RU2771059C1 - Space vehicle electrication regulator - Google Patents

Abstract

FIELD: electrical engineering.SUBSTANCE: invention relates to the field of electrical engineering, in particular to the field of space technology, and can be used to eliminate the accumulated static surface charges (electrons) on the elements of the outer structure of spacecraft. The electrification regulator (ER) of the spacecraft contains sensors for controlling the level of electrification of the spacecraft surface fragments and plasma engines with systems for power supply, control and supply of the working fluid to the engine chambers. Spacecraft ER also contains accelerator channels (AC) of plasma engines, which in the zone of low-temperature plasma creation are connected by side inlet gas lines, which have semi-cylindrical intake shells at the beginning in the AC, and side outlet gas lines with electrovalves located on them with neutralizers limited from all sides and communicating with the external environment through only gas lines. The composition of the gas mains includes metallized fragments of the spacecraft surface and thermoelements rigidly placed on them and having a galvanic connection with them.EFFECT: emission of electrons of differential charging from the surfaces of the spacecraft that are in contact with the PD plasma is improved.1 cl, 2 dwg

Description

The invention relates to the field of electrical engineering, in particular to the field of space technology, and can be used for the emission of electrons (static charges) from elements of the outer surfaces of the spacecraft (SC) structure.

A technical problem in the operation of a spacecraft is the electrical surface and volumetric charging of the structural materials of its surfaces. As a result of the interaction of the spacecraft with the surrounding space plasma, its structure acquires a negative charge due to the influx of space plasma electrons onto the surface of the spacecraft. The plasma parameters for the worst case charging in geostationary orbit are (Λ):

- electron concentration 1.12 cm ^-3 ;

- average electron energy 1.20⋅10 ⁴ eV;

- concentration of ions (mainly H ⁺ ) 0.236 cm ^-3 ;

- average ion energy 2.95⋅10 ⁴ eV.

The charge of the spacecraft surfaces will increase under conditions of geomagnetic storms in the magnetosphere, when the concentration of electrons with an energy of about 20 keV, which make the main contribution to the charge of the spacecraft surfaces, increases several times.

Due to the complex geometry of the spacecraft surfaces, differences in illumination and in the electrical parameters of various surface fragments, a fundamentally uneven charge distribution (differential charging) arises on the spacecraft surfaces interacting with the environment (electrons) and also uneven volumetric charging of structural materials (ions, in overwhelming concentration H ⁺ solar wind). A certain influence will also be exerted by the plasma created during the operation of propulsion systems.

The consequence of the occurrence of differential charging of the spacecraft is the occurrence of an electrostatic discharge (ESD). Its characteristics:

- current in the discharge pulse 100 A;

- discharge pulse duration 10 ^-7 - 10 ^-6 s;

- energy in the discharge pulse 0.2 J;

- discharge pulse voltage 20 kV;

- discharge frequency 10 Hz.

The number of non-conductive elements of subsystems should be kept to a minimum. Shielding of cable assemblies and blocks is mandatory.

It is believed that large total static charges on the surface and under the surface of the spacecraft in space conditions is an inevitable problem. It can be indirectly minimized by competently engaging in ESD prevention (shielding, metallization, selection of surface materials, installation of noise immunity filters, "grounding", that is, the organization of galvanic connections to the spacecraft body). The absolute sum of electric charge carriers in this case will not exceed the level determined for the environmental characteristic (Λ). Over time, the same space plasma contributes to the compensation of surface charges (the interaction of plasma ions and static electrons), and the accumulation of a positive volume charge contributes to the rejection of positive space plasma ions that again penetrate into the body of the spacecraft. However, the practice of preventing electrical breakdowns does not always work for long periods of SC existence. To forcefully minimize the sum of negative electric charge carriers means to directly reduce differential charging. It is more profitable and safer to control the process of neutralization (de-electrification) of the spacecraft body than simply to observe it, without the possibility of interfering in this process.

