RU2420669C1 - Liquid propellant engine with controlled thrust vector, and nozzle block of roll - Google Patents

Abstract

FIELD: engines and pumps. ^ SUBSTANCE: liquid propellant engine with controlled thrust vector includes power frame, combustion chamber having head, cylindrical part and nozzle, which is fixed on power frame by means of suspension assembly providing the possibility of rolling in two planes by means of drives attached to power ring made on combustion chamber, gas generator and turbo pump unit which in its turn contains turbine, pump of oxidiser, fuel pump, gas line connecting the turbine outlet to head of combustion chamber through suspension assembly, according to invention the roll nozzles are grouped into blocks of roll nozzles in pairs and mounted on annular header installed in lower part of nozzle, which has diametre which is more than that of nozzle section; gas generator gas removal pipeline is connected to annular header, and the other end of the above pipeline is connected to gas passage; annular header is attached by means of two articulated rods to turbo pump unit and by means of two articulated rods to damper bracket installed on gas generator gas removal pipeline. Roll nozzle block contains two roll nozzles combined into one assembly; at that, pair of roll nozzles is opposite installed and equipped with three-way valve installed between them; three-way valve has the drive connected to shaft. ^ EFFECT: improved control of thrust vector and control of rocket as to roll. ^ 5 cl, 3 dwg

Description

The invention relates to rocket technology, specifically to liquid-propellant rocket engines made in a closed circuit with afterburning of gas-generating gas, and is intended to control the thrust vector of the engine.

The well-known liquid rocket engine according to the patent of the Russian Federation for invention No. 2095607, intended for use as a part of space booster blocks, stages of rocket launchers and as a main engine of space vehicles, includes a combustion chamber with a regenerative cooling path, pumps for supplying components - fuel and an oxidizer with a turbine one shaft into which the capacitor is inserted. The condenser outlet through the refrigerant line is connected to the entrance to the combustion chamber and to the entrance to the regenerative cooling path of the combustion chamber.

The disadvantage of this engine is the lack of thrust vector control.

Known liquid rocket engine according to the patent of the Russian Federation for the invention No. 2187684. The method of operation of a liquid-propellant rocket engine is to supply fuel components to the combustion chamber of the engine, gasify one of the components in the cooling path of the combustion chamber, supply it to the turbine of the turbopump unit, and then discharge it into the nozzle head of the combustion chamber. Part of the flow rate of one of the fuel components is directed to the combustion chamber, and the remaining part is gasified and directed to turbines of turbopump units. The gaseous component spent on the turbines is mixed with the liquid component entering the engine at a pressure higher than the saturated vapor pressure of the resulting mixture.

The disadvantage of this scheme is that the thermal energy removed during cooling of the combustion chamber may not be enough to drive a turbopump engine unit of very high power.

Known LRE according to the patent of the Russian Federation for the invention No. 2190114, IPC 7 F02K 9/48, publ. 09/27/2002 This LPRE includes a combustion chamber with a regenerative cooling path, a TNA turbopump unit with oxidizer and fuel pumps, the output lines of which are connected to the head of the combustion chamber, the main turbine and the drive circuit of the main turbine. The main turbine drive circuit includes a fuel pump and a regenerative cooling path of the combustion chamber connected in series with each other and connected to the main turbine inlet. The exit from the turbine TNA is connected to the input of the second stage of the fuel pump.

This engine has a significant drawback. Bypassing the fuel combustion chamber heated in the regenerative cooling path to the entrance to the second stage of the fuel pump will lead to cavitation. Most LREs use fuel components such that the oxidizer consumption is almost always greater than the fuel consumption. Therefore, for powerful rocket engines with great thrust and high pressure in the combustion chamber, this scheme is not acceptable, because fuel consumption will not be enough to cool the combustion chamber and drive the main turbine. In addition, the LRE launch system, the ignition system of the fuel components and the LRE shutdown system and its cleaning of fuel residues in the regenerative cooling path of the combustion chamber have not been developed.

