JPH11291995A - Guidance control device - Google Patents

Abstract

PROBLEM TO BE SOLVED: To shorten the time required for setting a thether artificial satellite to an intended orbit by regulating the tension of a thether on the basis of the time change rate of the length of the thether. SOLUTION: In a damping controller 6, a tension Tp is calculated according to a prescribed expression on the basis of the time change rate of length of a thether 3. The result of the addition of the tension Tp to the output T from a mission function controller 5 is outputted as a control input Tt . A reel mechanism 4 generates a tension according to the control input Tt and performs the winding or feeding of the thether 3. When the control input Tt is applied, the reducing speed dM/dτ of a mission function M is dM/dτ=-(kΛM+c)(Λ')<2> , wherein γ: dimensionless time, Λ: dimensionless thether length, k: regulation parameter for mission function control, and c: regulation parameter for damping control. Accordingly, dM/dτ is smaller than -c(Λ')<2> , and the lowest reducing speed determined by the value of the constant (c) is ensured regardless of M.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a guidance control device for guiding and controlling a tether artificial satellite.

[0002]

2. Description of the Related Art FIG. 4 is a view for explaining conventional guidance control of a tether artificial satellite. What is tether satellite 1?
As shown in FIG. 4, it is a child satellite pulled by a tether (string) 3 from a large parent satellite 2 such as a space station. The orbit of the tether artificial satellite 1 viewed from the parent satellite 2 is uniquely determined from the length of the tether 3 in the settling state. Therefore, the guidance control of the tether artificial satellite 1 only needs to adjust the length of the tether 3 by the reel mechanism 4 if only the settling state is considered.

However, since the tether artificial satellite 1 moves like a pendulum suspended from the parent satellite 2, the tether artificial satellite 1 vibrates simply by keeping the tether 3 at a length corresponding to a desired orbit. It does not settle to the target trajectory. In outer space, the atmosphere is sparse, and there is no friction with the atmosphere, which is a damping force of the pendulum motion. Therefore, the pendulum motion cannot be stopped unless artificial damping is applied. Therefore, the guidance control of the tether artificial satellite 1 requires a function of preventing vibration.

[0004] A typical guidance control method for a tether satellite includes a mission function method (Hironori Fujii and Shinta).
ro Ishijima, “Mission Function Control for
Deployment and Retrieval of a Subsatellit
e. ”Journal of guidance, vol.12, March-April 1
989.pp.243-247).

[0005] This mission function method is minimized only when the tether satellite is in the target orbit, and its value is 0.
A mission function M is defined as follows, and a control input is applied so that the mission function M decreases over time. Hereinafter, the conventional mission function method will be briefly described using mathematical expressions.

FIG. 5 is a diagram for explaining guidance control of a tether artificial satellite by a conventional mission function method. When the tether artificial satellite 1 is modeled as shown in FIG. 5 and the tether tension T is used as a control input, the motion of the tether artificial satellite 1 around the parent satellite 2 is described by the following equation (1).

[0007]

(Equation 1)

Here, () ′ = d () / dτ τ: dimensionless time (= Ωt) t: real time [sec] Ω: angular velocity of the parent satellite orbiting the earth [rad / sec] Λ: l / ls (Dimensionless tether length) ls: Target length [km]

[0009]

(Equation 2)

M: mass of the tether satellite [kg] T: tether tension [N] l: tether length φ: inclination with respect to the perpendicular of the tether. The target orbit of the tether artificial satellite is represented by the following equation (2).

L = ls, l ′ = φ = φ ′ = 0 (2) For the tether satellite described by the above equation (1), a mission function M is defined as the following equation (3).

[0012]

(Equation 3) Here, a1, a2, and b are positive real numbers used as control adjustment parameters. M does not take a negative value as long as a1, a2, and b are selected as positive real numbers. Further, the minimum value of M is 0 only in the target trajectory of Expression (2).

[0013]

(Equation 4) Here, k is a positive real number used as an adjustment parameter.

FIG. 6 is a block diagram of a control system using the equation (4). Differentiating equation (3) with respect to time τ, equation (1),
By substituting (4), the following equation (5) is obtained. dM / dτ = −kΛ (Λ ′) ² M (5) Obviously, the right side of Expression (5) has a negative value except for the target trajectory. Thus, using the control input of equation (4), M monotonically decreases with time until its minimum value becomes zero.

[0015]

(Equation 5) Since M = 0 corresponds to the target orbit, the tether satellite can be put into the target orbit by using the control input of Expression (4).

[0016]

FIG. 7 is a view showing a simulation result of orbit insertion by the conventional mission function method. The parent satellite 2 is taken at the origin (X = Y = 0 km), and It is the figure which plotted the locus.
The target trajectory is X = -100 km and Y = 0 km. Although the tether satellite reaches stability near the target orbit,
In the vicinity of the target trajectory, the target trajectory orbit is not set.

