RU2383770C1 - Liquid propellant rocket engine - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: invention relates to rocketry, particularly to liquid propellant rocket engines operating on cryogenic oxidiser and hydrocarbon fuel. Proposed engine comprises two frame-mounted combustion chambers with jet nozzle, regenerative cooling system, one integrated turbopump unit which comprises main turbine, oxidiser and fuel pumps and gas generator mounted below said main turbine. Note here that lower bearing plate accommodates nozzle extensions arranged concentric about jet nozzles. Said lower bearing plate can move along engine lengthwise axis over guide cylinder fitted perpendicular to upper bearing plate communicated, via drive, with combustion chambers. Aforesaid cooled nozzle extensions are made from composite and comprises cooling jackets arranged concentrically about nozzle extensions to make gaps communicated, via flexible pipelines, with fuel pump and fuel manifold. Afore mentioned drive can comprise actuating mechanism and at least two actuating mechanisms. The latter are arranged regularly along nozzle circumference. Actuating mechanism represents either pneumatic or hydraulic cylinder.

EFFECT: higher reliability, power output and simplified pneumohydraulic circuit.

Description

The invention relates to rocket technology, specifically to liquid rocket engines running on the components of rocket fuel: oxidizer and fuel.

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 space vehicles, includes a combustion chamber with a regenerative cooling path, pumps for supplying components - fuel and oxidizer, with a turbine on 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 exit from the condenser through the coolant is connected to the inlet to the pump of one of the components. The output from the pump of the same component is communicated with the condenser inlet through the refrigerant line. The second input of the condenser is in communication with the output of the turbine. The output of the pump of the other component is communicated with the entrance to the combustion chamber.

The disadvantage of the engine is the deterioration of the cavitation properties of the pump during condensate bypass.

A known method of operation of the liquid propellant rocket engine and a 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. A liquid propellant rocket engine contains a combustion chamber with a regenerative cooling path, fuel component feed pumps, and a turbine. The engine comprises a booster turbopump pump and a mixer installed in series in front of the feed pump of one of the fuel components of the main turbopump assembly. The pump outlet of the main turbopump assembly is connected both to the nozzle head of the combustion chamber and to the regenerative cooling path of the combustion chamber. The regenerative cooling circuit, in turn, is connected with the turbines of the main and booster turbopump units, the outputs of which are connected to the mixer.

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 unacceptable, 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. 09/10/2003 g (prototype), 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. Another input of the regulator is connected to the starting tank with standard fuel. The output from the regulator is connected to a gas generator. The outlet of the gas generator is connected to the inlet of the turbine pump unit, the outlet of which is connected to the mixing head. The flow regulator is equipped with a hydraulic actuator of the preliminary stage, which is connected through the cavitating nozzle and hydraulic relay to the starting tank with standard fuel. The hydraulic relay is connected to the second stage of the fuel pump. The throttle installed at the output of the first stage of the fuel pump is made in conjunction with a controlled valve of the preliminary stage.

The main element of the output device of a jet engine and a combined thrust engine is a jet nozzle. In the jet nozzle, gas expands from the turbine or afterburner of the gas turbine engine or from the combustion chamber (or other device for heating the working fluid) of the rocket engine, accompanied by an increase in its speed and kinetic energy. The expansion of gas in the jet nozzle occurs to ambient pressure at the design mode of the nozzle and to a pressure different from the ambient pressure at off-design nozzle modes. The rate of gas outflow from the jet nozzle of the jet engine in the nozzle design mode determines at a given flight speed the specific thrust of the engine. The rate of gas outflow from the jet nozzle of a rocket engine in the nozzle design mode determines the specific thrust of the engine, regardless of flight speed. The speed of gas outflow from the jet nozzle of modern jet engines under terrestrial static conditions reaches 1000 m / s and more for air-jet engines and up to 3000 m / s and more for rocket engines. Distinguish between adjustable and unregulated jet nozzle. The adjustable nozzle is equipped with a device for changing its cross section during engine operation. In a subsonic tapering nozzle, the regulation consists, as a rule, in changing the area of the exit section of the nozzle. In a supersonic nozzle, both the critical section area and the nozzle exit section area are subject to regulation. The adjustable nozzle is used in turbojet engines with an afterburner, as well as in some other gas turbine, jet and rocket engines. The nozzle of a rocket engine is also called the nozzle of the engine chamber or simply a nozzle and its regulation is practically not applied due to the very high temperature of the expiration of the combustion products.

