is “x after application of the rule”.

In the seventh embodiment, thruster 1 is used for the sake of explanation.
Although the number of 8's is 14 and the pattern matching is performed twice, the number of the thrusters 18 and the number of pattern matching are not particularly limited. Further, here, a lookup table is used to obtain "x before applying rule", but the method for obtaining "x before applying rule" is not particularly limited.

Example 8. Next, an eighth embodiment of the present invention will be described with reference to the drawings. FIG. 9 is a flowchart showing an algorithm for matching offset / substitution injection patterns in a thruster control method according to an embodiment of the present invention. Step ST21 to step ST3
3 is each block in the flow chart, and step S
T21 is the start point and step ST33 is the end point.

Next, the operation will be described. Here, the total number of patterns is m, and the number of thrusters 18 is n. First, in step ST21, initial setting is performed, and the pattern number k for matching is set to 0. Next, in step ST22, the pattern number k is incremented by 1. In step ST23, it is determined whether or not matching of all patterns has been tried. If k> m, the process proceeds to step ST31, and if k ≦ m, the process proceeds to step ST24. In step ST24, an initial value is set to check whether or not the k-th offset / replacement injection pattern λ _k can be applied to x. That is, the conformity α of λ _k to x
_The element j that matches _k is set to 0. Note that this λ
_k is a vector quantity.

Next, in step ST25, the element number j for matching is incremented by 1, and then step ST26.
Proceed to and determine whether or not all elements of λ _k have been matched. If j ≦ n, the process proceeds to step ST27, and if j> n, the process returns to step ST22. Step ST
In 27, the lambda _k of j-th element lambda _k (j) determines the positive or not. If it is positive, it is necessary for the j-th element of x to be positive in order to apply the k-th offset / replacement injection pattern λ _k , so the process proceeds to step ST28 to judge it, and otherwise. Is step ST25
Return to. In step ST28, the j-th element x of x
It is determined whether or not (j) is correct. If it is positive, the process proceeds to step ST29, and if it is not positive, the k-th cancellation / replacement injection pattern λ _k cannot be matched with x, so the process returns to step ST22 to try the next rule.

Next, in steps ST29 and ST30, the fitness α _k of the k-th offsetting / substitution injection pattern λ _k to x is determined. α _k is the largest positive number in which all elements satisfy non-zero when x−α _k λ _k is calculated. Algorithmically, all j satisfying λ _k (j) ≠ 0
Is equivalent to setting the minimum value of the ratio of x (j) to λ _k (j) as α _k . Also, λ _k
Since scaling is performed so that the value obtained by adding j from 1 to n in (j) is 1, the value of the fitness α _k is expected when the offsetting / substitution injection pattern λ _k is applied to x. It is equal to the reduction of the total injection amount.

In step ST31, k = k _max giving the maximum value of α _k is calculated. This corresponds to finding a pattern that maximizes the decrease in the total injection amount. Next, in step ST32, the k _max th offset / replacement injection pattern is applied to x to reduce the total injection amount. Then, it progresses to step ST33 and complete | finishes a series of processes.

As described above, there are two types of injection patterns, offset and replacement, but when the pattern is described, the sign is used to distinguish between the offset and replacement patterns. Therefore, when performing pattern matching. , Can be performed in the same way without any computational distinction between cancellation and permutation. In FIG. 9, the pattern matching is set to 1 for the sake of explanation.
It has been repeated, but it does not limit the number of times. Considering the computer load and the like, a method of performing pattern matching by fixing the number of times such as 2-3 times, a method of performing pattern matching until the expected reduction amount of the total injection amount becomes a certain threshold value or less, etc. This is a modification of the eighth embodiment.

Example 9. Next, a ninth embodiment of the present invention will be described with reference to the drawings. FIG. 10 is a flowchart showing an algorithm for extracting a cancellation / substitution injection pattern in the thruster control method according to the embodiment of the invention described in claim 9. Steps ST41 to ST55 are blocks in the flowchart, with step ST41 being the start point and step ST55 being the end point.

Next, the operation will be described. Here again, the number of thrusters 18 is n. First, in step ST41, the initial value of the combination number m of the thrusters 18 is set.
Next, in step ST42, m is selected from the n thrusters 18.
After setting the initial value when considering the combination of the thrusters 18 of the book and setting the pattern number k to 0, step ST
Do not allow duplication from n thrusters 18 at 43 m
Make a combination of books. In addition, the selected thruster 18
The number _{{i 1, i 2, ···} , i m} and.

