How a liquid-propellant jet engine works and works. Liquid rocket engine

What is the first thing that comes to mind when you hear the phrase “rocket engines”? Of course, the mysterious space, interplanetary flights, the discovery of new galaxies and the alluring glow of distant stars. At all times, the sky has attracted people to itself, while remaining an unsolved mystery, but the creation of the first space rocket and its launch opened up new horizons of research for humanity.

Rocket engines are essentially ordinary jet engines with one important feature: they do not use atmospheric oxygen as a fuel oxidizer to create jet thrust. Everything that is needed for its operation is located either directly in its body or in the oxidizer and fuel supply systems. It is this feature that makes it possible to use rocket engines in outer space.

There are a lot of types of rocket engines and they all differ strikingly from each other not only in their design features, but also in their operating principle. That is why each type must be considered separately.

Among the main operating characteristics of rocket engines, special attention is paid to specific impulse - the ratio of the amount of jet thrust to the mass of the working fluid consumed per unit time. The specific impulse value represents the efficiency and economy of the engine.

Chemical rocket engines (CRE)

This type of engine is currently the only one that is widely used for launching spacecraft into outer space; in addition, it has found application in the military industry. Chemical engines are divided into solid and liquid fuels depending on the physical state of the rocket fuel.

History of creation

The first rocket engines were solid fuel, and they appeared several centuries ago in China. At that time, they had little to do with space, but with their help it was possible to launch military rockets. The fuel used was a powder similar in composition to gunpowder, only the percentage of its components was changed. As a result, during oxidation, the powder did not explode, but gradually burned, releasing heat and creating jet thrust. Such engines were refined, refined and improved with varying success, but their specific impulse still remained small, that is, the design was ineffective and uneconomical. Soon, new types of solid fuel appeared, allowing for greater specific impulse and greater thrust. Scientists from the USSR, USA and Europe worked on its creation in the first half of the twentieth century. Already in the second half of the 40s, a prototype of modern fuel was developed, which is still used today.

The RD-170 rocket engine runs on liquid fuel and an oxidizer.

Liquid rocket engines are the invention of K.E. Tsiolkovsky, who proposed them as a power unit for a space rocket in 1903. In the 20s, work on the creation of liquid rocket engines began to be carried out in the USA, and in the 30s - in the USSR. Already by the beginning of World War II, the first experimental samples were created, and after its end, liquid-propellant rocket engines began to be mass-produced. They were used in the military industry to equip ballistic missiles. In 1957, for the first time in human history, a Soviet artificial satellite was launched. A rocket equipped with Russian Railways was used to launch it.

Design and principle of operation of chemical rocket engines

A solid fuel engine contains fuel and an oxidizer in a solid aggregate state in its housing, and the container with fuel is also a combustion chamber. The fuel is usually shaped like a rod with a central hole. During the oxidation process, the rod begins to burn from the center to the periphery, and the gases resulting from combustion exit through the nozzle, forming draft. This is the simplest design of all rocket engines.

In liquid rocket engines, the fuel and oxidizer are in a liquid aggregate state in two separate tanks. Through the supply channels they enter the combustion chamber, where they mix and the combustion process occurs. Combustion products exit through the nozzle, forming draft. Liquid oxygen is usually used as an oxidizer, and the fuel can be different: kerosene, liquid hydrogen, etc.

Pros and cons of chemical RDs, their scope of application

The advantages of solid fuel rocket engines are:

• simplicity of design;
• comparative safety in terms of ecology;
• low price;
• reliability.

Disadvantages of solid propellant rocket engines:

• operating time limitation: fuel burns very quickly;
• impossibility of restarting the engine, stopping it and regulating traction;
• small specific gravity within 2000-3000 m/s.

Analyzing the pros and cons of solid propellant rocket motors, we can conclude that their use is justified only in cases where a medium-power power unit is needed, fairly cheap and easy to implement. The scope of their use is ballistic, meteorological missiles, MANPADS, as well as side boosters space rockets(American missiles are equipped with them; they were not used in Soviet and Russian missiles).

Advantages of liquid RDs:

• high specific impulse (about 4500 m/s and above);
• the ability to regulate traction, stop and restart the engine;
• lighter weight and compactness, which makes it possible to launch even large multi-ton loads into orbit.

Disadvantages of rocket engines:

• complex design and commissioning;
• In conditions of weightlessness, liquids in tanks can move chaotically. For their deposition it is necessary to use additional energy sources.

The scope of application of liquid propellant engines is mainly in astronautics, since these engines are too expensive for military purposes.

Despite the fact that so far chemical rocket engines are the only ones capable of launching rockets into outer space, their further improvement is practically impossible. Scientists and designers are convinced that the limit of their capabilities has already been reached, and to obtain more powerful units with a high specific impulse, other energy sources are needed.

Nuclear rocket engines (NRE)

This type of rocket engine, unlike chemical ones, produces energy not by burning fuel, but as a result of heating the working fluid by the energy of nuclear reactions. Nuclear rocket engines are isotopic, thermonuclear and nuclear.

History of creation

The design and operating principle of the nuclear propulsion engine were developed back in the 50s. Already in the 70s, experimental samples were ready in the USSR and the USA, which were successfully tested. The Soviet solid-state engine RD-0410 with a thrust of 3.6 tons was tested on a bench base, and the American NERVA reactor was to be installed on the Saturn V rocket before sponsorship lunar program was stopped. At the same time, work was carried out on the creation of gas-phase nuclear propulsion engines. Currently, scientific programs are underway to develop nuclear rocket engines, and experiments are being conducted at space stations.

Thus, there are already working models of nuclear rocket engines, but so far none of them have been used outside laboratories or scientific bases. The potential of such engines is quite high, but the risk associated with their use is also considerable, so for now they exist only in projects.

Device and principle of operation

Nuclear rocket engines are gas-, liquid- and solid-phase, depending on the state of aggregation of the nuclear fuel. The fuel in solid-phase nuclear propulsion engines is fuel rods, the same as in nuclear reactors. They are located in the engine housing and during the decay of fissile material they release thermal energy. The working fluid - hydrogen gas or ammonia - in contact with the fuel element, absorbs energy and heats up, increasing in volume and compressing, after which it exits through the nozzle under high pressure.

The operating principle of a liquid-phase nuclear propulsion engine and its design are similar to solid-phase ones, only the fuel is in a liquid state, which makes it possible to increase the temperature, and therefore the thrust.

Gas-phase nuclear propulsion engines operate on fuel in a gaseous state. They usually use uranium. Gaseous fuel can be held in the housing by an electric field or located in a sealed transparent flask - a nuclear lamp. In the first case, there is contact of the working fluid with the fuel, as well as partial leakage of the latter, therefore, in addition to the bulk of the fuel, the engine must have a reserve for periodic replenishment. In the case of a nuclear lamp, there is no leakage, and the fuel is completely isolated from the flow of the working fluid.

Advantages and disadvantages of nuclear powered engines

Nuclear rocket engines have a huge advantage over chemical ones - this is a high specific impulse. For solid-phase models, its value is 8000-9000 m/s, for liquid-phase models – 14,000 m/s, for gas-phase – 30,000 m/s. At the same time, their use entails contamination of the atmosphere with radioactive emissions. Now work is underway to create a safe, environmentally friendly and efficient nuclear engine, and the main “contender” for this role is a gas-phase nuclear engine with a nuclear lamp, where the radioactive substance is in a sealed flask and does not come out with a jet flame.

Electric rocket engines (ERM)

Another potential competitor to chemical thrusters is an electric thruster that operates using electrical energy. The electric propulsion can be electrothermal, electrostatic, electromagnetic or pulsed.

History of creation

The first electric propulsion engine was designed in the 30s by the Soviet designer V.P. Glushko, although the idea of ​​​​creating such an engine appeared at the beginning of the twentieth century. In the 60s, scientists from the USSR and the USA actively worked on the creation of electric propulsion engines, and already in the 70s the first samples began to be used in spacecraft as control engines.

Design and principle of operation

An electric rocket propulsion system consists of the electric propulsion engine itself, the structure of which depends on its type, working fluid supply systems, control and power supply. An electrothermal RD heats the flow of the working fluid due to the heat generated by the heating element or in an electric arc. The working fluid used is helium, ammonia, hydrazine, nitrogen and other inert gases, less often hydrogen.

Electrostatic RDs are divided into colloidal, ionic and plasma. In them, charged particles of the working fluid are accelerated due to the electric field. In colloidal or ionic RDs, gas ionization is provided by an ionizer, a high-frequency electric field, or a gas-discharge chamber. In plasma RDs, the working fluid - the inert gas xenon - passes through the annular anode and enters a gas-discharge chamber with a cathode compensator. At high voltage, a spark flashes between the anode and cathode, ionizing the gas, resulting in plasma. Positively charged ions exit through the nozzle at high speed, acquired due to acceleration by the electric field, and electrons are removed outward by the compensator cathode.

Electromagnetic thrusters have their own magnetic field - external or internal, which accelerates charged particles of the working fluid.

Pulse thrusters operate by evaporating solid fuel under the influence of electrical discharges.

Advantages and disadvantages of electric propulsion engines, scope of use

Among the advantages of ERD:

• high specific impulse, the upper limit of which is practically unlimited;
• low fuel consumption (working fluid).

Flaws:

• high level of electricity consumption;
• design complexity;
• slight traction.

Today, the use of electric propulsion engines is limited to their installation on space satellites, and solar batteries are used as sources of electricity for them. At the same time, it is these engines that can become the power plants that will make it possible to explore space, so work on creating new models of them is actively underway in many countries. It was these power plants that science fiction writers most often mentioned in their works dedicated to the conquest of space, and they can also be found in science fiction films. For now, electric propulsion is the hope that people will still be able to travel to the stars.

And propulsion systems of various spacecraft are the primary area of ​​application of liquid propellant engines.

The advantages of liquid rocket engines include the following:

• The highest specific impulse in the class of chemical rocket engines (over 4500 m/s for the oxygen-hydrogen pair, for kerosene-oxygen - 3500 m/s).
• Thrust control: by adjusting fuel consumption, you can change the amount of thrust over a wide range and completely stop the engine and then restart it. This is necessary when maneuvering a vehicle in outer space.
• When creating large rockets, for example, launch vehicles that launch multi-ton payloads into low-Earth orbit, the use of liquid propellant engines makes it possible to achieve a weight advantage compared to solid propellant engines (solid propellant engines). Firstly, due to the higher specific impulse, and secondly, due to the fact that the liquid fuel on the rocket is contained in separate tanks, from which it is supplied to the combustion chamber using pumps. Due to this, the pressure in the tanks is significantly (tens of times) lower than in the combustion chamber, and the tanks themselves are thin-walled and relatively light. In a solid propellant rocket engine, the fuel container is also a combustion chamber and must withstand high pressure (tens of atmospheres), and this entails an increase in its weight. The larger the volume of fuel in the rocket, the larger the size of the containers for storing it, and the greater the weight advantage of the liquid propellant rocket engine compared to the solid propellant rocket engine, and vice versa: for small rockets, the presence of a turbopump unit negates this advantage.

Disadvantages of rocket engines:

• A liquid propellant engine and a rocket based on it are much more complex and more expensive than solid propellant engines with equivalent capabilities (despite the fact that 1 kg of liquid fuel is several times cheaper than solid fuel). It is necessary to transport a liquid rocket with greater precautions, and the technology for preparing it for launch is more complex, labor-intensive and time-consuming (especially when using liquefied gases as fuel components), therefore, for military rockets, preference is currently given to solid fuel engines due to their higher reliability, mobility and combat readiness.

• In zero gravity, the components of liquid fuel move uncontrollably in the space of the tanks. For their deposition It is necessary to apply special measures, for example, turn on auxiliary engines running on solid fuel or gas.

• At present, for chemical rocket engines (including liquid propellant engines), the limit of the energy capabilities of the fuel has been reached, and therefore, theoretically, the possibility of a significant increase in their specific impulse is not foreseen, and this limits the capabilities of rocket technology based on the use of chemical engines, already mastered in two directions :
  1. Space flights in near-Earth space (both manned and unmanned).
  2. Space exploration within the Solar System using automatic vehicles (“Voyager”, “Galileo”).

If a short-term manned expedition to Mars or Venus on a liquid-propellant rocket engine still seems possible (although there are doubts about the feasibility of such flights), then for travel to more distant objects of the Solar System, the size of the rocket required for this and the duration of the flight look unrealistic.

Liquid rocket engines are in demand and will be in demand for a very long time, because no other technology is capable of more reliably and economically lifting cargo from the Earth and placing it into low-Earth orbit. They are safe from an environmental point of view, especially those that run on liquid oxygen and kerosene. But liquid rocket engines, of course, are completely unsuitable for flights to stars and other galaxies. The mass of the entire metagalaxy is 10 56 grams. In order to accelerate on a liquid-propellant rocket engine to at least a quarter of the speed of light, you will need an absolutely incredible amount of fuel - 10 3200 grams, so it’s stupid to even think about it. Liquid rocket engines have their own niche - propulsion engines. Using liquid engines, you can accelerate the carrier to the second escape velocity, fly to Mars, and that’s it.

Fuel system

The fuel system of a liquid-propellant rocket engine includes all elements used to supply fuel to the combustion chamber - fuel tanks, pipelines, a turbopump unit (TNA) - a unit consisting of pumps and a turbine mounted on a single shaft, an injector head, and valves that regulate the flow fuel.

Pump feed fuel allows you to create high pressure in the engine chamber, from tens of atmospheres to 250 atm (LPRE 11D520 RN "Zenit"). High pressure provides a greater degree of expansion of the working fluid, which is a prerequisite for achieving a high specific impulse. In addition, at high pressure in the combustion chamber, a better value of the engine's thrust-to-weight ratio is achieved - the ratio of the thrust to the weight of the engine. The higher the value of this indicator, the smaller the size and weight of the engine (with the same amount of thrust), and the higher the degree of its perfection. The advantages of the pump system are especially noticeable in high-thrust liquid rocket engines - for example, in the propulsion systems of launch vehicles.

In Fig. 1, exhaust gases from the TNA turbine enter through the nozzle head into the combustion chamber along with the fuel components (11). Such an engine is called a closed-cycle engine (otherwise known as a closed-cycle engine), in which the entire fuel flow, including that used in the TPU drive, passes through the combustion chamber of the liquid-propellant rocket engine. The pressure at the turbine outlet in such an engine should obviously be higher than in the combustion chamber of the liquid-propellant rocket engine, and at the inlet to the gas generator (6) feeding the turbine, it should be even higher. To meet these requirements, the same fuel components (at high pressure) that the liquid-propellant rocket engine itself operates on are used to drive the turbine (with a different ratio of components, usually with excess fuel, to reduce the thermal load on the turbine).

An alternative to a closed cycle is an open cycle, in which the turbine exhaust is released directly into the environment through an exhaust pipe. The implementation of an open cycle is technically simpler, since the operation of the turbine is not connected with the operation of the liquid propellant engine chamber, and in this case, the TPU can generally have its own independent fuel system, which simplifies the procedure for starting the entire propulsion system. But closed-cycle systems have slightly better specific impulse values, and this forces designers to overcome the technical difficulties of their implementation, especially for large launch vehicle engines, which have particularly high requirements for this indicator.

In the diagram in Fig. 1 one pump pump pumps both components, which is acceptable in cases where the components have comparable densities. For most liquids used as propellant components, the density varies in the range of 1 ± 0.5 g/cm³, which allows the use of one turbo drive for both pumps. The exception is liquid hydrogen, which at a temperature of 20 K has a density of 0.071 g/cm³. Such a light liquid requires a pump with completely different characteristics, including a much higher rotation speed. Therefore, in the case of using hydrogen as a fuel, an independent fuel pump is provided for each component.

With low engine thrust (and, therefore, low fuel consumption), the turbopump unit becomes too “heavy” an element, worsening the weight characteristics of the propulsion system. An alternative to the pump fuel system is a displacement fuel system, in which the supply of fuel to the combustion chamber is ensured by the boost pressure in the fuel tanks, created by compressed gas, most often nitrogen, which is non-flammable, non-toxic, non-oxidizing and relatively cheap to produce. Helium is used to pressurize tanks with liquid hydrogen, since other gases condense at the temperature of liquid hydrogen and turn into liquids.

When considering the operation of an engine with a displacement fuel supply system from the diagram in Fig. 1, the TNA is excluded, and the fuel components are supplied from the tanks directly to the main valves of the rocket engine (9, 10). The pressure in the fuel tanks during positive displacement must be higher than in the combustion chamber, and the tanks must be stronger (and heavier) than in the case of a pump fuel system. In practice, the pressure in the combustion chamber of an engine with displacement fuel supply is limited to 10-15 at. Typically, such engines have a relatively low thrust (within 10 tons). The advantages of the displacement system are the simplicity of the design and the speed of the engine's response to the start command, especially in the case of using self-igniting fuel components. Such engines are used to perform maneuvers of spacecraft in outer space. The displacement system was used in all three propulsion systems of the Apollo lunar spacecraft - service (thrust 9760 kgf), landing (thrust 4760 kgf), and takeoff (thrust 1950 kgf).

