RU2447313C1 - Restartable liquid-propellant engine (versions) - Google Patents

Abstract

FIELD: power industry.

SUBSTANCE: in the first version of engine design output of intermediate fuel container is connected to input of fuel receiver and through pressure control valve to input of engine chamber cooling jacket while output of intermediate oxidiser container is connected to input of oxidiser receiver and through pressure control valve to input of injector head of engine chamber; besides inputs of intermediate container exchange-heaters through cut-off valves are connected to output of engine chamber cooling jacket and outputs of heat-exchangers are connected to input of fuel to injector head of engine chamber. In the second version cooling jacket of engine chamber is divided into fuel-cooled section and oxidiser-cooled section, at that outputs of intermediate fuel and oxidiser containers are connected to respective receivers of fuel components and through pressure control valves are connected to inputs of sections of chamber cooling jackets by fuel and oxidiser respectively; besides inputs of heat-exchangers for fuel and oxidiser intermediate containers through cut-off valves are connected to outputs of respective sections of engine chamber cooling while outputs of heat-exchangers for fuel and oxidiser intermediate containers are connected to respective inputs of these components to injector head of engine chamber.

EFFECT: invention provides higher reliability, simplicity of design, reduction of cost and power consumption and improved conditions for storage of cryogenic fuel components.

2 cl, 3 dwg

Description

The invention relates to rocket and space technology (RKT) and can be used in propulsion systems (DU) and power plants of prospective booster blocks (RB) and interorbital tugs (MB) designed to launch spacecraft (SC) from low reference orbits to various high-energy orbits - highly elliptical and high circular orbits, including geosynchronous and geostationary (GSO) orbits, as well as to take-off trajectories (to the Moon, Mars, etc.).

The high cost, limited reliability and increased risk of failures due to the high-stress operation of the main components are organic disadvantages of modern marching liquid propellant rocket engines (LRE). For RBs and MBs, one of the solutions to overcome these shortcomings may be the use of multi-turn interorbital flight schemes with a large number of marching rocket engine rockets (up to 100 times or more), which can dramatically (by two orders of magnitude) reduce its thrust. Compared to traditional high thrust engines, such a low thrust rocket engine will have the following advantages:

- significantly smaller dimensions and mass of the engine;

- significantly reduced tension of the internal motor parameters;

- a simplified combustion chamber of the engine with the supply to the combustion chamber of both components of the fuel only in a gaseous state;

- the possibility of self-boosting fuel tanks RB and MB due to heating from the Sun.

The feasibility of using multiple-launch marching rocket engines and low thrust (hundreds ... thousands of newtons) is most obvious when switching to the use of new oxygen + hydrogen fuel, which has not yet been used in domestic RBs. The creation of a small-sized and simple, reliable and inexpensive, unified marching oxygen-hydrogen multiple-launch rocket engine, which could be effectively used as part of a number of booster blocks for launch vehicles (LV) of various payload classes and for performing various transport tasks, would significantly facilitate and accelerate the transition to the use of highly efficient and environmentally friendly fuels “oxygen + hydrogen”.

Two versions of highly efficient multiple-injection rocket engines are known [1].

In the first version, considered as an analogue, a multiple-injection LRE contains tanks with liquid fuel and an oxidizing agent, liquid fuel and oxidizer supply systems with electric pumps, an oxidizer receiver with an electric heater, an oxidizer heat exchanger-gasifier, a fuel receiver with a compressor and an engine chamber, as well as a source power supply, including an electrochemical storage battery (AB) with a charge-discharge device (ZRU) and an electrochemical generator (ECG) based on a fuel cell (TE) battery, operating x on the main components of the fuel.

