RU2813564C1 - Method of operation of liquid-propellant engine with afterburner - Google Patents

Abstract

FIELD: rocketry.

SUBSTANCE: method of operation of a liquid-propellant rocket engine with afterburner after the combustion chamber reaches a stable combustion mode is carried out by opening valves in the afterburner supply lines to supply fuel components under pressure to its fuel injectors, ignition of the supplied components in the afterburner is provided by the temperature of the gases flowing from the critical part of the nozzle, as a result of which there is an increase in temperature and speed of gases in the nozzle behind the afterburner, wherein afterburner operates in constant mass flow rate of fuel components, which leads to constantly increasing value of specific pulse until combustion chamber is switched off, or operates in an economical mode of maintaining an increased constant value of the specific pulse due to throttling in the fuel components supply lines for continuous reduction of second flow through injectors in proportion to degree of reduction of atmospheric pressure until the combustion chamber is switched off.

EFFECT: invention provides increased specific impulse of the engine.

1 cl, 1 dwg

Description

Field of technology

The invention relates to the field of rocket technology and can find application in the creation of promising economical liquid rocket engines.

State of the art

A known liquid rocket engine (LPRE) manufactured in the USSR RD-180 (RD-180-Wikipedia) is a closed cycle with afterburning of oxidizing generator gas after the turbine, running on the components liquid oxygen and kerosene, containing 2 combustion chambers with thrust vector control due to the swing of each chamber in two planes, a gas generator and a single-shaft turbopump unit consisting of a turbine, a two-stage fuel pump and a single-stage oxidizer pump. The specific impulse of the engine at sea level is 311.9 s, in vacuum - 338.4 s. The pressure in the combustion chambers is 26.7 MPa, the degree of gas expansion is 36.87. The dry weight of the engine is 5330 kg, height 3580 mm, diameter 3200 mm. Its method of operation is based on the constant consumption of the mass of fuel components over time. Even today, this engine amazes specialists with its high performance. At the same time, all rocket engines installed on the first stages of launch vehicles with fixed and limited dimensions of the nozzles of their combustion chambers lose specific impulse during flight when rising to altitude, and in this case the RD-180 engine loses 26.5 s due to the impossibility of eliminating the possibility of overexpansion of the jet of combustion products when rising to a height with fixed dimensions of the expanding part of the nozzle.

Known is the Raptor liquid rocket engine developed by the US private company SpaceX (ru.wikipedia.org/wiki/Raptor) of a closed type with complete gasification of components, designed for installation on both stages of a two-stage Starship Heavy launch vehicle for manned flights to the Moon and Mars and powered by components liquid oxygen and liquid methane. The engine thrust in the second version at sea level is 230 tf, in the void 250 tf. Specific impulse at sea level is 330 s, in vacuum 375 s. The pressure in the combustion chamber is 35 MPa, the degree of gas expansion is 40 at sea level and 200 in vacuum. The dry weight of the engine is 1500 kg, height 3100 mm, diameter 1300 mm. This is the most advanced methane rocket engine to date. According to its characteristics, it approaches the maximum characteristics of this type of engine. Engine developers are only hoping for some further improvement in terms of increasing the pressure in the combustion chamber to increase engine thrust. During the flight, when climbing to altitude, the engine also loses 45 s in specific impulse. Therefore, a version of the vacuum engine is used for use in the second stage of the launch vehicle.

There is a well-known project of a liquid-propellant rocket engine with a retractable nozzle (Patent No. 2612691, RU), which is taken as an analogue containing a chamber with a two-part nozzle. One part is mounted motionlessly with the combustion chamber, equipped with an extension mechanism in the form of a drive, an actuator and units for direction and fixation in the final position. The second part is made, in turn, of two kinematically connected individual parts with the ability to move along the engine axis, one of which is connected by means of an extension, direction and fixation mechanism with a stationary mounted part of the nozzle. The project is distinguished by the fact that along a cylindrical contour on the periphery of the stationary nozzle shell, profile multi-start screw guides are made, along screw trajectories that are identical around the circumference and equidistant from each other and the longitudinal axis of the engine. In this case, on the body of the retractable part of the nozzle with the maximum diameter, with the possibility of rotation and axial fixation, there is an annular shell equipped with two groups of trunnions with spherical bearings directed towards the longitudinal axis of the nozzle and in the other direction from it: one - interacting with its bearings with the internal profiles of the screw guides, and a second group of trunnions, equipped with spherical bearings through the connecting rods, with a group of trunnions located on the outer part of the nozzle of maximum diameter. The advantage of the project with a retractable nozzle is the ability to eliminate losses in the specific impulse of the combustion chamber when rising to a height and thereby significantly increase the payload mass of the launch vehicle. Its disadvantage is the complexity of the extension mechanism, which must be carried out while the engine is running, which reduces the reliability of the liquid-propellant rocket engine as a whole.

