RU2539315C1 - Liquid-propellant rocket engine turbopump unit - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: this turbopump unit comprises turbine, oxidiser pump and fuel pump with impellers. In compliance with this invention, this turbine represents a birotatory design with two impellers. The latter have no nozzle diaphragms and run in opposite directions, each being communicated with oxidiser pump impeller and fuel pump impeller. Turbopump unit can incorporated an extra fuel pump, impellers of extra fuel pump and fuel pump being fitted on one shaft.

EFFECT: decreased centrifugal loads at turbine rotor.

2 cl, 2 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 the main engine of space vehicles, includes 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 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 contains a combustion chamber with a regenerative cooling path, fuel component feed 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 (prototype), 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.

Objectives of the invention: reducing the size and weight of the TNA and reducing the centrifugal loads on the turbine rotor.

The solution of this problem was achieved in a turbopump unit of a liquid propellant rocket engine containing a turbine and oxidizer and fuel pumps with impellers, in that the turbine is made birotative and contains two impellers made without nozzle devices capable of rotation in opposite directions, each of which is connected respectively with the impeller of the oxidizer pump and the fuel pump. A turbopump unit of a liquid propellant rocket engine may comprise an additional fuel pump, while the impellers of the additional fuel pump and the fuel pump are mounted on the same shaft.

The invention is illustrated in figures 1 and 2, where:

- figure 1 shows a diagram of the first embodiment of the TNA,

- figure 2 shows a diagram of a second variant of the TNA with an additional fuel pump.

The turbopump unit of a TNA liquid-propellant rocket engine (FIG. 1) contains a turbine 1 with impellers of a turbine 2 and 3, without nozzle devices, its housing 4 and inlet pipe 5, an oxidizer pump 6 with an impeller 7, a fuel pump 8 with an impeller 9, two shafts 10 and 11 mounted on bearings 12, 13 and 14. Impellers 3 and 7 are mounted on shaft 11, and impellers 2 and 9 (rotor parts) are installed on shaft 10.

A second TNA variant is possible (FIG. 2), which further comprises an additional fuel pump 15 having an impeller 16 of an additional fuel pump 15, with the impeller 16 mounted on the shaft 10 together with the impeller 9 of the fuel pump 8.

As a result, a real opportunity appeared to design TNA and, first of all, a turbine of smaller dimensions and weight.

The application of the invention allowed:

1. Reduce the dimensions and weight of the TNA.

2. Provide modular design of the TNA.

3. Design all components of the TNA: turbine and pump, for optimal parameters, including rotational speeds, and coordinate rotational speeds through the use of a two-shaft scheme.

4. To increase the reliability of the TNA by reducing the action of centrifugal forces on the details of the turbine rotors.

Claims (2)

1. A turbopump unit of a liquid propellant rocket engine containing a turbine and oxidizer and fuel pumps with impellers, characterized in that the turbine is made biotative and contains two impellers made without nozzle devices capable of rotation in opposite directions, each of which is connected respectively to the working oxidizer pump and fuel pump wheels. 2. The turbopump unit of a liquid propellant rocket engine according to claim 1, characterized in that it comprises an additional fuel pump, while the impellers of the additional fuel pump and fuel pump are mounted on the same shaft.

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