Nozzle head- a unit in which nozzles are mounted, designed to inject fuel components into the combustion chamber. (You can often find the incorrect name for this unit “mixing head”. This is an inaccurate translation, a copy of English-language articles. The essence of the error is that the mixing of fuel components occurs in the first third of the combustion chamber, and not in the injector head.) The main requirement for injectors is - mixing the components as quickly and thoroughly as possible when entering the chamber, because the rate of their ignition and combustion depends on this.
Through the nozzle head of the F-1 engine, for example, 1.8 tons of liquid oxygen and 0.9 tons of kerosene enter the combustion chamber every second. And the residence time of each portion of this fuel and its combustion products in the chamber is calculated in milliseconds. During this time, the fuel should burn as completely as possible, since unburned fuel means a loss of thrust and specific impulse. The solution to this problem is achieved by a number of measures:

• Maximum increase in the number of nozzles in the head, with proportional minimization of the flow rate through one nozzle. (The F-1 engine's injector head has 2,600 oxygen injectors and 3,700 kerosene injectors.)
• Special geometry of the nozzles in the head and the order of alternating fuel and oxidizer nozzles.
• The special shape of the nozzle channel, due to which rotation is imparted when the liquid moves through the channel, and when it enters the chamber it is scattered to the sides by centrifugal force.

Cooling system

Due to the rapidity of the processes occurring in the combustion chamber of the liquid-propellant rocket engine, only an insignificant part (fractions of a percent) of the total heat generated in the chamber is transferred to the engine structure, however, due to the high combustion temperature (sometimes over 3000 K), and a significant amount of heat generated, even small part of it is enough for thermal destruction of the engine, so the problem of protecting the material part of the liquid-propellant rocket engine from high temperatures is very relevant. To solve this problem, there are two fundamental methods that are often combined - cooling and thermal protection.

For liquid-propellant rocket engines with pumped fuel supply, one cooling method is mainly used together with one method of thermal protection of the walls of the liquid-propellant rocket engine chamber: flow cooling And wall layer [unknown term] . Often used for small engines with positive displacement fuel systems. ablative cooling method.

Flow cooling consists in the fact that in the wall of the combustion chamber and the upper, most heated part of the nozzle, a cavity is created in one way or another (sometimes called a “cooling jacket”), through which one of the fuel components (usually fuel) passes before entering the nozzle head, cooling thus the wall of the chamber.

If the heat absorbed by the cooling component is returned to the chamber along with the coolant itself, then such a system is called “ regenerative", if the rejected heat does not enter the combustion chamber, but is dumped outside, then this is called “ independent» by flow cooling method.

Various technological methods have been developed to create a cooling jacket. The chamber of the V-2 liquid-propellant rocket, for example, consisted of two steel shells, an inner one (the so-called “fire wall”) and an outer one, repeating the shape of each other. The cooling component (ethanol) passed through the gap between these shells. Due to technological deviations in the thickness of the gap, uneven fluid flow arose, resulting in the creation of local zones of overheating of the inner shell, which often burned out in these zones with catastrophic consequences.

In modern engines, the inner part of the chamber wall is made of highly thermally conductive bronze alloys. Narrow thin-walled channels are created in it by milling (15D520 RN 11K77 Zenit, RN 11K25 Energia) or acid etching (SSME Space Shuttle). From the outside, this structure is tightly wrapped around a load-bearing sheet shell made of steel or titanium, which absorbs the force load of the internal pressure of the chamber. The cooling component circulates through the channels. Sometimes the cooling jacket is assembled from thin heat-conducting tubes, sealed with a bronze alloy for tightness, but such chambers are designed for lower pressure.

Launch of the rocket engine

Launching a liquid propellant rocket engine is a responsible operation, fraught with serious consequences in the event of emergency situations during its execution.

If the fuel components are self-igniting, that is, they enter into a chemical combustion reaction upon physical contact with each other (for example, heptyl/nitric acid), initiation of the combustion process does not cause problems. But in the case where the components are not such (for example oxygen/kerosene), an external ignition initiator is required, the action of which must be precisely coordinated with the supply of fuel components to the combustion chamber. An unburned fuel mixture is an explosive of great destructive power, and its accumulation in the chamber threatens a serious accident.

After ignition of the fuel, maintaining a continuous process of its combustion occurs by itself: the fuel newly entering the combustion chamber is ignited due to the high temperature created during the combustion of previously introduced portions.

For the initial ignition of fuel in the combustion chamber when starting a liquid-propellant rocket engine, different methods are used:

• The use of self-igniting components (usually based on phosphorus-containing starting fuels, self-igniting when interacting with oxygen), which at the very beginning of the engine starting process are introduced into the chamber through special, additional nozzles from the auxiliary fuel system, and after the start of combustion, the main components are supplied. The presence of an additional fuel system complicates the design of the engine, but allows it to be restarted several times.
• An electrical igniter located in the combustion chamber near the injector head, which, when turned on, creates an electric arc or a series of high voltage spark discharges. This igniter is disposable. Once the fuel is ignited, it burns.
• Pyrotechnic igniter. Near the nozzle head, a small pyrotechnic incendiary bomb is placed in the chamber, which is ignited by an electric fuse.

Automatic engine starting coordinates the action of the igniter and the fuel supply in time.

The launch of large liquid-propellant rocket engines with a pump fuel system consists of several stages: first, the pump starts and accelerates (this process can also consist of several phases), then the main valves of the liquid-propellant rocket engine are turned on, usually in two or more stages with a gradual increase in thrust from stage to stage. steps up to normal.

For relatively small engines, it is practiced to start the rocket engine immediately at 100% thrust, called “cannon”.

LRE automatic control system

A modern liquid-propellant rocket engine is equipped with rather complex automation, which must perform the following tasks:

• Safe starting of the engine and bringing it to the main mode.
• Maintaining stable operating conditions.
• Thrust change in accordance with the flight program or at the command of external control systems.
• Turning off the engine when the rocket reaches a given orbit (trajectory).
• Regulating the ratio of component consumption.

Due to the technological variation in the hydraulic resistance of the fuel and oxidizer paths, the ratio of component flow rates in a real engine differs from the calculated one, which entails a decrease in thrust and specific impulse in relation to the calculated values. As a result, the rocket may never complete its task, having completely used up one of the fuel components. At the dawn of rocket science, they struggled with this by creating a guaranteed reserve of fuel (the rocket is filled with more than the calculated amount of fuel, so that it would be enough for any deviations of real flight conditions from the calculated ones). The guaranteed fuel supply is created at the expense of the payload. Currently, large rockets are equipped with an automatic control system for the ratio of component consumption, which makes it possible to maintain this ratio close to the calculated one, thus reducing the guaranteed fuel supply, and accordingly increasing the payload mass.
System automatic control The propulsion system includes pressure and flow sensors at different points in the fuel system, and executive bodies it is the main valves of the liquid propellant engine and the turbine control valves (in Fig. 1 - positions 7, 8, 9 and 10).

Fuel components

The choice of fuel components is one of major decisions when designing a liquid propellant engine, which predetermines many details of the engine design and subsequent technical solutions. Therefore, the choice of fuel for a liquid-propellant rocket engine is made with a comprehensive consideration of the purpose of the engine and the rocket on which it is installed, the conditions of their operation, production technology, storage, transportation to the launch site, etc.

One of the most important indicators characterizing the combination of components is specific impulse, which has especially important when designing launch vehicles for spacecraft, since the ratio of the mass of fuel and payload, and therefore the size and mass of the entire rocket (see Tsiolkovsky Formula), which, if the specific impulse is not high enough, may turn out to be unrealistic, greatly depends on it. Table 1 shows the main characteristics of some combinations of liquid fuel components.

Table 1

Oxidizer	Fuel	Average density fuel, g/cm³	Chamber temperature combustion, K	Void specific impulse, s
Oxygen	Hydrogen	0,3155	3250	428
	Kerosene	1,036	3755	335
	Unsymmetrical dimethylhydrazine	0,9915	3670	344
	Hydrazine	1,0715	3446	346
	Ammonia	0,8393	3070	323
Dianitrogen tetroxide	Kerosene	1,269	3516	309
	Unsymmetrical dimethylhydrazine	1,185	3469	318
	Hydrazine	1,228	3287	322
Fluorine	Hydrogen	0,621	4707	449
	Hydrazine	1,314	4775	402
	Pentaborane	1,199	4807	361

Jet engines that run on compressed cold gas (for example, air or nitrogen) are also single-component. Such engines are called gas jet engines and consist of a valve and a nozzle. Gas jet engines are used where the thermal and chemical effects of the exhaust jet are unacceptable, and where the main requirement is simplicity of design. These requirements must be met, for example, by individual cosmonaut movement and maneuvering devices (UPMK), located in the backpack behind the back and intended for movement when working outside the spacecraft. UPMK operate from two cylinders with compressed nitrogen, which is supplied through solenoid valves into a propulsion system consisting of 16 engines.

Three-component rocket engines

Since the early 1970s, the USSR and the USA have been studying the concept of three-propellant engines that would combine a high specific impulse when using hydrogen as fuel, and a higher average fuel density (and, therefore, smaller volume and weight of fuel tanks), characteristic of hydrocarbon fuel. When starting, such an engine would run on oxygen and kerosene, and at high altitudes it would switch to using liquid oxygen and hydrogen. This approach may make it possible to create a single-stage space launch vehicle. Russian example The three-component engine is the liquid propellant rocket engine RD-701, which was developed for the reusable transport and space system MAKS.

It is also possible to use two fuels simultaneously - for example, hydrogen - beryllium - oxygen and hydrogen - lithium - fluorine (beryllium and lithium burn, and hydrogen is mostly used as a working fluid), which makes it possible to achieve specific impulse values ​​in the region of 550-560 seconds, however technically very difficult and has never been used in practice.

Rocket control

In liquid rockets, engines often, in addition to the main function of generating thrust, also serve as flight controls. Already the first guided ballistic missile V-2 was controlled using 4 graphite gas-dynamic rudders placed in the engine jet stream along the periphery of the nozzle. By deflecting, these rudders deflected part of the jet stream, which changed the direction of the engine thrust vector and created a moment of force relative to the center of mass of the rocket, which was the control action. This method significantly reduces engine thrust; moreover, graphite rudders in a jet stream are subject to severe erosion and have a very short service life.
Modern missile control systems use PTZ cameras Liquid rocket engines, which are attached to the load-bearing elements of the rocket body using hinges that allow the camera to be rotated in one or two planes. The fuel components are supplied to the chamber using flexible pipelines - bellows. When the camera deviates from an axis parallel to the axis of the rocket, the thrust of the camera creates the required control torque. The cameras are rotated by hydraulic or pneumatic steering machines, which execute commands generated by the rocket control system.
In the Russian space launch vehicle Soyuz-2, in addition to 20 main, fixed cameras of the propulsion system, there are 12 smaller rotating (each in its own plane) control cameras. The steering chambers share a common fuel system with the main engines.
Of the 11 propulsion engines (all stages) of the Saturn 5 launch vehicle, nine (except for the central 1st and 2nd stages) are rotary, each in two planes. When using the main engines as control engines, the operating range of camera rotation is no more than ±5°: due to the high thrust of the main camera and its location in the aft compartment, that is, at a considerable distance from the center of mass of the rocket, even a small deflection of the camera creates a significant control moment.

In addition to PTZ cameras, motors are sometimes used that serve only the purposes of controlling and stabilizing the aircraft. Two chambers with oppositely directed nozzles are rigidly fixed to the body of the apparatus in such a way that the thrust of these chambers creates a moment of force around one of the main axes of the apparatus. Accordingly, to control the other two axes, their own pairs of control motors are also installed. These motors (usually single-component) are turned on and off at the command of the vehicle control system, turning it in the required direction. Such control systems are usually used for orientation aircraft in outer space.

• World famous rocket engines

S-IC engines and Von Braun.jpg

Propulsion system North American Rockwell, Rocketdyne F-1. 5 engines are installed on the 1st stage of the Saturn 5 space launch vehicle. These engines powered man's flight to the Moon. Thrust at sea level - 691 tf. First flight - 1967

see also

• ORM (engine), ORM-1, ORM-12, ORM-4, ORM-5, ORM-52, ORM-65, ORM-8, ORM-9
• RD-0120, RD-107, RD-108, RD-170, RD-701

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Links

• A. A. Dorofeev.. MSTU im. N. E. Bauman. M., 1999.
• I. I. Shuneyko.. M., 1973.
• . The plot of the Roscosmos television studio.

Notes

Reciprocating Number of bars Location cylinders Piston types Way ignition Rotary Combined Main types Non-compressor Modifications and hybrid systems Chemical Liquid Nuclear Electrical Other By type of working fluid By design features Asynchronous Synchronous Other

An excerpt characterizing the Liquid Rocket Engine

- To His Majesty with an errand.
- Here he is! - said Boris, who heard that Rostov needed His Highness, instead of His Majesty.
And he pointed him to the Grand Duke, who, a hundred paces away from them, in a helmet and a cavalry guard's tunic, with his raised shoulders and frowning eyebrows, was shouting something to the white and pale Austrian officer.
- Yes, this is Grand Duke“And I should go to the commander-in-chief or to the sovereign,” Rostov said and started to move his horse.
- Count, count! - shouted Berg, as animated as Boris, running up from the other side, - Count, I was wounded in my right hand (he said, showing his hand, bloody, tied with a handkerchief) and remained in the front. Count, holding a sword in my left hand: in our race, the von Bergs, Count, were all knights.
Berg said something else, but Rostov, without listening to him, had already moved on.
Having passed the guards and an empty gap, Rostov, in order not to fall into the first line again, as he came under attack by the cavalry guards, rode along the line of reserves, going far around the place where the hottest shooting and cannonade was heard. Suddenly, in front of him and behind our troops, in a place where he could not possibly suspect the enemy, he heard close rifle fire.
"What could it be? - thought Rostov. - Is the enemy behind our troops? It can’t be, Rostov thought, and a horror of fear for himself and for the outcome of the entire battle suddenly came over him. “Whatever it is, however,” he thought, “there’s nothing to go around now.” I must look for the commander-in-chief here, and if everything is lost, then it’s my job to perish along with everyone else.”
The bad feeling that suddenly came over Rostov was confirmed more and more the further he drove into the space occupied by crowds of heterogeneous troops, located beyond the village of Prats.
- What's happened? What's happened? Who are they shooting at? Who's shooting? - Rostov asked, matching the Russian and Austrian soldiers running in mixed crowds across his road.
- The devil knows them? Beat everyone! Get lost! - the crowds of people running and not understanding, just like him, what was happening here, answered him in Russian, German and Czech.
- Beat the Germans! - one shouted.
- Damn them - traitors.
“Zum Henker diese Ruesen... [To hell with these Russians...],” the German grumbled something.
Several wounded were walking along the road. Curses, screams, moans merged into one common roar. The shooting died down and, as Rostov later learned, Russian and Austrian soldiers were shooting at each other.
"My God! what is this? - thought Rostov. - And here, where the sovereign can see them at any moment... But no, these are probably just a few scoundrels. This will pass, this is not it, this cannot be, he thought. “Just hurry up, pass them quickly!”
The thought of defeat and flight could not enter Rostov’s head. Although he saw French guns and troops precisely on Pratsenskaya Mountain, on the very one where he was ordered to look for the commander-in-chief, he could not and did not want to believe it.