In the second version, taken as a prototype, a multiple-rocket engine [1] also contains an electric power source, tanks with liquid fuel and an oxidizer, a liquid fuel and oxidizer supply system with pumps, an oxidizer receiver with an electric heater, an oxidizer heat exchanger-gasifier, a fuel receiver with a compressor and engine chamber. The power supply here also uses an electrochemical battery with switchgear and ECG based on a battery of fuel cells operating on the main components of the fuel, and the inputs of the fuel and oxidizer in the ECG are connected, respectively, to the receivers of the fuel and oxidizer through pressure and flow controllers, and the electrochemical battery through the switchgear connected to the electric output of the ECG. Unlike the first LRE variant, here in the systems for supplying liquid components of the fuel, not electric pumps are installed, but multiplier pumps with pneumatic drives, the inputs of which are connected to the corresponding outputs of the fuel and oxidizer from the heat exchanger-gasifier of the oxidizer, and the outputs of the pneumatic drives are connected to the corresponding inputs of the fuel and oxidizer mixing head of the chamber.

The advantage of the prototype over the analogue is the elimination of electric drives at the pumps for the supply of liquid oxidizer and fuel. Electric drives have a large net weight and are the most powerful consumers of electricity in a propulsion system. The transition from electronic to pneumatic pump drives fed by gaseous fuel and oxidizer from the heat exchanger-gasifier of the oxidizer can significantly reduce the mass of pumping units, as well as significantly reduce the required power and mass of the electrochemical generator, and reduce the cost of fuel components for generating electricity in ECG.

The disadvantage of the prototype, which is also the case with previous analogues, is the use of new rocket engines in their composition, which are relatively complex and expensive in the development of low-cost means of supplying piston-type cryogenic fuel components - pumps and compressor. The practical experience of developing these units for one of the analog engines shows that the most difficult to develop are a liquid hydrogen feed pump and a hydrogen compressor.

The complexity and high cost of a hydrogen pump with the required performance is determined, first of all, by the large number of operation cycles (many tens ... hundreds of thousands cycles per flight) of their moving elements (pistons and valves) at extremely low operating temperatures (~ 20 K). The piston type of cryogenic fuel and oxidizer pumps also determines problems such as the formation of stagnant zones under the pistons, which reduce their efficiency, as well as pressure pulsations at the pump outlet that occur during relocation of the pistons, which require special damping cylinders to be installed after the pumps.

The complexity and high cost of the electric compressor used in the prototype and in the previous analogues is determined by the high compression ratio (~ 50 times or more) required for periodic “charging” of the hydrogen receiver with a supply of gaseous hydrogen consumed at the engine start mode in each of its starts . This determines the large number of compression stages in the compressor (at least 3 stages), as well as the relatively high temperature of the compressed hydrogen at the compressor outlet, which determines the large volume (~ 200 l or more) and the mass of the hydrogen receiver to accommodate the required hydrogen supply.

With its novelty, structural complexity, complexity of working conditions and the required long service life, cryogenic pumps and a high-pressure compressor will be critical elements in terms of the overall reliability of the prototype and analogues.

The purpose of the present invention is to create a highly efficient multiple-injection rocket engine with a simplified, pump-free fuel supply system while providing:

- higher reliability, structural simplicity and low cost of rocket engines;

- reducing power consumption of the liquid propellant rocket engine, reducing the power and mass of the ECG with a decrease in fuel losses for the generation of electricity in it;

- improving storage conditions for cryogenic fuel components in the tanks of the Republic of Belarus and MB.

The goal is achieved in two engine options.

In the first embodiment, a multiple-injection rocket engine containing an engine chamber (CD) with a cooling jacket, an electric power source in the form of an electrochemical AB with an indoor switchgear and ECG based on a TE battery, tanks with liquid fuel and oxidizer, a fuel supply system with fuel and oxidizer receivers, for supply each component of the fuel (fuel and oxidizer) from the tank to the engine chamber uses an intermediate cylinder with shut-off valves at the inlet and outlet, inside the intermediate cylinder there is a measuring tank with an expansion nozzle and a heat exchanger, moreover, the output of the intermediate cylinder of fuel is connected to the entrance to the receiver of the fuel and through the pressure regulator to the inlet of the cooling jacket of the engine chamber, and the output of the intermediate cylinder of the oxidizer is connected to the entrance to the receiver of the oxidizer and through the pressure regulator to the inlet of the KD nozzle, in addition to the inputs of the intermediate heat exchangers cylinders through shut-off valves are connected to the outlet of the CD cooling jacket, and the exits of the heat exchangers are connected to the fuel inlet to the nozzle of the CD.