Next, in step ST44, the set of thrusters 18 selected in step ST43 {i ₁ , i ₂ , ...
.., i _m } is checked to see if it includes the set of thrusters 18 already registered in the list as the offset / replacement injection pattern. If included, the process proceeds to step ST52, and if not included, the process proceeds to step ST45. At step ST45, each thruster {i ₁ , i ₂ , ..., i
Vectors a _i1 , a representing the control force / torque generated by _m }
_i2, · · ·, side by side a _im constituting the thruster matrix A _m. The dimension of this thruster matrix A _m is 6 × m.

Next, in step ST46, the eigenvalues of the thruster matrix A _m are obtained, and then the process proceeds to step ST47. In step ST47, it is determined whether or not 0 exists in the eigenvalue of A _m obtained in step ST46. If 0 is present, the process proceeds to step ST48, and if 0 is not present, the process proceeds to step ST52. Step ST4
At 8, the eigenvector λ ₀ corresponding to the 0 eigenvalue is obtained.

Here, the fact that the injection pattern of the thruster 18 has "waste" is equivalent to that the "useless" set of thrusters 18 does not generate a new control force / torque component. It means that the combination of such thrusters 18 is in a first-order dependent relationship. Therefore, when the thrusters 18 have a first-order dependency relationship, the following (1
6) A coefficient vector e (e ₁ , e ₂ , ...
,, e _m ) exists.

E ₁ a _i1 + e ₂ a _i2 + ... + e _{m aim} = 0 (16)

Here, the coefficient vector e is a duty ratio vector representing a cancellation / substitution injection pattern. On the other hand, this is also the definition of the 0 eigenvector. That is, for the 0 eigenvector λ ₀ of the thruster matrix A _m, the following (1
Equation 7) holds.

A _m λ ₀ = 0 (17)

Expanding this equation (17), the following (18)
The equation is then obtained, so the 0 eigenvector and e are the same except for scheduling.

Λ ₀ (1) a _i1 + λ ₀ (2) a _i2 + ... + λ ₀ (m) _aim = 0 (18)

Next, in step ST49, it is determined whether the eigenvector λ ₀ corresponds to the cancellation injection pattern or the replacement injection pattern with the sign of the non-zero component.
If all non-zero components of λ ₀ have the same sign, the process proceeds to step ST50, and if they do not have the same sign, the process proceeds to step ST51. In steps ST50 and ST51, the following (1
Schedule λ _{0 according} to equation 9).

[0127]

[Equation 4]

Next, in step ST52, it is determined whether or not all combinations nCm in the case of selecting m thrusters 18 have been checked. If the check is completed, the procedure proceeds to step ST53, and if not completed, the step ST53
Return to 43. In step ST53, the number m of combinations of thrusters 18 is increased by 1, and then the process proceeds to step ST54. In step ST54, the size relationship between the number m of combinations of the thrusters 18 and the number of components 6 is compared. If m is 6 or less, the process returns to step ST42 to continue the pattern check. When m exceeds 6, the process proceeds to step ST55 to end the series of processes.

Example 10. Next, Example 10 of the present invention
Will be described with reference to FIG. FIG. 11 is a flowchart showing an algorithm for matching offset / substitution injection patterns in a thruster control method according to another embodiment of the present invention. Steps ST21 to ST3
0 and steps ST32 to ST35 are blocks in the flowchart, where step ST21 is the start point and step ST33 is the end point. Incidentally, step ST35
Is a database of weighting w for each injection pattern in consideration of the disturbance and contamination of the target due to the residual fuel / injected gas.

Next, the operation will be described. Here, FIG.
The description of the parts common to the above is omitted. In step ST34, k = k _max is _{obtained when} the maximum value of w _k × α _k is given. By using such a weighting w _k , more desirable thruster 1 is taken into consideration not only the total injection amount but also the disturbance or pollution of the target due to the residual fuel and the injected gas.
Eight distributions can be realized. As a method of creating w _k , for example, when the use of the thrusters 18 of # 13 and # 14 is not preferable in the sense of increasing the disturbance to the target, the injection amount of the thrusters 18 of # 13 and # 14 is decreased. It is conceivable to increase the weight w _k of the cancellation / replacement rule.