Near the village of Praca, Rostov was ordered to look for Kutuzov and the sovereign. But here not only were they not there, but there was not a single commander, but there were heterogeneous crowds of frustrated troops.
He urged his already tired horse to get through these crowds as quickly as possible, but the further he moved, the more upset the crowds became. The high road on which he drove out was crowded with carriages, carriages of all kinds, Russian and Austrian soldiers, of all branches of the military, wounded and unwounded. All this hummed and swarmed in a mixed manner to the gloomy sound of flying cannonballs from the French batteries placed on the Pratsen Heights.
- Where is the sovereign? where is Kutuzov? - Rostov asked everyone he could stop, and could not get an answer from anyone.
Finally, grabbing the soldier by the collar, he forced him to answer himself.
- Eh! Brother! Everyone has been there for a long time, they have fled ahead! - the soldier said to Rostov, laughing at something and breaking free.
Leaving this soldier, who was obviously drunk, Rostov stopped the horse of the orderly or the guard of an important person and began to question him. The orderly announced to Rostov that an hour ago the sovereign had been driven at full speed in a carriage along this very road, and that the sovereign was dangerously wounded.
“It can’t be,” said Rostov, “that’s right, someone else.”
“I saw it myself,” said the orderly with a self-confident grin. “It’s time for me to know the sovereign: it seems like how many times I’ve seen something like this in St. Petersburg.” A pale, very pale man sits in a carriage. As soon as the four blacks let loose, my fathers, he thundered past us: it’s time, it seems, to know both the royal horses and Ilya Ivanovich; It seems that the coachman does not ride with anyone else like the Tsar.
Rostov let his horse go and wanted to ride on. A wounded officer walking past turned to him.
-Who do you want? – asked the officer. - Commander-in-Chief? So he was killed by a cannonball, killed in the chest by our regiment.
“Not killed, wounded,” another officer corrected.
- Who? Kutuzov? - asked Rostov.
- Not Kutuzov, but whatever you call him - well, it’s all the same, there aren’t many alive left. Go over there, to that village, all the authorities have gathered there,” said this officer, pointing to the village of Gostieradek, and walked past.
Rostov rode at a pace, not knowing why or to whom he would go now. The Emperor is wounded, the battle is lost. It was impossible not to believe it now. Rostov drove in the direction that was shown to him and in which a tower and a church could be seen in the distance. What was his hurry? What could he now say to the sovereign or Kutuzov, even if they were alive and not wounded?
“Go this way, your honor, and here they will kill you,” the soldier shouted to him. - They'll kill you here!
- ABOUT! what are you saying? said another. -Where will he go? It's closer here.
Rostov thought about it and drove exactly in the direction where he was told that he would be killed.
“Now it doesn’t matter: if the sovereign is wounded, should I really take care of myself?” he thought. He entered the area where most of the people fleeing from Pratsen died. The French had not yet occupied this place, and the Russians, those who were alive or wounded, had long abandoned it. On the field, like heaps of good arable land, lay ten people, fifteen killed and wounded on every tithe of space. The wounded crawled down in twos and threes together, and one could hear their unpleasant, sometimes feigned, as it seemed to Rostov, screams and moans. Rostov started to trot his horse so as not to see all these suffering people, and he became scared. He was afraid not for his life, but for the courage that he needed and which, he knew, would not withstand the sight of these unfortunates.
The French, who stopped shooting at this field strewn with the dead and wounded, because there was no one alive on it, saw the adjutant riding along it, aimed a gun at him and threw several cannonballs. The feeling of these whistling, terrible sounds and the surrounding dead people merged for Rostov into one impression of horror and self-pity. He remembered his mother's last letter. “What would she feel,” he thought, “if she saw me now here, on this field and with guns pointed at me.”
In the village of Gostieradeke there were, although confused, but in greater order, Russian troops marching away from the battlefield. The French cannonballs could no longer reach here, and the sounds of firing seemed distant. Here everyone already clearly saw and said that the battle was lost. Whoever Rostov turned to, no one could tell him where the sovereign was, or where Kutuzov was. Some said that the rumor about the sovereign’s wound was true, others said that it was not, and explained this false rumor that had spread by the fact that, indeed, the pale and frightened Chief Marshal Count Tolstoy galloped back from the battlefield in the sovereign’s carriage, who rode out with others in the emperor’s retinue on the battlefield. One officer told Rostov that beyond the village, to the left, he saw someone from the higher authorities, and Rostov went there, no longer hoping to find anyone, but only to clear his conscience before himself. Having traveled about three miles and having passed the last Russian troops, near a vegetable garden dug in by a ditch, Rostov saw two horsemen standing opposite the ditch. One, with a white plume on his hat, seemed familiar to Rostov for some reason; another, unfamiliar rider, on a beautiful red horse (this horse seemed familiar to Rostov) rode up to the ditch, pushed the horse with his spurs and, releasing the reins, easily jumped over the ditch in the garden. Only the earth crumbled from the embankment from the horse’s hind hooves. Turning his horse sharply, he again jumped back over the ditch and respectfully addressed the rider with the white plume, apparently inviting him to do the same. The horseman, whose figure seemed familiar to Rostov and for some reason involuntarily attracted his attention, made a negative gesture with his head and hand, and by this gesture Rostov instantly recognized his lamented, adored sovereign.
“But it couldn’t be him, alone in the middle of this empty field,” thought Rostov. At this time, Alexander turned his head, and Rostov saw his favorite features so vividly etched in his memory. The Emperor was pale, his cheeks were sunken and his eyes sunken; but there was even more charm and meekness in his features. Rostov was happy, convinced that the rumor about the sovereign’s wound was unfair. He was happy that he saw him. He knew that he could, even had to, directly turn to him and convey what he was ordered to convey from Dolgorukov.
But just as a young man in love trembles and faints, not daring to say what he dreams of at night, and looks around in fear, looking for help or the possibility of delay and escape, when the desired moment has come and he stands alone with her, so Rostov now, having achieved that , what he wanted more than anything in the world, did not know how to approach the sovereign, and he was presented with thousands of reasons why it was inconvenient, indecent and impossible.
"How! I seem to be glad to take advantage of the fact that he is alone and despondent. An unknown face may seem unpleasant and difficult to him at this moment of sadness; Then what can I tell him now, when just looking at him my heart skips a beat and my mouth goes dry?” Not one of those countless speeches that he, addressing the sovereign, composed in his imagination, came to his mind now. Those speeches were mostly held under completely different conditions, they were spoken for the most part at the moment of victories and triumphs and mainly on his deathbed from his wounds, while the sovereign thanked him for his heroic deeds, and he, dying, expressed his love confirmed in fact my.
“Then why should I ask the sovereign about his orders to the right flank, when it is already 4 o’clock in the evening and the battle is lost? No, I definitely shouldn’t approach him. Shouldn't disturb his reverie. It’s better to die a thousand times than to receive a bad look from him, a bad opinion,” Rostov decided and with sadness and despair in his heart he drove away, constantly looking back at the sovereign, who was still standing in the same position of indecisiveness.
While Rostov was making these considerations and sadly driving away from the sovereign, Captain von Toll accidentally drove into the same place and, seeing the sovereign, drove straight up to him, offered him his services and helped him cross the ditch on foot. The Emperor, wanting to rest and feeling unwell, sat down under an apple tree, and Tol stopped next to him. From afar, Rostov saw with envy and remorse how von Tol spoke for a long time and passionately to the sovereign, and how the sovereign, apparently crying, closed his eyes with his hand and shook hands with Tol.
“And I could be in his place?” Rostov thought to himself and, barely holding back tears of regret for the fate of the sovereign, in complete despair he drove on, not knowing where and why he was going now.
His despair was all the greater because he felt that his own weakness was the cause of his grief.
He could... not only could, but he had to drive up to the sovereign. And this was the only opportunity to show the sovereign his devotion. And he didn’t use it... “What have I done?” he thought. And he turned his horse and galloped back to the place where he had seen the emperor; but there was no one behind the ditch anymore. Only carts and carriages were driving. From one furman, Rostov learned that the Kutuzov headquarters was located nearby in the village where the convoys were going. Rostov went after them.
The guard Kutuzov walked ahead of him, leading horses in blankets. Behind the bereytor there was a cart, and behind the cart walked an old servant, in a cap, a sheepskin coat and with bowed legs.
- Titus, oh Titus! - said the bereitor.
- What? - the old man answered absentmindedly.
- Titus! Go threshing.
- Eh, fool, ugh! – the old man said, spitting angrily. Some time passed in silent movement, and the same joke was repeated again.
At five o'clock in the evening the battle was lost at all points. More than a hundred guns were already in the hands of the French.
Przhebyshevsky and his corps laid down their weapons. Other columns, having lost about half of the people, retreated in frustrated, mixed crowds.
The remnants of the troops of Lanzheron and Dokhturov, mingled, crowded around the ponds on the dams and banks near the village of Augesta.
At 6 o'clock only at the Augesta dam the hot cannonade of the French alone could still be heard, who had built numerous batteries on the descent of the Pratsen Heights and were hitting our retreating troops.
In the rearguard, Dokhturov and others, gathering battalions, fired back at the French cavalry that was pursuing ours. It was starting to get dark. On the narrow dam of Augest, on which for so many years the old miller sat peacefully in a cap with fishing rods, while his grandson, rolling up his shirt sleeves, was sorting out silver quivering fish in a watering can; on this dam, along which for so many years the Moravians drove peacefully on their twin carts loaded with wheat, in shaggy hats and blue jackets and, dusted with flour, with white carts leaving along the same dam - on this narrow dam now between wagons and cannons, under the horses and between the wheels crowded people disfigured by the fear of death, crushing each other, dying, walking over the dying and killing each other only so that, after walking a few steps, to be sure. also killed.
Every ten seconds, pumping up the air, a cannonball splashed or a grenade exploded in the middle of this dense crowd, killing and sprinkling blood on those who stood close. Dolokhov, wounded in the arm, on foot with a dozen soldiers of his company (he was already an officer) and his regimental commander, on horseback, represented the remnants of the entire regiment. Drawn by the crowd, they pressed into the entrance to the dam and, pressed on all sides, stopped because a horse in front fell under a cannon, and the crowd was pulling it out. One cannonball killed someone behind them, the other hit in front and splashed Dolokhov’s blood. The crowd moved desperately, shrank, moved a few steps and stopped again.
Walk these hundred steps, and you will probably be saved; stand for another two minutes, and everyone probably thought he was dead. Dolokhov, standing in the middle of the crowd, rushed to the edge of the dam, knocking down two soldiers, and fled onto the slippery ice that covered the pond.
“Turn,” he shouted, jumping on the ice that was cracking under him, “turn!” - he shouted at the gun. - Holds!...
The ice held it, but it bent and cracked, and it was obvious that not only under a gun or a crowd of people, but under him alone it would collapse. They looked at him and huddled close to the shore, not daring to step on the ice yet. The regiment commander, standing on horseback at the entrance, raised his hand and opened his mouth, addressing Dolokhov. Suddenly one of the cannonballs whistled so low over the crowd that everyone bent down. Something splashed into the wet water, and the general and his horse fell into a pool of blood. No one looked at the general, no one thought to raise him.
- Let's go on the ice! walked on the ice! Let's go! gate! can't you hear! Let's go! - suddenly, after the cannonball hit the general, countless voices were heard, not knowing what or why they were shouting.
One of the rear guns, which was entering the dam, turned onto the ice. Crowds of soldiers from the dam began to run to the frozen pond. The ice cracked under one of the leading soldiers and one foot went into the water; he wanted to recover and fell waist-deep.
The nearest soldiers hesitated, the gun driver stopped his horse, but shouts could still be heard from behind: “Get on the ice, let’s go!” let's go! And screams of horror were heard from the crowd. The soldiers surrounding the gun waved at the horses and beat them to make them turn and move. The horses set off from the shore. The ice holding the foot soldiers collapsed in a huge piece, and about forty people who were on the ice rushed forward and backward, drowning one another.
The cannonballs still whistled evenly and splashed onto the ice, into the water and, most often, into the crowd covering the dam, ponds and shore.

On Pratsenskaya Mountain, in the very place where he fell with the flagpole in his hands, Prince Andrei Bolkonsky lay, bleeding, and, without knowing it, moaned a quiet, pitiful and childish groan.
By evening he stopped moaning and became completely quiet. He didn't know how long his oblivion lasted. Suddenly he felt alive again and suffering from a burning and tearing pain in his head.
“Where is it, this high sky, which I did not know until now and saw today?” was his first thought. “And I didn’t know this suffering either,” he thought. - Yes, I didn’t know anything until now. But where am I?
He began to listen and heard the sounds of approaching horses and the sounds of voices speaking French. He opened his eyes. Above him was again the same high sky with floating clouds rising even higher, through which a blue infinity could be seen. He did not turn his head and did not see those who, judging by the sound of hooves and voices, drove up to him and stopped.
The horsemen who arrived were Napoleon, accompanied by two adjutants. Bonaparte, driving around the battlefield, gave the last orders to strengthen the batteries firing at the Augesta Dam and examined the dead and wounded remaining on the battlefield.
- De beaux hommes! [Beauties!] - said Napoleon, looking at the killed Russian grenadier, who, with his face buried in the ground and the back of his head blackened, was lying on his stomach, throwing one already numb arm far away.
– Les munitions des pieces de position sont epuisees, sire! [There are no more battery charges, Your Majesty!] - said at that time the adjutant, who arrived from the batteries that were firing at Augest.
“Faites avancer celles de la reserve, [Have it brought from the reserves,” said Napoleon, and, having driven off a few steps, he stopped over Prince Andrei, who was lying on his back with the flagpole thrown next to him (the banner had already been taken by the French, like a trophy) .
“Voila une belle mort, [This is a beautiful death,”] said Napoleon, looking at Bolkonsky.
Prince Andrei realized that this was said about him, and that Napoleon was saying this. He heard the one who said these words called sire. But he heard these words as if he heard the buzzing of a fly. Not only was he not interested in them, but he did not even notice them, and immediately forgot them. His head was burning; he felt that he was emanating blood, and he saw above him the distant, high and eternal sky. He knew that it was Napoleon - his hero, but at that moment Napoleon seemed to him such a small, insignificant person in comparison with what was now happening between his soul and this high, endless sky with clouds running across it. He didn’t care at all at that moment, no matter who stood above him, no matter what they said about him; He was only glad that people were standing over him, and he only wished that these people would help him and return him to life, which seemed so beautiful to him, because he understood it so differently now. He mustered all his strength to move and make some sound. He weakly moved his leg and produced a pitying, weak, painful groan.
- A! “He’s alive,” said Napoleon. – Raise this young man, ce jeune homme, and take him to the dressing station!
Having said this, Napoleon rode further towards Marshal Lan, who, taking off his hat, smiling and congratulating him on his victory, drove up to the emperor.
Prince Andrei did not remember anything further: he lost consciousness from the terrible pain that was caused to him by being placed on a stretcher, jolts while moving, and probing the wound at the dressing station. He woke up only at the end of the day, when he was united with other Russian wounded and captured officers and carried to the hospital. During this movement he felt somewhat fresher and could look around and even speak.
The first words he heard when he woke up were the words of the French escort officer, who hurriedly said:
- We must stop here: the emperor will pass by now; it will give him pleasure to see these captive gentlemen.
“There are so many prisoners these days, almost the entire Russian army, that he probably got bored with it,” said another officer.
- Well, however! This one, they say, is the commander of the entire guard of Emperor Alexander,” said the first, pointing to a wounded Russian officer in a white cavalry uniform.
Bolkonsky recognized Prince Repnin, whom he had met in St. Petersburg society. Next to him stood another, 19-year-old boy, also a wounded cavalry officer.
Bonaparte, galloping up, stopped his horse.
-Who is the eldest? - he said when he saw the prisoners.
They named the colonel, Prince Repnin.
– Are you the commander of the cavalry regiment of Emperor Alexander? - asked Napoleon.
“I commanded a squadron,” answered Repnin.
“Your regiment honestly fulfilled its duty,” said Napoleon.
“The praise of a great commander is the best reward for a soldier,” said Repnin.
“I give it to you with pleasure,” said Napoleon. -Who is this young man next to you?
Prince Repnin named Lieutenant Sukhtelen.
Looking at him, Napoleon said, smiling:
– II est venu bien jeune se frotter a nous. [He came to compete with us when he was young.]
“Youth doesn’t stop you from being brave,” Sukhtelen said in a breaking voice.
“Excellent answer,” said Napoleon. - Young man, you will go far!
Prince Andrei, who, to complete the trophy of the captives, was also put forward, in full view of the emperor, could not help but attract his attention. Napoleon apparently remembered that he had seen him on the field and, addressing him, used the same name of the young man - jeune homme, under which Bolkonsky was reflected in his memory for the first time.
– Et vous, jeune homme? Well, what about you, young man? - he turned to him, - how do you feel, mon brave?
Despite the fact that five minutes before this, Prince Andrei could say a few words to the soldiers carrying him, he now, directly fixing his eyes on Napoleon, was silent... All the interests that occupied Napoleon seemed so insignificant to him at that moment, so petty seemed to him his hero himself, with this petty vanity and joy of victory, in comparison with that high, fair and good heavens, which he saw and understood - that he could not answer him.
And everything seemed so useless and insignificant in comparison with the strict and majestic structure of thought that was caused in him by the weakening of his strength from the bleeding, suffering and the imminent expectation of death. Looking into the eyes of Napoleon, Prince Andrei thought about the insignificance of greatness, about the insignificance of life, the meaning of which no one could understand, and about the even greater insignificance of death, the meaning of which no one living could understand and explain.
The emperor, without waiting for an answer, turned away and, driving away, turned to one of the commanders:
“Let them take care of these gentlemen and take them to my bivouac; let my doctor Larrey examine their wounds. Goodbye, Prince Repnin,” and he, moving his horse, galloped on.
There was a radiance of self-satisfaction and happiness on his face.
The soldiers who brought Prince Andrei and removed from him the golden icon they found, hung on his brother by Princess Marya, seeing the kindness with which the emperor treated the prisoners, hastened to return the icon.
Prince Andrei did not see who put it on again or how, but on his chest, above his uniform, suddenly there was an icon on a small gold chain.
“It would be good,” thought Prince Andrei, looking at this icon, which his sister hung on him with such feeling and reverence, “it would be good if everything were as clear and simple as it seems to Princess Marya. How nice it would be to know where to look for help in this life and what to expect after it, there, beyond the grave! How happy and calm I would be if I could now say: Lord, have mercy on me!... But to whom will I say this? Either the power is indefinite, incomprehensible, which I not only cannot address, but which I cannot express in words - the great all or nothing, - he said to himself, - or this is the God who is sewn up here, in this palm, Princess Marya? Nothing, nothing is true, except the insignificance of everything that is clear to me, and the greatness of something incomprehensible, but most important!
The stretcher started moving. With each push he again felt unbearable pain; the feverish state intensified, and he began to become delirious. Those dreams of his father, wife, sister and future son and the tenderness that he experienced on the night before the battle, the figure of the small, insignificant Napoleon and the high sky above all this, formed the main basis of his feverish ideas.
A quiet life and calm family happiness in Bald Mountains seemed to him. He was already enjoying this happiness when suddenly little Napoleon appeared with his indifferent, limited and happy look at the misfortune of others, and doubts and torment began, and only the sky promised peace. By morning, all the dreams mixed up and merged into the chaos and darkness of unconsciousness and oblivion, which, in the opinion of Larrey himself, Doctor Napoleon, were much more likely to be resolved by death than by recovery.
“C"est un sujet nerveux et bilieux," said Larrey, "il n"en rechappera pas. [This is a nervous and bilious man, he will not recover.]
Prince Andrey, among other hopelessly wounded, was handed over to the care of the residents.