During the production of spacecraft, there are requirements for the prevention of ESD and maintaining the operability of onboard equipment of service systems and payloads, which make it possible to minimize the impact of ESD:

- all materials used on surfaces in contact with the external environment must be tested for electrification and meet the requirements for resistance to the effects of electrification factors and for the effect on the deterioration of the interference environment (in electronic circuits);

- shielding (Faraday grids) of cable assemblies and blocks must be made in accordance with the requirements of the guiding technical materials;

- housings of blocks and assemblies located outside the spacecraft must be made of conductive materials;

- cable assemblies must be metallized;

- conductive structural elements should be metallized, the number of non-conductive elements on external surfaces should be minimized;

- the cable network located outside the spacecraft must be shielded; on the block and cable parts of the connectors of the cable system, shielding cases must be installed, forming a continuous screen with the cases of external units and the spacecraft case on one side and with the screens of cable trunks on the other side;

- control circuits should not have galvanic connection with the body;

- sensor equipment, including cables, must be shielded;

- in order to protect electronic networks and the spacecraft as a whole from the static potential difference induced between the power buses and the case, it is necessary to install a filter.

All these requirements are a standard set of measures to ensure electrostatic protection (ESPS) on a spacecraft. These requirements are reduced to two, which have the properties of generalizing and mandatory features: shielding, that is, the prevention of static charges behind the shields, and grounding, or, if this is not possible, the creation of a reliable system of galvanic connections of structural elements within the entire spacecraft with a conductive spacecraft case.

For example, the following inventions belong to active methods of dielectricization of spacecraft.

There is a known method of protecting spacecraft from static electricity and a device for its implementation (RU 2612474 C1, IPC B64G1 / 52, B64G1 / 50), in part of the method including the metallization of all units, equipment and spacecraft housing, as well as collecting, measuring, converting to thermal energy of the accumulated surface charge and its dissipation in outer space, characterized by the fact that the equipment of service systems is metallized on the bus, and the complex of target and / or scientific equipment - on another bus, output to the bodies of two parts of the spacecraft, isolated from each other by a non-conductive farm, moreover, both tires are connected to the plates of a large-capacity electric capacitor, which is automatically discharged upon reaching a certain potential difference on command from the control system to ohmic resistance with the release of the last heat on the surface, which is removed through the radiator-cooler to the surrounding space.

Analog has disadvantages:

1. There is no need to divide all equipment into two unequal groups: service system modules, on the one hand, and target and scientific equipment, on the other hand. This is hardly realistically possible, since the equipment on board the spacecraft is placed compactly, which is one of the reasons for the reduction in the mass of the spacecraft. The solution must be found in optimizing the placement of on-board equipment on two tires isolated from each other of a completely arbitrary design without sharp bends;

2. Electromagnetic energy (heat) is removed into the surrounding space, thereby equalizing the electrical potentials as much as possible over the entire surface and internal structure of the spacecraft. Over time, this unequivocally leads to a more frequent occurrence of ESD, as the density of static charges and the density of convection currents inside the spacecraft gradually increase, the cause of which is solar illumination and solar wind. And it would be necessary to divert static electrons. The verified work on the removal of static electrons guarantees the complete absence of ESD on the spacecraft during the entire period of its active existence.

But, accepting these shortcomings, the method and the device corresponding to it are industrially applicable, and they should be used in the design of new spacecraft, since this is a step forward compared to only SKMOES.

There is a known method for regulating the electrization of spacecraft and a device for its implementation (RU 2020776 C1, IPC H05F3/04), in terms of the device (with an acceptable explanation of the method) containing sensors for monitoring the level of electrization of fragments of the surface of the apparatus and neutralizers in the form of ion-plasma engines (IPD) with systems of power supply, control and supply of the working substance to the engine chambers, characterized in that, in order to increase the efficiency of regulation while reducing the level of dynamic and electromagnetic interference, at least two conjugated sensors of the electrization level and one pair of PPDs with symmetrically coaxially oriented opposite or opposite to the outlet nozzles, while the channels for supplying the working substance to the engine chambers and the power supply and engine control circuits are made conjugated, while the latter are connected to a common control unit, the outputs of which are connected to conjugated sensors for controlling fragment neutralization levels in.

Based on the signals from the sensors, two counter- or opposite-directed flows are generated with mutual compensation of their reactive force impact on the spacecraft, which creates an artificial external own atmosphere near the surface to be neutralized, which fends off the effect of electronic flows and prevents the potentials from increasing in individual areas, due to their discharge through the surface adjacent to the surface. denser plasma. This device is taken as a prototype.

It is important here that such an engine develops thrust due to the injection of ions of the working fluid (RT), regardless of the type of engine. It can also be a stationary plasma engine (hereinafter referred to as a plasma engine (PD).