The objectives of the invention: to ensure optimal operation of the rocket engine in a wide range of modes at different heights, simplifying the pneumohydraulic circuit, increasing reliability, increasing the power and characteristics of the rocket engine.

The solution to these problems was achieved due to the fact that a liquid rocket engine containing two combustion chambers mounted on the frame with a jet nozzle having a regenerative cooling system, one turbopump assembly made in the form of a single unit containing the main turbine, oxidizer and fuel pumps, gas generator, installed under the main turbine, characterized in that the nozzle nozzles are installed concentrically to the jet nozzles, made along the profile as an extension of the nozzle and having the ability to move in The longitudinal axis of the combustion chambers by means of a drive. Nozzle nozzles are made of carbon-carbon composite material. Nozzle nozzles are made cooled and contain a cooling jacket mounted concentrically with the nozzle nozzle, with the formation of a gap, the cavity of which is connected by flexible pipelines to the fuel pump and fuel manifold. The drive contains one or at least two actuators. Actuators (if there are two or more) can be placed evenly around the circumference around the nozzles. Actuators can be mounted on the frame. Actuators can be mounted on the combustion chamber. Actuators can be connected by a synchronization mechanism. An electric motor is used as an actuator. A pneumatic cylinder is used as an actuator. A hydraulic cylinder was used as an actuator.

The invention is illustrated by drawings, where:

figure 1 shows the design of the rocket engine,

figure 2 shows a view,

figure 3 shows a section along aa

figure 4 shows a section along BB

figure 5 shows a view B for nozzle nozzles made of carbon-carbon composite material,

figure 6 shows a view B for a nozzle nozzle made cooled.

A liquid-propellant rocket engine (FIGS. 1 ... 6) comprises a frame 1, two combustion chambers 2 with jet nozzles 3, and one turbopump assembly 4. The turbopump assembly 4, in turn, contains an oxidizer pump 5, a fuel pump 6, an additional fuel pump 7, and the main turbine 8, made in the upper part of the turbopump unit 4. The gas generator 9 is installed under the main turbine 8 coaxially with the turbopump unit 4. The combustion chamber 2 contains a jet nozzle 3 made of two shells 10 and 11 with a gap "G" between them, and a head 12 combustion chambers 2. Inside head 12 of the combustion chamber 2 are installed a fuel and an oxidant nozzle (not shown).

The engine contains two nozzle nozzles 13, made along the profile as a continuation of the nozzle 3 and having the ability to move along the axis of the combustion chamber 2 using the drive 14. The drive 14 may contain one or more actuators 15. It is preferable (Fig. 2) to use two actuators 15 installed symmetrically along the axis of the combustion chambers 2 between them, or three installed around the circumference (figure 2), and connect them with a synchronization mechanism. This mechanism can be made electric, or hydraulic, or mechanical (the design of this mechanism is not shown). Along the axis of symmetry of the engine, it is advisable to install a guide cylinder 16 with a travel stop 17 and a damper 18 in the lower part (figure 2).

Between the nozzle nozzle 13 and the nozzle 3 is installed conical high-temperature seal 19.

The actuators can, for example, be made in the form of pneumatic cylinders 20 having rods 21 connected to the lower power plate 22. Two nozzle nozzles 13 are mounted on the lower power plate 22 and rigidly fixed thereon. The pneumatic cylinders 20 themselves are connected to the upper power plate 23. Pneumatic cylinders 20 are connected by pipelines 24 through valves 25 to a cylinder of compressed gas 26. Instead of pneumatic cylinders, it is possible to use hydraulic cylinders. In the power plates 22 and 23, holes "D" are made to facilitate.

The output from the fuel pump 6 by a pipe 27 containing the fuel valve 28 is connected to the fuel manifold 29, and the output from the oxidizer pump 5 by the pipe 30 is connected to the inlet of the gas generator 9. The output from the additional pump 7 by the pipe 31 is also connected to the input of the gas generator 9.