Example 11. Next, Example 11 of the present invention
Will be described with reference to FIG. FIG. 12 is an explanatory diagram showing a thruster selection rule of the thruster selection unit 16 in the thruster control method according to the embodiment of the invention described in claim 10. The thruster selection unit 16 in the eleventh embodiment performs thruster distribution according to the thruster selection rule shown in FIG. 12, which is characterized by the function redundancy of applying the replacement injection pattern when the thruster 18 fails.

Next, the operation will be described. The thruster selection unit 16 refers to the look-up table for the control signal b received from the control unit 23 and generates the duty ratio vector x. This is "x before rule application". Here, for the sake of explanation, it is assumed that the thruster # 13 fails and cannot be used. Since "x before applying rule" is premised on using the thruster of # 13, the replacement injection pattern at the time of failure is applied first. That is, it is replaced with thruster distribution that produces the same control force and torque as the resultant force distributed to the thruster # 13. In FIG. 12, the thruster of # 9 and the thruster of # 12 replace the thruster of # 13. With respect to the thus obtained “x after application of the replacement injection pattern at the time of failure”, the “offset / replacement injection pattern” that does not include the thruster # 13 is applied to reduce the total injection amount. The finally obtained x is output to the thruster modulator 17 as “x after rule application”.

Example 12. Next, Example 12 of the present invention
Will be described with reference to FIG. FIG. 13 is a flowchart showing an algorithm for constructing a replacement injection pattern at the time of failure in the thruster control method according to an embodiment of the invention described in claim 10. Steps ST61 to ST69 are blocks in the flowchart, and step ST61 is the start point and step ST69 is the end point. FIG. 14 is an explanatory diagram showing the concept of creating the matrix A _i in step ST63.

Next, the operation will be described. First, in step ST61, the number i of the thruster 18 is initially set to 1. In this case also, the number of thrusters 18 is n. Further, the value of the weighting vector c is set in proportion to the magnitude of the maximum thrust of each thruster 18. Next, in step ST62, the control force / torque a _i generated when the i-th thruster 18 is injected at a duty ratio of 1
Is substituted into the target control force / torque b. Next, in step ST63, all the elements a _i corresponding to the thruster #i of the thruster matrix A are replaced with 0 as shown in FIG. 14 to form a matrix A _i . This operation corresponds to the inability to generate the control force / torque when the thruster #i fails.

Next, in step ST64, among the combinations of other thrusters that generate the same control force / torque as the thruster of #i as the resultant force, the solution μ _i with the smallest total injection amount is obtained by the linear programming method. Ask. Note that this μ _i
Is a vector quantity. Next, in step ST65,
In order to have the same data format as the replacement / offset injection pattern according to the invention described in claim 9, the sign is inverted and 1 is substituted for the i-th element of μ _i . Then, in step ST66, i is incremented and the process proceeds to the next thruster # (i + 1). In step ST67, all thrusters 18
It is determined whether or not the replacement injection pattern at the time of failure is obtained. If i ≦ n, the process returns to step ST62 and i>
In the case of n, it progresses to step ST68. Step ST6
In 8, the obtained replacement injection pattern at the time of failure {μ ₁ , ...
· After tabling mu _n}, and ends at step ST69.

For the sake of explanation, μ _i corresponds to the injection quantity of the duty ratio 1 of the thruster of failure #i in FIG. 13, but there is also a method of scaling it as in the following expression (20) in step ST68. .

[0137]

[Equation 5]

Example 13 Next, Example 13 of the present invention
Will be described with reference to FIG. FIG. 15 is an explanatory diagram showing a thruster selection rule in the thruster control method according to an embodiment of the invention described in claim 11, and FIG. 16 is an explanatory diagram showing division of the vector b and the thruster matrix A. The thruster selecting unit 16 in the thirteenth embodiment distributes the thrusters 18 according to the thruster selection rule shown in FIG. 15, which is characterized by using a linear programming method during operation.