At the beginning of 1806, Nikolai Rostov returned on vacation. Denisov was also going home to Voronezh, and Rostov persuaded him to go with him to Moscow and stay in their house. At the penultimate station, having met a comrade, Denisov drank three bottles of wine with him and, approaching Moscow, despite the potholes of the road, he did not wake up, lying at the bottom of the relay sleigh, near Rostov, which, as it approached Moscow, came more and more to impatience.
“Is it soon? Soon? Oh, these unbearable streets, shops, rolls, lanterns, cab drivers!” thought Rostov, when they had already signed up for their holidays at the outpost and entered Moscow.
- Denisov, we’ve arrived! Sleeping! - he said, leaning forward with his whole body, as if by this position he hoped to speed up the movement of the sleigh. Denisov did not respond.
“Here is the corner of the intersection where Zakhar the cabman stands; Here he is Zakhar, and still the same horse. Here is the shop where they bought gingerbread. Soon? Well!
- To which house? - asked the coachman.
- Yes, over there at the end, how can you not see! This is our home,” said Rostov, “after all, this is our home!” Denisov! Denisov! We'll come now.
Denisov raised his head, cleared his throat and did not answer.
“Dmitry,” Rostov turned to the footman in the irradiation room. - After all, this is our fire?
“That’s exactly how daddy’s office is lit up.”
– Haven’t gone to bed yet? A? How do you think? “Don’t forget to get me a new Hungarian at once,” Rostov added, feeling the new mustache. “Come on, let’s go,” he shouted to the coachman. “Wake up, Vasya,” he turned to Denisov, who lowered his head again. - Come on, let's go, three rubles for vodka, let's go! - Rostov shouted when the sleigh was already three houses away from the entrance. It seemed to him that the horses were not moving. Finally the sleigh took to the right towards the entrance; Above his head, Rostov saw a familiar cornice with chipped plaster, a porch, a sidewalk pillar. He jumped out of the sleigh as he walked and ran into the hallway. The house also stood motionless, unwelcoming, as if it did not care about who came to it. There was no one in the hallway. "My God! is everything alright? thought Rostov, stopping for a minute with a sinking heart and immediately starting to run further along the entryway and familiar, crooked steps. The same door handle of the castle, for the uncleanness of which the countess was angry, also opened weakly. One tallow candle was burning in the hallway.
Old man Mikhail was sleeping on the chest. Prokofy, the traveling footman, the one who was so strong that he could lift the carriage by the back, sat and knitted bast shoes from the edges. He looked at the opened door, and his indifferent, sleepy expression suddenly transformed into an enthusiastically frightened one.
- Fathers, lights! Young Count! – he cried out, recognizing the young master. - What is this? My darling! - And Prokofy, shaking with excitement, rushed to the door to the living room, probably to make an announcement, but apparently changed his mind again, returned back and fell on the young master’s shoulder.
-Are you healthy? - Rostov asked, pulling his hand away from him.
- God bless! All glory to God! We just ate it now! Let me look at you, Your Excellency!
- Is everything all right?
- Thank God, thank God!
Rostov, completely forgetting about Denisov, not wanting to let anyone warn him, took off his fur coat and ran on tiptoe into the dark, large hall. Everything is the same, the same card tables, the same chandelier in a case; but someone had already seen the young master, and before he had time to reach the living room, something quickly, like a storm, flew out of the side door and hugged and began to kiss him. Another, third, same creature jumped out of another, third door; more hugs, more kisses, more screams, tears of joy. He couldn’t make out where and who dad was, who was Natasha, who was Petya. Everyone was screaming, talking and kissing him at the same time. Only his mother was not among them - he remembered that.
- I didn’t know... Nikolushka... my friend!
- Here he is... ours... My friend, Kolya... He has changed! No candles! Tea!
- Yes, kiss me!
- Darling... and then me.
Sonya, Natasha, Petya, Anna Mikhailovna, Vera, the old count, hugged him; and people and maids, filling the rooms, muttered and gasped.
Petya hung on his legs. - And then me! - he shouted. Natasha, after she had bent him to her and kissed his whole face, jumped away from him and holding onto the hem of his Hungarian jacket, jumped like a goat all in one place and squealed shrilly.
On all sides there were eyes shining with tears of joy, loving eyes, on all sides there were lips seeking a kiss.
Sonya, red as red, also held his hand and was all beaming in the blissful gaze fixed on his eyes, which she was waiting for. Sonya was already 16 years old, and she was very beautiful, especially at this moment of happy, enthusiastic animation. She looked at him without taking her eyes off, smiling and holding her breath. He looked at her gratefully; but still waited and looked for someone. The old countess had not come out yet. And then steps were heard at the door. The steps are so fast that they couldn't be his mother's.
But it was she in a new dress, still unfamiliar to him, sewn without him. Everyone left him and he ran to her. When they came together, she fell on his chest, sobbing. She could not raise her face and only pressed it to the cold strings of his Hungarian. Denisov, unnoticed by anyone, entered the room, stood right there and, looking at them, rubbed his eyes.
“Vasily Denisov, a friend of your son,” he said, introducing himself to the count, who was looking at him questioningly.
- Welcome. I know, I know,” said the count, kissing and hugging Denisov. - Nikolushka wrote... Natasha, Vera, here he is Denisov.
The same happy, enthusiastic faces turned to the shaggy figure of Denisov and surrounded him.
- Darling, Denisov! - Natasha squealed, not remembering herself with delight, jumped up to him, hugged and kissed him. Everyone was embarrassed by Natasha's action. Denisov also blushed, but smiled and took Natasha’s hand and kissed it.
Denisov was taken to the room prepared for him, and the Rostovs all gathered in the sofa near Nikolushka.
The old countess, without letting go of his hand, which she kissed every minute, sat next to him; the rest, crowding around them, caught his every movement, word, glance, and did not take their rapturously loving eyes off him. The brother and sisters argued and grabbed each other's places closer to him, and fought over who should bring him tea, a scarf, a pipe.
Rostov was very happy with the love that was shown to him; but the first minute of his meeting was so blissful that his present happiness seemed not enough to him, and he kept waiting for something else, and more, and more.
The next morning, the visitors slept from the road until 10 o'clock.
In the previous room there were scattered sabers, bags, tanks, open suitcases, and dirty boots. The cleaned two pairs with spurs had just been placed against the wall. Servants brought washbasins, hot water for shaving, and cleaned dresses. It smelled of tobacco and men.
- Hey, G"ishka, t"ubku! – Vaska Denisov’s hoarse voice shouted. - Rostov, get up!
Rostov, rubbing his drooping eyes, raised his confused head from the hot pillow.
- Why is it late? “It’s late, it’s 10 o’clock,” Natasha’s voice answered, and in the next room the rustling of starched dresses, the whispering and laughter of girls’ voices was heard, and something blue, ribbons, black hair and cheerful faces flashed through the slightly open door. It was Natasha with Sonya and Petya, who came to see if he was up.
- Nikolenka, get up! – Natasha’s voice was heard again at the door.
- Now!
At this time, Petya, in the first room, saw and grabbed the sabers, and experiencing the delight that boys experience at the sight of a warlike older brother, and forgetting that it was indecent for sisters to see undressed men, opened the door.
- Is this your saber? - he shouted. The girls jumped back. Denisov, with frightened eyes, hid his furry legs in a blanket, looking back at his comrade for help. The door let Petya through and closed again. Laughter was heard from behind the door.
“Nikolenka, come out in your dressing gown,” said Natasha’s voice.
- Is this your saber? - Petya asked, - or is it yours? - He addressed the mustachioed, black Denisov with obsequious respect.
Rostov hastily put on his shoes, put on his robe and went out. Natasha put on one boot with a spur and climbed into the other. Sonya was spinning and was just about to puff up her dress and sit down when he came out. Both were wearing the same brand new blue dresses - fresh, rosy, cheerful. Sonya ran away, and Natasha, taking her brother by the arm, led him to the sofa, and they began to talk. They did not have time to ask each other and answer questions about thousands of little things that could only interest them alone. Natasha laughed at every word that he said and that she said, not because what they said was funny, but because she was having fun and was unable to contain her joy, which was expressed by laughter.
- Oh, how good, great! – she condemned everything. Rostov felt how, under the influence of the hot rays of love, for the first time in a year and a half, that childish smile blossomed on his soul and face, which he had never smiled since he left home.
“No, listen,” she said, “are you completely a man now?” I'm terribly glad that you are my brother. “She touched his mustache. - I want to know what kind of men you are? Are they like us? No?
- Why did Sonya run away? - Rostov asked.
- Yes. That's another whole story! How will you talk to Sonya? You or you?
“As it will happen,” said Rostov.
– Tell her, please, I’ll tell you later.
- So what?
- Well, I’ll tell you now. You know that Sonya is my friend, such a friend that I would burn my hand for her. Look at this. - She rolled up her muslin sleeve and showed a red mark on her long, thin and delicate arm under the shoulder, much above the elbow (in a place that is sometimes covered by ball gowns).
“I burned this to prove my love to her.” I just lit the ruler on fire and pressed it down.
Sitting in his former classroom, on the sofa with cushions on his arms, and looking into those desperately animated eyes of Natasha, Rostov again entered that family, children's world, which had no meaning for anyone except for him, but which gave him some of the best pleasures in life; and burning his hand with a ruler to show love did not seem useless to him: he understood and was not surprised by it.
- So what? only? - he asked.
- Well, so friendly, so friendly! Is this nonsense - with a ruler; but we are forever friends. She will love anyone, forever; but I don’t understand this, I’ll forget now.
- Well, what then?
- Yes, that’s how she loves me and you. - Natasha suddenly blushed, - well, you remember, before leaving... So she says that you forget all this... She said: I will always love him, and let him be free. It’s true that this is excellent, noble! - Yes Yes? very noble? Yes? - Natasha asked so seriously and excitedly that it was clear that what she was saying now, she had previously said with tears.
Rostov thought about it.
“I don’t take back my word on anything,” he said. - And then, Sonya is such a charm that what fool would refuse his happiness?
“No, no,” Natasha screamed. “We’ve already talked about this with her.” We knew you would say this. But this is impossible, because, you know, if you say that - you consider yourself bound by the word, then it turns out that she seemed to say it on purpose. It turns out that you are still forcibly marrying her, and it turns out completely different.
Rostov saw that all this was well thought out by them. Sonya amazed him with her beauty yesterday too. Today, having caught a glimpse of her, she seemed even better to him. She was a lovely 16-year-old girl, obviously loving him passionately (he did not doubt this for a minute). Why shouldn’t he love her now, and not even marry her, Rostov thought, but now there are so many other joys and activities! “Yes, they came up with this perfectly,” he thought, “we must remain free.”
“Well, great,” he said, “we’ll talk later.” Oh, how glad I am for you! - he added.
- Well, why didn’t you cheat on Boris? - asked the brother.
- This is nonsense! – Natasha shouted laughing. “I don’t think about him or anyone else and I don’t want to know.”
- That's how it is! So what are you doing?
- I? – Natasha asked again, and a happy smile lit up her face. -Have you seen Duport?
- No.
– Have you seen the famous Duport the dancer? Well, you won't understand. That's what I am. “Natasha took her skirt, rounding her arms, as they dance, ran a few steps, turned over, did an entreche, kicked her leg against the leg and, standing on the very tips of her socks, walked a few steps.
- Am I standing? after all, she said; but couldn’t help herself on her tiptoes. - So that’s what I am! I will never marry anyone, but will become a dancer. But do not tell anyone.
Rostov laughed so loudly and cheerfully that Denisov from his room became envious, and Natasha could not resist laughing with him. - No, it’s good, isn’t it? – she kept saying.
- Okay, don’t you want to marry Boris anymore?
Natasha flushed. - I don’t want to marry anyone. I'll tell him the same thing when I see him.
- That's how it is! - said Rostov.
“Well, yes, it’s all nothing,” Natasha continued to chatter. - Why is Denisov good? – she asked.
- Good.
- Well, goodbye, get dressed. Is he scary, Denisov?
- Why is it scary? – asked Nicholas. - No. Vaska is nice.
- You call him Vaska - strange. And that he is very good?
- Very good.
- Well, come quickly and drink tea. Together.
And Natasha stood on tiptoe and walked out of the room the way dancers do, but smiling the way only happy 15-year-old girls smile. Having met Sonya in the living room, Rostov blushed. He didn't know how to deal with her. Yesterday they kissed in the first minute of the joy of their date, but today they felt that it was impossible to do this; he felt that everyone, his mother and sisters, looked at him questioningly and expected from him how he would behave with her. He kissed her hand and called her you - Sonya. But their eyes, having met, said “you” to each other and kissed tenderly. With her gaze she asked him for forgiveness for the fact that at Natasha’s embassy she dared to remind him of his promise and thanked him for his love. With his gaze he thanked her for the offer of freedom and said that one way or another, he would never stop loving her, because it was impossible not to love her.

The R-7 ICBM was created, equipped with liquid rocket engines RD-107 and RD-108, at that time the most powerful and advanced in the world, developed under the leadership of V.P. Glushko. This rocket was used as the carrier of the world's first artificial earth satellites, the first manned spacecraft and interplanetary probes.

In 1969, the first Apollo series spacecraft was launched in the United States, launched onto a flight path to the Moon by the Saturn 5 launch vehicle, the first stage of which was equipped with 5 F-1 engines. The F-1 is currently the most powerful among single-chamber liquid propellant engines, inferior in thrust to the four-chamber engine RD-170, developed by the Energomash Design Bureau in the Soviet Union in 1976.

Currently, the space programs of all countries are based on the use of liquid rocket engines.

Scope of use, advantages and disadvantages

Katorgin, Boris Ivanovich, academician of the Russian Academy of Sciences, former head of NPO Energomash

Design and principle of operation of a two-component liquid propellant rocket engine

Rice. 1 Scheme of a two-component rocket engine
1 - oxidizer line
2 - fuel line
3 - oxidizer pump
4 - fuel pump
5 - turbine
6 - gas generator
7 - gas generator valve (oxidizer)
8 - gas generator valve (fuel)
9 - main oxidizer valve
10 - main fuel valve
11 - turbine exhaust
12 - mixing head
13 - combustion chamber
14 - nozzle

There is quite a wide variety of liquid propellant rocket engine design schemes, with the same main principle of their operation. Let us consider the design and principle of operation of a liquid-propellant rocket engine using the example of a two-component engine with pumped fuel supply, as the most common, the design of which has become classic. Other types of liquid propellant rocket engines (with the exception of the three-component one) are simplified versions of the one under consideration, and when describing them it will be enough to indicate the simplifications.

In Fig. 1 schematically shows the liquid propellant rocket engine device.

Fuel system

The fuel system of a liquid-propellant rocket engine includes all elements used to supply fuel to the combustion chamber - fuel tanks, pipelines, turbopump unit(TNA) - a unit consisting of pumps and a turbine mounted on a single shaft, an injector head, and valves that regulate the fuel supply.

Pump feed fuel allows you to create high pressure in the engine chamber, from tens of atmospheres to 250 atm (LPRE 11D520 RN "Zenit"). High pressure provides a greater degree of expansion of the working fluid, which is a prerequisite for achieving a high specific impulse. In addition, with high pressure in the combustion chamber, a better value is achieved thrust-to-weight ratio engine - the ratio of the amount of thrust to the weight of the engine. The higher the value of this indicator, the smaller the size and weight of the engine (with the same amount of thrust), and the higher the degree of its perfection. The advantages of the pump system are especially noticeable in high-thrust liquid-propellant engines - for example, in the propulsion systems of launch vehicles.

