RU2734686C1 - Control method of solid-propellant propulsion system of spacecraft and device for its implementation - Google Patents

Abstract

FIELD: aerospace engineering.

SUBSTANCE: invention can be used in development of propulsion systems for controlled spacecraft. Method for control of solid-propellant propulsion system consists in change of solid-propellant propulsion unit pulse and vector of thrust using a combination of nozzle electromagnetic control units opening and closing by commands from the control system. Gas flow under pressure created by working medium sources passing through collector is provided through said units. Manifold pressure monitoring is performed in manifold with the help of sensors. Note here that gas pressure accumulation and stabilization in storage receiver is carried out at outlet of working medium sources. After that, gas pressure is reduced by means of step-down mechanical reduction gear separating the header into high and low pressure circuits. Control system controls pressure in both circuits. In accordance with the control algorithm, commands are sent to actuate the second and subsequent working body sources. Commands are supplied at reduction of pressure in high pressure circuit below allowable for correct operation of reduction gear, as well as to open all nozzle electromagnetic control units with accidental exceeding of maximum allowable pressure in low pressure circuit for discharge of excess gas into environment, without affecting the control of the spacecraft.

EFFECT: technical result is higher efficiency of control of spacecraft propulsion due to smoothing dynamics and efficient use of working medium, as well as increase in the device operation reliability.

10 cl, 3 dwg

Description

The invention relates to the field of rocket and space technology, and can be used to create propulsion systems (PS) for guided spacecraft (SC).

It is known that various propulsion systems are used to control the spacecraft in flight, providing orientation and holding the spacecraft on the flight path, as well as changing the spacecraft position.

The last 10 years of space exploration is characterized by the activation of countries and private companies - participants in various projects to create spacecraft using various class of propulsion systems, using chemical fuel (liquid, solid, gel and gaseous), electric, plasma, ionic, etc. as an energy source. (see the review of the FSUE "Scientific and Technical Center of the Defense Complex" Compass ", Moscow, 2018) It is known to use a propulsion system for controlling a spacecraft in the form of a gas-jet attitude control system (GRSO) installed on a space rocket (see patent 2025645 dated 30.12. 1992 - prototype). In the patent under consideration, the GDSO is located on the nozzle block of the third sustainer stage and contains (see Fig. 3) gas generators (28), distributors (29), nozzle blocks (30) and collectors (31). can be made with solid fuel, liquefied or compressed gas. la.

The disadvantage of the GDSO control method, taken as a prototype, is that when the gas generators are used in turn, fluctuations in the flow of combustion products occur in the system, which negatively affects the performance of the regulating and control devices.

It should be noted that the manufacture of a gas distribution station, taken as a prototype, is carried out by element-by-element installation of the components of a gas distribution station on the nozzle block, which reduces the operational characteristics and complicates the installation and acceptance tests.

In addition, the disadvantages of propulsion systems using liquid or gaseous (liquefied or compressed gas) fuel in comparison with TDU are short guaranteed shelf life, reduced reliability and performance characteristics.

The objective of the proposed invention is to create a control method and a TDU design that will smooth out dynamic fluctuations and more efficiently use the working fluid reserve due to preliminary accumulation and stabilization of the gas pressure in the storage receiver, as well as increase the accuracy, reliability and operational characteristics of the TDU.

This is achieved due to the fact that in the known method of controlling the TDU, which consists in changing the impulse and thrust vector of the TDU using a combination of nozzle electromagnetic control units (ECUs), which open and close the outlet section on commands from the control system (CS) through which gas flows out under pressure, created by the sources of the working fluid and passing through the manifold, in which the pressure telemetry is monitored using a sensor, at the exit from the sequentially activated sources of the working fluid, the gas pressure in the reservoir-accumulator is accumulated and stabilized, after which the gas pressure is reduced by means of a lowering a mechanical reducer dividing the collector into high pressure (HP) and low pressure (LP) circuits, while the control system monitors the pressure in both circuits and, in accordance with the control algorithm (AC), issues commands: to activate the second and subsequent sources of the working fluid when loop pressure re HP below the permissible for the correct operation of the reducer, to open all ECMUs in case of accidentally exceeding the maximum permissible pressure in the LP circuit to discharge excess gas into the environment, without affecting the control of the spacecraft.

In addition, the gas pressure can be reduced with the possibility of changing the pressure level in the LP circuit due to the use of a reduction gear with an electromagnetic drive, the operating mode of which is changed according to the commands of the control system.

Control of permissible pressure levels in the HP and LP circuits can be carried out according to the readings of three sensors in each circuit with subsequent comparison and exclusion of deliberately false readings in the event of failure of one of the sensors, according to the AU.

Stabilization of pressure in the LP circuit directly in front of the ECM can be done using a receiver-damper.