Distinctive features are that in the supply system of each component an intermediate cylinder is used with shut-off valves at the inlet and outlet, inside the intermediate cylinder there is a measuring tank with an expansion nozzle and a heat exchanger, and the output of the intermediate fuel cylinder is connected to the inlet to the fuel receiver and through a pressure regulator with the entrance to the KD cooling jacket, and the output of the intermediate oxidizer cylinder is connected to the entrance to the oxidizer receiver and through the pressure regulator with the entrance to the KD nozzle head; in addition, the inputs of the heat exchangers of the intermediate cylinders through the shut-off valves are connected to the outlet of the CD cooling jacket, and the outputs of the heat exchangers are connected to the fuel inlet to the nozzle of the CD.

The differences adopted for this option allow you to implement:

1) gasification in the intermediate cylinders of the entire mass of fuel components with heat taken from the CD cooling jacket (with the refusal to use a special oxidizer-gasifier heat exchanger and electric heater in the oxidizer receiver);

2) the simplest and most reliable way of supplying fuel components to the chamber at the steady-state mode of operation of the liquid-propellant rocket engine is a gas balloon by displacing the components from the intermediate cylinders with internal pressure (with the abandonment of the use of piston pumps and damping cylinders that smooth pressure pulsations after the pumps);

3) the easiest way to repeatedly “charge” the receivers for starting the engine — by displacing the gaseous components with internal pressure directly from the intermediate cylinders (with the hydrogen compressor and the injection and gasification of oxygen in the receiver electrically heated).

Thus, the adopted differences can significantly simplify, lower cost and increase engine reliability due to the possibility of abandoning a number of complex, critical in terms of reliability and expensive units, which also have a large total mass, which makes up the bulk of the mass of the prototype engine:

- piston pumps;

- cylinder dampers at the outlet of the pumps;

- high-pressure compressor with electric drive;

- heat exchanger-gasifier oxidizer;

- an electric heater in an oxidizer receiver.

With the elimination of the compressor, there is no need to use a special thermodynamic system in the hydrogen tank. This system should guarantee the selection from the tank and the supply to the compressor of fully gasified hydrogen at any time both with the engine running and in passive sections of the flight of the Republic of Belarus (i.e., in the absence of axial overload when liquid hydrogen bubbles occupy an arbitrary position in the tank) . The thermodynamic system includes an intake device located inside the tank, an expansion nozzle, and a large-sized heat exchanger.

In addition to this, elimination of the electric compressor and electric heater in the oxidizer receiver leads to a decrease in the electric power consumption in the engine to the minimum possible level, when only electric valves, the igniter of the ignition device, the regulators of the operating modes and sensors remain the consumers of electric energy. This makes it possible to minimize the required power and mass of the ECG, as well as to minimize the unproductive (from the point of view of creating thrust LRE) costs of the components of the fuel to generate electricity in the ECG.

In addition, elimination of the pumps allows eliminating the additional parasitic heat inflows to the cryogenic fuel components in the tanks associated with them in the prototype and analogous engines by placing the pumps inside the tanks in order to ensure that the pumps are in a constantly cold state (which is necessary to prevent large total losses of cryogenic fuel components to cool the engine when it is turned on repeatedly).