Known liquid rocket engine according to the patent of the Russian Federation for the invention No. 2232915, publ. September 10, 2003, which contains a chamber, a turbopump unit, a gas generator, a launch system, means for igniting fuel components and fuel lines. The output of the oxidizer pump is connected to the inlet of the gas generator. The output of the first stage of the fuel pump is connected to the channels of regenerative cooling of the chamber and to the mixing head. The output of the second stage of the fuel pump is connected to an electric flow regulator.

The disadvantage is that the engine does not have a thrust vector control system and roll control.

Known liquid rocket engine and TNA according to the patent of the Russian Federation for the invention No. 2161263, prototype.

This engine contains a power frame, a combustion chamber made with the possibility of swinging in two planes, a gas generator and a turbopump unit, coupled to the gas generator by means of a gas duct, which in turn contains a turbine, an oxidizer pump, a fuel pump and an additional fuel pump, a gas duct connecting the turbine exit with a combustion chamber, and a rocket assembly of a rocket engine combustion chamber installed between the gas duct and the combustion chamber, more precisely, the head of the combustion chamber. This unit is made in the form of a bellows and a universal joint, which together provide the swing of the combustion chamber and the sealing of the supply of gas-generating gas having high pressure and temperature. In addition, a bellows cooling system is provided, since its performance under such extreme conditions is in doubt.

The turbopump assembly comprises a turbine with an impeller and oxidizer, fuel and additional fuel pumps mounted coaxially to the pump.

The disadvantages of this engine and the suspension unit of the combustion chamber included in its composition: low unreliability of the suspension unit of the combustion chamber of the rocket engine due to the presence of a large number of parts, low strength of thin-walled bellows operating at high pressure and temperature. Gimbal bearings, transmitting the thrust of the combustion chamber, reaching 200 ... 1000 tf, also work at high temperatures (from 500 to 800 ° C), while the grease burns out, the bearings are destroyed, the thrust vector control is difficult.

The use for cooling this unit of fuel, intended for feeding into the combustion chamber, not only complicates the design of this unit and the engine as a whole, but also makes its operation extremely dangerous, since when the bellows breaks, the fuel and gas-generating gas containing excess oxidizer will come into contact that will inevitably lead to a fire in the engine compartment of the rocket and the cessation of fuel supply to the combustion chamber.

The thrust vector control is unreliable, and there is no control over roll angles.

The objective of the invention is to ensure the reliability of the thrust vector control of the rocket engine and rocket control over the roll.

The solution to these problems was achieved due to the fact that a liquid-propellant rocket engine with a controlled thrust vector containing a power frame, a combustion chamber having a head, a cylindrical part and a nozzle, which is mounted on the power frame using a suspension unit that allows swinging in two planes by means of drives attached to a power ring made on the combustion chamber, a gas generator and a turbopump assembly, which in turn contains a turbine, an oxidizer pump, a fuel pump, a gas duct connecting the outlet of turbines with a combustion chamber head through a suspension unit, while the bank nozzles are grouped into bank nozzle blocks in pairs and are mounted on an annular manifold installed in the lower part of the nozzle and having a diameter larger than the nozzle cut diameter; a gas generation gas extraction pipe is connected to the annular manifold, the other end of which connected to the gas duct, the annular manifold is mounted using two hinged rods to the turbopump assembly, two hinged rods to the damping bracket mounted on the pipe Gadfly selection gasification gas. In the vicinity of the critical section of the combustion chamber, an additional force ring is mounted concentrically to it and with a gap, to which a gas-generating gas extraction pipe is attached through a damping support and two articulated rods are attached, the ends of which are connected to the annular collector.

The solution of these problems was achieved in a block of nozzles of a bank containing two nozzles of a bank combined into one node, while a pair of nozzles of a bank was installed opposite, equipped with a three-way valve installed between them and having a drive connected to it by a shaft.

A fuel pipe having a start-off valve is connected to each roll nozzle from the fuel pump. All roll nozzles are equipped with an ignition device. The invention is illustrated in figure 1 ... 3, where:

- figure 1 shows a diagram of a liquid rocket engine,

- figure 2 shows a top view,

- figure 3 shows the design of the block nozzle roll.

A liquid-propellant rocket engine (FIGS. 1 ... 3) contains a power frame 1, a combustion chamber 2, which is designed to swing in two planes, a gas generator 3 and a turbopump 4, coupled to the gas generator 3 by means of a gas duct 5, which in turn contains a turbine 6, a pump oxidizer 7, fuel pump 8. The turbopump assembly may include an additional fuel pump 9.

