RU2441170C1 - Liquid propellant rocket engine with controlled nozzle, and unit of bank nozzles - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: proposed engine comprises structural frame, combustion chamber with head, cylindrical section and controlled nozzle with extending tip, suspension assembly to allow driven swinging in to planes, drives secured to power ring arranged on combustion chamber, gas generator and turbo pump assembly including, in its turn, turbine, oxidiser pump, fuel pump, and gas duct communicating turbine outlet with combustion chamber head via suspension assembly. In compliance with this invention, is differs from known designs in that bank nozzles are grouped into pairs and fitted on lower power ring mounted on nozzle lower section and connected with nozzle edge. Gas generator gas fee pipelines are connected via gas and fuel three-way valves to bank nozzles. Opposite ends of gas generator are connected by gas bleed pipeline and fuel pipelines. Note here that bank nozzle units are secured on lower power ring attached to extending tip. Bank nozzle unit comprises two bank nozzles integrated in one unit. Note here that said unit includes three-way gas and fuel valves fitted between said nozzles and provided with common drive. All nozzles are equipped with ignition device.

EFFECT: higher reliability of thrust vector and rocket bank control.

3 cl, 5 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.

Known liquid rocket engine according to the patent of the Russian Federation for invention No. 2095607, intended for use as part of space booster blocks, stages of rocket launchers, and as the main engine of spacecraft 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, containing, in turn, a turbine, an oxidizer pump, a fuel pump and an additional fuel pump, a gas duct connecting the outlet from a turbine with a combustion chamber, and a rocker assembly of the rocket chamber of the LRE 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 sealing the supply of gas-generating gas, which has 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 of these problems was achieved in a liquid-propellant rocket engine with an adjustable nozzle containing a power frame, a combustion chamber having a head, a cylindrical part and an adjustable nozzle with a retractable nozzle, a suspension unit that allows swinging in two planes by means of actuators attached to a power ring made on a 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 suspension unit, characterized in that the nozzle rolls are grouped into blocks of nozzle nozzles in pairs and are mounted on the lower power ring installed in the lower part of the nozzle and connected to the nozzle exit, respectively, gas supply gas supply pipelines are connected to the nozzle nozzles through three-way gas and fuel valves , the other ends of which are connected first by a gas extraction pipe and fuel pipelines, while the bank nozzle blocks are fixed on the lower power ring, which is fixed on the extension nozzle.

The solution of these problems was achieved in a block of nozzle rolls containing two roll nozzles combined into one unit, characterized in that the pair of nozzle rolls are equipped with three-way gas and fuel valves installed between the nozzle rolls and having a common drive. All roll nozzles are equipped with an ignition device

The invention is illustrated in figure 1 ... 5, where:

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

- figure 2 shows the nozzles,

- figure 3 shows the nozzles, top view,

- figure 4 shows the design of the block nozzle roll,

- figure 5 shows a section aa of figure 3.

A liquid-propellant rocket engine (FIGS. 1 ... 5) comprises a power frame 1, a combustion chamber 2, capable of swinging in two planes, a gas generator 3 and a turbopump unit 4, coupled to the gas generator 3 by means of a gas duct 5, which, in turn, comprises a turbine 6 oxidizer pump. 7, a fuel pump 8. The turbopump assembly may comprise 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 19 to the power frame 1 and having rods 19. On the combustion chamber 2, for example, on its cylindrical part 12 , made the main power ring 20, to which the rods 18 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, comprising a start-off valve 22 and a main fuel bellows 23, the outlet of this pipeline is connected to the main manifold 24 of the combustion chamber 2. Exit from the pump the oxidizer 7 by the oxidizer pipe 25 containing the start-off valve of the oxidizer 26 is connected to the gas generator 3. Also, the output of the additional fuel pump 9 by the fuel pipe 27 containing the start-off valve fuel 28, connected to the gas generator 3. On the gas generator 3 and on the combustion chamber 2 are installed, 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 34 made hollow inside.

