RU2458245C1 - Liquid-propellant rocket engine and turbopump unit - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: proposed engine comprises bearing frame, combustion chamber with cylindrical part and nozzle secured on bearing frame by suspension assembly to swing in two planes and driven by drives attached to force ring made on combustion chamber. gas generator and turbopump unit including turbine, oxidiser and fuel pump, extra fuel pump and gas line communicating turbine outlet with combustion chamber head via suspension assembly. Note here that fuel pump comprises two independent stages with independent interconnected shafts coupled with extra fuel pump by device designed to transfer torque and vary rpm. Said device is arranged in housing with feed and discharge channels. Note also that turbine shafts and those of fuel pumps of both stages are arranged concentrically to make annular chamber feed channel there between while discharge channel is communicated with pump inlet. Note that said device designed to transfer torque and vary rpm is made up of magnetic gearing comprising discs with permanent magnets arranged in circle with constant pitch.

EFFECT: higher reliability.

2 cl, 3 dwg

Description

The group of inventions relates to rocket technology, specifically to liquid-propellant rocket engines made in a closed circuit, with afterburning of gas-generating gas, and to turbo-pumping units of the LRE.

Known liquid rocket engine according to the patent of the Russian Federation for invention No. 2095607, intended for use in 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 LRE according to the patent of the Russian Federation for the invention No. 2190114, IPC ⁷ F02K 9/48, publ. September 27, 2002. This liquid-propellant rocket engine 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 №2161263, prototype rocket engine.

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 rocking unit of the 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 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 not reliable, and there is no control at all in the corners of the roll.

A known turbopump assembly with a gearbox lubrication system, where the gearbox is partially filled with grease so that one of the gears of the gearbox is immersed in the grease and spread the grease to the remaining gears and spray it (see UK patent No. 1281362, N. CL F1C, 1972).

A disadvantage of the known turbopump unit is the low efficiency of lubrication and cooling at high loads.

Closest to the invention is a turbopump assembly comprising a turbine and a pump, the shafts of which are interconnected by a gearbox placed in a housing with inlet and outlet channels (see US patent No. 3269317, N. CL. 417-405, 1966).

A disadvantage of the known turbopump unit is the low efficiency of lubrication and cooling of the gearbox.

Known turbopump assembly according to the patent of the Russian Federation for the invention No. 2318129 (prototype).

This turbopump assembly comprises a turbine and an oxidizer pump, a fuel pump, an additional fuel pump, the shafts of which are interconnected using a device for transmitting torque and changing the speed.

The disadvantage is the low reliability and fire hazard of the fuel pump design due to heating of the fuel cooling the gearbox.

The objective of the invention is to ensure the reliability of the turbopump engine rocket engine.

 The solution to this problem was achieved in a liquid-propellant rocket engine 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 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, an additional fuel pump, a gas duct connecting the turbine outlet to the head of the combustion chamber through the suspension unit, in that according to the invention, the fuel pump is made two-stage, comprising independent stages with its own shafts, which are connected to each other and to the shaft of the additional fuel pump using a device for transmitting torque and changing the speed.

The solution to this problem was achieved in a turbopump assembly containing a turbine and an oxidizer pump, a fuel pump, an additional fuel pump, the shafts of which are interconnected using a device for transmitting torque and changing the speed, so that according to the invention the fuel pump is made two-stage, the device for transmission of torque and changes in speed are located in a housing with inlet and outlet channels, while the shafts of the turbine and both stages of the fuel pump are located trically with the formation of an annular volost between them, which serves as the supply channel, and the discharge channel is in communication with the pump inlet, while a magnetic transmission containing disks with permanent magnets placed around a circle with a constant pitch is used as a device for transmitting torque and changing the speed.

The invention is illustrated in figure 1 ... 3, where

- figure 1 shows the layout of the liquid rocket engine in the side rocket unit,

- figure 2 is a diagram of a fuel pump and an additional fuel pump,

- figure 3 shows a diagram of the node of the swing of the combustion chamber.

A liquid-propellant rocket engine (FIGS. 1 ... 3) comprises a combustion chamber 1, capable of swinging in two planes, a gas generator 2 and a turbopump assembly 3, coupled to the combustion chamber 1 by means of a gas duct 4. The turbopump assembly, in turn, comprises a turbine 5, oxidizer pump 6, fuel pump 7. Turbopump assembly 3 also contains an additional fuel pump 8.

The output of the fuel pump 7 is connected by a pipe 9 to the entrance to the additional fuel pump 8. The combustion chamber 1 contains a head 10, a cylindrical part 11 and a nozzle 12. The gas generator 2 is mounted on the power frame 13 using a hinge 14, and the TNA 3 using two hinged rods 15. Between the gas duct 4 and the combustion chamber 1, or rather its head 10, a suspension assembly 16 of the combustion chamber 1 is mounted. It provides the swing of the combustion chamber 1 in one plane relative to the center of the suspension assembly 16 for controlling the thrust vector R, in order to control the launch vehicle in angle m pitch and yaw.

For this, each liquid rocket engine contains actuators 17, made, for example, in the form of hydraulic cylinders 18, attached to the power frame 13 by means of a hinge 19 and having rods 20. On the combustion chamber 1, for example on its cylindrical part 11, the main power ring 21 is made to which the rods 20 of the actuators 17 are pivotally attached. The actuators 17 serve to control the launch vehicle at pitch angles and yaw.