From the point of view of design, this is achieved due to the fact that the existing TDU containing solid fuel gas generators (TTGG) as a source of the working fluid, an ECM and a manifold connecting them with a pressure sensor installed in it, is equipped with a storage receiver in the form of a container installed at the outlet of TTGG, and a reduction gear installed after the storage receiver and dividing the TDU collector into LP and HP circuits, each of which has pressure sensors connected by electrical connection to the control system.

In the LP circuit, a receiver-damper in the form of a container can be additionally installed, the inner cavity of which is connected by gas ducts to the entrance to the ECM.

Receivers-dampers can be installed in front of each ECMU, and their minimum volume is determined by the formula:

n is the number of ECBUs associated with the receiver-damper;

k - permissible pressure spread in the receiver-damper;

- секундный объемный расход продуктов сгорания через один ЭМБУ;

- second volumetric flow rate of combustion products through one computer;

τ is the minimum time the computer is in the open state. Component parts of the TDU can be mounted on a base in the form of a plate equipped with attachment points for installation in the spacecraft, while ribs and local samples are made on the plate, providing an optimal ratio of stiffness and mass characteristics.

In the HP and LP circuits, sensor units are installed, each of which can accommodate 3 pressure sensors connected by electrical connection with the control system, and one sensor in each circuit connected with the telemetric system for recording the parameters of the spacecraft.

FIG. 1 shows a schematic diagram of the TDU.

FIG. 2 shows a variant of the layout of the TDU (top view).

FIG. 3 shows a variant of the layout of the TDU (front view).

The TDU consists of an ECMU (1), a TTGG (2), a storage receiver (3) installed at the outlet of the TTGG (2), a reducer (4) installed at the outlet of a storage receiver (3) and dividing the collector into HP circuits (6) and LP (7), sensor units (5) located in the HP (6) and LP (7) circuits, the receiver-damper (9) and the base (8), on which the TDU elements are installed.

TDU works as follows:

On command from the control system, the first TTGG (2) is triggered, the combustion products of which are fed into the storage receiver (3) of the HP circuit (6). Then they pass through the pressure reducer (4) and flow through the LP circuit manifold (7) to the ECMU (1), which open / close on the commands of the control system and create control forces to correct the spacecraft flight trajectory. The block of sensors (5) in the HP circuit (6) controls the gas pressure in the storage receiver (3), which gradually decreases as the combustion products are consumed through the ECM (1), and when the minimum pressure level allowed for the correct operation of the reducer (4) is reached gas, SU gives a command to start the second and subsequent TTGG (2). The sensor unit (5) in the LP circuit (7) monitors the gas pressure level, and if the gas pressure is accidentally exceeded, the control unit gives the command to simultaneously open all ECMUs (1) to release the pressure. Since the thrust vector of all ECMUs (1) is directed in the radial direction, and all of them are symmetrically pairwise relative to the SC axis, their simultaneous opening does not affect the SC control. The receiver-damper (9) stabilizes the pressure in the LP circuit (7), smoothing out the dynamic oscillations arising as a result of transient processes when opening / closing the gearbox (4), in order to maintain the highest possible accuracy of the thrust level created by the ECMU (1). The minimum volume Vmin of the receiver-damper (9) is determined by the formula:

n is the number of ECUs associated with the receiver-damper;

k - permissible pressure spread in the receiver-damper;

- секундный объемный расход продуктов сгорания через один ЭМБУ;

- second volumetric flow rate of combustion products through one computer;

τ is the minimum time the computer is in the open state.

Each block of sensors (5) contains three sensors for the control system and one sensor for telemetric pressure recording. The installation of three sensors for the control system makes it possible to compose the AU in such a way that in the event of a critical difference in the readings of one sensor relative to the other two due to the failure of the first one, the readings of the failed sensor do not interfere with the correct operation of the control system.

The reducer (4) can be made with an electromagnetic drive connected by electrical connection with the control system. In this case, the opening / closing of the flow section of the gearbox (4) is carried out by an electromagnet, the commands for which are received from the control system according to the readings of the HP and LP sensors.

The combination of the described method and design of the TDU allows to increase the accuracy and reliability of the TDU functioning, to minimize the gas-dynamic oscillatory processes. The use of TTGG in TDU makes it possible to maximize the warranty periods of the TDU, as well as to exclude refueling and control operations before the start, and due to the sequential switching on of the TSGG and multiple filling of the storage receiver with gas, to reduce the weight of the TDU. Sequential switching on of the TTGG followed by gas stabilization in the storage receiver allows meeting the requirements for the total readiness time of the TDU to create a thrust impulse through the ECM, while the total readiness time can significantly exceed the total active operation time.

The use of a reduction gear with an electromagnetic drive allows you to switch to a reduced thrust mode for more accurate control of small-sized objects at the finishing section, which ultimately increases control accuracy, maneuverability and accuracy of dosing the total thrust impulse of the TDU. A positive effect in terms of control accuracy is achieved due to the fact that the accuracy of the sensor (up to ± 3%) is significantly higher than the spread of the parameters of the regulating body, for example, a spring (≈ ± 10%).