In Fig. 1, exhaust gases from the TNA turbine enter through the nozzle head into the combustion chamber along with the fuel components (11). Such an engine is called an engine with closed loop(otherwise - with a closed cycle), in which the entire fuel flow, including that used in the TPU drive, passes through the combustion chamber of the liquid-propellant rocket engine. The pressure at the turbine outlet in such an engine should obviously be higher than in the combustion chamber of the liquid-propellant rocket engine, and at the inlet to the gas generator (6) feeding the turbine, it should be even higher. To meet these requirements, the same fuel components (under high pressure) that the liquid propellant engine itself operates on are used to drive the turbine (with a different ratio of components, usually with excess fuel to reduce the thermal load on the turbine).

An alternative to a closed loop is open loop, in which the turbine exhaust is produced directly into the environment through an outlet pipe. The implementation of an open cycle is technically simpler, since the operation of the turbine is not connected with the operation of the liquid propellant engine chamber, and in this case, the TPU can generally have its own independent fuel system, which simplifies the procedure for starting the entire propulsion system. But closed-cycle systems have slightly better specific impulse values, and this forces designers to overcome the technical difficulties of their implementation, especially for large launch vehicle engines, which have particularly high requirements for this indicator.

In the diagram in Fig. 1 one pump pump pumps both components, which is acceptable in cases where the components have comparable densities. For most liquids used as propellant components, the density varies in the range of 1 ± 0.5 g/cm³, which allows the use of one turbo drive for both pumps. The exception is liquid hydrogen, which at a temperature of 20°K has a density of 0.071 g/cm³. Such a light liquid requires a pump with completely different characteristics, including a much higher rotation speed. Therefore, in the case of using hydrogen as a fuel, an independent fuel pump is provided for each component.

With low engine thrust (and, therefore, low fuel consumption), the turbopump unit becomes too “heavy” an element, worsening the weight characteristics of the propulsion system. An alternative to a pump fuel system is repressive, in which the supply of fuel to the combustion chamber is ensured by the boost pressure in the fuel tanks, created by compressed gas, most often nitrogen, which is non-flammable, non-toxic, non-oxidizing and relatively cheap to produce. Helium is used to pressurize tanks with liquid hydrogen, since other gases condense at the temperature of liquid hydrogen and turn into liquids.

When considering the operation of an engine with a displacement fuel supply system from the diagram in Fig. 1, the TNA is excluded, and the fuel components are supplied from the tanks directly to the main valves of the liquid propellant engine (9) and (10). The pressure in the fuel tanks during positive displacement must be higher than in the combustion chamber, and the tanks must be stronger (and heavier) than in the case of a pump fuel system. In practice, the pressure in the combustion chamber of an engine with displacement fuel supply is limited to 10 - 15 at. Typically, such engines have a relatively low thrust (within 10 tons). The advantages of the displacement system are the simplicity of the design and the speed of the engine's response to the start command, especially in the case of using self-igniting fuel components. Such engines are used to perform maneuvers of spacecraft in outer space. The displacement system was used in all three propulsion systems of the Apollo lunar spacecraft - service (thrust 9,760 kG), landing (thrust 4,760 kG), and takeoff (thrust 1,950 kG).

Nozzle head- the unit in which they are mounted injectors, designed for injection of fuel components into the combustion chamber. The main requirement for injectors is the fastest and most thorough mixing of the components upon entering the chamber, because the rate of their ignition and combustion depends on this.
Through the nozzle head of the F-1 engine, for example, 1.8 tons of liquid oxygen and 0.9 tons of kerosene enter the combustion chamber every second. And the residence time of each portion of this fuel and its combustion products in the chamber is calculated in milliseconds. During this time, the fuel should burn as completely as possible, since unburned fuel means a loss of thrust and specific impulse. The solution to this problem is achieved by a number of measures:

• Maximum increase in the number of nozzles in the head, with proportional minimization of the flow rate through one nozzle. (The engine injector head contains 2600 oxygen injectors and 3700 kerosene injectors).
• Special geometry of the nozzles in the head and the order of alternating fuel and oxidizer nozzles.
• The special shape of the nozzle channel, due to which rotation is imparted when the liquid moves through the channel, and when it enters the chamber it is scattered to the sides by centrifugal force.

Cooling system

Due to the rapidity of the processes occurring in the combustion chamber of a liquid-propellant rocket engine, only an insignificant part (fractions of a percent) of the total heat generated in the chamber is transferred to the engine structure, however, due to the high combustion temperature (sometimes over 3000°K), and a significant amount of heat generated, even a small part of it is enough for thermal destruction of the engine, so the problem of cooling the liquid propellant engine is very relevant.

For liquid-propellant rocket engines with pumped fuel supply, two methods of cooling the walls of the liquid-propellant rocket engine chamber are mainly used: regenerative cooling And wall layer, which are often used together. Often used for small engines with positive displacement fuel systems. ablative cooling method.

Regenerative cooling consists in the fact that in the wall of the combustion chamber and the upper, most heated part of the nozzle, a cavity is created in one way or another (sometimes called a “cooling jacket”), through which one of the fuel components (usually fuel) passes before entering the mixing head, thus cooling the chamber wall. The heat absorbed by the cooling component is returned to the chamber along with the coolant itself, which justifies the name of the system - “regenerative”.

Various technological methods have been developed to create a cooling jacket. The liquid-propellant rocket engine chamber of the V-2 rocket, for example, consisted of two steel shells, internal and external, repeating the shape of each other. The cooling component (ethanol) passed through the gap between these shells. Due to technological deviations in the gap thickness, uneven fluid flow arose, resulting in the creation of local overheating zones of the inner shell, which often “burned out” in these zones, with catastrophic consequences.

In modern engines, the inner part of the chamber wall is made of highly thermally conductive bronze alloys. Narrow thin-walled channels are created in it by milling (15D520 RN 11K77 Zenit, RN 11K25 Energia), or acid etching (SSME Space Shuttle). From the outside, this structure is tightly wrapped around a load-bearing sheet shell made of steel or titanium, which absorbs the force load of the internal pressure of the chamber. The cooling component circulates through the channels. Sometimes the cooling jacket is assembled from thin heat-conducting tubes, sealed with a bronze alloy for tightness, but such chambers are designed for lower pressure.

Wall layer(boundary layer, Americans also use the term “curtain”) is a gas layer in the combustion chamber, located in close proximity to the chamber wall, and consisting mainly of fuel vapor. To organize such a layer, only fuel nozzles are installed along the periphery of the mixing head. Due to an excess of fuel and a lack of oxidizer, the chemical combustion reaction in the near-wall layer occurs much less intensely than in the central zone of the chamber. As a result, the temperature of the wall layer is significantly lower than the temperature in the central zone of the chamber, and it insulates the chamber wall from direct contact with the hottest combustion products. Sometimes, in addition to this, nozzles are installed on the side walls of the chamber, removing part of the fuel into the chamber directly from the cooling jacket, also with the aim of creating a wall layer.

Launch of the rocket engine

Launching a liquid propellant rocket engine is a responsible operation, fraught with serious consequences in the event of emergency situations during its execution.

If the fuel components are self-igniting, that is, entering into a chemical combustion reaction upon physical contact with each other (for example, heptyl / nitric acid), the initiation of the combustion process does not cause problems. But in the case where the components are not such, an external ignition initiator is required, the action of which must be precisely coordinated with the supply of fuel components to the combustion chamber. An unburned fuel mixture is an explosive of great destructive power, and its accumulation in the chamber threatens a serious accident.

After ignition of the fuel, maintaining a continuous process of its combustion occurs by itself: the fuel newly entering the combustion chamber is ignited due to the high temperature created during the combustion of previously introduced portions.

For the initial ignition of fuel in the combustion chamber when starting a liquid-propellant rocket engine, different methods are used:

• The use of self-igniting components (usually based on phosphorus-containing starting fuels, self-igniting when interacting with oxygen), which at the very beginning of the engine starting process are introduced into the chamber through special, additional nozzles from the auxiliary fuel system, and after the start of combustion, the main components are supplied. The presence of an additional fuel system complicates the design of the engine, but allows it to be restarted several times.
• An electrical igniter located in the combustion chamber near the mixing head which, when energized, creates an electric arc or series of high voltage spark discharges. This igniter is disposable. Once the fuel is ignited, it burns.
• Pyrotechnic igniter. Near the mixing head, a small pyrotechnic incendiary bomb is placed in the chamber, which is ignited by an electric fuse.

Automatic engine starting coordinates the action of the igniter and the fuel supply in time.

The launch of large liquid-propellant rocket engines with a pump fuel system consists of several stages: first, the pump starts and accelerates (this process can also consist of several phases), then the main valves of the liquid-propellant rocket engine are turned on, usually in two or more stages with a gradual increase in thrust from stage to stage. steps up to normal.

For relatively small engines, it is practiced to start the rocket engine immediately at 100% thrust, called “cannon”.

LRE automatic control system

A modern liquid-propellant rocket engine is equipped with rather complex automation, which must perform the following tasks:

• Safe starting of the engine and bringing it to the main mode.
• Maintaining stable operating conditions.
• Thrust change in accordance with the flight program or at the command of external control systems.
• Turning off the engine when the rocket reaches a given orbit (trajectory).
• Regulating the ratio of component consumption.

Due to the technological variation in the hydraulic resistance of the fuel and oxidizer paths, the ratio of component flow rates in a real engine differs from the calculated one, which entails a decrease in thrust and specific impulse in relation to the calculated values. As a result, the rocket can fail its task by completely consuming one of the fuel components. At the dawn of rocket science they fought against this by creating guaranteed fuel supply(the rocket is filled with more than the calculated amount of fuel, so that it is enough for any deviations of real flight conditions from the calculated ones). The guaranteed fuel supply is created at the expense of the payload. Currently, large rockets are equipped with an automatic control system for the ratio of component consumption, which makes it possible to maintain this ratio close to the calculated one, thus reducing the guaranteed fuel supply, and accordingly increasing the payload mass.

The automatic control system of the propulsion system includes pressure and flow sensors at different points of the fuel system, and its executive bodies are the main valves of the rocket engine and turbine control valves (in Fig. 1 - positions 7, 8, 9 and 10).

Fuel components

The choice of fuel components is one of the most important decisions when designing a liquid propellant engine, predetermining many details of the engine design and subsequent technical solutions. Therefore, the choice of fuel for a liquid-propellant rocket engine is made with a comprehensive consideration of the purpose of the engine and the rocket on which it is installed, the conditions of their operation, production technology, storage, transportation to the launch site, etc.

One of the most important indicators characterizing the combination of components is specific impulse, which is especially important in the design of spacecraft launch vehicles, since the ratio of the mass of fuel and payload, and therefore the size and mass of the entire rocket (see Tsiolkovsky Formula), which, if the specific value is not high enough, greatly depends on it impulses may turn out to be unrealistic. Table 1 shows the main characteristics of some combinations of liquid fuel components.

Table 1.

Oxidizer	Fuel	Average density fuel, g/cm³	Chamber temperature combustion, °K	Void specific impulse, s
Oxygen	Hydrogen	0,3155	3250	428
	Kerosene	1,036	3755	335
		0,9915	3670	344
	Hydrazine	1,0715	3446	346
	Ammonia	0,8393	3070	323
Dianitrogen tetroxide	Kerosene	1,269	3516	309
	Unsymmetrical dimethylhydrazine	1,185	3469	318
	Hydrazine	1,228	3287	322
Fluorine	Hydrogen	0,621	4707	449
	Hydrazine	1,314	4775	402
	Pentaborane	1,199	4807	361

Jet engines that run on compressed cold gas (for example, air or nitrogen) are also single-component. Such engines are called gas jet engines and consist of a valve and a nozzle. Gas jet engines are used where the thermal and chemical effects of the exhaust jet are unacceptable, and where the main requirement is simplicity of design. These requirements must be met, for example, by individual cosmonaut movement and maneuvering devices (UPMK), located in the backpack behind the back and intended for movement when working outside the spacecraft. UPMKs operate from two cylinders of compressed nitrogen, which is supplied through solenoid valves to a propulsion system consisting of 16 engines.

Three-component rocket engines

Since the early 1970s, the USSR and the USA have been studying the concept of three-propellant engines that would combine a high specific impulse when using hydrogen as fuel, and a higher average fuel density (and, therefore, smaller volume and weight of fuel tanks), characteristic of hydrocarbon fuel. When starting, such an engine would run on oxygen and kerosene, and at high altitudes it would switch to using liquid oxygen and hydrogen. This approach may make it possible to create a single-stage space launch vehicle. A Russian example of a three-component engine is the RD-701 liquid-propellant rocket engine, which was developed for the MAKS reusable transport and space system.

It is also possible to use two fuels simultaneously - for example, hydrogen-beryllium-oxygen and hydrogen-lithium-fluorine (beryllium and lithium burn, and hydrogen is mostly used as a working fluid), which makes it possible to achieve specific impulse values ​​in the region of 550-560 seconds, however technically very difficult and has never been used in practice.

Rocket control

In liquid rockets, engines often, in addition to their main function of generating thrust, also serve as flight controls. Already the first guided ballistic missile V-2 was controlled using 4 graphite gas-dynamic rudders placed in the engine jet stream along the periphery of the nozzle. By deflecting, these rudders deflected part of the jet stream, which changed the direction of the engine thrust vector and created a moment of force relative to the center of mass of the rocket, which was the control action. This method significantly reduces engine thrust; moreover, graphite rudders in a jet stream are subject to severe erosion and have a very short service life.
Modern missile control systems use PTZ cameras Liquid rocket engines, which are attached to the load-bearing elements of the rocket body using hinges that allow the camera to be rotated in one or two planes. The fuel components are supplied to the chamber using flexible pipelines - bellows. When the camera deviates from an axis parallel to the axis of the rocket, the thrust of the camera creates the required control torque. The cameras are rotated by hydraulic or pneumatic steering machines, which execute commands generated by the rocket control system.
In the domestic space launch vehicle Soyuz (see photo in the title of the article), in addition to 20 main, fixed cameras of the propulsion system, there are 12 rotating (each in its own plane), smaller control cameras. The steering chambers share a common fuel system with the main engines.
Of the 11 propulsion engines (all stages) of the Saturn-5 launch vehicle, nine (except for the central 1st and 2nd stages) are rotary, each in two planes. When using the main engines as controls, the operating range of camera rotation is no more than ±5°: due to the high thrust of the main camera and its location in the aft compartment, that is, at a considerable distance from the center of mass of the rocket, even a small deflection of the camera creates a significant control

Classification, schemes and types of rocket engines

Topic 2. LIQUID ROCKET ENGINES

Lecture No. 3

Questions for the seminar.

1. Concept and characteristics of insurance legal relations.

2. The difference between insurance legal relations and related relations.

3. Object of the insurance legal relationship.

4. Insurable interest in insurance.

5. Subjects of insurance legal relations.

Developed by Head of the Department of Civil Law, Doctor of Law, Professor M.V. Rybkina

Without pretending to be a complete and comprehensive account of modern liquid-propellant engines, the classification of the most common types of engines is presented in the figure (see Fig. 2.12.).

The proposed scheme is based on the principle of dividing all circuit and technical solutions into two large groups, differing in the principles of ensuring the supply of fuel components to the combustion chamber of the rocket engine. These are engines with pumping system submissions and engines with displacement feeding system components.

The first group mainly includes propulsion engines of launch vehicles, intercontinental ballistic missiles, and reusable space systems. The use of the second group of liquid propellant engines is, as a rule, limited to propulsion systems of spacecraft, large-sized modules of manned orbital complexes and transport ships, as well as propulsion systems of inter-orbital transportation vehicles.

Rice. 2.12. General classification of liquid rocket engines

An important classification feature of liquid propellant engines is also the method of utilization of the working fluid (fuel combustion products) obtained at the outlet of the engine's turbopump unit. According to this criterion, all motors are fundamentally divided into “open” circuit motors and “closed” circuit motors. In an “open” liquid-propellant rocket engine, after activation at the turbine, the generator gas is discharged either without additional use or is utilized in additional devices. In a liquid-propellant rocket engine of a “closed” design, the generator gas that enters the turbine enters the combustion chamber and is burned, due to the additional supply of one or two components entering the combustion chamber.

Depending on the type of gas generator, liquid-propellant rocket engines can be classified into engines with gas generators on the main or auxiliary fuel components, and also have a generator-less design, when the working fluid necessary to drive the pump is obtained by gasifying one of the fuel components in the cooling path of the chamber.

To improve efficiency and ratio useful action In the case of a turbopump unit, schemes with separate fuel pumps along the fuel and oxidizer lines are sometimes used, as well as schemes in which the turbopump unit also contains booster pumps necessary to create the required pressure at the inlet of the engine, especially when starting it.