The project of a liquid propellant rocket engine with afterburner (LPRE) is also known (Patent No. 2789943, RU), taken as a prototype, made according to a closed circuit with complete gasification of the components, containing two turbopump units, each of which contains a high-pressure pump sitting on the same shaft with the gas generator. In this case, the fuel passes through a reduction gas generator, and the oxidizer passes through an oxidation gas generator. The most important element is the combustion chamber for burning gaseous components using fuel nozzles in the head of the combustion chamber, which passes into the nozzle and then after the critical section into the expanding part. The project contains a cooling system for the combustion chamber and nozzles with a fuel component, a system for purging the engine before starting, an engine starting system, and an engine operation control system. The project is distinguished by the fact that the engine is equipped with an afterburner device created in the expanding part of the nozzle behind the critical section. The afterburner device consists of an annular belt that carries out a local stepwise increase in the diameter of the nozzle, fuel injectors installed along the perimeter of the annular belt, and two pipelines through which some of the gasified components of the fuel and oxidizer are supplied to the fuel injectors. Ignition of the supplied components in the afterburner device is ensured by the high temperature of the gases escaping from the critical part of the nozzle, resulting in an increase in the temperature and speed of gas movement in the nozzle behind the afterburner device. To reduce the gas pressure at the nozzle exit, the length of the nozzle has been increased. The advantage of the project is that its implementation would make it possible to increase the payload mass of the launch vehicle by increasing the thrust of the combustion chamber, which covers the loss of specific impulse when rising to altitude. The disadvantage of the project is that it does not contain a description of the method of operation of the liquid propellant rocket engine, which can ensure an increase in the specific impulse of the engine when rising to a height to improve engine efficiency.

Thus, the known technical solution for compensating for the loss of specific impulse when rising to a height for a liquid propellant rocket engine, the combustion chamber of which is equipped with a retractable nozzle nozzle, does not provide the necessary reliability of this liquid propellant engine as a whole, and the design of the liquid propellant rocket engine does not contain a description of the method of its operation, which can realize an increase in the specific impulse engine impulse when climbing to a height to improve engine efficiency.

Disclosure of the invention

A liquid rocket engine with afterburner, operating on a fuel pair of kerosene plus liquid oxygen and having a closed cycle with afterburning of oxidizing generator gas after the turbine, contains a gas generator, a turbopump unit, which consists of a turbine, a two-stage fuel pump and a single-stage oxidizer pump. Each of these pumps is a high pressure pump and is mounted on the same shaft as the turbine. The LRDF contains a system for purging the engine before starting, an engine starting system, an engine operation control system, a combustion chamber with fuel injectors installed in its head for burning fuel components. The most important element of the engine is the nozzle, which passes after the critical section into the expanding part. The combustion chamber is cooled with liquid kerosene. In the expanding part of the nozzle behind the critical section, an afterburner device is installed, consisting of an annular belt that carries out a local stepwise increase in the diameter of the nozzle, fuel injectors installed along the perimeter of the annular belt, and two pipelines for supplying some of the fuel and oxidizer components to the fuel injectors. To reduce the gas pressure at the nozzle exit to atmospheric pressure, the nozzle dimensions in length and diameter are increased. The method of operation of the liquid fuel combustion engine includes the launch of a turbopump unit to supply fuel components under pressure to the nozzles installed in the head of the combustion chamber, atomization, mixing, ignition of the fuel components, and the combustion chamber entering the mode of stable combustion of the fuel components. Next, the valves in the power lines of the afterburner device open to supply fuel components under pressure to its fuel injectors. Ignition of the supplied components in the afterburner device is ensured by the temperature of the gases escaping from the critical part of the nozzle. As a result, the temperature and speed of gas movement in the nozzle behind the afterburner increases. In this case, the afterburner operates in the mode of consuming a constant mass of fuel components, which leads to a constantly increasing specific impulse until the combustion chamber is turned off. In this case, when rising to a height, the gas flow overexpands. To increase the efficiency of the liquid propellant engine and prevent overexpansion of the gas flow when rising to a height, the afterburner operates in the mode of maintaining an increased constant value of the specific impulse due to throttling in the supply lines of fuel components to continuously reduce the second flow rate through the nozzles in proportion to the degree of reduction in atmospheric pressure until the combustion chamber is turned off.

