RU2429369C2 - Single- or multi-shaft turbo pump of liquid-propellant rocket engine - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: proposed turbo pump comprises fuel and oxidise pumps and multistage turbine (separate turbines) running on gas produced in gas generator in combusting fuel with excess amount of one of fuel components. Gas from turbine first stage is forced to second turbine second stage. Generated gas heater is arranged between turbine stages. Note here that said heater is made up of extra gas generator with generated gas making its one component and fuel making its another component that increases generator gas in combusting said components. Besides, said heater may be made up of heat exchanger with gas heat carrier generated by gas generator and heated body represents of off gas of turbine first stage.

EFFECT: higher power performances.

5 cl, 4 dwg

Description

Technical field

The invention relates to the field of mechanical engineering and, in particular, to a turbopump unit of a liquid propellant rocket engine.

State of the art

In liquid rocket engines, turbopump units are widely used, made both according to the single-shaft scheme, in which the turbines work on a common shaft, driving both oxidizer and fuel pumps, and according to the multi-shaft scheme, in which each turbine brings its own pump.

A single-shaft turbo pumping unit of a liquid propellant rocket engine is known, consisting of centrifugal single-stage oxidizer pumps and a fuel and axial two-stage turbine, with a separation between the steps of the total differential pressure (see the American liquid rocket engine F-1, developed by Rocketdyne, "Cosmonautics" encyclopedia, chap. Editor V.P. Glushko, M .: Soviet Encyclopedia, p. 420, 1995). This technical decision is taken as an analogue of the invention.

The disadvantage of the analogue is expressed in a drop in the gas temperature between the turbine stages (between the turbines of the turbopump unit of the two-shaft scheme), which leads to a drop in the turbine power, which must be compensated by an increase in other parameters, such as gas flow and pressure drop, as well as an increase in turbine size. In addition, this is accompanied by a deterioration in the performance of the liquid propellant rocket engine itself.

Also known is a turbopump unit of a liquid propellant rocket engine, including two single shaft turbopump units having a pump of the corresponding fuel component and two-stage turbines. In this scheme of a turbopump assembly, the gas produced by the gas generator is used sequentially to drive a turbopump fuel supply unit and a turbopump oxidizer supply unit (see the F-1 American liquid rocket engine developed by Rocketdyne Encyclopedia, editor-in-chief V.P. Glushko, Moscow: Soviet Encyclopedia, p. 420, 1995). This technical solution is taken as a prototype of the invention.

The disadvantages of the prototype are similar to the disadvantages of the analogue. However, the prototype has an improved layout that improves the reliability of the turbopump due to the exclusion of the installation of seals in the cavities between the pumps.

Disclosure of invention

The objective of the present invention is to increase the energy characteristics of multi-stage turbines of turbopump units of a liquid rocket engine by restoring (increasing) the gas temperature at the inlet of the next stage.

This problem is solved due to the fact that in a single-shaft or multi-shaft turbopump unit of a liquid propellant rocket engine, including fuel and oxidizer pumps, and a multi-stage turbine (separate turbines), operating (working) on gas received in the gas generator when burning fuel with an excess of one of the components while gas from the exit of the first stage of the turbine (first turbine) enters the input of the second stage of the turbine (second turbine), characterized in that between the two stages of the turbine (first and second turbines) is installed roystvo for heating the gas generator.

Other differences of the invention are:

- an additional gas generator is used as a device for heating gas, one of the fuel components of which is generator gas, and the other is a fuel component, which ensures that the temperature of the generator gas increases during their combustion;

- the device for heating the gas is made in the form of a heat exchanger, the heating body of which (heat carrier) is the gas produced by the gas generator, and the heated body is the exhaust gas of the first stage of the turbine (first turbine);

- the number of stages of a multistage turbine is equal to two;

- the number of turbines connected in series and installed on autonomous shafts is equal to two.

The technical result consists in lowering the temperature of the generator gas at the inlet of the first stage of the turbine (first turbine) while maintaining the turbine power of the turbopump unit or, if necessary, in increasing the turbine power while maintaining the existing temperature of the generator gas.

Brief Description of the Drawings

Figure 1 presents a schematic diagram of a single-shaft turbopump unit with gas heating using an additional gas generator.

Figure 2 presents a schematic diagram of a twin-shaft turbopump unit with gas heating using an additional gas generator

Figure 3 presents a schematic diagram of a single-shaft turbopump unit with gas heating in the heat exchanger.

Figure 4 presents the T-S (temperature-entropy) diagram illustrating the thermodynamics of the processes occurring in the turbine, with the original version (process ED) and with the present invention (process ABCD).

An example embodiment of the invention

Turbopump unit 1 (Figure 1) is made according to a single-shaft scheme. It includes a coaxially mounted fuel pump 2 with a booster stage 3, an oxidizer pump 4, a gas turbine 5 with the first and second stages 6 and 7. With its supply manifold 8, the turbine 5 is connected to the gas generator 9, and the output manifold 10 is connected to the nozzle head of the engine chamber ( not shown).

