RU2574192C1 - Liquid-propellant rocket engine fuel feed plant - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: claimed rocket engine comprises the gas generator with the head and turbopump unit arranged there under. Said turbopump comprises the main turbine and oxidiser and fuel pumps, and starting turbine with at least one high-pressure source including the pyro charge. Oxidiser pump outlet is communicated via pipeline equipped with oxidiser valves with gas generator head. High-pressure source has the end wall with holes, their number corresponding to that of pyro charges. Note here that at least two pyro charges are used. Said starting turbine incorporates at least two nozzle diaphragms closed by the flap. Said flap can open in turns said holes and communicate them with one of nozzle diaphragms. The pipe at the joint with the gas generator is arranged radially. Aforesaid flap is coupled with the drive. Aforesaid flap is coupled with the drive via a mechanical gearing. Two or more high-pressure sources can be used.

EFFECT: multiple in-flight engine and turbopump starting.

5 cl, 9 dwg

Description

The invention relates to rocket technology, specifically to liquid-propellant rocket engines LRE, operating on a cryogenic oxidizer and on hydrocarbon 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 spacecraft, includes a combustion chamber with a regenerative cooling path of the turbopump unit - TNA. TNA contains pumps for the supply of components - fuel and oxidizer with a turbine on the same 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 line 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 engine TNA is the deterioration of the cavitation properties of the pump during condensate bypass. Such a property of the pump will inevitably lead to a decrease in the consumption of one of the fuel components through the ТНА, to a drop in rocket thrust by several times, and to disrupt the missile’s flight program or to disaster.

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 comprises a combustion chamber with a regenerative cooling path, fuel component supply pumps, and a turbine. Pumps and turbines are arranged in two TNAs: main and booster. The engine contains installed in series in front of the feed pump of one of the fuel components of the main turbopump assembly, a booster turbopump pump and a mixer. 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 and the consequences indicated above. 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 and method of starting it according to the patent of the Russian Federation for invention No. 2232915, publ. 07/20/2004, which contains a combustion 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 (additional 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 disadvantage of this scheme is a fire or explosion of TNA and missiles at launch or in flight due to the low reliability of the seal between the turbine and the oxidizer pump, between the oxidizer pump and the fuel, and also between the fuel pump and the additional fuel pump due to the large pressure drop acting on them: 300 ... 400 kgf / cm ² for modern rocket engines. For example, when using hydrogen and oxygen as components of rocket fuel, the smallest leaks of these components lead to the formation of an “explosive mixture” and almost always result in a rocket explosion.

Known fuel supply unit of the Internet site http://www.lpre.de/sntk/NK-33/index.htm, prototype.

This unit contains a gas generator and a turbopump unit located beneath it, which, in turn, contains a main turbine and oxidizer and fuel pumps, an additional fuel pump and a start turbine with at least one high pressure source containing pyro charge, while the outlet from the oxidizer pump is connected a pipeline containing an oxidizer valve with a gas generator head.

The disadvantage of this unit is the impossibility of multiple starting of the TNA and the engine in flight due to the fact that the starting turbine has one pyrotechnic charge, and the gas generator ignition system is made chemical in the form of a TEA block of self-igniting fuel - triethyl aluminum in an ampoule and adapted for one start. The engine can only be restarted at the stand after replacing the pyrotechnic charge and the TEA block.

Objectives of the invention: providing multiple launch of the TNA and the engine in flight.

The solution to this problem was achieved in the fuel supply unit of a liquid propellant rocket engine containing a gas generator and a turbopump unit located below it, which, in turn, contains a main turbine and oxidizer and fuel pumps, an additional fuel pump and a start turbine with at least one high pressure source, containing pyrocharge, while the outlet of the oxidizer pump is connected by a pipeline containing the valve of the oxidizer with the head of the gas generator, the fact that according to the invention on a source of high The end wall was made with holes, the number of which corresponds to the number of pyro charges, at least two pyro charges were installed, the starting turbine was made with at least two nozzle devices, closed by a shutter, which had the possibility of opening the holes one by one and combining them with one of the nozzle devices, and the pipeline at the junction with the gas generator is made radially.

The damper can be connected to the actuator. The damper can be connected to the actuator via a mechanical transmission. Two or more high pressure sources may be installed.

The invention is illustrated in FIG. 1 ... 9, where:

- in FIG. 1 shows a diagram of the unit,

- in FIG. 2 shows a diagram of the starting turbine,

- in FIG. 3 shows nozzle devices and flaps,

- in FIG. 4 shows view A,

- in FIG. 5 shows the shutter,

- in FIG. 6 shows the end wall with two holes.

- in FIG. 7 shows TNA with two pressure accumulators,

- in FIG. 8 shows view B,

- in FIG. 9 shows a second embodiment of the oxidizer pipeline.

