JPS59128100A - Orbit charging method of space ship - Google Patents

Abstract

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

Description

DETAILED DESCRIPTION OF THE INVENTION The present invention relates to a method for inserting an artificial satellite, an artificial planet, or an interplanetary spacecraft into a desired orbit from another spacecraft.

Conventionally, the methods for inserting this type of spacecraft into orbit include using a large launch rocket, equipping the spacecraft itself with a propulsion device such as a perige motor or aboge motor and inserting it into a high orbit, or using a space shuttle or space platform. There is a method of launching using a propulsion device called an upper stage based on the form, etc., but all of these require a large propulsion device and its control device to be installed on the spacecraft to be launched, resulting in an increase in weight and volume. This method has disadvantages in that it inevitably involves impact loads during combustion and complexity of the structure.

In order to eliminate these drawbacks, the present invention utilizes stability due to gravitational gradient, and connects a mother spacecraft (first spacecraft) to a receiving spacecraft (second spacecraft) with a wire. It is characterized by inserting it into the desired orbit by cutting the wire when the required orbital radius, tangential velocity, etc. are obtained, and the spacecraft to be inserted (second spacecraft) itself Since the propulsion device for orbit insertion is unnecessary or simplified, the load upon orbit insertion is reduced, and
This makes it possible to reduce weight and simplify the structure, and the drawings will be explained below.

1, 2, 3, and 4 are conceptual diagrams showing embodiments of the present invention. For simplicity, the initial orbit is a circular orbit, and the spacecraft is placed in a higher orbit. An explanation will be given using an example of input.

Figure 1 shows the initial situation, where 1 is the star, 2 is the initial circular orbit, 3 is the mother ship spacecraft (first spacecraft), and in the initial state, the inserted spacecraft (second spacecraft). 4, R is the radius of orbit 2, and ■ is the tangential velocity of spacecraft 3.

Figure 2 shows a state in which a second spacecraft 4 is connected to a first spacecraft 3 by a wire rod 5, and 4 is the second spacecraft which is the inserted spacecraft, and 3' is the second spacecraft. The first spacecraft after separating the spacecraft 4, 5 indicates a wire rod.

Figure 3 shows a state in which the second spacecraft 4 is stabilized by the gravitational gradient after the wire 5 has been drawn out to the desired radius of rotation, and Ra5V
a is the radius of rotation and tangential speed of the second spacecraft 4, respectively,
Rb and Vb are the radius of rotation of the first spacecraft 3', respectively,
Shows tangential velocity. FIG. 4 shows the state when the wire 5 is cut, 6 shows the orbit into which the second spacecraft 4 has been inserted, and 7 shows the orbit taken by the first spacecraft 3'.

For the sake of simplicity, we assume that disturbances such as atmospheric resistance and solar radiation pressure can be ignored.
As shown in Figure 2, when the second spacecraft 4 on board is pulled out from the spacecraft 3 by the wire 5, tension acts on the wire 5 due to the gravitational gradient, and the first spacecraft 3 The straight line connecting ' and the second spacecraft 4 is stabilized so that it points toward the center of gravity of star 1, and orbits around the star 1. 〇For further simplicity, the line is made so that the original angular velocity does not change. When the wire 5 is combed out under speed control and the state shown in FIG. 3 is reached, the tangential velocity Va of the second spacecraft 4 and the first spacecraft 3'.

vb is Va=Ka@V, KaHRa/R)IVb=
Kb-V, KbE is expressed as Rb/R<1, but Ra5Rb, or KaSKb, is a value determined by the length of the wire 5, the linear density, and the mass ratio of the second spacecraft 4 and the first spacecraft 3'. be.

At this point, if the wire 5 is separated from the second spacecraft 4 as shown in FIG. 4, it will be inserted into a desired orbit.

In general, if the orbital radius at the perige point of a mass point moving around a star with mass M1 and universal gravitational constant G is r1 and the tangential velocity is 1, then the eccentricity e of the orbit is expressed as G, and orbit 2 in Figure 1 is Since it is a circular orbit, G, MG=RV2, and the eccentricity C of the orbit 6 into which the second spacecraft 4 is inserted.
GaVa2 = Ka3-1>0, and depending on the KaO value, 1<Ka<v': Elliptical orbit Ka = J2: Parabolic orbit f(Ka
: Draw a trajectory like a hyperbolic trajectory.

By appropriately selecting Ka, the second spacecraft 4 can be inserted into a desired orbit.

In this embodiment, for the first spacecraft 3', if the mass of the wire 5 is negligible, or if the wire 5 is to be separated and discarded in space, the trajectory after separation is 7.
In the case where the orbit is an ellipse such that the aboriginal point is the point where the wire is separated, or when the wire 5 whose mass cannot be ignored is attached, or when it is recovered by being rolled into the first spacecraft 3'. , the elliptical orbit will be disturbed, but if necessary, it can be re-entered into orbit by installing an appropriate propulsion device.

Similarly, from an elliptical orbit, a parabolic orbit, or a non-orbital orbit such as a hyperbolic orbit, the gravitational gradient of the star is used to eject the inserted spacecraft from the mother ship's spacecraft with a wire rod. Once the required orbital radius, tangential velocity, etc. are obtained, the wire can be separated and the wire can be inserted into a new orbit. When inserting into a new orbit, it is projected while utilizing the stability provided by the gravitational gradient, so the orbit insertion propulsion device of the inserted spacecraft is unnecessary or simple, and the accompanying equipment and structures are unnecessary. Alternatively, it becomes simpler, the impact load caused by the propulsion device is eliminated or reduced, and the weight of the spacecraft is also reduced.

This method involves launching a spacecraft such as an artificial satellite into a higher orbit from a space platform or space vehicle such as a space ferries orbiting in a low-altitude orbit around a star such as the Earth, or conversely from a high orbit to a lower orbit. In addition, when a spacecraft flying between planets is taken into the gravitational field of a star, it is separated from the spacecraft. It can also be applied to cases where a spacecraft is placed in a low or high orbit.

[Brief explanation of the drawing]

Figure 1 is a diagram showing the spacecraft orbiting in an initial circular orbit, Figure 2 is a diagram showing how the second spacecraft is extended from the first spacecraft with a wire, and Figure 3 The figure shows the situation when the second spacecraft has the required orbital radius and tangential velocity and is just before the wires are separated, and Figure 4 shows the situation when the wires connecting the spacecrafts are separated. It is a diagram.

Claims (1)

[Claims] The first orbit that orbits or passes through the star's gravitational field.
A method of inserting a second spacecraft from a spacecraft into Yuzu's orbit, in which the first spacecraft and the second spacecraft are released while being connected by a wire in the direction of gravity of the star, and the gravitational gradient is The wire is pulled out in a stabilized state, and when the required orbital radius, tangential velocity, etc. for the second spacecraft are obtained, the first
A method for inserting a spacecraft into an orbit, the method comprising the step of projecting the second spacecraft in a tangential velocity direction and inserting it into a desired orbit by cutting off a wire connecting the spacecraft and the second spacecraft.