In the second version of the design, as well as the first version, the multiple-rocket engine contains a CD with a cooling jacket, an electric power source in the form of an electrochemical battery with a switchgear and ECG based on a TE battery, tanks with liquid fuel and oxidizer, a fuel supply system with fuel receivers and oxidizer, the KD cooling jacket is divided into a fuel cooling section and an oxidizer cooling section, and for supplying each component of fuel from the tanks to the KD, an intermediate cylinder with shut-off valves at the inlet and outlet is used, inside There is a measuring tank with an expansion nozzle and a heat exchanger, while the outputs of the intermediate cylinders of fuel and oxidizer are connected to the corresponding receivers of the fuel components and, through pressure regulators, are connected to the inputs to the sections of the cooling jacket of the fuel and oxidizer, respectively, in addition to the inputs of the heat exchangers of the intermediate cylinders of fuel and oxidizer through the shut-off valves are connected to the outputs of the respective sections of the cooling CD, and the outputs of the heat exchangers of the intermediate cylinders burn it and oxidant are connected to respective inputs of these components in the nozzle head KD.

This embodiment of the multiple-injection rocket engine, as well as the first option described above, uses intermediate cylinders with shut-off valves at the inlet and outlet to supply the fuel components from the tanks to the CD, having an internal measuring tank with an expansion nozzle and a heat exchanger, which provides all of these same advantages in comparison with the prototype [1]. However, unlike the first option, the CD cooling jacket here consists of two separate sections (separately cooled by the fuel and oxidizer), and the heat exchanger of the oxidizer intermediate cylinder uses an oxidizer (rather than fuel, as in the first embodiment) with its own separate source heat - section cooling jacket KD oxidizer. This allows you to maximize the spread and make independent the contours of the supply of fuel components, minimizing the possibility of contact of the oxidizer and fuel inside the engine, which will provide higher reliability and increased safety of its operation.

The proposed technical solutions are illustrated by the drawings presented in figure 1 - figure 3. Figure 1 shows a diagram of the first embodiment of multiple-engagement rocket engines (LRE No. 1). Figure 2 shows a diagram of a second embodiment of multiple rocket engines (LRE No. 2). Figure 3 shows a diagram of an intermediate cylinder with shut-off valves and with a measuring container located inside the nozzle and heat exchanger, which is the same for both the oxidizer and fuel, and for the variants of the liquid-propellant rocket No. 1 and No. 2 (for different cylinder volumes )

In both versions (Fig. 1, 2), the multiple-rocket engine contains an ECG 1, an electrochemical battery 2 with a charge-discharge device 3, pressure regulators of the oxidizer 4 and fuel 5 in front of the ECG, a tank with a liquid oxidizer 6, a tank with liquid fuel 7, oxidizer supply system with an intermediate cylinder 8, pressure regulator 9, valves, pipelines, etc., a fuel supply system with an intermediate cylinder 10, pressure regulator 11, valves, pipelines, etc., a gas fuel receiver 12, an oxidizer gas receiver 13, ka Yeru motor 14 with the nozzle head 15. In the first embodiment (Figure 1) LRE multiple inclusion has a single cooling jacket 16, and in the second embodiment (Figure 2) - shirt separate cooling sections oxidant 17 and fuel 18.

The diagram of the intermediate cylinder (Figure 3) shows the main tank with a liquid oxidizing agent 6 (or with liquid fuel 7), equipped with a sampling device 19, an intermediate cylinder 8 (or 10) with shut-off valves at the inlet 20 and at the outlet 21 and located inside the cylinder measuring tank 22 with an expansion nozzle 23 and a heat exchanger 24 having an inlet 25 and an outlet 26 nozzles.

In both versions of the proposed LRE (FIGS. 1, 2), the intermediate cylinders of the oxidizing agent 8 and fuel 10 function in the same way. Consider their work on the example of an oxidizer cylinder.

In the initial position, the LRE starts to work on the gaseous components of the fuel from the receivers 12 and 13 (see Fig.1, 2); wherein the intermediate cylinder 8 (FIG. 3) is evacuated (i.e. pressure p _int = 0), the shutoff valves 20 and 21 are closed. The liquid oxidizer is located in the main tank 6 in the precipitated position (at the intake device 19 on the bottom of the tank) due to the action of axial overload from the thrust of the rocket engine.