The fuel supply unit of the TNA liquid propellant rocket engine (Fig. 1 ... 9) comprises a gas generator 1 having a head 2 and a fuel supply manifold 3, and a turbopump assembly 4.

The turbopump assembly 4 contains, in turn, a main turbine 5 having a housing 6, an input housing 7 and an output housing 8. A nozzle apparatus 9, an impeller 10 with impellers 11 are installed inside the housing 6. The impeller 10 is mounted on the shaft 12. The main the turbine 1 has a support 13 (Fig. 1 and 2).

In addition, the TNA contains an oxidizer pump 14 with an impeller 15 mounted on the shaft 12. The oxidizer pump 14 has a support 16.

In addition, the TNA includes a fuel pump 17 with an impeller 18 and supports 19 and 20.

From the end of the TNA opposite the main turbine 5, a starting turbine 21 is installed with a high pressure source 22 and an exhaust pipe 23.

TNA 4 contains in the lower part an additional fuel pump 24, which is connected to the shaft 12 using a multiplier 25.

The outlet of the oxidizer pump 14 by the oxidizer pipe 26 having an oxidizer valve 27 is connected to the head 2 of the gas generator 1. (Fig. 1). Moreover, to reduce the axial dimension of the TNA 3 and the engine, the output of the pipe 26 is connected radially to the head 2 (perpendicular to the longitudinal axis of the gas generator 1).

The start-up turbine 21 (Fig. 2) contains an input case 28, an output case 29. In the input case 28, the first and second nozzle devices 30 and 31 are installed, in front of which there is a shutter 32 that covers the openings 33 and 34 made in the end face 35. K the shutter 32 through a mechanical transmission 36 is connected to the actuator 37.

An impeller 38 with impellers 39 is mounted on the shaft 12.

A high pressure source 40 is connected to the input building 28, comprising a building 41, inside of which at least two pyrotechnic charges are placed, in our example the first 42 and second 43 separated by a partition 44. At least two pyroinitiators 45 and 46 are installed on the housing 41.

In the end face 35, one hole 47 is made (Figs. 5 and 6) for alternating with the holes 33 and 34.

A TNA version with two high pressure sources 48 and 49 (Fig. 7 and 8), providing two starts, is possible.

The gas generator 1 is equipped with at least one refillable igniter 50. (Laser spark plug, plasma, ion or electric spark.)

An embodiment of a gas generator 1 with an oxidizer collector 51 on its side surface and an oxidizer pipe supply 26 perpendicular to the longitudinal axis of the gas generator 1 is possible (Fig. 9).

TNA works as follows.

For the first start-up (Fig. 1), the shutter 32 is shifted by the actuator 37 and the first opening 33 is opened. Then, voltage is supplied to the first pyro initiator 45 and the first pyrotechnic charge 42 is ignited. The combustion products exit through the first nozzle apparatus 30 to the impeller blades 39 of the impeller 38. The working the wheel 38 spins the shaft 12 and the impellers 15 and 18 of the oxidizer pumps 14 and fuel 17. Then, the generator gas is supplied to the main turbine 5 through the inlet housing 7, which passes through the impeller 11 of the impeller 10. The main the rib 5, and the starting turbine 21 is idle.

To restart, the TNA rotates the shutter 32 and aligns the hole 47 with the second hole 34 and gives a command to the second pyroinitiator 46. The second pyrotechnic charge 43 is ignited and the TNA 4 and the engine are restarted.

The application of the invention allowed:

1. To ensure multiple launch of the TNA and rocket engine in flight.

2. Reduce the dimensions and weight of the TNA.

3. To ensure the modularity of the design of the TNA.

4. Design all TNA units: two turbines and two pumps for optimal parameters.

Claims (5)

1. A fuel supply unit of a liquid propellant rocket engine, comprising a gas generator and a turbopump assembly located beneath it, comprising, in turn, a main turbine and oxidizer and fuel pumps, an additional fuel pump and a starting turbine with at least one high pressure source containing pyrolysis, wherein the outlet of the oxidizer pump is connected by an oxidizer pipe containing an oxidizer valve with the head of the gas generator, characterized in that the end wall is made on the high-pressure source holes, the number of which corresponds to the number of pyro charges, at least two pyro charges are installed, the starting turbine is made with at least two nozzle devices, closed by a shutter that can open the holes one by one and combine them with one of the nozzle devices, and the pipeline at the junction with gas generator installed radially. 2. The unit according to claim 1, characterized in that the damper is connected to the actuator. 3. The assembly according to claim 2, characterized in that the damper is connected to the actuator via a mechanical transmission. 4. Unit under item 1 or 2, characterized in that two or more sources of high pressure are installed. 5. The unit according to claim 3, characterized in that two or more sources of high pressure are installed.

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