RU2115008C1 - Spacecraft motion control device - Google Patents

Abstract

FIELD: space engineering; change or stabilization of orbit parameters and orientation of spacecraft. SUBSTANCE: proposed device includes control systems for control of motion about center of mass of spacecraft and motion of center of mass of spacecraft; used as actuating members of both systems are the same electric rocket engines with deflectable vectors of thrust. Functional decoupling between these systems is effected by means of engine-flywheels which also integrate disturbing moments; for unloading the engines-flywheels, provision is made for adjustment (through control unit) of direction of thrust of electrical rocket engines mainly made in form of plasma accelerators with closed drift of electrons and provided with control magnetizing coils connected to this unit and additional control poles. EFFECT: reduction of mass and enhanced reliability of engine plant. 4 cl, 2 dwg

Description

 The proposed device relates to space technology and is used to change or stabilize the parameters of the orbit and the angular position of the spacecraft (SC). It is especially advisable to use the device in high orbits and, in particular, in a geostationary orbit.

 Known devices using two or more electric rocket engines (ERE), for simultaneous control of the orbital parameters of the geostationary spacecraft and its orientation using electrostatic changes in the direction of the thrust vector [1].

 The disadvantages of the device are: the need for the orientation mode of short-term, pulsed operation of the electric propulsion, in which the traction characteristics of the engine are significantly reduced; the need to compensate not only constant, but also alternating disturbing moments; the interaction of control operations with orientation parameters and orbit parameters.

 Also known are motion control devices separately around the center of mass and the center of mass of the spacecraft [2, 3, 4, 5].

 The closest in technical essence to the proposed invention is a device consisting of two or more electric propulsion engines controlling the motion of the center of mass of the spacecraft, and a system that controls movement around the center of mass of the spacecraft, which in each of the stabilization channels contains: a signal generator of the angular deviation of the spacecraft, a control unit A spacecraft, an executive body control unit and an executive body in the form of an electric flywheel engine (DM) with an unloading executive body (IOR) of the kinetic moment of this DM [2].

 The disadvantage of this device is the presence of a separate IOR unit, which reduces the reliability and increases the mass of the device.

 The aim of the invention is the exclusion of the block IOR.

 The essence of the invention lies in the fact that in the proposed device, the functions of the electric propulsion and electronic ignition are combined and they are performed by one or more electric propulsors with directionally controlled thrust. Traction control by means of electromagnetic effects on the acceleration process is more easily implemented in a plasma accelerator (PU) with a closed electron drift [6, 7, 8].

 When one or more electric propulsion engines are located on the spacecraft so that it is possible to pass the thrust vectors in at least one of the vector orientations outside the center of mass of the spacecraft, this new quality ensures the simultaneous creation of thrust and controlled unloading moment for the DM, or more generally for the hydraulic power stabilizer. This eliminates the need for pulsed operation of the electric propulsion, integrates disturbing moments, and removes the mutual influence of operations controlling the orientation parameters and orbit parameters.

 The technical result that is achieved in the proposed device is to reduce the mass due to the high specific impulse of the thrust of the electric propulsion operating in a continuous mode, and to increase the reliability due to the exclusion of the IOR unit from the device.

 This result is achieved by the fact that in the spacecraft’s motion control device containing actuators in the form of flywheel engines, a flywheel engine control unit with angular velocity vector meters, as well as one or more electric rocket engines with directionally controlled thrust, while electric rocket engines the engines are installed with the possibility of the passage of the lines of action of their rods outside the center of mass of the spacecraft the control unit for the direction of thrust of electric rockets is included vigateley, wherein measuring the angular velocity of each flywheel motor (DM) connected with the direction of traction control unit (BUNT) electric propulsion.

 The device also differs in that in it the electric rocket engine is made in the form of a plasma accelerator (PU) with a closed electron drift, in which at least one source of magnetomotive force in the form of a magnetization control coil is connected to the thrust direction control unit.

 It is also new that in the magnetic circuit of a plasma accelerator at least one source of magnetomotive force is added one or more control magnetization coils installed at the control pole.

 Another distinctive feature of the proposed device is that between at least one control magnetization coil and one or more additional electromagnetic control coils, at least one additional control pole is installed.

 Comparison of the claimed solution not only with the prototype, but also with other close technical solutions in space technology did not allow us to identify in them signs that distinguish, in the aggregate, the claimed solution from the prototype, which allows us to conclude that the criterion of "significant differences" is met.

 To illustrate the invention in FIG. 1 shows a functional diagram of the connections of the blocks and the device, and in FIG. 2 is a structural diagram of a PU with a closed electron drift and with a deflected thrust vector.