The objective of this invention is to develop a method of operation of a liquid propellant propulsion motor that provides an increase in the specific impulse of the engine in order to increase its efficiency.

A method of operation of a liquid-propellant rocket engine with afterburner, running on a fuel pair of kerosene plus liquid oxygen and having a closed cycle with afterburning of oxidizing generator gas after the turbine, containing a gas generator, a turbopump unit consisting of a turbine, a two-stage fuel pump and a single-stage oxidizer pump, each of which is a high-pressure pump mounted on the same shaft with the turbine, an engine purge system before starting, an engine starting system, an engine control system, a combustion chamber with fuel injectors installed in its head for burning fuel components, a nozzle that passes after the critical section into an expanding part, cooling system of the combustion chamber and nozzle with the liquid component of the fuel, an afterburner device installed in the expanding part of the nozzle behind the critical section, consisting of an annular belt that carries out a local stepwise increase in the diameter of the nozzle, fuel injectors installed along the perimeter of the annular belt and two pipelines through which to the fuel injectors some of the fuel and oxidizer components are supplied, and to reduce the gas pressure at the nozzle exit to atmospheric pressure, the dimensions of the nozzle in length and diameter are increased, sequential launch of the turbopump unit to supply fuel components under pressure to the nozzles installed in the head of the combustion chamber, atomization, mixing, ignition fuel components, the combustion chamber reaches the mode of stable combustion of the fuel components, according to the invention, after the combustion chamber reaches the mode of stable combustion, the valves in the power lines of the afterburner device are opened to supply fuel components under pressure to its fuel injectors, ignition of the supplied components in the afterburner device is ensured by temperature gases flowing out of the critical part of the nozzle, resulting in an increase in the temperature and speed of gas movement in the nozzle behind the afterburner, while the afterburner operates in the mode of consuming a constant mass of fuel components, which leads to a constantly increasing specific impulse until the combustion chamber is turned off, or operates in an economical mode of maintaining an increased constant value of the specific impulse due to throttling in the supply lines of fuel components to continuously reduce the second flow rate through the injectors in proportion to the degree of reduction in atmospheric pressure until the combustion chamber is turned off.

The essence of the invention is illustrated by the drawing.

The drawing (Fig. 1) shows a diagram of the main element of the LRDF - the combustion chamber and nozzle with an afterburner device located inside the nozzle.

In this drawing:

1 - nozzle head of the combustion chamber;

2 - pipelines for supplying components

fuel to the injector head;

3 - combustion chamber;

4 - critical section of the nozzle;

5 - pipelines for supplying components

fuel to the fuel injectors of the afterburner;

6 - ring belt;

7 - fuel injectors;

8 - LRDF nozzle;

9 - liquid-propellant rocket engine nozzle.

Carrying out the invention

An example of a possible implementation of the proposed technical solution.

A liquid rocket engine with afterburner contains (Fig. 1) an injector head 1, pipelines 2 supplying gasified fuel components to the combustion chamber 3, a nozzle with a critical section 4, pipelines 5 supplying some of the fuel components to the afterburner device, consisting of an annular belt 6 with fuel injectors 7 and carrying out a local stepwise increase in the diameter of the nozzle 8. For comparison, the dotted line shows the nozzle 9 of the engine without the use of an afterburner device. When using the engine at sea level, the length of the nozzle must provide a pressure at the nozzle exit equal to one atmosphere, and when using the engine in a vacuum, the pressure at the nozzle exit must be zero.

In the engine, with a total second fuel consumption of 625.5 kg/s, including kerosene 168.1 kg/s and oxygen 457.4 kg/s, and with a specific impulse at sea level of 311.9 s, the engine develops a thrust of 195.1 ts at sea level. The degree of gas expansion in the nozzle is 36.87. The diameter at the critical section of the nozzle is 235.5 mm. The length of the nozzle from the critical section to the nozzle exit is assumed to be 1.78 m. The diameter of the nozzle in the cut plane is 1.43 m. With an increase in the productivity of the gas generator and turbopump unit by 20%, an additional 125.1 kg of fuel is supplied to the injectors of the afterburner device. A gas-dynamic calculation of the parameters of the flow of combustion products along the sections of the combustion chamber and nozzle was carried out. The calculation results are shown in the table.