When the shut-off valve 20 is opened under the action of a small pressure inside the tank 6 (pressure p _tank ≤0.2 MPa), the liquid oxidizer is displaced into the measuring tank 22. After filling it, the valve 20 closes and the heat carrier heated in the cooling jacket of the working engine is supplied to the input of the heat exchanger 25 (fuel from the jacket 16 in the variant LRE No. 1 in FIG. 1 or the oxidizing agent from the section 17 of the cooling jacket in the variant LRE No. 2 in FIG. 2). Under the action of heat input, the oxidizing agent evaporates and, passing through the expansion nozzle 23 (in FIG. 3), is completely gasified, filling the entire volume of the intermediate cylinder 8; while the pressure in the cylinder 8 increases rapidly. When the operating pressure level in the intermediate cylinder 8 is reached, the outlet valve 21 opens and the oxidizing agent enters from the cylinder 8 either into the nozzle head of the engine 15 in the LRE No. 1 variant (Fig. 1) or in the cooling jacket section 17 in the LRE No. 2 variant (Fig. 2), at the same time, the oxidizer enters the engine from the receiver 13. At the same time, the engine enters the steady state oxidant supply mode. Similar and simultaneous processes in the intermediate fuel cylinder 10 are completed by reaching the steady state power supply of the engine fuel from the cylinder 10 and the cessation of the flow of fuel into the engine from the receiver 12.

At a steady engine operating mode, the receivers 13 and 12 are “charged” with gaseous oxidizer and fuel, which is carried out by squeezing them into the receivers with excessive internal pressure from the intermediate cylinders 8 and 10, respectively. The body of the intermediate cylinders 8 and 10 is made of high pressure, able to withstand the internal pressure p _point = 40 MPa (taking into account possible pressure spikes during gasification of the oxidizer and fuel). If necessary, the intensity of gasification of the oxidizer and fuel in cylinders 8 and 10 at a steady engine operating mode can be reduced by temporarily (partially or completely) stopping the flow of coolant into the heat exchangers of the intermediate cylinders.

With the use of portions of the oxidizer and fuel, the pressure in the intermediate cylinders 8 and 10 drops below the operating level and the engine either turns off or switches to power from the fuel receivers 12 and the oxidizer 13 with a further repetition of the refueling and operation of the intermediate cylinders described above. For this, the remains of the gaseous oxidizer and fuel from the intermediate cylinders 8 and 10 are effectively used (for example, partially to feed the receivers of the auxiliary RB flight control remote control, and the rest is discharged through the axial nozzles creating thrust) and the cavity of the cylinders are evacuated to bring the engine to its original state.

In both versions of the proposed liquid propellant rocket engine, as in the prototype [1], ECG 1 (Figs. 1, 2) based on fuel cells operating on the main components of the fuel is used. The supply of gaseous components of the fuel in the ECG 1 is carried out by their displacement from the receivers of fuel 12 and the oxidizing agent 13 under the influence of internal pressure, which at any time is always higher than the operating pressure of the components in the fuel cell. The supply of oxidizer and fuel to the ECG 1 is carried out through pressure regulators 4 and 5, respectively. In flight, the ECG 1 operates continuously, energizing the electrochemical storage battery 2. Since in both versions of the proposed LRE the composition of the internal consumers of electricity (with the elimination of the compressor with electric drive and electric heater in the oxidizer receiver) is made extremely minimal, the main consumers of electricity generated in the ECG are onboard service systems of the upper stage. Thus, the proposed liquid propellant rocket engine performs not only its own functions (as a propulsion unit), but also the functions of the power plant of the Republic of Belarus in a long multi-turn flight. This gives an additional effect, allowing us to simplify the composition of the RB, reduce the mass of its structure and, accordingly, increase the mass of the output payload.

The proposed rocket engine allows you to implement the same effective multi-turn schemes of interorbital flights (with the main engine in apsidal points of intermediate orbits) and the same total duration of the flight as the prototype [1]. For example, here it is realistic to achieve the spacecraft launch duration from the low reference orbit to the GSO, not exceeding 4 ... 7 days, which is at least no worse than the performance indicators for the delivery of geostationary spacecraft by means traditional for world practice - the upper stages of launch vehicles and apogee DU KA.