Figures 5 and 6 show the junction of the nozzle nozzle 13 and the nozzle 3, while figure 5 shows the nozzle nozzle 13 made of graphite-graphite composite material, only the upper ring 32 is made of metal and can be combined with the lower power plate 22 The docking nozzle nozzle 13 with the lower part of the nozzle 3 is made to ensure tightness on the conical surface of the conical high-temperature seal 19, made, for example, of metal rubber. Figure 6 shows a cooled nozzle nozzle, which contains a cooling jacket formed by the gap "G", which is connected by flexible piping 33 and 34, respectively, with the exit of the fuel pump 6 and the fuel manifold 29. The liquid propellant rocket engine contains a control unit 35, which is electrically connected 36 connected with actuator 14 and valve 28.

When starting the rocket engine from the control unit 35, a signal is issued to start the engine (Means of ignition of the components of rocket fuel in the combustion chambers and in the gas generator are not shown in FIGS. 1 ... 6). The main turbine 8 spins the rotor TNA 4. The pressure of the oxidizing agent and fuel at the outlet of the oxidizing pump 5, the fuel pump 6 and the additional fuel pump 7 increases. The oxidizing agent and fuel enter the combustion chamber 2 and the gas generator 9. The engine started. After the rocket has climbed the altitude, the control unit 35 via electrical connections 36 provides a signal to the actuators 15, which move the nozzle nozzles 13 to an extremely low position. The length of the nozzle and the degree of expansion of the combustion products in it increases. The travel stop 17 prevents the nozzles 3 from being pulled out excessively, and the damper 18 reduces the impact loads on the nozzles 3 when the nozzles 13 extend to an extremely low position. Conical high-temperature seals 19 prevent the leakage of hot combustion products into the gaps between the nozzles 3 and the nozzle nozzles 13. The combustion products flowing from the nozzle expand further in the nozzle nozzles 13 to the ambient pressure, creating an additional traction force (up to 10 ... 15%) of the nominal thrust of the rocket engine without increasing fuel consumption. This leads to an improvement in the specific characteristics of the LRE at high altitudes.

The application of the invention allowed:

1. To provide high technical characteristics of the rocket engine in a wide range of modes of operation at various heights.

2. Ensure reliable operation of the nozzle nozzle at high temperatures.

3. To exclude distortion and shock loads when extending nozzle nozzles to the lower position.

4. To ensure the tightness of the joints of the nozzle nozzles with nozzles.

Claims (8)

1. A liquid rocket engine containing two combustion chambers mounted on the frame with a jet nozzle having a regenerative cooling system, one turbopump unit made in the form of a single unit containing a main turbine, oxidizer and fuel pumps, a gas generator installed under the main turbine, characterized in that the nozzle nozzles are arranged concentrically to the jet nozzles on the lower power plate, made along the profile as an extension of the nozzle, the lower power plate is made with the ability to move in The longitudinal axis of the engine along the guide cylinder mounted perpendicular to the upper power plate connected to the combustion chambers by means of a drive. 2. The liquid rocket engine according to claim 1, characterized in that the nozzle nozzles are made of carbon-carbon composite material. 3. The liquid rocket engine according to claim 1, characterized in that the nozzle nozzles are made cooled and contain cooling jackets installed concentrically with the nozzle nozzles, with the formation of gaps, the cavities of which are connected by flexible pipelines to the fuel pump and fuel manifold. 4. The liquid rocket engine according to claim 1 or 2, or 3, characterized in that the drive contains an actuator. 5. The liquid rocket engine according to claim 1 or 2, or 3, characterized in that the drive contains at least two actuators. 6. The liquid rocket engine according to claim 5, characterized in that the actuators are placed evenly around the circumference around the nozzle. 7. The liquid rocket engine according to claim 1 or 2, or 3, characterized in that a pneumatic cylinder is used as an actuator. 8. The liquid rocket engine according to claim 1 or 2, or 3, characterized in that a hydraulic cylinder is used as an actuator.

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