Next, the operation will be described. The thruster selection unit 16 first sets the control signal b received from the control unit 23 to 2
Divide into _two vectors b ₁ and b ₂ . For the sake of explanation, b ₁
Is the control force 3 component, and b ₂ is the control torque 3 component. Corresponding to the division of the vector b, the thruster matrix A is also divided into two 3 × n dimensional submatrices A ₁ and A ₂ . next,
In each case of b ₁ = A ₁ x ₁ and b ₂ = A ₂ x ₂ , the linear programming method is executed so that the total injection amount is minimized under the constraint that the duty ratio is 0 or more. The solutions x ₁ and x ₂ thus obtained are added together to obtain the final duty ratio vector x.

[0140] As a method of dividing a vector b, the remainder was a b ₁ 3 single Pick in descending order of the absolute value of each component of b and how to b _2, the control precision required for each component of b from weighted by multiplying the inverse of the b ₁ 3 single pick in descending order of the absolute value a method of the remainder and b _2. In addition, x is calculated by linear programming with the three major components being b _1.
1 obtains, seek x ₂ table look up the remaining three components as b _2, there is also a method of combining them. In these cases, as shown in FIG. 16, the vector b
The thruster matrix A is also divided according to.

Example 14 Next, Example 14 of the present invention
Will be described with reference to FIG. FIG. 17 is an explanatory diagram showing a thruster selection rule in a thruster control method according to another embodiment of the invention described in claim 11, and FIG. 18 is an explanatory diagram showing division of the vector b and the thruster matrix A. The thruster selection unit 16 in the fourteenth embodiment distributes the thrusters 18 according to the thruster selection rule shown in FIG. 17, which is characterized in that the linear programming method is used when the thruster fails during operation.

Next, the operation will be described. The thruster selection unit 16 first sets the control signal b received from the control unit 23 to 2
Divide into _two vectors b ₁ and b ₂ . Corresponding to the division of the vector b, the thruster matrix A is also divided into two 3 × n dimensional submatrices A ₁ and A ₂ . However, here A
18 uses a corrected thruster matrix A in which all the elements a _i corresponding to the failed thruster #i are replaced with 0, as shown in FIG.

Next, in each case of b ₁ = A ₁ x ₁ and b ₂ = A ₂ x ₂ , the linear injection is performed so as to minimize the total injection amount under the constraint that the duty ratio is 0 or more. Carry out the planning method. The solution x thus obtained
_The final duty ratio vector x is obtained by adding ₁ and x ₂ . Therefore, in this fourteenth embodiment,
With the configuration similar to that of 3, the distribution of the thrusters 18 that makes the total injection amount suboptimal can be obtained even when the thruster fails.

Example 15. Next, Embodiment 15 of the present invention
Will be described with reference to FIG. FIG. 19 is an explanatory diagram showing an injection pattern database in which the thruster selection unit 16 according to the embodiment of the invention described in claim 12 is incorporated. The thruster selecting unit 16 in the fifteenth embodiment generates the desired control force / torque according to the look-up table corresponding to the thruster failure when the thruster 18 fails during operation. Is provided, and thruster distribution is performed by using it as a thruster selection rule.

Next, the operation will be described. If the thruster of #i fails and cannot be used, the thruster selector 16
Selects the look-up table at the time of thruster failure of #i from the built-in injection pattern database for the control signal b received from the control unit 23 to obtain the desired resultant force.
The thrusters 18 are distributed so as to generate a combined torque.
Further, in FIG. 19, a method of preparing a look-up table at the time of failure for each thruster has been described for the sake of convenience of explanation. Good luck
This method does not significantly increase the required memory capacity.

Example 16. Next, Embodiment 16 of the present invention
Will be described with reference to FIG. FIG. 20 is an explanatory diagram showing an algorithm for creating a look-up table corresponding to each thruster failure of the injection pattern database in the fifteenth embodiment.

Next, the operation will be described. Here, for the sake of explanation, an algorithm for creating a lookup table corresponding to the thruster failure of #k is shown. In this case also, the number of thrusters is n. First, i is initialized to 1,
In the distribution of the thruster 18 that substitutes 1 only for the i-th element of the control signal b and generates a unit amount of force or torque in the positive direction about one axis, linearly so as to minimize the total injection amount. The duty ratio vector μ _i is calculated by the programming method. A is a thruster matrix having a size of 6 × n.
The j-th column vector a _{j of} is composed of 6 components of the control force and the control torque generated when the j-th thruster injects at a duty ratio of 1. In this case, as indicated by the modification of the thruster matrix when the thruster of #k fails, #k
Replace all elements a _k corresponding to the thruster of 0 with 0. Further, the weighting vector c is an n-dimensional vector, and is assumed to be all components 1 here for the sake of explanation.