The roll angle control system (FIGS. 1 and 4) comprises at least four roll nozzles 35 installed in the form of blocks 36 of roll nozzles 35. From 2 to 4 blocks of roll nozzles can be used. The roll nozzle blocks 36 contain two roll nozzles 35 mounted on the lower power ring 37. The lower power ring 37 is installed in the region of the nozzle 13 cutoff and is rigidly connected to the nozzle nozzle 38. This power ring serves to transmit torque from the roll nozzle blocks 36 36 on the power frame 1, for this, each block 36 of nozzles of the roll 35 is attached to the lower power ring 37. At least two actuators 17 are connected to the lower power ring 17. The actuators 17 are made in the form of hydraulic cylinders 18 attached by hinges 14 to the power frame 1 or to main with silt ring 20 and having rods 19. These actuators 17 serve to control the degree of expansion of the nozzle and are similar in design to actuators 17 for controlling the rocket at pitch and yaw angles (the rocket is not shown in FIGS. 1 ... 5).

Generator gas supply pipelines 39 are connected to the roll nozzles 35, containing a bellows 40, the other ends of which are connected to the gas duct 5. In each block 36 of the roll nozzles 35, a three-way gas valve 41 is installed between them, which is connected to the gas-supply gas supply pipe 40, and a three-way valve fuel 42, to which a fuel pipe 43 is connected, extending from the main manifold 24 and containing a bellows 44. On three-way valves 41 and 42, a common drive 45 is installed on each block. Thus, every two nozzle roll 35, three-way valves 41 and 42 and the common drive 45 form one node: block 36 nozzle roll 35.

Roll nozzles 35 (FIGS. 4 and 5) are made with two walls 46 and 47 and collectors 48 for passage of cooling fuel. In each nozzle of the roll 35 installed nozzles of fuel 49, oxidizing agent 50 and ignition device 51.

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 to the oxidizer and fuel rocket valves via electric communication channels 31 (rocket valves are not shown in FIG. 1). After priming the oxidizer 7 and fuel 8 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. Gas generator gas and fuel are fed into 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 a regenerative cooling path (figure 1), goes into the inner cavity of the chambers s of combustion 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 ... 5), 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 through the bellows 40 through three-way valves 41 enters the block 36 of the nozzle 35.

To control the thrust vector R using the drive 17, acting on the main force ring 20 by the rod 18, the combustion chamber 2 is rotated relative to the point “O” by an angle of 5 ... 7 °. In this case, 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 ... 4).

To control the rocket on which the engine is mounted, a command is sent from the control unit 30 to the common drives 45, while one roll nozzle 35 from each pair is turned on and their jet thrust creates a torque that is transmitted through the lower power ring 37 to the nozzle 13 first , then - to the power frame 1 and further to the rocket body (the rocket in figure 1 ... 5 is not shown).

When climbing, for example, more than 20 km, a command is sent to the drives 17 connected to the lower power codec 37, which together with the nozzle 38 moves to the lower position, while the degree of expansion of the nozzle 13 increases and the thrust force increases by 7 ... 10%.

The application of the invention allowed:

1. To provide reliable control of the thrust vector of the rocket engine, the degree of expansion of the nozzle depending on the height of the flight of the rocket, and the control of the rocket along the roll angle by using two blocks of roll nozzles containing two opposite mounted roll nozzles, and their rational mounting on the engine on the annular manifold the use of four inclined rods, providing torque transmission to the engine nozzle and further to the power frame with the minimum weight of structural elements transmitting the torque.

2. Significantly improve the reliability of the rocket control system by roll through the use of two three-way valves: gas and fuel and a common drive for them. This design prevents the inclusion of one of the nozzles of the roll, for example, due to a failure of the start-off fuel valve.

Claims (3)

1. A liquid-propellant rocket engine with an adjustable nozzle, comprising a power frame, a combustion chamber having a head, a cylindrical part and an adjustable nozzle with a retractable nozzle, a suspension unit that allows swinging in two planes by means of actuators attached to a power ring made on the combustion chamber, a gas generator and a turbopump assembly, comprising, in turn, 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 suspension unit The fact that the nozzle rolls are grouped into blocks of nozzle rolls in pairs and are mounted on the lower power ring installed in the lower part of the nozzle and connected to the nozzle cut, to the nozzles of the roll through three-way gas and fuel valves is connected respectively to the gas supply gas pipelines, the other ends of which are connected first, a gas extraction pipe and fuel pipelines, while the roll nozzle blocks are fixed to the lower power ring, which is fixed to the extension nozzle. 2. A block of roll nozzles comprising two roll nozzles integrated into one assembly, characterized in that the pair of roll nozzles is equipped with three-way gas and fuel valves installed between the roll nozzles and having a common drive. 3. The roll nozzle block according to claim 2, characterized in that all the roll nozzles are equipped with an ignition device.

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