In this case, both versions of the proposed LRE have similar advantages with the prototype [1]. Here, at least, the same pressure in the combustion chamber and specific impulse of thrust can be realized. The presence in the LRE (Figs. 1, 2) of receivers 12 and 13 provides a simple and reliable multiple launch in zero gravity conditions on the gaseous components of the fuel, and their complete gasification at steady state operation — high completeness of fuel combustion, low loss of specific impulse of thrust and simplified regulation of operating modes. For both variants of the proposed liquid-propellant rocket engine, the problem of its unification is also simply and effectively solved both for RBs of various dimensions and for solving various transport problems by them. This is achieved by varying the number of rocket engine starts - the higher the load capacity of the carrier used (i.e., the higher the mass of the orbital block) and the more energy-intensive the transport task to be performed, the greater the number of orbital turns and the number of engine starts.

Evaluation of the comparative (in comparison with the prototype [1]) energy-ballistic efficiency of the use of the proposed multiple-injection rocket engine in the RB was carried out in relation to oxygen + hydrogen fuel using the Soyuz-2-1b launch vehicle as an example of the SC delivery task from Baikonur. It was believed that both versions of the proposed rocket engine (version with a single cooling jacket and the option with separate sections of the cooling jacket for fuel and oxidizer) have the same weight and size parameters. A liquid propellant rocket engine having characteristics similar to the prototype [1] is proposed: pressure in the combustion chamber p _k = 1.5 MPa, thrust R = 750 N, specific thrust impulse I _y = 4500 m / s with a ratio of oxygen and hydrogen consumption K _m = 5 , 0.

It was found that for the prototype [1] the required level of power generated in ECG 1 is N _el ≈1 kW. The mass of the remote control, including multiple rocket engines (without ECG 1, battery 2 and charge-discharge device 3), steering gears, flight control engine blocks, was obtained equal to 107 kg, including a mass of 83 kg of fuel supply system. As a result of ballistic calculations, it was found that the use of a RB with a rocket propellant rocket engine in combination with the Soyuz-2-16 rocket makes it possible to output 1,500 kg to the GSO spacecraft in no more than 5.7 days. The total number of LRE inclusions in the spacecraft launch process from the low reference orbit to the GSO was n _on = 40. At the same time, the generation of electricity to power the electric drive of the compressor and electric heater gasifying liquid oxygen in the receiver required the consumption of fuel components in the ECG equal to 22.5 kg. Emissions of fuel components from the tanks through the drainage safety valves (WPC) are assumed to be zero (taking into account the possibility of using a compressor to suck out excess hydrogen vapor from the hydrogen tank and the possibility of using these vapors to cool the oxygen tank).

In the variant of the proposed LRE with intermediate cylinders (Fig.1, 2) there are similar indicators for the flight time on the GSO, the total number of LRE inclusions and the mass of the orbital block displayed on the GSO. Based on the consumption of 0.25 kW onboard RB systems in standby mode (on passive flight sections), the ECG 1 power here amounted to 0.35 kW, which is ~ 3 times less than in the version with the LRE prototype. The corresponding reduction in the mass of the power supply system of the Republic of Belarus was 13 kg. The total mass of fuel system elements withdrawn, compared with the prototype [1], (pneumatic pumps with damping cylinders, an electric compressor, an oxygen gas heat exchanger and an electric heater from an oxygen receiver) was 47 kg, and the total mass of added elements (intermediate cylinders 8 and 10 with a measured capacity, an expansion nozzle and a heat exchanger, as well as additional pressure regulators 9 and 11) amounted to 46 kg. In this case, the masses of the power shells of the intermediate cylinders of the oxidizer 8 (volume 50 l) and fuel 10 (volume 100 l) are respectively 11 kg and 22 kg (according to research conducted at the enterprise) under the assumption that the power shells are made of carbon fiber based of the maximum internal working pressure equal to p _point = 40 MPa.