The exit from the fuel pump 8 is connected by a pipe 10 to the entrance to the additional fuel pump 9 (if any). The combustion chamber 2 contains a head 11, a cylindrical part 12 and a nozzle 13. The gas generator 3 is mounted on the power frame 1 with a hinge 14, and the TNA 4 with two hinged rods 15. Between the gas duct 5 and the combustion chamber 2, more precisely, its head 11 is installed the suspension assembly 16 of the combustion chamber 2. It provides the swing of the combustion chamber 2 in two planes relative to the point "O", to control the thrust vector R.

For this, the engine contains two drives 17 mounted in mutually perpendicular planes of the combustion chamber 2, made, for example, in the form of hydraulic cylinders 18, hinged 14 to the power frame 1, and having rods 19. On the combustion chamber 2, for example, on its cylindrical part 12, the main power ring 20 is made, to which the rods 19 of the actuators 17 are pivotally attached. The actuators 17 serve to control the rocket at the pitch and yaw angles.

A possible pneumohydraulic scheme of the liquid propellant rocket engine is shown in Fig. 1 and contains a fuel pipe 21 connected at one end to the outlet of the fuel pump 8, containing a start-off valve 22 and a bellows 23, the outlet of this pipeline is connected to the main manifold 24 of the combustion chamber 2. Exit from the oxidizer pump 7 the oxidizer pipe 25 containing the start-off valve of the oxidizer 26 is connected to the gas generator 3. Also, the output from the additional fuel pump 9 the fuel pipe 27 containing the start-off valve of the fuel 28 is connected to gas generator 3. On the gasifier 3 and the combustion chamber 2 are set, at least one ignition device 29.

The engine is equipped with a control unit 30, which is electrically connected 31 to the ignition devices 29 and to the shut-off valves 22, 26 and 28.

A feature of the engine (FIGS. 1 and 2) is that the TNA 4 is rigidly fixed to the power frame 1 using at least three articulated rods 15, and the combustion chamber 2 has the ability to rotate relative to the point “O”.

The suspension assembly 16 of the combustion chamber 2 of the LRE contains two parts: the fixed 32 and the movable 33. The fixed part 32 is rigidly connected to the gas duct 5, and the movable part 33 is rigidly connected to the head 11 of the combustion chamber 2 due to the fact that both parts form a spherical joint, executed hollow.

The roll angle control system (FIGS. 1 and 3) comprises four roll nozzles 35 mounted as a roll nozzle block. The roll nozzle blocks contain two roll nozzles 35 mounted on the lower power ring 36. The lower power ring 36 is installed in the region of the nozzle exit 13 and is rigidly connected to it. This power ring serves to transmit torque from the nozzles of the roll 36 to the power frame 1, for this, each block of nozzles of the roll 35 is connected to the lower power ring using two inclined rods 37. Gas supply gas supply pipelines 38, the other ends of which are connected to the nozzles of the roll 35 connected to the gas sampling pipe 39 and then to the gas duct 5. In the block of nozzles of the roll 35, a three-way valve 40 is installed between them, which is connected by a pipe 39 to the supply pipe of the gas-generating gas 38. A th tube is connected to each nozzle of the roll 35 the throttle 41, coming from the fuel pump 8 and containing a start-off valve 42. Actuators 43 are installed on the three-way valves 40. Thus, each two nozzles of the bank 35, the three-way valve 40 and the actuator 43 form one unit: the block of nozzles of the bank. The nozzle roll 35 (figure 3) is made with two walls 44 and 45 and manifolds (not shown) for the passage of cooling fuel. In each nozzle of the roll 35, fuel nozzles 47, an oxidizer 48 and an ignition device 49 are installed. Start-off valves 42 can be installed on the manifolds (not shown in FIGS. 1–3).

The engine starts as follows.

