RU2136550C1 - Spin-stabilized spacecraft - Google Patents

Abstract

FIELD: space engineering; spin-stabilized space vehicles. SUBSTANCE: spacecraft has casing, sensor showing angular rate relative to axis of rotation, solar batteries, current collecting power circuits, battery converters, current-carrying wires both shielded and-non-shielded from external magnetic fields, comparator, sawtooth voltage generator, biassed blocking oscillator, symmetrical flip-flop and rectifiers. Time of connection to circuits of shielded or nonshielded wires is controlled in flight by means of rectifiers whose control inputs are connected with output of angular-rate sensor through flip-flop and comparator. EFFECT: possibility of maintenance of angular rate at any preset accuracy due to control of currents of interaction of solar batteries and earth's magnetic field; reduced losses of electric power; enhanced efficiency of control. 2 cl, 3 dwg

Description

 The present invention relates to space technology and can be used in spacecraft stabilized by rotation.

 Known spacecraft with rotation stabilization, comprising a vehicle body, deployable solar panels placed symmetrically about the axis of maximum moment of inertia of the device and equipped with collector power circuits combined with groups of series-connected photoelectric battery converters, as well as connected to collector circuits shielded from an external magnetic field current-carrying wires, while collector power circuits and photoelectric converters about edineny secured according to the direction of current relative to the housing unit [1].

 The disadvantage of this spacecraft is the need to control the speed of the vehicle’s own rotation when its optimal value is exceeded, especially when the batteries are operating in overload or short circuit conditions, when the moment of Ampere force applied to the device as a result of the interaction of solar cell currents with the Earth’s magnetic field will significantly exceed the effect of inhibitory factors.

 The aim of the proposed solution is time-pulse control of maintaining the angular velocity of the spacecraft’s own rotation within specified limits by controlling the compensating effect resulting from the interaction of solar cell currents and the Earth’s magnetic field on the influence of disturbing factors.

 This goal is achieved by controlling the time of connecting wires shielded from the Earth’s external magnetic field to the current collecting power circuits or unshielded current-carrying wires by means of valves whose control inputs are connected through a symmetric trigger, the comparison device is connected to the output of the angular velocity sensor of the spacecraft, comprising: a casing; angular velocity sensor relative to the axis of rotation; deployable solar panels placed symmetrically with respect to the axis of maximum moment of inertia of the spacecraft and equipped with collector power circuits combined with groups of series-connected photoelectric transducers of batteries, providing according to the direction of their currents, relative to the body of the spacecraft; current-carrying wires shielded from an external magnetic field are connected to current-collecting circuits by means of gates; unshielded current-carrying wires are connected to collector circuits by means of valves; comparison device; sawtooth generator; braked blocking generator; symmetric trigger; valves.

 In FIG. 1 is a drawing of the proposed spacecraft, consisting of a housing 1, an angular velocity sensor relative to the axis of rotation 2, solar panels 3, placed symmetrically about the axis of maximum moment of inertia of the spacecraft and equipped with current-collecting power circuits 4, combined with groups of series-connected photoelectric converters 5 of batteries with providing a consonant direction of their currents relative to the spacecraft body 1, shielded from an external magnetic field, then non-shielded current-carrying wires 6, 14 connected to current collection circuits 4 by means of valves 7, while unshielded current-carrying wires 8, 15 are connected to current-collecting power circuits 4 through valves 9, which control inputs of valves 7, 9 through a symmetrical trigger 10, which is inhibited a blocking generator 11, a comparison device 12, to which a sawtooth voltage generator 13 is connected, are connected to the output of the angular velocity sensor relative to the axis of rotation 2.