Next, only the i-th element of the control signal b is -1.
And the duty ratio vector η _i is obtained by linear programming so as to minimize the total injection amount in the thruster distribution that generates a unit amount of force or torque in the negative direction on one axis. Thus, in addition to the operation for obtaining the duty ratio vector μ _i , the duty ratio vector η
_The operation for obtaining _i is necessary because the ejection direction of the thruster 18 is unidirectional rather than bidirectional. Since it is unidirectional, 0 ≦ μ _i and 0 ≦ η _i are constraint conditions.

Here, since the control force / torque have a total of 6 components (12 components when the sign is taken into consideration), the above processing is executed for each axis while incrementing i. When the processing for all the axes is completed, appropriate scaling is performed, and then μ _i and η _i are arranged and the lookup table S
To create. As a scaling method, it is possible to perform scaling so that the generated control force / torque becomes a unit amount. When the thrusts of the thrusters 18 are equal, the weighting vectors c may all be 1, but when the thrusts are different, the vector c is weighted by a size proportional to the maximum thrust.

Example 17 Next, Example 17 of the present invention
Will be described with reference to FIG. FIG. 21 is an explanatory diagram showing a thruster selection rule in a thruster control method according to another embodiment of the present invention. The thruster selecting unit 16 in the seventeenth embodiment is configured to operate the thruster 1 during operation.
In the case where 8 fails, thruster distribution is performed according to the thruster selection rule shown in FIG. 21, which is characterized in that a desired control force / torque is generated according to a look-up table corresponding to the thruster failure.

Next, the operation will be described. The thruster selection unit 16 generates a duty ratio vector x with respect to the control signal b received from the control unit 23 by referring to a normal look-up table. This is "x before replacement". For the sake of explanation, it is assumed that the thruster # 13 is out of order and cannot be used. This "x before replacement" is based on the premise that the thruster of # 13 is used. Therefore, referring to the replacement table (lookup table) at the time of thruster failure of # 13, the control force distributed to the thruster of # 13 is referenced. Replace with a distribution of thrusters 18 that produce the same torque and resultant force. The “x after replacement” thus obtained is output to the thruster modulator 17.

In the seventeenth embodiment, instead of giving up the sub-optimization of the total injection amount, the required memory capacity and the calculation amount are minimized. This is applicable when "Pattern" is not applied.

Example 18. Next, an eighteenth embodiment of the present invention will be described with reference to the drawings. 22 is a flowchart showing an algorithm for increasing or decreasing the redundant number of thrusters 18 according to the thirteenth embodiment of the invention according to the embodiment of the invention. Steps ST71 to ST80 are blocks in the flowchart, with step ST71 being the start point and step ST80 being the end point.

Next, the operation will be described. First, in step ST71, i is set to an initial value of 1. In this case also, the number of thrusters 18 is n. In addition, each thruster 18
The value of the weighting vector c is set in proportion to the magnitude of the maximum thrust. Next, in step ST72, the control force / torque a _i generated when the i-th thruster injects at the duty ratio of 1 is substituted into the target control force / torque b. Next, in step ST73, all the elements a _i corresponding to the thruster #i in the thruster matrix A are replaced with 0 as shown in FIG. 14 to form a matrix A _i . This operation corresponds to the inability to generate the control force / torque when the thruster #i fails.

Next, in step ST74, among the combinations of other thrusters that generate the same control force / torque as the thruster of #i as the resultant force, the solution μ _i with the smallest total injection amount is calculated by the linear programming method. Ask. Then step S
In T75, the replacement according to the invention described in claim 9
In order to have the same data format as the offset injection pattern, the sign is inverted and 1 is substituted for the i-th element of μ _i . Next, in step ST76, i is incremented and the process proceeds to the next thruster 18. In step ST77, it is determined whether or not replacement injection patterns at the time of failure have been obtained for all thrusters 18, and if i ≦ n, the process returns to step ST72.
If i> n, the process proceeds to step ST78.