This analogue has the following disadvantages:

1. If we use the PD used to correct the motion of the center of mass of the spacecraft, then the jets directed along the neutralized surface will not create their own atmosphere for the spacecraft, but will simply burn everything that gets in their way. The plasma temperature at the engine outlet is about 1000 K;

2. If you use special PDs with a reduced second consumption of the working substance, then it is necessary to reduce the cross section of the inlet jet, which leads to stricter requirements for the purity of the working substance in fuel tanks. What is economically unacceptable;

3. It is possible that it will be necessary to build a power supply system for electrification motors-regulators, parallel to a similar power supply system for the engines of the spacecraft center of mass movement correction system: the same tanks, the same fuel lines, the same mass. Mass issues are among the main ones in spacecraft design.

For the claimed invention, the following essential features in common with the prototype have been identified: a spacecraft electrification controller containing sensors for controlling the level of electrization of spacecraft surface fragments and plasma engines with systems for power supply, control and supply of the working fluid to the engine chambers.

The technical problem of the invention is the creation of a device that regulates the level of electrization of the spacecraft. This technical problem is solved due to the fact that the electrification controller (RE) of the spacecraft, containing sensors for monitoring the level of electrization of fragments of the surface of the spacecraft and plasma engines with power supply systems, control and supply of the working fluid to the engine chambers, differs from the prototype in that it contains acceleration channels (UK) plasma engines, which in the zone of low-temperature plasma creation are connected by side inlet gas lines, which have semi-cylindrical intake shells at the beginning in the UK, and side outlet gas lines with electrovalves located on them, limited from all sides and communicating with the external environment through only gas lines with neutralizers, which include metallized fragments of the surface of the spacecraft and thermoelements rigidly placed on them and having a galvanic connection with them.

The idea of the proposed invention is to create on board the spacecraft, in contrast to covering most of the spacecraft with a cloud of its own atmosphere, as in the prototype, a closed flow of the so-called low-temperature plasma from the UK PD adjacent to the surface of the spacecraft, which then flows back to the UK for further regeneration and acceleration , that is, without RT losses for spacecraft de-electrification. The idea is new, and its implementation is fundamentally possible.

The invention is directed to the technical result - the emission of differential charging electrons from the surfaces of the spacecraft in contact with the PD plasma.

The technical result is achieved due to the fact that thermionic emission and the PD plasma flow picking up electrons create conditions for the final deportation of differential charge electrons into outer space. The end result is a reduction in the levels of electrization of the hull parts of the spacecraft.

In FIG. Figure 1 shows a functional-kinematic diagram of the RE device in cross section: one PD - one neutralizer. In FIG. 2 shows an exemplary schematic diagram of networks of gas mains, neutralizers (their minimum is given) and PD in the RE structure to eliminate differential charging.

We note once again that SKMOES is prescribed (quite rightly) to metallize the conductive elements of the structure, and to minimize the number of non-conductive elements on the outer surfaces.

RE contains:

- PD 9 (according to the staffing table for the spacecraft correction system) with systems for power supply, control and supply of RT to the engine chambers (UK 10);

- sensors for monitoring the level of electrization of 16 fragments of the surface of the spacecraft 14;

- UK 10, each of which has lateral outlet and inlet openings of gas lines 5 connecting this UK 10 with a network of neutralizers 4 located on the surface 1 of the spacecraft;

- a network of gas pipelines 5 with electrovalves 6 delivering RT plasma in the presence of thrust and semi-cylindrical intake shells 11 to neutralizers 4 to neutralize surface electrons 2 and discharge the end product of neutralization into UK 10;

- solenoid valves 6 that regulate the access of RT plasma to neutralizers 4: they can be opened during operation of PD 9 during correction of the parameters of the orbital motion of the spacecraft and closed when PD 9 is turned off;

- semi- cylindrical intake shells 11, which contribute to the forced selection of RT from UK 10 to neutralizers 4;

- neutralizers 4, representing spaces limited on all sides, communicating with the external environment by gas pipelines 5, the composition of neutralizers 4 includes metallized fragments of the surface 14 of the spacecraft with thermoelements 13 rigidly placed on them and having galvanic connections with them.

The main parts of the RE are PD 9 and neutralizers 4. UK 10 PD having RT inputs/outputs through gas lines 5 to neutralizers 4, using thermionic emission in operation - essential features that meet the criterion of "novelty". No analogue device has such a set of components.

The gas lines 5 are heat resistant.