In the initial position, all engine valves are closed. When starting a liquid propellant liquid propellant rocket from a control unit 30, an instruction is sent via electric communication channels 31 to the rocket valves of the oxidizer and fuel (rocket valves are not shown in FIG. 1). After priming the oxidizer 7 and fuel 18 pumps, the start-off valves 22, 26 and 28 are installed, installed behind the oxidizer 7, after the fuel pump 8 and after the additional fuel pump 9. The oxidizing agent and fuel enter the gas generator 3, where they are ignited using the igniter 29. The gas generator gas and fuel are supplied to the combustion chamber 2. The fuel cools the combustion chamber 2, passing through the gap, between the shells of its nozzle 13 and the cylindrical part 12 forming the regenerative cooling path (Fig. 1), enters the inner cavity of the chamber Combustion rods 2 for afterburning the gas generating gas coming from the gas generator 3. Ignition of these components is also carried out by the ignition device 29 mounted on the combustion chamber 2

After the start of the turbopump unit 4, the gas-generating gas is supplied from the gas-generator 3 to the turbine 6, the TNA rotor is untwisted (not shown in FIGS. 1 ... 3), the pressure at the outputs of the pumps 7, 8 and 9 increases. Further, through the gas duct 5 and through the suspension assembly 16, the gas-generating gas is supplied to the head 11 of the combustion chamber 2. A part of the gas-generating gas is taken through the gas sampling pipe 39 and then through the pipelines 38 and through three-way valves enters the bank nozzle blocks.

To control the thrust vector R, using the actuator 17 acting on the power ring 20 by the rod 19, rotate the combustion chamber 2 relative to the point "O" at an angle of 5 ... 7 °. The direction of the thrust vector R1 deviates relative to the initial position R1 of the longitudinal axis of symmetry of the combustion chamber 2 and relative to the rocket on which this engine is mounted (the rocket is not shown in FIGS. 1 ... 3).

To control the rocket on which the engine is mounted, a command is sent from the control unit 30 to the actuators 43 and start-off valves 42, while one roll nozzle 35 from each pair is turned on and their jet thrust generates a torque that through the lower power ring 36 and through four inclined thrusts 37 are transmitted first to the nozzle 13, then to the power frame 1 and then to the rocket body (the rocket is not shown in FIGS. 1 ... 3).

The application of the invention made it possible to ensure reliable control of the thrust vector of the rocket engine and roll control over the angle of roll due to the use of two blocks of roll nozzles containing two opposed roll nozzles, and their rational mounting on the engine on the annular manifold and the use of four inclined rods providing torque transmission to the engine nozzle and further to the power frame with a minimum weight of structural elements transmitting the torque.

Claims (5)

1. A liquid-propellant rocket engine with a controlled thrust vector, comprising a power frame, a combustion chamber having a head, a cylindrical part and a nozzle, which is mounted on the power frame using a suspension unit that allows swinging in two planes by means of actuators attached to the power ring made on the combustion chamber, a gas generator and a turbopump assembly, which in turn contains a turbine, an oxidizer pump, a fuel pump, a gas duct connecting the outlet of the turbine to the head of the combustion chamber through a unit under suspensions, characterized in that the nozzle rolls are grouped into blocks of nozzle rolls and mounted on an annular collector installed in the lower part of the nozzle, which has a diameter greater than the diameter of the nozzle cut, a gas-extraction gas extraction pipe is connected to the annular collector, the other end of which is connected to the gas duct The annular manifold is mounted using two hinged rods to the turbopump assembly and two hinged rods to a damping bracket mounted on the gas-generating extraction pipe th gas. 2. The liquid rocket engine according to claim 1, characterized in that in the critical section of the combustion chamber, an additional force ring is mounted concentrically to it and with a gap, to which a gas generation gas extraction pipe is attached through a damping support and two articulated rods are connected, the ends of which are connected to ring collector. 3. The liquid propellant rocket engine according to claim 1 or 2, characterized in that a fuel feed pipe is connected to each roll nozzle from the fuel pump, each of which is equipped with a start-and-shut off valve. 4. The liquid rocket engine according to claim 1 or 2, characterized in that all the nozzle roll are equipped with an ignition device. 5. The block of nozzles of the roll, containing two nozzles of the roll, combined in one node, characterized in that the pair of nozzles of the roll installed opposite and equipped with a three-way valve installed between them, a three-way valve has a drive connected to the shaft.

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