At step ST78, two parameters are calculated as an index for determining the redundant number of thrusters 18. That is, “increment of total injection amount” and “number of thrusters required for replacement”. This total injection amount increment is μ
_The value of j of _i (j) is added from 1 to n, and the number of thrusters required for replacement is given as the number of j in 1 to n satisfying μ _i (j)> 0. Next, in step ST79, the redundant number is preferentially assigned in the descending order of the indexes obtained in step ST78, and then step S
At T80, a series of processing ends.

As a modification of the eighteenth embodiment, there is a method in which the fact that the difficulty obtained in step ST78 varies little depending on the thruster 18 is used as an index for determining the quality of the placement of the thruster 18.

[0158]

As described above, according to the first aspect of the invention, the position control system is constructed from the deviation between the position target value and the position detector, and similarly, the position control system is formed from the deviation between the attitude target value and the attitude detector. The attitude control system is configured to eliminate the interference of the manipulated variables of the two on the dynamics by decoupling calculation, and by the feedforward calculation from the target value and the detector output, the acceleration of the target value and the inertial force generated on the dynamics. The position control system and the attitude control system can be configured independently even if the control amount of the position control system is not the center of mass position, and the degree of freedom in selecting the control amount of the position control system is increased. There is an effect that the influence of the detector error on the control amount can be reduced as well as the increase.

Further, according to the invention described in claim 2, since the inertial force on the dynamics is compensated in advance by the feedforward calculation, the characteristic of the control system such as the tracking characteristic with respect to the target value is improved. The effect is that you can.

According to the invention described in claim 3, a camera type sensor (proximity sensor) for detecting the relative position and relative attitude between the spacecraft and the target is used for the position detector and the attitude detector. Since the relative position and relative attitude between the spacecraft and the target are independently controlled, there is an effect that the relative 6 degrees of freedom between the spacecraft and the target can be accurately controlled.

According to the invention described in claim 4, a camera type sensor (proximity sensor) for detecting the relative position and relative attitude between the spacecraft and the target is used for the position detector and the attitude detector. Also, the angular velocity of the spacecraft is detected by the gyro, the relative attitude and the angular velocity of the target are estimated by the output of the proximity sensor and the gyro, and the results are used for position / attitude control calculation and feedforward calculation. Since it is configured to control the relative position and relative attitude between the target and the target, the effect of noise can be reduced by the estimator even when the noise of the nearby sensor poses a problem. Relative between 6
The effect is that the degree of freedom can be controlled accurately.

According to the invention described in claim 5, the position detector is provided with a distance (range) between the spacecraft and the target.
And a sensor that detects the angle (LOS angle) at which the spacecraft looks into the target, and a sensor that detects the attitude of the spacecraft with respect to the inertial space, and the range and LOS angle between the spacecraft and the target. Since it is configured to control the attitude of the spacecraft in the inertial space independently, it is possible to accurately execute the control so that the spacecraft faces the target and approaches the target while keeping the attitude with respect to the inertial space constant. There is an effect.

Further, according to the invention described in claim 6, in the control device for a spacecraft according to claim 5, the position target value is set to a position such that it rotates at a constant angular velocity with respect to the inertial space. Since it is configured to control the attitude and attitude independently, it is possible for the spacecraft to fly around the target while facing the target on the surface with inertial space without using an acceleration sensor etc. to detect the position. There is an effect that various controls can be executed accurately.

According to the invention described in claim 7, in the control device for a spacecraft according to claim 5, the attitude target value is set such that the attitude angular velocity fluctuates at a certain ratio with respect to the inertial space. Since the position and attitude are independently controlled, the spacecraft turns around the target while facing the target on the surface with inertial space without using an acceleration sensor or the like to detect the position. In addition, there is an effect that the control such that the orbiting surface rotates in the inertial space can be accurately executed.

According to the eighth aspect of the present invention, in the thruster selection problem in the thruster control method, the duty ratio vector x is corrected in the direction of decreasing the total injection amount by matching the cancellation / replacement injection patterns. Since it is configured, there is an effect that efficient thruster distribution can be realized with a relatively small amount of calculation.

According to the ninth aspect of the invention, since the offset / replacement injection pattern of the thruster selection in the thruster control method is distinguished by the sign and the appropriate scaling is performed, the duty ratio vector and x And has the effect of facilitating pattern matching.