Depending on the type of gas generator, liquid-propellant rocket engines can be classified into engines with gas generators on the main or auxiliary fuel components, and also have a generator-less design, when the working fluid required to drive the pump is obtained by gasifying one of the fuel components in the cooling path of the chamber.

To increase the efficiency and efficiency of a turbopump unit, schemes with separate fuel pumps and oxidizer pumps are sometimes used, as well as schemes in which the turbopump unit also contains booster pumps necessary to create the required pressure at the engine inlet, especially when its launch.

Relatively simple schemes are typical for liquid propellant engines with a displacement fuel supply system.

In a scheme with displacement fuel supply (see Fig. 2.13.), gas is supplied to the tanks with the oxidizer and fuel from a compressed gas cylinder (for example, nitrogen), while its pressure in the fuel component tanks is maintained constant using a reducer. The pressure in the gas cushion of the fuel tanks ensures the displacement of liquid-phase components into the combustion chamber of the liquid-propellant rocket engine. It is quite obvious that the pressure in the chamber cannot be higher than the pressure in the tanks. Shut-off valves are used to ensure that the engine starts and stops. The undoubted advantage of the scheme presented above is its simplicity and, as a result, reliability. However, with a displacement system, the compressed gas cylinder is heavy and the fuel tanks become significantly heavier. In general:

(2.18.)

Gas pressure in fuel tanks;

Pressure in the combustion chamber of the rocket engine;

Pressure losses in hydraulic paths and automation elements between the tanks and the engine chamber.

The pressure in the gas cushion of the fuel tanks ensures the displacement of liquid components into the combustion chamber of the liquid-propellant rocket engine. It is quite obvious that the pressure in the chamber cannot be higher than the pressure in the tanks. Shut-off valves are used to ensure that the engine starts and stops. The undoubted advantage of the above scheme is its simplicity and reliability. Since with increasing pressure in the chamber, the efficiency of the engine increases, the desire to increase it, for this liquid-propellant rocket engine scheme, is associated with an increase in the mass of all elements of the supply system and, above all, fuel tanks. Similar disadvantages apply to the displacement fuel supply system with two-component LGGs. However, less gas consumption is required to pressurize the fuel and oxidizer tanks. In this version of the circuit, the injection is carried out by combustion products obtained in the LGG, and the performance of the “heated” gas is much higher than that of the “cold” one.

The effect on the mass characteristics of a propulsion system with a liquid-propellant rocket engine can be clearly illustrated in the following example. If the propulsion system of the second stage of the Saturn - 5 launch vehicle were replaced with an installation with a displacement supply system at the same pressure in the combustion chamber of the liquid-propellant rocket engine, then the increase in the mass of such a propulsion system would be equal to the mass of the Apollo spacecraft, which made it would be impossible to implement the lunar program.

For the displacement circuit option (see Fig. 2.14.), one can expect a slight reduction in losses, since the displacement of components will be carried out by heated combustion products produced in the LGG.

From the explanations it follows why a displacement feed system with a balloon feed system is used exclusively in low-thrust engines with a pressure in the combustion chamber of a liquid-propellant rocket engine of no more than 10-12 · 10 5 Pa.

Practical application of low-thrust liquid propellant rocket engines (LPRE) is found in the creation of integrated propulsion systems (UPS) for artificial earth satellites (AES), spacecraft (SV) and spacecraft (SC). Being in orbit, when the pressure outside the aircraft is close to zero, the specific impulse can have a fairly high value, even with a low pressure in the chamber. It should be remembered that the specific impulse increases from the ratio of pressure in the combustion chamber to the pressure at the nozzle exit (see Fig. 2.10.).

Quite a lot of circuit solutions for ODEs using liquid propellant rocket engines can be considered. First of all, the difference in scheme options will depend on the requirements determined by the purpose of the aircraft. These can be engines running on both single-component and dual-component fuels. The schemes will differ in the principles of traction regulation and stabilization. Other factors may also influence the determination of the circuit design. However, in all variants of the circuit, the pressure in the gas accumulators must be higher than the pressure in the chambers, which determines the features of the displacement system for supplying components.

The presentation of all, or at least most, possible schemes of propulsion systems with displacement feed systems in this textbook is not included in the plans of the authors. Therefore, to illustrate possible circuit options, as an example, a diagram of an integrated propulsion system (ODS) for an artificial earth satellite (AES) powered by two component fuels is given (see Fig. 2.15.).

Rice. 2.15. Scheme of an ODE with a two-component liquid propellant rocket engine for an IC.

1. Pressure reducer, 2. Maneuvering rocket engine (Each with 22 N thrust),

3. Apogee liquid rocket engine (thrust 490 N)

The designs and fundamental features of the operation of liquid propellant rocket engines are very diverse. Among the most important problems in the creation of liquid propellant rocket engines is ensuring the operability of the combustion chambers. Especially when you consider that the resources required for a liquid propellant rocket engine significantly exceed the resources for the chambers of conventional liquid propellant rocket engines.

The list of similar ones can include: launching, organizing the work process, choosing a system to counteract the temperature effect on the chamber walls and a number of others. Most of the problems that are difficult to solve are associated, first of all, with the extremely low operating costs of the components. So for some chambers the consumption of oxidizer and fuel does not exceed 0.5 and 0.3 g/s, respectively. A similar circumstance, for example, determines the impossibility of using regenerative cooling of the walls (as the most effective), and choosing refractory metals for the manufacture of chamber walls, using heat-resistant heat-protective coatings, much lower than the shells

For propulsion systems, one of the diagrams of which is shown in Figure 2.15., used as part of a transport spacecraft or other aircraft and flying for a long time, fuel tanks must be refueled. Options for refueling systems are shown in the figure (see Fig. 2.16.).

Rice. 2.16. Diagrams of fuel tanks for aircraft refueled in flight.

1. Tank walls; 2. Pressure pipe; 3. Piston; 4. Fuel intake; 5. Bellows;

6. Elastic bag; 7. A rod with holes for pressurization; 8. Plastic diaphragm; 9. Plastic wetted partitions; 10. Central pipe for fuel intake.

A - with a piston; B - with a bellows displacement device (fuel outside the bellows); B - with a bellows displacement device (fuel inside the bellows); G - with a displacement bag (fuel outside the bag); D - with a displacement bag (fuel inside the bag); E - with a plastic diaphragm; F - with a capillary intake device.

More complete information on refueling systems can be found in the training manual, which is mentioned in the bibliography.

To implement liquid-propellant rocket engines of medium, high and ultra-high thrust, it is necessary to create engines with the highest possible increase in pressure in the combustion chamber. In such engine options, circuits with a turbopump system for supplying fuel components are used.

The figure (see Fig. 2.17.) shows a block diagram of a liquid-propellant rocket engine with a pump system for supplying components. A characteristic feature of the scheme under consideration should be considered that the gas exhausted from the turbine is simply discharged into the surrounding atmosphere. It should be noted that the combustion products after the turbine still have significant efficiency and not using them negatively affects the efficiency of the engine. Nevertheless, such schemes can be implemented.

Rice. 2.17. Pneumohydraulic circuit of a liquid propellant engine, with a turbopump supply of components to the combustion chamber.

A component of unitary rocket fuel (for example, hydrogen peroxide - H 2 O 2), from the tank, is supplied to the liquid gas generator. Gas generator is a unit designed to produce high-temperature generator gas used to drive the turbocharger turbine. The turbine provides torque to the fuel and oxidizer pumps. The main components of the fuel are supplied by pumps into the engine chamber, and the fuel, as a rule, is used to cool the chamber, for which it is supplied to the gap between its walls, usually called the cooling “jacket”. The oxidizer is supplied directly to the nozzle head of the chamber, where it is mixed with fuel heated in the cooling path. The process of interaction of fuel components occurs in the combustion chamber. The resulting high-temperature combustion products pass through the critical section of the chamber and expand in the nozzle to supersonic speeds. The outflow of combustion products is the final phase of the liquid rocket engine operation and forms the thrust of the rocket engine.

Schemes of this type, which are called “open circuits,” can be more effective if, after triggering on the turbine, the generator gas can be discharged through additional devices that ensure utilization of the energy of the discharged gas...

In the general case, the magnitude of the thrust of an “open” rocket engine can consist of a value equal to the sum of the thrusts produced by the main chamber and the additional after-turbine device. A similar effect can be obtained by ensuring the removal of the generator gas into the auxiliary nozzle; implementation into the supercritical part of the main nozzle, in different design options for the main nozzle.

The figure (see Fig. 2.18) shows diagrams of devices in which the generator gas, after realizing part of its energy in the turbine, is used to create additional thrust.

Fig. 2.18 Diagrams of devices that utilize post-turbine gas

In any of the presented options, the additional thrust implemented in the device must be taken into account.

Those. there is a relation:

where: is the total thrust of the liquid-propellant rocket engine of the “open” design;

Thrust produced by the main chamber of the rocket engine;

Thrust produced in auxiliary devices.

Using the previously given dependencies to determine the specific impulse (see equations 2.11, 2.12. and 2.13), we transform expression 2.19. to view 2.20.

(2.20.)

where: - effective specific impulse of the liquid-propellant rocket engine of the “open” circuit;

Specific impulses provided by the main chamber and auxiliary devices, respectively;

Mass fuel consumption in the gas generator and total mass fuel consumption in the liquid propellant engine.

Dependency Analysis 2.20. shows that the value of the effective specific impulse is greater, the smaller the proportion of fuel consumed through the gas generator and the more efficiently the generator gas is utilized after activation at the turbine. There is a well-defined dependence characterizing the influence of pressure in the chamber of an “open” liquid-propellant rocket engine on the value of the specific impulse. In contrast to the monotonic increase in . In the general case discussed above, with an increase in pressure in the chambers of liquid propellant engines operating according to the scheme without afterburning of the generator gas, a clearly defined region corresponding to the optimal value is observed (see Fig. 2.19.).

Fig.2.19. Dependence of specific impulse on chamber pressure

open circuit engine

The appearance of an extremum in the dependence is explained by the necessary increase in fuel consumption through the gas generator with increasing pressure in the combustion chamber. An increase in flow rate is required to increase turbine power to meet the increased need of pumps for more torque. This situation leads to an increase in the proportion of inefficiently used fuel and, as a consequence, to a decrease in the specific impulse of the rocket engine.

It is permissible to provide for the release of gas generator gas through special rotary nozzles used to control the flight of the rocket

In order to make maximum use of the capabilities of rocket fuel, through the efforts of Russian scientists and engineers, a scheme for organizing the working process of a liquid-propellant rocket engine was developed, which provides for afterburning of the generator gas in the combustion chamber after its activation on the TNA turbine, the so-called “schemes with afterburning of the generator gas” (see Fig. 2.20 .).

Rice. 2.20. Structural diagrams of liquid-propellant rocket engines with afterburning of generator gas

1. and 2. Tanks with fuel and oxidizer, 3. LGG, 4. and 5. fuel and oxidizer pumps, 7., 8. and 9. valves, 10. combustion chamber.

The main feature of the “closed” circuit, made according to the variant Fig. 2.20 is as follows. All the oxidizer necessary for the operation of the CS is supplied to the gas generator. There is minimal supply there. required amount fuel. The ratio of fuel components supplied to the gas generator is dictated solely by the need to obtain gas at a temperature acceptable to ensure the thermomechanical loads of the turbine. After the generator gas is activated at the turbine, which in the case under consideration has an excess of the oxidizing component, the gas is supplied to the compressor station. An additional amount of fuel necessary to maintain the optimal ratio of fuel components is also supplied there. In this version, the liquid-propellant rocket engine operates according to the “gas (oxidizer) – liquid (fuel)” scheme. It is also possible to organize a work process when an excess amount of fuel is supplied to the gas generator with a lack of oxidizer. In the first case they speak of an oxidizing gas generator, in the second - a reducing gas generator.

Both methods have their advantages and disadvantages. In the case of a reducing gas generator, the issues of ensuring thermal stability are much easier to solve, since at high temperatures of the working process in a gas generator it is much easier to protect structural materials (mainly metals and their alloys) from fire in the presence of a reducing environment. At the same time, an excess of fuel with an insufficient amount of oxidizer is fraught with a number of negative consequences associated with incomplete combustion of the fuel, which in the case of carbon-containing components leads to the precipitation of the solid phase of carbon and, as a consequence, to erosive wear of turbine blades and other elements of the pump.

The oxidative gas generation scheme does not have these disadvantages, but it has its own characteristics. They consist in the need to use refractory structural materials that are resistant to fire in an oxidizing environment, which leads to an increase in the cost of engines, a potential decrease in their stability when exposed to microparticles in the oxidizing gas flow entering the turbine blades, which makes it difficult to create highly reliable liquid propellant engines.

In practice, the reduction gas generation scheme is used, most often, in oxygen-hydrogen liquid rocket engines, where the fuel (liquid hydrogen) does not contain carbon and, therefore, there is fundamentally no danger of soot formation. In the future, the possibility of using the first member of the homologous series of saturated hydrocarbons, methane (CH 4), as a rocket fuel, is being considered, the carbon content of which is minimal, which makes its effective use in gas generators of a reduction scheme fundamentally possible.

The liquid-propellant rocket engine scheme presented above is implemented according to the “gas-liquid” scheme. This variant of the scheme provides for the organization of a work process with afterburning of generator gas.

In another option, afterburning of generator gas can be built according to the “gas-to-gas” scheme. The main difference between this scheme is the presence of two gas generators. One gas generator operates according to an oxidative scheme, the second – a reduction scheme. It is preferable to use hydrogen or a hydrocarbon fuel with a minimum mass content of carbon (kerosene, etc.) for the reduction gas generator, and liquid oxygen as an oxidizer. Thus, the introduction of liquid hydrogen into the composition of rocket fuel makes it possible to significantly reduce the release of the condensed phase of carbon (soot), thereby ensuring the possibility of more reliable operation of the reduction gas generator.

The gas generation products enter the oxidizing and reducing gas turbines, and then, after passing through the turbines, into the combustion chamber, where their final interaction occurs, with the required ratio of components (see Fig. 2.21.).

Rice. 2.21. Pneumohydraulic circuit of a liquid propellant engine with afterburning of generator gases.

1. and 2. Tanks with fuel and oxidizer, 3. and 4. LGG gas with excess fuel and LGG gas with excess oxidizer, 5. and 6. Fuel and oxidizer pumps, 7. and 8. Turbines for fuel gas and oxidizing gas, 9. and 10. Valves, 11. Combustion chamber.

A similar scheme can be in a slightly different design, when there are two gas generators. LGG with excess fuel ensures pressurization of the fuel tank. The second gas generator produces oxidizing high-temperature gas, one part of which enters the turbine and after the turbine into the main combustion chamber. The second - smaller part in the mixer is supplemented with an additional amount of oxidizer and is used to inflate the oxidation tank.

For a hydrogen-oxygen engine, a gas-free generator circuit is usually used (see Fig. 2.22.).

Fig.2. 22. Gas-free liquid-propellant rocket engine scheme

1. Combustion chamber, 2. draft regulator, 3. Liquid hydrogen pump. 4. Liquid oxygen pump, 5. Speed ​​reducer, 6. turbine, 7. 8. and 9. shut-off valves, 10. ignition system valve..

In the pneumatic-hydraulic gas-generator-free operation scheme of the liquid-propellant rocket engine, the following order of operation is provided. Components from the tanks enter the pump inlet through inlet valves. The THA engine has a two-shaft design with parallel shafts and a gear reducer. This important feature of this TNA. The centrifugal hydrogen pump is mounted on the same shaft as the turbine, has two stages and an axial inlet. The first stage of the pump is screw-centrifugal. The screw centrifugal oxygen pump is single-stage. The turbine is axial two-stage, jet.

Liquid oxygen, through a valve block with an electromechanical component ratio regulator, enters the cavity of the mixing head from the pump. In flight, according to signals from the tank emptying system, the ratio of components can vary within ± 10%. Hydrogen from the pump is supplied through a pipeline to the inlet manifold of the chamber's cooling path.

Liquid hydrogen from the pump enters the collector located in the critical section of the nozzle. From the collector, along part of the tubes, hydrogen is directed to the nozzle exit, then, through another part of the tubes, it moves to the collector near the head. From this collector, hydrogen gas, heated in the cooling circuit to a temperature of 200K, is directed from the draft regulator to the turbine. The draft regulator operates on the principle of bypassing part of the hydrogen to the outlet of the turbine. From the turbine, waste hydrogen passes through a start-up and shut-off valve through a gas duct into the mixing head. All main valves are controlled by helium gas with using control valves.

The diagram also shows valves that ensure the operation of the engine cooling system before starting. Such an operation is necessary for the normal start of an engine using cryogenic components. what is necessary for hydraulic systems. The tanks are pressurized with helium gas, the reserve of which is in a special cylinder.