For the proposed rocket engine, in comparison with the prototype, were also evaluated and taken into account:

- an increase of 36 kg in the total mass of the receivers 12 and 13 of fuel, due to a four-fold increase in the mass of their one-time “charging” of fuel components and operating time (from ~ 15 to ~ 60 seconds) to ensure not only the initial heating of the chamber, but also gasification of the guaranteed number of components fuel in intermediate cylinders 8 and 10 for their next inclusion in the work;

- additional hydrogen losses in the amount of 23.5 kg due to emissions through the tank’s duct during the flight of the Republic of Belarus (due to the removal of the compressor from the engine);

- a decrease in the mass of the hydrogen tank by 32 kg due to the removal of the thermodynamic system (which included a sampling device, an expansion nozzle, and a large-sized heat exchanger for the prototype [1]).

In General, it was found that for the proposed LRE, in comparison with the prototype [1], there is a small total decrease in the mass of the design of the RB, comprising:

Δm _Σ = -13-47 + 46 + 36-32 = -10 kg,

where Δm _Σ is the total decrease in the mass of the RB structure.

Given this and almost equal unproductive fuel losses, the prototype [1] (due to the increased consumption in ECG 1) and the proposed LRE (due to emissions through the WPC), we can assume that the proposed LRE, at least, is not inferior to the prototype [ 1] by mass of the spacecraft displayed on the GSO. Moreover, as shown above, the proposed liquid propellant rocket engine has such important advantages as higher reliability, structural simplicity, low cost, reduced power consumption and improved storage conditions for cryogenic fuel components due to the absence of engine assemblies built into the tanks.

References

1. RF patent for the invention No. 2364742 "Liquid multiple-launch rocket engine (options)". Priority 04/17/2008, publ. 08/20/2009

Claims (2)

1. A liquid multiple-launch rocket engine comprising an engine chamber with a cooling jacket, an electric power source in the form of an electrochemical storage battery with a charge-discharge device and an electrochemical generator based on a fuel cell battery, liquid fuel and oxidizer tanks, fuel supply systems with fuel receivers and an oxidizer, characterized in that for the supply of each component of the fuel from the tank to the engine chamber, an intermediate cylinder with shut-off valves at the inlet and outlet is used, inside the intermediate cylinder there is a measuring tank with an expansion nozzle and a heat exchanger, and the output of the intermediate fuel cylinder is connected to the inlet of the fuel chamber and through the pressure regulator to the cooling jacket of the engine chamber, and the output of the intermediate oxidizer cylinder is connected to the inlet of the oxidizer and through the pressure regulator with the entrance to the nozzle head of the engine chamber, in addition, the inputs of the heat exchangers of the intermediate cylinders through the shut-off valves are connected to the output of the cooling jacket engine, and the outputs of the heat exchangers are connected to the fuel inlet to the nozzle head of the engine chamber. 2. A multiple-launch liquid propellant rocket engine containing an engine chamber with a cooling jacket, an electric power source in the form of an electrochemical battery with a charge-discharge device and an electrochemical generator based on a fuel cell battery, liquid fuel and oxidizer tanks, fuel supply systems with fuel receivers and oxidizer, characterized in that the cooling jacket of the engine chamber is divided into a fuel cooling section and an oxidizing cooling section, and for supplying each component nta fuel from the tanks into the engine chamber uses an intermediate cylinder with shut-off valves at the inlet and outlet, inside the intermediate cylinder there is a measuring tank with an expansion nozzle and a heat exchanger, while the outputs of the intermediate cylinders of fuel and oxidizer are connected to the corresponding receivers of the fuel components and are connected to pressure regulators the entrances to the sections of the jacket of the cooling chamber of the engine fuel and oxidizer, respectively, in addition, the inputs of the heat exchangers of the intermediate cylinders of fuel and oxide of Tell through check valves are connected to the outputs of the respective sections of the cooling chamber of the engine, and outputs the intermediate heat exchangers of fuel and oxidizer tanks are connected to respective inputs of these components in the nozzle chamber of the engine head.

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