Further, according to the invention of claim 10,
Since the replacement injection logic at the time of thruster failure is applied to the thruster selection rule in the thruster control method, the function redundancy of the thruster can be ensured by software and the increase in total injection amount can be minimized. There is.

According to the invention described in claim 11,
6 in the thruster selection problem in the thruster control method
Since it is configured so that the axial components are not simultaneously solved but divided into three axial components + three axial components, it is possible to execute the linear programming method onboard, and as a result, the total injection amount is small in the sense that There is an effect that an optimal solution can be obtained.

According to the invention described in claim 12,
The thruster selection rule in the thruster control method is configured to prepare a table corresponding to the thruster failure in addition to the lookup table in the normal state. The torque can be easily obtained with a small amount of calculation.

According to the invention described in claim 13,
Since the redundant number is preferentially assigned to a thruster that is difficult to replace with another thruster, the total weight of the thrusters to be deployed can be reduced while ensuring the redundancy of the thrusters.

[Brief description of drawings]

FIG. 1 is a block diagram showing a control device for a spacecraft according to a first embodiment of the present invention.

FIG. 2 is a block diagram showing a control device for a spacecraft according to a second embodiment of the present invention.

FIG. 3 is a block diagram showing a control device for a spacecraft according to a third embodiment of the present invention.

FIG. 4 is a block diagram showing a control device for a spacecraft according to a fourth embodiment of the present invention.

FIG. 5 is a flow chart showing an algorithm of processing of a posture target value generator according to Embodiment 5 of the present invention.

FIG. 6 is a flow chart showing an algorithm of processing of a posture target value generator according to Embodiment 6 of the present invention.

FIG. 7 is a configuration diagram showing a thruster control method according to a seventh embodiment of the present invention.

FIG. 8 is an explanatory diagram showing a thruster control law in the embodiment.

FIG. 9 is a flowchart showing an algorithm for matching offset / replacement injection patterns according to an eighth embodiment of the present invention.

FIG. 10 is a flowchart showing an algorithm for extracting a cancellation / substitution injection pattern according to a ninth embodiment of the present invention.

FIG. 11 is a flowchart showing an algorithm for matching offset / replacement injection patterns according to a tenth embodiment of the present invention.

FIG. 12 is an explanatory diagram showing a thruster selection rule in the thruster control unit according to the eleventh embodiment of the present invention.

FIG. 13 is a flowchart showing an algorithm for extracting a replacement injection pattern at the time of failure according to the twelfth embodiment of the present invention.

FIG. 14 is an explanatory diagram showing a concept of creating a matrix A _i in the above embodiment.

FIG. 15 is an explanatory diagram showing a thruster selection rule in a thruster selection unit according to a thirteenth embodiment of the present invention.

FIG. 16 is an explanatory diagram showing division of a vector b and a thruster matrix A in the above embodiment.

FIG. 17 is an explanatory diagram showing a thruster selection rule in the thruster selection unit according to the fourteenth embodiment of the present invention.

FIG. 18 is an explanatory diagram showing division of a vector b and a thruster matrix A in the above embodiment.

FIG. 19 is an explanatory diagram showing an ejection pattern database in which a thruster selection unit according to a fifteenth embodiment of the present invention is incorporated.

FIG. 20 is an explanatory diagram showing a method of extracting a lookup table of an injection pattern database according to Embodiment 16 of the present invention.

FIG. 21 is an explanatory diagram showing a thruster selection rule in a thruster selection unit according to the seventeenth embodiment of the present invention.

FIG. 22 is a flow chart showing an algorithm for redundant placement of thrusters according to the eighteenth embodiment of the present invention.

FIG. 23 is a block diagram showing a conventional spacecraft controller.

FIG. 24 is a configuration diagram showing a conventional thruster control method.

FIG. 25 is a perspective view showing the thruster arrangement.

FIG. 26 is an explanatory diagram showing a jetting state of thrusters corresponding to the respective control axes.