A number of liquid-propellant rocket engine schemes were discussed above, in which TNAs are used to supply components to the CS. At low pressures in the inlet pipes, stall modes may occur, characterized by the onset of cavitation in the inter-blade cavities of pumps. In all the presented pneumohydraulic schemes of liquid propellant rocket engines equipped with pump pumps, gas is supplied to the tanks with components from the cylinders through gearboxes, which supercharges them. In this case, one could count on obtaining the required pressure at the inlet to the pumps. At the same time, the pressure in the tanks required for normal operation of the centrifugal screw pump is often unacceptably high, which leads to a noticeable increase in the wall thickness and weight of the tanks. This disadvantage can be avoided if an additional pumping (booster) pumping unit (BPU) is installed at the outlet of the tanks. Installing a BPU, which ensures the operation of the main pump of the TPU, can significantly reduce the amount of pressurization of the tanks and, consequently, their weight. Therefore, the design of a modern pump is unthinkable without the sequential use of various pumps arranged in a multi-stage scheme. The role of boosters can be performed by an axial blade (auger) or a jet pump (ejector).

Pumping booster pump units (BPU), which are usually called pre-pumps, are located in close proximity to the tank with the component, which eliminates hydraulic losses when supplying the component from the tank to the entrance to the BPU pump. In the figure (see Fig. 2.30).

Rice. 2.30. Booster device circuits

Option a). 1. Tank with a component, 2. centrifugal prepump, 3. liquid turbine of the prepump unit, 4. turbine of the main pump, 5. pump.

Option b). 1. Tank with a component, 2. pre-pump, 3. gas turbine of the pre-pump unit, 4. pump of the main heat pump.

Option c). 1. Tank with the component, 2. jet pre-pump (ejector), 3. ejector nozzle, 4. pump of the main pump., 5. Component supply line to the ejector nozzle.

In the scheme of option “a”, the BNA hydraulic turbine is driven by liquid high pressure, taken from the TNA pump. After activation on the turbine, the liquid returns to the pressure line. In the scheme of option “b”, the gas turbine operates on the gas of the main LGG, and in option “c”, the jet pre-pump-ejector, as well as in the option of scheme “a”, is powered by a component from the pump of the main HPU.

As follows from the above brief analysis of the effectiveness of possible variants of liquid-propellant rocket engine schemes, increasing the pressure in the chamber does not in all cases lead to an increase in the specific impulse. The discussed features of the construction of liquid-propellant rocket engine schemes, to a greater extent, relate to the schemes of high- and ultra-high-thrust engines, and also, to a certain extent, to medium-thrust engines. The figure (see Fig. 2.31.) shows the qualitative dependence of the specific impulses of the chamber and the liquid propellant rocket engine, made according to the displacement circuit, according to the “open” circuit and according to the “closed” circuits of various options.

Rice. 2.31. Dependence of specific impulse on chamber pressure

From the analysis of the graph it follows that in engines running on a liquid-liquid circuit, with increasing pressure the specific impulse of the chamber increases monotonically. However, in the future, due to an increase in gas consumption for the TPU drive (see Fig. 2.26.), the specific impulse of the engine increases only to a certain limit. The increase in specific impulses of engines built according to closed circuits increases with increasing pressure in the chamber, although it is not very significant.

When choosing a liquid propellant engine option for a newly designed aircraft, in addition to using the data obtained from the analysis of the graph presented in Figure 2.18, you should consider the dependence called the altitude characteristic (Figure 2.32.).

Rice. 2.32. Altitude characteristics.

On the image. 2.32. changes in the main engine parameters with changes in back pressure are presented. As can be seen from the figure, the flow of the altitude characteristics of the liquid-propellant rocket engine with a change in ambient pressure environment can be divided into two sections: the section of operation of the nozzle without shock wave I and the section of operation of the nozzle with shock wave P.

In the section with non-shock operation of the nozzle, thrust and specific thrust decrease linearly with increasing pressure environment. In this case, the working process in the chamber and its nozzle is independent of ambient pressure. At a certain pressure pk a shock wave enters the chamber nozzle - the linearity of the change in thrust and specific thrust is disrupted. The nature of the change in thrust and specific thrust in the operating mode of the nozzle with a shock wave is determined by the pattern of movement of the shock wave into the depth of the nozzle and the recovery of pressure behind the shock wave. In Figure 2.33. The dotted lines show the nature of the change in the main parameters of the liquid-propellant rocket engine, for the case if the shock wave did not enter the nozzle and the usual gas expansion occurred at all nozzle pressures. From the moment the shock wave enters the nozzle, the pressure behind the shock increases as the shock wave penetrates deep into the nozzle. A similar operating mode is observed in the first-stage liquid-propellant rocket engines of intercontinental missiles, the pressure at the nozzle exit of which is selected sufficiently small from the condition of obtaining the average maximum specific thrust in the active section of the rocket's trajectory. or for rockets. For this type of rocket, the engine parameters are selected from the condition of obtaining the average maximum specific thrust in the air portion of the trajectory. Therefore, for these rockets the pressure at the nozzle exit is quite low and atmospheric pressure it is enough for the shock wave to enter the depth of the nozzle. The figure shows that in specified conditions The mode of operation of the nozzle with a shock wave improves the characteristics of the liquid-propellant rocket engine.

For the rocket version, for which it is necessary for the thrust to change during flight, the liquid-propellant rocket engine must be made with a throttle characteristic (see Fig. 2.33.).

Rice. 2.33. Throttle characteristics of liquid propellant rocket engines.

As it follows from the figure, to change the magnitude of the traction force, a change in the costs of the components is required. However, it should be remembered that the change in flow rate is ensured by correcting the difference across the injectors in accordance with the following expression.

, (2.21.)

where G is the flow rate of the component through the nozzle,

Nozzle flow coefficient,

F f – nozzle nozzle exit cross-sectional area,

Component density,

Pressure drop across the nozzle.

In addition to the presented options, another direction for circuit improvement is three-component liquid propellant rocket engines. In a liquid rocket engine of this type, some kind of hydrocarbon (for example, kerosene) and liquid hydrogen are simultaneously used as fuel, and liquid oxygen is used as an oxidizer. Three-component engines also make it possible to fully realize the possibility of effectively using different rocket fuels on board the same aircraft. Ballistic and mass calculations of the efficiency of using various fuels in the propulsion systems of launch vehicles, ballistic missiles, and reusable space systems are largely determined by the characteristics of the rocket fuel used. As shown earlier, fuels determine the value of the specific impulse of the liquid-propellant rocket engine, which is especially important for the engines of the upper stages of the launch vehicle, while the first stages can be equipped with liquid-propellant rocket engines with not so much high value, but the fuel density should be maximum.

Three-component engines make it possible to ensure the operation of the first stages with a minimum hydrogen content in the rocket fuel. That is, it indicates the feasibility of using fuel with a higher density. At subsequent stages of rocket flight, hydrogen, as a more energy-intensive and lower-density fuel, is more preferable, since its use will lead to an increase in the specific impulse of the rocket engine, and, consequently, the efficiency of the entire aircraft.

A liquid propellant engine can provide the required parameters and characteristics, provided that automation and engine control units are included in the pneumohydraulic circuit (PGS). The most important functions performed by ASG units include:

· stabilization of the ratio of components supplied to the combustion chamber;

· maintaining the required level or regulating traction;

· ensuring control and management of the operation of the engine and its main units (combustion chamber, pump pump, gas generator and, possibly, some others), which determine its overall performance.

For specific types of engines, the presented list can be expanded.

As has been noted more than once, for this textbook, observing the conditions of brevity of the presented materials, it is not possible to present possible options for ASG with descriptions of the circuits included in the engines of automation and control units. You can only indicate in the list of literary sources a list of special teaching aids on this issue.

However, diagrams and design features of the main units will be presented.

By using the word “main” units, the authors mean units that provide the most important functional parameters and characteristics of the liquid-propellant rocket engine. These include combustion chambers, turbopump units, and gas generators. These units will determine the type of rocket engine. Work on their creation requires the greatest time and financial costs. At the same time, it is necessary to emphasize that the degree of importance in determining the performance of liquid propellant engines, and sometimes reliability, of units not mentioned among the main ones (valves, regulators, etc.) require no less attention to their design and development.

2.5.1. Combustion chambers of liquid propellant rocket engines

The combustion chamber is designed in a certain sequence. Initially, unless specifically specified in the technical specifications, the components and the optimal pressure in the combustor are selected. The design of the combustor is determined after performing gas-dynamic calculations. Based on the results of these calculations, the geometric dimensions and gas-dynamic profile of the CS are established (see Fig. 2.34.).

Rice. 2.34. Gas-dynamic profile of the combustion chamber.

The liquid-propellant rocket engine combustor experiences extremely high thermal loads. For engines of medium, high and very high thrusts, for almost all types of components, the cooling system is performed with external cooling. For low-thrust cameras, issues of temperature resistance are resolved taking into account the service life, geometric contours of the camera, traction force and other specific features of each camera option. The main structural elements of the CS, made with external cooling, are shown in the figure (see Fig. 2.35.)

Rice. 2.35. Combustion chamber with associated shells

1. Chamber body, 2. Mixing head, 3. Cylindrical part of the chamber, 4. Nozzle, 5. Chamber jacket, 6. Power bracket.

A. Curtain belt knot, b. Coolant (fuel) supply unit, c. Camera Mount Brackets

In Figure 2.35, the cooling component is introduced into the chamber jacket in the section of the outer diameter of the nozzle. This is not the only solution. The designer usually chooses the option of installing a component input manifold, depending on a number of reasons (the degree of nozzle expansion, the desire to reduce resistance along the path, strength, etc.).

The figure (see Fig. 2.36) shows options for the location of the input sections.

Rice. 2.36. Options for the location of sections for introducing the cooling component into the intershell gap of the chamber “jacket”.

A- at the exit section of the nozzle. b.- at the outlet section and in the middle section of the nozzle, V– in the middle section of the nozzle

In modern high-thrust engines, to increase the thermal resistance of the chamber, a number of design measures are used aimed at reducing the temperature of the most heat-stressed elements of the combustion chamber.

Such measures include:

· organization of regenerative cooling by pumping relatively cold fuel components through the cooling “jacket”;

· the use of so-called “cooling curtains”, which are special zones of heat-stressed areas of the chamber, equipped with devices for supplying an additional amount of one of the fuel components (usually fuel) in order to reduce local heat flows;

· application of special measures in the most thermally loaded critical section of the chamber (reducing the intershell gap, inserts of refractory materials in the critical part of the nozzle).

To organize external cooling, the size of the gap is regulated by special spacers - connections. They also ensure the strength of the chamber and the stability of the inner shell of the chamber when the pressure of the cooling component in the gap of the “jacket” exceeds the pressure in the chamber. The figure (see Fig. 2.30.) shows the types of spacers used in modern CS designs. The spacers, outer and inner shells are connected by soldering, the composition of the solder is stable in the component and retains its strength characteristics when the walls are heated.

Rice. 2.37. Types of connections of CS shells.

A. corrugated spacer, b. ribbing of the inner shell, V. tubular chamber.

There is another important circumstance for increasing the performance of the CS, which is ensured by introducing connections into the design of the CS. The LRE chamber body experiences significant force loading. The combustion process can take place at product pressures of several tens of MPa. In this case, the pressure of the cooling component in the intershell gap must always be greater than the pressure in the chamber. Otherwise, the component will not be able to enter the KS. Consequently, the inner shell of the chamber, being under an external pressure difference equal to the difference between the supply pressure and the pressure in the chamber, can collapse and lose stability. And if, during the ongoing process in the chamber, it is heated up, then the mechanical characteristics of the shell material have a reduced value. On the first samples of engines, the outer and inner shells worked independently of one another (see Fig. 2.38.), which excluded the possibility of increasing the pressure in the combustor.

Rice. 2.38. Combustion chamber of the RD-1100 engine

1. Injector unit with ignition system, 2. independently operating (without connections) chamber shells. 3 nozzle block.

In modern liquid propellant engines, as noted earlier, the combustors are made with connected shells. When a cooling component is introduced into the “between-jacket” gap at the nozzle exit (the most commonly executed scheme) (see Fig. 2.39.), the greatest pressure drop acting on the inner shell is determined. In this section, the pressure of the component is maximum, and the pressure in the chamber is close to zero. An assessment of the strength reliability of chamber shells (shell strength, stability of the inner shell, strength of connections and other positions) should be made taking this circumstance into account.

Rice. 2.39. Distribution of loads along the length of the chamber

The following notations are used on the graph: p g - pressure in the chamber, p g - pressure of the cooling component in the “intershell” gap, t g - gas temperature in the chamber, t avg.o. – average, according to the thickness of the inner shell, temperature, - pressure drop across the nozzle, m cool. – mass flow rate of the cooling component, L – chamber length..

It should be noted that the connection options presented in this manual, as the most frequently used in modern CS designs, have been tested by a large number of experiments and have proven themselves well in the operation of numerous samples of railway motors of various dimensions.

Another means of reducing the thermal impact on the inner wall of the chamber is the introduction of curtain units into the design. The figure (see Fig. 2.40) shows options for design solutions for curtain units through which flammable substances are introduced to ensure the creation of a gas-liquid film on the inner surface of the “jacket” shell.

Fig.2.40. Options for camera curtain components.

A – with holes , b – with slot gap

The combustion chambers of liquid propellant rocket engines are characterized by two types of operating modes (see Fig. 3.7.). For a chamber with a steady state of operation, the cooling system of the inner wall can be chosen according to the principle of chambers that have just been disassembled. The LPRE version operating in pulse mode can use a chamber with a “capacitive system” for protecting the chamber wall. This option provides for the design of a single shell (without a “cooling jacket”) of increased thickness and with additional stiffening rings (see Fig. 2.41.).

Rice. 2.41. Combustion chamber of a low-thrust liquid-propellant rocket engine.

1. Fuel valve block, 2. Combustion chamber, 3. Nozzle attachment mounting unit, 4. Nozzle attachment, 5. Igniter, 6. Fuel valve unit.

Such a solution is acceptable, since during breaks between the operation of the chamber, the wall “rests” from the effects of combustion products and its heating decreases.

A particularly important unit is the KS head. At the bottom of the head there are nozzles through which components enter the chamber. Types of nozzles vary significantly in design. In the figure (see Fig. 2.42). Listed are some warrants of jet, centrifugal and two-component injectors that are used in liquid-liquid engines.

Rice. 2.42. Liquid injector options.

1. Front bottom, 2. Middle bottom, 3. Two-piece jet-jet nozzle, 4. One-piece nozzle with swirler, 5. One-piece centrifugal-jet nozzle, 6. Two-piece centrifugal nozzle with tangential holes, 7. Spacer sleeve.

For engines operating according to schemes with afterburning of generator gas, the chamber heads are equipped with gas-liquid nozzles (Fig. 2.43.).

Rice. 2 43. Options for gas-liquid injectors.

1. Front bottom, 2. Middle bottom, 3. Jet-jet nozzle, 4. Jet-centrifugal nozzle, 5. Jet-centrifugal nozzle with a screw swirler, 6. Two-stage (combined) nozzle: the first stage is gas-liquid jet-jet, the second cascade is liquid centrifugal with tangential holes.

The injector option for the mixing head is selected by the designer based on previously gained experience in testing the engine chamber - prototype and performing calculations. The location of the nozzles on the bottoms of the head is dictated by the designer’s desire to obtain the best complete combustion of the components and the need to create an effective wall layer of fuel. The last of the mentioned positions should ensure acceptable heating of the inner wall of the chamber (see Fig. 2.44).

Rice. 2.44. Layout of injectors on KS heads

A - Honeycomb arrangement of injectors.

1. Jet-centrifugal nozzles, 2. Centrifugal nozzles.

b – Staggered arrangement of nozzles

1. Oxidizer nozzle 2. Fuel nozzle.

V– Arrangement of nozzles in concentric circles

1. Two-component nozzle, 2. Single-component nozzle

From examination of the figures it follows that, regardless of the arrangement of nozzles on the bottoms of the mixing head, it is necessary to form a reliable curtain of fuel injectors located on the outer diameter.

The CS liquid-propellant engine has a large number of components necessary for the normal functioning of the engine. These are manifolds for input and output of components, nodes of curtain belts, nodes for connecting parts of the chamber (mixing head, cylindrical and nozzle sections), starting and stopping nodes, brackets that transmit traction force to the aircraft, etc. All of the listed nodes must be designed and evaluated calculations, and also subjected to tests confirming their performance. The desire of the authors to highlight such features of the creation of a CS is not consistent with the need to ensure the brevity of the presented textbook.

The assessment of the perfection of the CS is characterized by the specific impulse completeness coefficient, determined by the following expression:

, (2.22.)

where: - specific impulse completeness coefficient,

I sp.p - experimentally measured specific impulse,

Theoretical specific impulse,

Process perfection factor in the chamber,

Process perfection coefficient in the chamber nozzle,

The design coefficient is determined based on statistical data obtained from testing engines operating on similar components. Typically, the value of this coefficient is 0.96...0.99.