The installation of three sensors in each loop and the use of the comparison algorithm makes it possible to exclude knowingly false readings in the event that one of the sensors fails.

Installing a receiver-damper in front of the EMBU makes it possible to smooth out oscillations in the LP circuit that occur when opening / closing the flow section of the gearbox, which ultimately ensures the EMBU thrust accuracy up to ± 2%.

Installation of all elements of the TDU on a single base simplifies its installation in the compartment as much as possible, allowing in advance to perform outside the spacecraft compartment such operations as bending and fixing the position of gas ducts, checking the tightness of the TDU, etc., and also simplifies bench testing and acceptance tests ... The base is made in the form of a plate equipped with attachment points for installation in the spacecraft. The plate has ribs and local grooves that provide the optimal ratio of stiffness and mass characteristics.

The TDU of the proposed design is planned to be used in the creation of advanced spacecraft.

Claims (15)

1. A method for controlling a solid-propellant propulsion system of a spacecraft, which consists in changing the impulse and thrust vector of a solid-propellant propulsion system using a combination of nozzle electromagnetic control units that open and close on commands from the control system, through which gas flows out under pressure created by sources of the working fluid passing through the manifold, in which the telemetric pressure control is carried out using sensors, characterized in that at the outlet from the sequentially used sources of the working fluid, the gas pressure in the reservoir is accumulated and stabilized, after which the gas pressure is reduced using a mechanical reducing reducer separating the manifold to the high and low pressure circuits, while using the control system, the pressure in both circuits is monitored and, in accordance with the control algorithm, commands are given to activate the second and subsequent working fluid sources when the pressure in the high pressure circuit drops below the permissible for the correct operation of the reducer, as well as on the opening of all nozzle electromagnetic control units in case of accidentally exceeding the maximum permissible pressure in the low pressure circuit to discharge excess gas into the environment, without affecting to control the spacecraft. 2. A method for controlling a solid-propellant propulsion system of a spacecraft according to claim 1, characterized in that the gas pressure is reduced with the possibility of changing the pressure level in the low pressure circuit by using a reduction gearbox with an electromagnetic drive, the operating mode of which is changed according to the commands of the control system. 3. A method for controlling a solid-propellant propulsion system of a spacecraft according to claim 1, characterized in that the control of the permissible pressure levels in the high and low pressure circuits is carried out according to the readings of three sensors in each circuit and subsequent comparison and exclusion of deliberately false readings in the event of failure of one of the sensors, according to the control algorithm. 4. A method for controlling a solid-propellant propulsion system of a spacecraft according to claim 1, characterized in that pressure stabilization in the low-pressure circuit directly in front of the electromagnetic control units is performed using a receiver-damper. 5. Solid-propellant propulsion system of the spacecraft, containing as a source of the working fluid solid-propellant gas generators, nozzle electromagnetic control units, a manifold connecting them with pressure sensors installed in it and a control system that provides opening and closing of the outlet sections of the nozzles of electromagnetic control units, characterized in that The solid-fuel propulsion system is equipped with a storage receiver in the form of a container installed at the outlet of the solid-fuel gas generators and a reduction gear installed after the storage receiver and dividing the manifold into low and high pressure circuits, each of which has pressure sensors connected by electrical connection to the system control, which provides control of the propulsion system according to the readings of pressure sensors and the algorithm of the control system, while the thrust vector of all nozzles of the electromagnetic control units is directed in the radial direction and all they are in pairs symmetrically relative to the axis of the spacecraft. 6. The solid-propellant propulsion system of the spacecraft according to claim 5, characterized in that a receiver-damper in the form of a container is additionally installed in the low pressure circuit, the internal cavity of which is connected by gas ducts to the entrance to the nozzle electromagnetic control units. 7. Solid-propellant propulsion system of the spacecraft according to claim 5, characterized in that the receiver-dampers are installed in front of each nozzle electromagnetic unit and their minimum volume (V _min ) is determined by the formula:

n is the number of electromagnetic control units associated with the receiver-damper; w - second volumetric flow rate of combustion products through one nozzle electromagnetic control unit; τ is the minimum time the nozzle electromagnetic unit is in the open state; k - permissible pressure spread in the receiver-damper. 8. Solid-propellant propulsion system of the spacecraft according to claim 5, characterized in that the gearbox is made with an electromagnetic drive connected by electrical connection with the control system. 9. Solid-propellant propulsion system of the spacecraft according to claim 5, characterized in that its component parts are mounted on a base in the form of a plate equipped with attachment points for installation in a spacecraft, while ribs and local samples are made on the plate, providing an optimal ratio of stiffness and mass characteristics. 10. Solid-propellant propulsion system of the spacecraft according to claim 5, characterized in that sensor units are installed in the high and low pressure circuits, in each of which there are 3 pressure sensors connected electrically to the control system, and one sensor in each circuit connected with the telemetric system for recording the parameters of the propulsion system.

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