[Explanation of symbols]

 1 Spacecraft main body 2 Position detector 3 Position target value generation unit 4 Position control calculation unit 5 Attitude detector 6 Attitude target value generation unit 7 Attitude control calculation unit 8 Decoupling calculation unit 9 Feedforward calculation unit 10 Camera type sensor (Nearby sensor) 11 Gyro 12 Estimator 13 Attitude estimator 16 Thruster selector 17 Thruster modulator 18 Thruster 19 Spacecraft dynamics

Claims (13)

[Claims] 1. A position detector for detecting the position of a spacecraft whose position and attitude are controlled in six degrees of freedom, a position target value generator for generating a target value for the position of the spacecraft, and the position detector. Based on the deviation between the output of the position target value and the position target value generation unit, a position control calculation unit that performs a calculation for controlling the position of the spacecraft, a posture detector that detects the posture of the spacecraft, and a posture of the spacecraft An attitude target value generator that generates a target value, an attitude control calculator that performs an operation to control the attitude of the spacecraft, based on a deviation between outputs of the attitude detector and the attitude target value generator,
A control device for a spacecraft, comprising: a decoupling calculation unit that removes interference on dynamics from the calculation results of the position control calculation unit and the attitude control calculation unit. 2. The acceleration of the target value of the position and attitude of the spacecraft and the inertial force on the dynamics based on the outputs of the position detector, the position target value generator, the attitude detector and the attitude target value generator. The control device for a spacecraft according to claim 1, further comprising a feedforward calculation unit that compensates in advance. 3. The camera-type sensor for detecting the relative position and relative attitude between the spacecraft and the target is used as the position detector and the attitude detector. The control device for the spacecraft described in. 4. The relative velocity and the angular velocity of the target based on the relative attitude of the gyro that detects the angular velocity of the spacecraft, the detection value detected by the gyro, and the target detected by the camera type sensor. Estimate,
The control device for a spacecraft according to claim 3, further comprising an estimator that sends an estimation result to the attitude control calculation unit and the feedforward calculation unit. 5. The position detector detects a distance between the spacecraft and a target and an angle at which the spacecraft looks at the target, and the attitude detector detects an attitude of the spacecraft in an inertial space. The control device for a spacecraft according to claim 1 or 2, wherein the control device detects the spacecraft. 6. The spacecraft control device according to claim 5, wherein the attitude target value generation unit provides a attitude that rotates at a constant angular velocity with respect to the inertial space as a target value. . 7. The spacecraft according to claim 5, wherein the attitude target value generation unit gives a attitude such that the attitude angular velocity changes at a certain ratio with respect to the inertial space, as a target value. Control device. 8. A position provided on a spacecraft for generating a control force / torque obtained by a control calculation.
In a thruster control method of selecting a thruster for attitude control and distributing an injection amount to drive the thruster, first, the selection of the thruster and the distribution of the injection amount are performed based on a look-up table, and then the selection is performed. A canceling injection logic for removing a canceling injection pattern in which the resultant force and the resultant torque of each thruster become 0, or a replacement injection for replacing the thruster injection combination that produces the same resultant force and resultant torque but requires a small total injection amount A thruster control method comprising applying logic to perform more efficient jet distribution of the thruster. 9. When applying the offset injection logic and the replacement injection logic, the components of the offset injection pattern and the replacement injection pattern that match are described with opposite signs, and appropriate scheduling is performed based on the signs. The thruster control method according to claim 8, wherein the thruster control method is performed. 10. In the replacement injection logic, a replacement replacement injection pattern at the time of failure is created in which each thruster is replaced with another remaining thruster in the sense that the resultant force and the resultant torque are equal, and the failure is performed when the thruster fails. 9. The thruster control method according to claim 8, wherein the failed thruster is replaced with a normal thruster according to the temporal replacement injection pattern. 11. A control force obtained by control calculation
In the thruster control method of selecting a thruster for position / attitude control provided in the spacecraft and distributing the injection amount to drive the torque, the control force / torque to be generated Is divided into a 3-axis component and a 3-axis component, and the thrust distribution of the thruster is determined by on-board linear programming. 12. A look-up table for thruster failure and a look-up table for thruster failure are prepared in addition to the look-up table for thruster selection and injection amount distribution. And use the normal thruster to
9. The thruster control method according to claim 8, wherein a combined torque is obtained. 13. A control force obtained by control calculation
In the thruster control method of selecting the thrusters for position / attitude control provided in the spacecraft and distributing the injection amount to drive the torque in order to generate the torque, other remaining thrust The thruster control method is characterized in that the total injection amount in the case of replacing the thruster is calculated, and the redundant number is preferentially assigned to the thruster having a large degree.