The movement (thrust) of the plasma-like RT in the neutralizer 4 is created due to the fact that:

1. When PD 9 is turned on, a working thrust is created, and, therefore, the intake device (shell) 11 will effectively divert part of the formed low-temperature plasma into the corresponding gas main 5. Further, the plasma, passing through the distribution unit 7, enters the network of gas mains 5 and, finally, - in all neutralizers 4 RE.

2. The side outlet of the gas line 5 in UK 10 PD 9 is located closer to the anode (RT inlet from the RT supply unit (BPRT) 12) than the side inlet of the other gas line 5 in UK 10 PD 9. The pressure of the RT in stationary PD ( SPT) 9 drops rapidly from about 2–3 kgf/cm ² at the engine inlet to almost zero at the output of the SPT: SPT100 thrust (outlet diameter 100 mm) is 8.3⋅10 ^-2 N; it is 8.5⋅10 ^-3 kgf; we divide by the area of the outlet SPD100, we get the outlet pressure RT 1.1⋅10 ^-4 kgf/cm ² , that is, almost zero. With an AC length of about 0.1 m. Even adjacent openings of gas pipelines 5 give a significant circulation of plasma-like RT in the interval of about 1.5 – 2 hours of the correction time for the motion parameters of the center of mass of the spacecraft. The quality of the torch 8 will not suffer from the presence of gas pipelines at the wall of the UK 10 when entering the operating mode of the RE. And the output to the operating mode with the opening of the solenoid valve 6 occurs almost instantly.

Solenoid valves 6 on highways 5, along with the sensor 16 for monitoring the level of electrification, are necessary for flexible coupling of the work of the correction system and the electrical protection complex.

In order to maximize the efficiency of active de-electrification, the RE contains thermoelements 13, which contribute to the forced emission of electrons 2 into the RT plasma. The plasma in the neutralizers becomes more neutral, since 3 RT ions, combining with electrons 2, turn into neutral RT particles.

The surfaces 14 of the spacecraft should be any surfaces 1 of the spacecraft design, if possible, free from elements of service systems and target and scientific equipment, or other surfaces, the state of the elements of which is not critical at elevated temperatures.

Within the contact surfaces 14, thermoelements 13 should be placed on them (above them) in full and tight contact. That is, it is necessary to organize a galvanic connection so that the electrons 2, leaving the thermoelements 13, create an electric current in the conductive surfaces of the spacecraft adjacent to this neutralizer 4, in the direction of its surface 14. The number of thermoelements 13 is calculated based on the energy balance on board the spacecraft and the practical efficiency of their work.

Each neutralizer 4 has a connection with all PD 9 engines, and there are not two gas pipelines 5, but a whole network with distribution input and output nodes 7, and neutralizers 4 in the RE on board the spacecraft from 4 pieces according to the number of available PD 9 involved in the regular process maintaining the orbital position of the spacecraft. It is required to have seats on the body 1 of the spacecraft that most effectively assemble, as shown in Fig. 1 arrows 15 negative charges of differential charging of spacecraft. However, the coverage areas of each neutralizer 4 can be large enough to satisfactorily cope with the task of minimizing the electrostatic charge on board the spacecraft. The process of de-electrification of contact surfaces leads to de-electrification of the entire spacecraft.

The supply unit RT 12 is designed to provide the supply of RT (xenon) with a given pressure to PD 9.

BPRT 12 is made in the form of a monoblock, which includes the following main elements: a high pressure sensor, filters, a frame for screen-vacuum thermal insulation (EVTI), pipelines and connection elements, electrical connectors and cables, and two lines (main and backup) of pressure reduction RT and its supply to the engines. Each line consists of a pressure reducer, a receiver, a low pressure sensor, a pressure alarm and two solenoid valves (high and low pressure). The BPRT also contains a heater RT, which is necessary to prevent the ingress of the liquid phase of the RT into the reducer. To ensure the thermal regime, the block is closed from the outside with an EVTI casing.

Claims (1)

A spacecraft (SC) electrification regulator containing sensors for monitoring the level of electrization of spacecraft surface fragments and plasma engines with systems for power supply, control and supply of the working fluid to the engine chambers, characterized in that it contains accelerating channels (AC) of plasma engines, which are in the zone of creating a low-temperature plasmas are connected by side inlet gas lines, which have semi-cylindrical intake shells at the beginning in the AC, and side outlet gas lines with electrovalves located on them with neutralizers limited on all sides and communicating with the external environment through only gas lines, which include metallized fragments of the spacecraft surface and thermoelements rigidly placed on them and having a galvanic connection with them.

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