A possible pneumohydraulic scheme of the liquid propellant rocket engine is shown in Fig. 1 and contains a fuel pipe 22 connected at one end to the outlet of the fuel pump 7, containing a start-shut-off valve 23 and a bellows 24, the output of this pipe is connected to the main manifold 25 of the combustion chamber 1. Exit from the pump the oxidizer 6 by the oxidizer pipe 26 containing the start-off valve of the oxidizer 27 is connected to the gas generator 2. Also, the output of the additional fuel pump 8 by the fuel pipe 28 containing the start-off valve of the fuel 29 is connected to gas a generator 2. On the gas generator 2 and on the combustion chamber 1 are installed at least one ignition device 30.

The engine is equipped with a control unit 31, which is connected by electrical connections 32 to the ignition devices 30 and to the shut-off valves 23, 27 and 29.

A feature of the engine (Fig. 1 ... 3) is that the TNA 3 is rigidly fixed to the power frame 13 with at least three articulated rods 15, and the combustion chamber 1 has the ability to rotate relative to the center of the suspension assembly 16 in one plane.

A feature of the turbopump assembly 3 (Fig. 2) is that the fuel pump 7 is made of a two-stage, containing two independent (with different speeds rotating) stages 33 and 34 mounted respectively on the outer and inner shafts 35 and 36. The fuel pump 7 has an inlet pipe 37, a housing 38, inside which a device is installed for transmitting torque and changing the rotational speed 39, for example a gearbox, and an opening 40 connecting the internal cavity 41 of the housing 38 with the input cavity 42. In addition, the device is connected to a fuel fuel pump 8, more precisely, with its centrifugal wheel 43. An annular gap 44 is made between the shafts 35 and 36 to supply fuel to the internal cavity 41 for cooling the device for transmitting torque and changing the rotational speed 39. As a device for transmitting torque and changes in the rotational speed 39 it is proposed to use a magnetic transmission containing instead of gears cylinders 45 with permanent magnets 46, placed around a circle with a constant pitch, which will reduce the heating of fuel cooling for transmitting torque and changing rotational speed 39 and will reduce fire hazard.

The suspension assembly 16 of the combustion chamber 1 of the LRE (Fig. 3) contains two parts: the stationary 47 and the movable 48. The fixed part 47 is rigidly connected to the gas duct 4, and the movable part 48 is rigidly connected to the head 10 of the combustion chamber 1, due to the fact that both the parts form a spherical joint 49 made hollow inside.

The engine starts as follows.

In the initial position, all engine valves are closed. When starting a liquid propellant liquid propellant rocket engine from a control unit 31 through an electric communication channel 32, a command is given to the oxidizer and fuel rocket valves (rocket valves are not shown in FIGS. 1 ... 3). After filling the oxidizer pumps 6 and fuel 7 and 8, the start-off valves 23, 27 and 29 (Fig. 13) are opened, installed behind the oxidizer pump 6, after the fuel pump 7 and after the additional fuel pump 8. The oxidizing agent and fuel enter the gas generator 2, where ignited using igniters 30. Gas-generating gas and fuel is supplied to the combustion chamber 1. The fuel cools the combustion chamber 1, passing through the gap, between the shells of its nozzle 12 and the cylindrical part 11 forming the regenerative cooling path (Fig. 1), goes into the inner awn combustion chamber 1 for afterburning gasification gas coming from the gasifier 2. The ignition of these components are also ignition devices 30 mounted on the combustion chamber 1.

After starting the turbopump unit 3 (Fig. 1), the gas-generating gas is supplied from the gas-generator 2 to the turbine 5, the internal shaft 36 is untwisted, the pressure at the outputs of the pumps 6, 7 and 8 increases. Next, through the gas duct 4 and through the suspension unit 16, the gas-generating gas is supplied to the head 10 of the combustion chamber 1.

To control the launch vehicle on which the proposed liquid propellant rocket engines are installed, a command is sent from the control unit 31 (Fig. 3) to the drive 17 at the pitch and roll angles.

The application of the invention allowed:

1. Significantly increase the reliability and fire safety of the TNA and the engine as a whole through the use of magnetic transmission instead of the gearbox.

2. To provide reliable control of the thrust vector of the rocket engine and control of the launch vehicle at the angles of pitch and yaw.

Claims (2)

1. A liquid rocket engine 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 provides the ability to swing in two planes by means of actuators attached to a power ring made on the combustion chamber, a gas generator and a turbopump unit, comprising, in turn, a turbine, an oxidizer pump, a fuel pump, an additional fuel pump, a gas duct connecting the outlet of the turbine to the head of the combustion chamber through a unit suspension, characterized in that the fuel pump is a two-stage, containing independent stages with its own shafts, which are connected to each other and to the shaft of the additional fuel pump using a device for transmitting torque and changing the speed. 2. A turbopump assembly comprising a turbine and an oxidizer pump, a fuel pump, an additional fuel pump, the shafts of which are interconnected using a device for transmitting torque and changing the rotational speed, characterized in that the fuel pump is made two-stage, a device for transmitting torque and changes in the rotational speed are housed in a housing with inlet and outlet channels, and the shafts of the turbine and both stages of the fuel pump are concentric with the formation of an annular cavity between them, which serves as the supply channel, and the discharge channel is in communication with the pump inlet, while a magnetic transmission is used as a device for transmitting torque and changing the rotational speed, comprising disks with permanent magnets arranged around a circle with a constant pitch.

Images