The following sections are used in the table: “ks” - combustion chamber (before entering the nozzle), “cr” - critical section of the nozzle, “ss” - nozzle section, “uv” - afterburner device. The cross section of the afterburner device is located 300 mm from the critical section of the nozzle. The adiabatic coefficient equal to 1.15 was used in the calculations.

According to the calculation results, the degree of expansion of gases in the nozzle from the afterburner device to the nozzle exit is 32.76. The cross-sectional diameter of the nozzle at its exit is 3.32 m, which indicates that the liquid propellant rocket engine is a vacuum version of the original liquid propellant rocket engine. The table shows that the speed of gas flow from the nozzle, and, consequently, the specific impulse increases by 10%. At the same time, with a second mass consumption of 750.6 kg/s, the thrust of the LRDF engine is 257.5 tf. With an increase in the second mass flow through the nozzles in the head of the combustion chamber of the liquid-propellant rocket engine, its thrust will be equal to 234.1 tf. An increase in the thrust of the LPRE engine by 23.4 tf not only compensates for the increase in the mass of the dry LPRE compared to the dry LPRE, as well as the increase in the mass of the elongated shell of the launch vehicle assembly compartment in which the extended LPRE is installed, but also makes it possible to increase the mass of the payload of the launch vehicle .

According to the second method of operation of the liquid propellant rocket engine, to increase its efficiency, the afterburner device uses the effect of increasing specific impulse when rising to a height. To do this, due to throttling in the supply lines of fuel components to the injectors of the afterburner device, the second consumption of fuel components is continuously reduced in proportion to the degree of decrease in atmospheric pressure. According to this method, at each moment of engine operation, a decrease in the second consumption of fuel components, leading to a decrease in specific impulse, is compensated by an increase in specific impulse due to a decrease in atmospheric pressure when ascending to altitude. As a result, the specific impulse remains constant, and engine efficiency increases due to a continuous reduction in the second consumption of fuel components through the injectors of the afterburner device.

As a result of the application of the present invention, a technical solution aimed at developing a method for operating a liquid-propellant rocket engine with afterburner is implemented due to a constant time flow of fuel components through the nozzles of the afterburner device, or due to a continuously decreasing flow rate of fuel components through the nozzles of the afterburner device, proportional to the decreasing atmospheric pressure when climbing to a height, resulting in increased engine efficiency.

Claims (1)

A method of operation of a liquid-propellant rocket engine with afterburner, running on a fuel pair of kerosene plus liquid oxygen and having a closed cycle with afterburning of oxidizing generator gas after the turbine, containing a gas generator, a turbopump unit consisting of a turbine, a two-stage fuel pump and a single-stage oxidizer pump, each of which is a high-pressure pump mounted on the same shaft with the turbine, an engine purge system before starting, an engine starting system, an engine control system, a combustion chamber with fuel injectors installed in its head for burning fuel components, a nozzle that passes after the critical section into an expanding part, cooling system of the combustion chamber and nozzle with the liquid component of the fuel, an afterburner device installed in the expanding part of the nozzle behind the critical section, consisting of an annular belt that carries out a local stepwise increase in the diameter of the nozzle, fuel injectors installed along the perimeter of the annular belt and two pipelines through which to the fuel injectors some of the fuel and oxidizer components are supplied, and to reduce the gas pressure at the nozzle exit to atmospheric pressure, the dimensions of the nozzle in length and diameter are increased, consisting of the sequential launch of a turbopump unit to supply fuel components under pressure to the nozzles installed in the head of the combustion chamber, atomization, mixing , arson of the fuel components, the combustion chamber entering the stable combustion mode of the fuel components, characterized in that after the combustion chamber enters the stable combustion mode, the valves in the power lines of the afterburner device are opened to supply fuel components under pressure to its fuel injectors, the supplied components are ignited in the afterburner device is provided by the temperature of the gases flowing from the critical part of the nozzle, as a result of which there is an increase in the temperature and speed of gas movement in the nozzle behind the afterburner device, while the afterburner device operates in the mode of consuming a constant mass of fuel components, which leads to a constantly increasing value of the specific impulse until turning off the combustion chamber, or operates in an economical mode of maintaining an increased constant value of the specific impulse due to throttling in the supply lines of fuel components to continuously reduce the second flow rate through the injectors in proportion to the degree of reduction in atmospheric pressure until the combustion chamber is turned off.

Images