FIG. 8 is a diagram showing a time change of the tether length at this time. Although the target 100 km is reached in about 3 × 10 ⁴ seconds, the vibration continues and the calculation is completed 9 ×
It can be seen that does not settle even I致to 10 ⁴ seconds.

As described above, the conventional method has a problem that the vibration continues near the target trajectory. The cause can be explained as follows. As the tether satellite approaches the target orbit, the value of the mission function M decreases, and the rate of decrease of the mission function given by equation (5) is equal to the mission function M
Is proportional to the value of Therefore, the rate of decrease of the mission function M decreases near the target trajectory, so that the settling to the target trajectory is delayed. An object of the present invention is to provide a guidance control device that reduces the time required for setting a tether artificial satellite to a target orbit.

[0019]

SUMMARY OF THE INVENTION In order to solve the above-mentioned problems and achieve the object, a guidance control device of the present invention is configured as follows. The guidance control device of the present invention is a guidance control device for guiding and controlling a tether artificial satellite based on a mission function method, comprising a control means for adjusting the tension of the tether based on a time change rate of the length of the tether.

[0020]

FIG. 1 is a block diagram of a control system of a guidance control device according to an embodiment of the present invention. The guidance control device 7 for a tether artificial satellite according to the present embodiment is characterized in that an attenuation controller 6 is added to the mission function controller 5 of the conventional guidance control device described above.

[0021]

(Equation 6) Here, c is a positive real number, and is used as an adjustment parameter for damping control.

[0022]

(Equation 7)

When the control input of equation (7) is applied, the equation (7)
The decreasing speed of the mission function M shown in 3) is given by the following equation (8). dM / dτ = − (kΛM + c) (Λ ′) ² (8) Therefore, the following equation (9) is obtained.

DM / dτ <−c (Λ ′) ² (9) Accordingly, the mission function M is guaranteed to have a minimum decreasing speed determined by the value of the constant c regardless of the value of M. Due to this property, even if the tether artificial satellite approaches the target orbit and the value of the mission function M becomes small, the decrease speed does not become 0 as shown in Expression (9), so that the settling to the target orbit is accelerated. be able to.

FIG. 2 is a diagram showing a simulation result of orbit insertion by the tether artificial satellite guidance control device 6 according to the present embodiment. Compared to the result of the conventional method shown in FIG. 7, the tether satellite has a target orbit (X = −100 km,
It can be seen that the target trajectory is settled without vibrating around (Y = 0 km).

FIG. 3 is a diagram showing the change over time of the tether length at that time. It can be seen that the settling time is about 3 × 10 ⁴ seconds for 100 km of the target trajectory. As described above, according to the guidance control device of the present embodiment, the time required for settling to the target trajectory can be significantly reduced, which is effective in operation.

It should be noted that the present invention is not limited to only the above-described embodiment, and can be appropriately modified and implemented without changing the gist. (Summary of Embodiment) The configuration, operation and effect shown in the embodiment are summarized as follows.

The guidance control device shown in the embodiment is a guidance control device 7 for guiding and controlling the tether artificial satellite 1 based on the mission function method, wherein the tension of the tether 3 is used as a control input and the length of the tether 3 is controlled. Control means (attenuation controller 6) for adjusting the tension of the tether 3 based on the time change rate is provided. As described above, according to the guidance control device, the tether artificial satellite 1 is settled in the target orbit in a short time without vibrating near the target orbit.

[0029]

According to the guidance control device of the present invention, the tether artificial satellite is settled in the target orbit in a short time without vibrating near the target orbit.

[Brief description of the drawings]

FIG. 1 is a block diagram of a control system of a guidance control device according to an embodiment of the present invention.

FIG. 2 is a diagram showing a simulation result of orbit insertion by the guidance control device of the tether artificial satellite according to the embodiment of the present invention.

FIG. 3 is a diagram showing a time change of a tether length according to the embodiment of the present invention.

FIG. 4 is a diagram for explaining guidance control of a tether artificial satellite according to a conventional example.

FIG. 5 is a diagram for explaining guidance control of a tether artificial satellite by a mission function method according to a conventional example.

FIG. 6 is a block diagram of a control system of a guidance control device according to a conventional example.

FIG. 7 is a diagram showing a simulation result of orbit insertion by a mission function method according to a conventional example.

FIG. 8 is a diagram showing a temporal change of a tether length according to a conventional example.

[Explanation of symbols]

 DESCRIPTION OF SYMBOLS 1 ... Tether artificial satellite 2 ... Parent satellite 3 ... Tether (string) 4 ... Reel mechanism 5 ... Mission function controller 6 ... Attenuation controller 7 ... Guidance control device

Claims (1)

[Claims] 1. A guidance control device for guiding and controlling a tether artificial satellite based on a mission function method, comprising: control means for adjusting the tension of the tether based on a time change rate of the length of the tether. Control device.