 The functional block diagram of the spacecraft motion control device for one of the channels for stabilizing the angular position and one of the directions of motion of the center of mass of the spacecraft (Fig. 1) contains serially connected blocks: an actuator in the form of DM1 with a DM control unit and an angular velocity vector meter 2. The output of the angular velocity vector meter 2 is connected to the thrust direction control input 3, which, in turn, is connected to the electric propulsion 4 'and 4' '.

 The device (Fig. 1) works as follows.

 When DM1 accumulates the limiting angular velocity, which is informed by the BKNT threshold level of the signal coming from the meter 2, information about the necessary direction of the thrust vector deviation is generated in the BUNT 3 block to create a discharge moment along this axis. Block 3 generates a command to turn on the control magnetizing coils of the electric propulsion 4 'and 4' ', which control the direction of the thrust vector.

 The signal for the effect of the unloading moment from block 2 enters block 3 until the kinetic moment of DM1 decreases to the insensitivity level of meter 2.

 According to the invention, the electric propulsion device can be made in the form of a control circuit (Fig. 2), in which there are external sources of magnetomotive force made in the form of electromagnetic magnetization coils 1 ', 1' 'and control magnetization coils 2', 2 '' with additional control poles 3 'and 3' '.

The operation of the device (Fig. 2)
The deviation of the PU thrust vector in a plane is made by reducing the currents flowing to the control magnetization coils 1 'located on one side of the axis of the PU in this plane and increasing the currents in the control magnetizing coils 1''located on the other side in this plane . In this case, the thrust vector deviates toward the control magnetization coils 1 ', in which the current decreases, due to the "bias" of the magnetic field in this direction.

 The magnetization coils 2 'and 2' 'play the same role. Coils 2 'reduce the magnetic field together with coils 1' on one side of the PU and increase it on the other side together with coils 1 ''.

 Additional control poles 3 'and 3' 'contribute to the "skew" of the fields in the working zone of acceleration.

Sources of information
1. Kanlan Marschall AJAA Pap., 1975, N 376, 1-12.

 Design and analysis of the functioning of control systems with zero kinematic moment for geostationary satellites. Express information. A and P N 38/75, VNIITI.

 2. Anuprienko G. E. A device for controlling the motion of a spacecraft around the center of mass.

SU 1819 A1, B 64 G 1/24 dated 07/31/90, Bull. N 21, 06/07/93
Anuprienko G. E. A way to control the orientation of a space object. SU 1811500 A3, B 64 G 1/28, 07/31/90, Bull. N 15, 04/23/93.

 4. Artsimovich L.A., Andronov I.M., Rylov Yu.P. and etc.

 Development of a stationary plasma engine (SPD) and its testing on the METEOR satellite. Space exploration. t. X11, no. 3., 1974. S. 451-468.

 5. Razygraev A. P. Fundamentals of flight control of spacecraft and ships., M.: Mechanical Engineering, 1977.

 6. Morozov B. and al. Demande de Brevet d'invention. Moteur a plasma a derive fermee d'electrons.

 FR N de pullication 2693770, 01/21/94.

 7. Egorov V. and al. Plasma accelerator with closed eltctron drift. U.S. Patent N 5, 218, 271, Jun. 8. 1993.

 8. Belyaev N.M., Belik N.P., Uvarov E.I., Reactive control systems for spacecraft. M .: Mechanical Engineering 1979

Claims (4)

 1. A device for controlling the motion of a spacecraft containing executive bodies in the form of flywheel engines, a control unit for flywheel engines with angular velocity vectors, and also one or more electric rocket engines with directionally controlled thrust, while electric rocket engines are installed with the possibility of the passage of the lines of action of their thrusts outside the center of mass of the spacecraft, characterized in that the device includes a control unit for the direction of electric propulsion thrust engines, and the angular rotation speed meter of each flywheel engine is connected to the thrust direction control unit of the electric propulsion engines.  2. The device according to claim 1, characterized in that in it the electric rocket engine is made in the form of a plasma accelerator with a closed electron drift, in which at least one source of magnetomotive force in the form of a magnetization control coil is connected to the thrust direction control unit.  3. The device according to claims 1 and 2, characterized in that in the magnetic circuit of the plasma accelerator at least one source of magnetomotive force is added one or more control magnetization coils connected to the thrust direction control unit and installed at the control pole.  4. The device according to claims 1 to 3, characterized in that between at least one control magnetization coil and one or more additional electromagnetic control coils, at least one additional control pole is installed.

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