The gas generator 9 is supplied with liquid fuel and an oxidizer from high-pressure lines 11 and 12, which are connected to the booster stage 3 of the fuel pump 2 and to the outlet of the oxidizer pump 4, respectively. The output manifold of the first stage 6 of the turbine 5 is connected by a gas duct 13 to the second stage 7 of the turbine 5. A device 14 for heating the generator gas is installed in this gas duct. As this device, an additional gas generator 15 is used, one of the fuel components of which is generator gas (gas with an excess of oxidizing agent), and the other is a fuel component - fuel, which ensures that the temperature of the generator gas increases during their combustion. The fuel supply to the additional gas generator 15 is carried out from the line 16.

In another embodiment (Figure 2), the turbopump unit 1 is made according to a two-shaft scheme, with the turbine 17 driving the oxidizer pump 4, and the turbine 18 driving the fuel pump 2. An additional gas generator 15 is also installed in the gas duct 13 connecting the turbines 17 and 18. from the manifold of the turbine 18 is connected to the entrance to the nozzle head (not shown).

In a further embodiment (FIG. 3), a single-shaft turbopump assembly for heating the generator gas uses a heat exchanger 19, which is installed in the gas duct 13 connecting the stages 6 and 7 of the turbine 5. The heating body of the heat exchanger 19 is the generator gas generated by the gas generator 15, and the heated body is the spent generator gas of the first stage of the turbine 5. In this case, the temperature of the generator gas at the outlet of the gas generator 15 should be higher than the gas temperature at the entrance to the first stage 6 of the turbine 5 by the overall operation of the heat exchanger 19.

Device operation

The liquefied oxidizer (oxygen) by gravity enters the pump 4, from which it is supplied through the high-pressure line 12 to the gas generator 15. Liquid fuel (kerosene) also flows by gravity from the pumping stage of the pump 3 through the high-pressure line 11 to the gas generator 15. As a result of burning of the fuel components in the gas generator 15, generating gas is generated, which enters the first stage 6 of the turbine 5 (FIG. 1 and FIG. 3) or turbine 17 (FIG. 2), then enters the gas duct 13, where it is heated in the additional gas generator 15 (FIG. 1 and FIG. .2) or in heat exchange nickel 19 (FIG. 3), after which it enters the second stage 7 of turbine 5 (FIG. 1 and FIG. 3) or turbine 18 (FIG. 2) and drives the fuel pumps 2 and oxidizer 4 through a common shaft (FIG. 1 and FIG. 3) or two shafts (FIG. 2).

Fig. 4 is a T-S diagram showing an increase in working gas temperature between two stages of a turbine. As shown in Fig. 4, the two-stage ABCD process with intermediate recovery of the working gas temperature due to isobaric heating (section BC) allows one to obtain work (the sum of the differences in the enthalpies of the beginning and end of the processes in sections AB and CD) equal to or close to the operation of the initial single-stage process ED ( the difference in the enthalpies of the beginning and end of the ED process) at a significantly lower temperature at the turbine inlet and, accordingly, at the outlet of the gas generator.

Thus, the introduction of intermediate gas heating between the turbine stages (between two sequentially operating turbines, for example, at the point of gas expansion process, halving the total pressure drop in a two-stage turbine (s)) allows reducing the gas temperature at the turbine inlet by 8-10% with respect to to the inlet temperature without intermediate heating.

Such a decrease in temperature makes it possible in some cases to ensure the requirements of strength standards and to guarantee the implementation of predetermined operating modes and operational reserves of liquid-propellant rocket engines.

Industrial application

The proposed turbopump unit is ready for use in rocket technology and, in particular, in liquid rocket engines.

Claims (5)

1. A single-shaft or multi-shaft turbopump unit of a liquid propellant rocket engine, including fuel and oxidizer pumps and a multi-stage turbine (separate turbines), operating (operating) on the gas received in the gas generator when burning fuel with an excess of one of the components, and gas from the outlet of the first stage turbine (first turbine) enters the input of the second stage of the turbine (second turbine), characterized in that between the two stages of the turbine (first and second turbines) a device for heating the generator gas. 2. A single-shaft or multi-shaft turbopump unit of a liquid propellant rocket engine according to claim 1, characterized in that an additional gas generator is used as a device for heating the gas, one of the fuel components of which is generator gas, and the other is a fuel component, which ensures an increase in temperature when burning generator gas. 3. A single-shaft or multi-shaft turbopump unit of a liquid propellant rocket engine according to claim 1, characterized in that the device for heating the gas is made in the form of a heat exchanger, the heating body of which (the heat carrier) is the gas produced by the gas generator, and the heated body is the exhaust gas of the first stage of the turbine ( first turbine). 4. A single-shaft or multi-shaft turbopump unit of a liquid propellant rocket engine according to claim 1, characterized in that the number of stages of a multi-stage turbine is two. 5. A single-shaft or multi-shaft turbopump unit of a liquid propellant rocket engine according to claim 1, characterized in that the number of turbines connected in series and installed on autonomous shafts is two.

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