RU2560656C1 - Turbo-pump unit of liquid missile engine - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: in turbo-pump unit of the liquid missile engine comprising main turbine and oxidant, fuel pumps, and start-up turbine with at least one HP source containing explosive charge, according to the invention on HP source the end wall with holes is made, number of holes corresponds to number of the explosive charges, at that at least two explosive charges are installed, the start-up turbine is made with at least two nozzles assemblies, covered by the shutter with possibility of sequential holes opening. The shutter can be connected with drive via the mechanical gear. Two or more HP sources can be installed. On shaft of the in turbo-pump unit via the multiplying gear the additional fuel pump can be connected.

EFFECT: invention ensures re-usable starting of the turbo-pump unit and engine.

6 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 a part of space booster blocks, stages of rocket launchers and as a main engine of space vehicles, including a combustion chamber with a regenerative cooling path, a turbopump assembly - 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 the thrust of the rocket by several times, and to the disruption of the flight program of the rocket 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 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 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 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 and method of starting it according to the patent of the Russian Federation for invention No. 2232915, publ. September 10, 2003, 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 and fuel pump, 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 turbopump engine rocket engine from the Internet site http://www.lpre.de/sntk/NK-33/index.htm, prototype.

This TNA contains a main turbine and oxidizer and fuel pumps and a starting turbine with at least one high pressure source containing pyrolysis.

The disadvantage of this TNA is the inability to repeatedly start the TNA and the engine in flight.

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

The task of creating the invention: providing multiple start TNA and engine.

The solution to this problem was achieved in a turbopump unit of a liquid propellant rocket engine containing a main turbine and oxidizer and fuel pumps and a starting turbine with at least one high pressure source containing a pyro charge, in that an end wall with openings is made on the high pressure source, the number of which corresponds to the number of pyro charges, at least two pyro charges were installed, the starting turbine is made with at least two nozzle devices closed by a shutter, which has the ability to alternate th opening holes and their alignment with one of the nozzle units.

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. An additional fuel pump can be connected to the TNA shaft via a multiplier.

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

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

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

- in FIG. 3 shows nozzle devices and flaps,

- in FIG. 4 shows a view of A,

- in FIG. 5 shows the end wall with two holes,

- in FIG. 6 shows the shutter,

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

- in FIG. 8 shows view B,

- in FIG. 9 shows TNA with an additional fuel pump.

The turbopump unit of a TNA liquid-propellant rocket engine (Fig. 1-9) contains a main turbine 1 having a casing 2, an input casing 3 and an output casing 4. Inside the casing 2, a nozzle apparatus 5 is installed, an impeller 6 with impellers 7. An impeller 6 is installed on the shaft 8. The main turbine 1 has a support 9.

In addition, the TNA contains an oxidizer pump 10 with an impeller 11 mounted on the shaft 8. The oxidizer pump 10 has a support 12.

In addition, the TNA includes a fuel pump 13 with an impeller 14 and supports 15 and 16.

From the end of the TNA opposite the main turbine 1, a starting turbine 17 is installed, which, in turn, contains a housing 18, an input housing 19, an output housing 20. In the input housing 19, the first and second nozzle apparatuses 21 and 22 are installed, in front of which a shutter is installed 23, closing the holes 24 and 25, made in the end face 26. To the valve 23 through a mechanical transmission 27 is connected to the actuator 28.

An impeller 29 with impellers 30 is mounted on the shaft 8.

A high pressure source 31 is connected to the input housing 19, comprising a housing 32, inside of which at least two pyrotechnic charges are placed, in our example the first 33 and second 34 separated by a partition 35. At least two pyroinitiators 36 and 37 are installed on the housing 32. An exhaust pipe 38 is connected to the output housing 20 to discharge the spent combustion products of the pyro charges 33 and 34.

In the end face 26, one hole 43 is made (Figs. 5 and 6) for alternating with the holes 24 and 25.

An embodiment of the TNA with two high-pressure sources 39 and 40 (Fig. 7 and 8), providing two starts, is possible.

It is possible to perform a TNA with an additional fuel pump 41 connected to the shaft 8 using a multiplier 42 (Fig. 9).

TNA works as follows.

For the first start-up, the first shutter 23 is shifted by the actuator 28 and the first hole is opened 24. Then, the first pyro initiator 36 is energized and the first pyrotechnic charge 33 is ignited. The combustion products exit through the first nozzle apparatus 21 to the impeller 30 of the impeller 29. The impeller 29 spins 8 and the impellers 11 and 14 of the oxidizer pumps 10 and fuel 13. Then, the generator gas is supplied to the main turbine 1 through the inlet housing 3, which passes through the impellers 7 of the impeller 6. The main tour enters bin 1, and the starting turbine 17 idles.

To restart the TNA, turn the shutter 23, align the hole 43 with the second hole 25 and give a command to the second pyroinitiator 37. The second pyrocharge 34 is ignited and the TNA is 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 (6)

1. A turbopump unit of a liquid propellant rocket engine containing a main turbine and oxidizer and fuel pumps and a starting turbine with at least one high pressure source containing a pyro charge, characterized in that an end wall with openings is made on the high pressure source, the number of which corresponds to the number of pyro charges at the same time, at least two pyro charges are installed, the starting turbine is made with at least two nozzle devices closed by a shutter having the ability to open from versts and their combination with one of the nozzle apparatus. 2. The turbopump unit of a liquid propellant rocket engine according to claim 1, characterized in that the damper is connected to the actuator. 3. The turbopump unit of a liquid propellant rocket engine according to claim 2, characterized in that the damper is connected to the actuator via a mechanical transmission. 4. The turbopump unit of a liquid propellant rocket engine according to claim 1 or 2, characterized in that two or more high pressure sources are installed. 5. The turbopump unit of a liquid propellant rocket engine according to claim 3, characterized in that two or more high pressure sources are installed. 6. The turbopump unit of a liquid propellant rocket engine according to claim 1 or 2, characterized in that an additional fuel pump is connected to the TNA shaft through a multiplier.

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