где B - индукция магнитного поля Земли;
I - ток солнечной батареи;
L1 - длина токосборного контура и проводов солнечной батареи от точки А до точки Б;
L - длина экранированного провода солнечной батареи от точки С до точки Д;
L2 - длина неэкранированного провода от точки Б до точки Д.When a spacecraft moves with stabilization of rotation in the orbit of an artificial Earth satellite, the Earth’s magnetic field will act according to Ampere’s law on its solar batteries with force (see Fig. 3)

where B is the induction of the Earth's magnetic field;
I is the current of the solar battery;
L1 - the length of the collector circuit and wires of the solar battery from point A to point B;
L is the length of the shielded wire of the solar battery from point C to point D;
L2 is the length of the unshielded wire from point B to point D.

Момент силы Ампера, приложенный к космическому аппарату, будет равен

где n - количество развертываемых солнечных батарей;
M - момент силы Ампера.We take L1 = L2, then

Ampere's moment of force applied to the spacecraft will be equal to

where n is the number of deployable solar panels;
M - Ampere moment of force.

где E - величина углового ускорения;
J - момент инерции космического аппарата.The magnitude of the angular acceleration of the spacecraft created by the action of the Earth's magnetic field and the currents of deployable solar cells, according to Newton’s second law for rotating bodies, will be

where E is the magnitude of the angular acceleration;
J is the moment of inertia of the spacecraft.

 It can be seen from expression (1) that if the shielded section of the conductor L is disconnected from the solar battery circuit, and an unshielded conductor is connected instead, the Ampere force will become zero. Thus, by adjusting the time of connecting a shielded conductor to the solar battery circuit, we can adjust the speed of the spacecraft according to the linear law with almost any accuracy.

 Consider the principle of controlling the angular velocity of rotation of the spacecraft shown in FIG. 1. An error signal arrives from the angular velocity sensor relative to the axis of rotation 2 to the comparison device 12, if the rotation frequency is less than the nominal (see diagram (a) of Fig. 2). The sawtooth voltage is also supplied to the comparison device 12 from the sawtooth voltage generator 13. When the voltage coincides, the comparison device 12 is closed, the blocking generator 11 is triggered, the first pulse of which throws the symmetric trigger 10 to the opposite state. From one output of the symmetric trigger 10 (see diagram (e) of Fig. 2), the signal enters the valves 9, and from the inverse output (see diagram (e) of Fig. 2) - to the valves 7. Valves 9 open and 7 close. When the voltage U threshold is reached. By the sawtooth voltage generator, the symmetric trigger 10 returns to its original state.

 Therefore (see diagram (c) of Fig. 2), a shielded conductor 6 is connected to the solar battery in section O - T1, the Ampere force acts according to expression (1) - the spacecraft is untwisted, and in section T1 - T2 - a shielded conductor is connected 14, the spacecraft is inhibited. This provides the equation of angular velocity relative to the axis of rotation within the required limits.

 The control accuracy is determined by the sensitivity threshold of the angular velocity sensor relative to the axis of rotation 2 and the pulse front of the blocking generator (see diagram (d) of Fig. 2).

 Tracking a given angular velocity of the spacecraft relative to the axis of rotation is carried out in key mode.

 Thus, compared with the prototype, the use of this technical solution allows you to maintain the angular velocity of the spacecraft relative to soybean rotation with a given accuracy.

Sources of information
1. O. Nikonov. Spacecraft with rotation stabilization. Application 93057506 \ 11 (057692) dated 12/28/93; Published on 10/20/95.

Claims (1)

 A spacecraft with rotation stabilization, containing solar panels symmetrically placed with respect to the axis of maximum moment of inertia of the device and equipped with collector power circuits combined with groups of series-connected photoelectric converters of solar batteries providing a consistent direction of their currents relative to the device’s body, shielded current-carrying wires connected to the collector power circuits, and unshielded current-carrying wires connected to the specified collector circuits parallel to the current-carrying wires, characterized in that it is equipped with an angular velocity sensor relative to the axis of rotation of the apparatus, while these shielded and unshielded current-carrying wires are connected to the indicated collector power circuits through valves, the control inputs of which are through a symmetric trigger, a braked blocking generator and a comparison device to which a sawtooth voltage generator is connected is connected to the output of said angular velocity sensor.

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