The nozzle coefficient () is calculated taking into account friction losses () and losses due to the unevenness of the flow velocity field at the nozzle exit (). In addition, additional losses () associated with cooling the flow in the nozzle, the degree of nonequilibrium, and others are taken into account:

. (2.23.)

In general, the numerical values ​​of the listed coefficients fall within the following limits: = 0.975... 0.999, = 0.98...0.99 and = 0.99...0.995. In this case, value = 0.945...0.975.

Taking into account the given values, the specific impulse completeness can range from 0.9 to 0.965.

2.5.2. Liquid gas generators (LGG).

Design solutions and features of intra-chamber processes largely depend on whether LGGs are installed on liquid propellant engines of “open” or “closed” circuits. For engines with an “open” design, LGGs are performed at a pressure close to the pressures of the main compressors. LGG engines of a “closed” circuit provide the working fluid (combustion products) of the turbine with a pressure significantly higher than the pressure in the main combustor. However, LGGs, both oxidative and reductive versions, operate at component ratios much lower than the values ​​established for CS. Consequently, the temperatures at which the process takes place in the gas generator chambers are also very different from the process temperatures in the compressor chamber.

Liquid propellant engines use two-component and one-component liquid gas engines. Two-component LGGs are most widely used. For engines with afterburning of generator gas, two-component LGGs are naturally used as the most natural ones. It can be noted that a significant part of the issues related to the features of the design and testing of this variant of the LGG are resolved according to the positions adopted for the CS. The mixing head of the nozzle and their location on the bottom of the head will be made according to the schemes used when choosing similar solutions for the CS. At the same time, given the relatively low temperature level in the LGG chamber, an uncooled version of the wall is usually used. The figure (see Fig. 2.45) shows the main part of a two-component ZhGG, one of the domestic engines.

Rice. 2.45. Two-component ZhGG

A similar version of the ZhGG was used as part of the RD-111 engine. The arrows in the figure show the component input fittings.

The development of single-component gas generators is carried out according to different principles. In the recent past, hydrogen peroxide (H 2 O 2) was used as a component for such gas generators. A special substance (catalyst) was located in the gas generator chamber, the interaction with which hydrogen peroxide led to the production of water vapor and gaseous oxygen at high temperatures (from 720 to 1030 K at a concentration of 80% and 90%, respectively). The figure (see Fig. 2.46) shows the gas generator (the so-called gas generator that produces steam as the working fluid of the turbine), developed by the Energomash enterprise for the RD-107 liquid rocket engine and its modifications.

Rice. 2.46. Single-component liquid gas generator.

1. Component inlet fitting, 2. catalyst packages, 3 steam outlet pipes

The component - hydrogen peroxide - is not the only component that can be gasified to obtain a working fluid for a turbine. Especially if you consider that high concentration hydrogen peroxide is not sufficiently stable during storage, it is advisable to use other components. Hydrazine and unsymmetrical dimethylhydrazine (UDMH) can be used as such, but for which, like hydrogen peroxide, special catalysts are required.

2.5.3. Turbopump unit (TNA),

The fuel pump largely determines the energy characteristics of the rocket engine. The degree of perfection of the main components of pumping pumps, turbines and pumps, in the process of creating modern models, is always under the close attention of engine developers. For CS and LGG designers, the issues of ensuring complete combustion of components, ensuring temperature resistance and strength of parts and assemblies determine the success of the subsequent operation of the created liquid-propellant rocket engine. For a specialist working on the creation of a pump pump, the main issues are: increasing the efficiency of turbines and pumps, the strength of their parts (turbine blades and disk, pump impellers, housings, shaft), the reliability of seals and a number of others that determine the reliability and perfection of pump pumps. Successful solution of the listed positions increases the specific impulse of thrust, reduces the specific mass of the pump and the engine. Upon further examination of the parameters and characteristics of the pump, it will be clear that the items listed above directly depend on such a parameter as rotor speed (system - “turbine, pumps, shaft”).

The initial data for the development of fuel pumps are the types of components, requirements for flow rates and pressures, service life and other data arising from the requirements for the rocket engine. Design studies allow us to draw a conclusion about the costs and parameters of the working fluid to create the required turbine power needed to drive the pumps. When performing this work, the following are determined: the fundamental layout of the pump, rotor speed, sealing systems and, ultimately, its mass characteristics.

When working on the creation of TNA, the developer takes into account the mandatory requirements that guide him:

· ensuring the basic parameters (dimensions, weight and parts of the TPU fastenings arising from the requirements for the engine layout) and characteristics for a given resource;

· ensuring the required flow rates and pressures of components installed for use in the engine;

· identifying positions that provide for the approximate cost of the sample being developed.

With further work on the creation of liquid propellant rocket engines, additional requirements may be established.

Among the main positions that determine the design and parameters of the TPU, the layout diagrams of the TPU should be considered. Possible circuit options are presented in the figure (see Fig. 2.47).

Rice. 2.47. TNA layout diagrams

a, b And V - single-rotor pumps, G. – multi-rotor pumps

Accepted designations: BUT - oxidizer pumps, NG – fuel pumps.

As follows from consideration of the figure, the options for layout schemes will differ whether a gearless scheme or a scheme with a gearbox is chosen for further development. With a gearless design, it is often not possible to select a single optimal speed for the turbine and each pump. However, a pump with a gearbox will always have worse mass characteristics. Modern liquid rocket engines of medium, large and very large ones, in which case the approximate mass of the fuel pump can be calculated using the following expression:

The figure (see Fig. 2.48) shows the structural diagrams of the pumps, with two-way pumps and one-way pumps. The diagrams show the nodes mentioned above.

Rice. 2.48. TNA block diagrams

1. Fuel pumps, 2. Turbines, 3. and 4. Internal seals of the pump and turbine, 5. Oxidizer pump, 6. Hydrodynamic seal, 7. Intermediate seal.

Liquid rocket engines of medium, high and very high thrust use gas turbines driven by centrifugal pumps. The layout options depend on the characteristics of the liquid rocket engine options, such as the type of components, the launch system of the pump, the characteristics of the product entering the turbine, and others. The design appearance of the TNA will also differ from private solutions determined by the designer at his own discretion. The figures (see Fig. 2.48 and 2.49) show types of TNA in which the supply of components is carried out by one-way and two-way inputs.

Rice. 2.42. TNA with pumps, with one-way component inputs

1. Exhaust manifold flange, 2. Turbine, 3. Inlet pipe with auger, 4. Inlet pipe of the fuel pump, 5. Spring, 6. Output flange of the outlet pipe of the fuel pump, 7. Housing of the oxidizer pump with auger, 8. Flange of the inlet pipe fuel pump.

In TNA, the pump casings are made with pre-pumps (screws) that provide increased inlet pressure in front of the main, one-sided impellers. This type of booster device eliminates the occurrence of cavitation during pump operation.

Rice. 2.50. TNA with pumps, with two-way component inputs

1. Flange of the fuel pump inlet pipe, 2. Inlet pipe of the oxidizer pump, 3. Pyrostarter, 4. Flange for supplying the working fluid to the turbine, 5. Turbine, 6. Turbine exhaust manifold.

The presented type of pump is made with a two-stage gas turbine and two centrifugal pumps. The pumps have two-way component inputs. The TNA design is designed with two shafts connected by a spring. A turbine and a centrifugal oxidizer pump are mounted on one shaft, with its two bearings and seals. On the second shaft, also with its own bearings and seals, there is a fuel pump. The performance of the bearings is maintained by grease, which is filled into the bearing cavities during the assembly of the pump. One and the second part of the rotor are installed in separate housings, connected by pins.

Centrifugal pumps are usually used in LRE pumps. Anti-cavitation properties are very important for TPU pumps, on which the erosive effect on the flow part of the pump depends, but also, what is especially important, the possibility of disrupting all parameters, the stability of which determines the performance of the required tasks of the entire LRE. An increase in the anti-cavitation properties of the pump is ensured by the use of special devices, some diagrams of which were previously presented in Figure 2.23. But most widely, in the practice of creating pumping pumps, screw centrifugal pumps are used.

For example, the figure (see Fig. 2.51) shows the design of an oxygen screw centrifugal pump.

Fig.2.51. Screw centrifugal pump.

1. Housing cover, 2. Bearing, 3. Pump impeller, 4. Pump housing. 5. Auger, 6. Bearing.

The efficiency of the pump depends on reducing losses, among which the main ones are:

· flow of the component from the high-pressure cavity (inlet from the impeller) into the inlet cavity;

· friction of the component against the walls of the internal cavities of the pump;

· friction in seals and bearings.

The listed pump efficiency losses are estimated:

Component density,

Volumetric flow rate of the component,

N – pressure developed by the pump,

N n - actual power consumed by the pump.

Typically, the efficiency of liquid rocket engine pumps ranges from 0.5...0.8,

In addition to the marked provisions, the figures (see Fig. 2.52.) show the designs of other booster devices - string prepumps (ejectors).

Fig.2.52. Design of the jet device (ejector).

A– an ejector with a row of holes. 1. Ejector body, 2. Component supply holes, equally spaced around the circumference, 3. Component supply pipe. b– ejector with a set of nozzles. 1. Component supply pipe, 2. Nozzles, 3. Ejector body.

Due to their low efficiency, jet pumps are advisable to use in engines with afterburning, since an increase in turbine power when supplying high-pressure active fluid to the ejector practically does not reduce the energy characteristics of the liquid-propellant rocket engine. On the image. 2.52, A shows the design of an ejector with twelve nozzles located around the circumference of the mixing chamber with an exit angle of 18°. When the ratio of active fluid flow to ejected fluid is up to 25 %, the pressure of the main flow increases significantly. However, the efficiency of such a device in optimal mode reaches no more than 0.15. The low pressure capacity of ejectors with an efficiency of 0.08 to 0.2 limits their use in modern liquid propellant pumps.

LIQUID ROCKET ENGINE (LPRE) - rocket engine running on liquid rocket fuel. The transformation of fuel into a jet gas stream that creates thrust occurs in camera. In modern LRE Both two-component rocket propellants are used, consisting of an oxidizer and fuel, which are stored in separate tanks, and single-component rocket propellants, which are liquids capable of catalytic decomposition. By type of oxidizing agent used LRE There are nitric acid, nitrogen tetroxide (oxidizing agent - nitrogen tetroxide), oxygen, hydrogen peroxide, fluorine, etc. Depending on the thrust value, they are distinguished LRE low, medium and high thrust. The conventional boundaries between them are 10 kN and 250 kN (the aircraft were installed LRE with thrust from tenths of N to 8 MN). LRE are also characterized specific thrust impulse, operating mode, dimensions, specific gravity, pressure in the combustion chamber, general structure and design of the main units. LRE is the main type space engines and is also widely used in high-altitude research rockets, long-range combat ballistic missiles, anti-aircraft guided missiles; limited - in combat missiles of other classes, on experimental aircraft, etc.

The main problems when creating LRE: rational choice of fuel that meets energy requirements and operating conditions; organization of the work process to achieve the calculated specific impulse; ensuring stable operation at specified modes, without developed low-frequency and high-frequency pressure fluctuations that cause destructive engine vibrations; cooling of a rocket engine exposed to aggressive combustion products at very high temperatures (up to 5000 K) and pressures up to many tens of MPa (this effect is aggravated in some cases by the presence of a condensed phase in the nozzle); supply of fuel (cryogenic, aggressive, etc.) at pressures reaching many tens of MPa for powerful engines and flow rates up to several t/s; ensuring a minimum mass of units and the engine as a whole, operating under very intense conditions; achieving high reliability.

LRE was proposed by K. E. Tsiolkovsky in 1903 as an engine for space flight. The scientist developed a schematic diagram LRE, indicated the most profitable rocket fuels, and investigated the design of the main units. Practical work on creating LRE were started in 1921 in the USA by R. Goddard. In 1922, he first recorded thrust while testing an experimental LRE, and in 1926 he launched a small liquid-propellant rocket. In the late 20s - early 30s. to development LRE started in Germany, the USSR and other countries. In 1931, the first Soviet LRE ORM and ORM-1, created by V. P. Glushko at the Gas Dynamics Laboratory. In 1933, the OR-2 propulsion system designed by F.A. Zander was tested, and engine 10, created by the Jet Propulsion Research Group, ensured the flight of a liquid-propellant rocket.

Before the start of the 2nd World War 1939-45. prototypes appeared in the USSR and the USA LRE with thrust up to several kN, intended for experimental aircraft. Intensive work in the field of rocket technology carried out in Germany during the war gave rise to a variety of types LRE for military purposes, many of which were mass-produced. The best were LRE designs by H. Walter (including HVK 109-509A (HWK 109-509A)) and H. Zborowski, LRE anti-aircraft guided missile "Wasserfall" (Wasserfall) and ballistic missile V-2 (V-2). Until the 2nd half of the 40s. the largest Soviet LRE were D-1-A-1100 and RD-1, developed by the Jet Research Institute. The first serial Soviet LRE The RD-1 and RD-1KhZ engines, created by the end of the war at the GDL-OKB, became the engines. There in 1947-53. the first powerful ones in the USSR were developed LRE: RD-100, RD-101, RD-103. During the same period, it was produced in the USA LRE with a thrust of ~ 350 kN for the Redstone ballistic missile.

Further development LRE and their current state was determined by the work that began in the mid-50s. in the USSR and the USA the development of ICBMs and launch vehicles. To implement them, it was necessary to create powerful, economical and compact LRE. The first among them were RD-107 and RD-108, with the advent of which the thrust LRE doubled, the propulsion thrust increased 10 times. Specific impulse LRE increased by almost 30%, the specific gravity decreased by more than 1.5 times. These results were made possible thanks to the development of a fundamentally new design LRE, which made it possible to switch from oxygen-ethyl alcohol fuel to oxygen-kerosene while simultaneously increasing the pressure in the combustion chamber by 2–2.5 times.

Since the beginning of the 60s. they also began to be used on launch vehicles (LVs) LRE operating on high-boiling fuels. The first of them was RD-214. The creation in the mid-60s was of great importance for the development of astronautics. oxygen-hydrogen LRE(intended for the upper stages of the launch vehicle), which in specific impulse exceed oxygen-kerosene ones by 30%. Because oxygen-hydrogen fuel, compared to oxygen-kerosene fuel, requires three times the volume for its placement at the same mass, and hydrogen tanks have to be equipped with thermal insulation, then the Tsiolkovsky number for oxygen-hydrogen fuel is 40% larger. This disadvantage is more than compensated for by the high efficiency of oxygen-hydrogen LRE. With equal launch mass of the launch vehicle, they are capable of delivering three times more payload into low-Earth orbit than oxygen-kerosene launch vehicles. LRE.

Mastering increasingly more efficient fuels, designers LRE strived simultaneously to convert the chemical energy of fuels into the kinetic energy of the jet stream with the greatest possible Efficiency. For this purpose, a scheme was developed LRE with afterburning of generator gas in the chamber. To implement this scheme, it was necessary to create cameras operating under conditions of high mechanical and thermal loads, as well as compact high-power power units. LRE with afterburning from the mid-60s. They are widely used on launch vehicles, in particular they are used on all stages of the Proton launch vehicle.

Along with powerful cosmic LRE numerous LRE medium and low thrust. The trouble-free operation of spacecraft (SC) engines is ensured to a large extent by the use of high-boiling single-component and self-igniting rocket fuels, the storage of which on board a spacecraft does not cause difficulties. remote control with LRE those using single-component fuel are simpler in design, but have a significantly lower specific impulse. By the mid 60s. in auxiliary LRE Hydrogen peroxide received the greatest use, which then began to be replaced by hydrazine and two-component fuels. The use of hydrazine made it possible to increase the specific impulse LRE on single-component fuel by approximately 40%.

Most Soviet space LRE created at the GDL-OKB V. P. Glushko, OKB A. M. Isaev and OKB S. A. Kosberg. Engines RD-107, RD-108, RD-214, RD-216, RD-253 and other GDL-OKB designs ensured the launch of all Soviet launch vehicles; the second stages of the launch vehicle series are also equipped with LRE GDL-OKB designs: RD-119, RD-219, etc. Kosberg OKB engines are installed on the upper stages of the Vostok, Voskhod (Soyuz) and Proton launch vehicles. Isaev Design Bureau engines are used mainly on artificial Earth satellites (AES), interplanetary spacecraft and spacecraft (KRD-61, KDU-414, TDU-1, KTDU-5A, etc.).

The largest foreign organizations engaged in development LRE, are located in the USA. The leading company is Rocketdyne, which created LRE Jay-2 (J-2), LR-79-NA (LR-79-NA), LR-89-NA (LR-89-NA), LR-105-NA (LR-105-NA), RS- 2701 (RS-2701), H-1 (H-1), F-1 (F-1), SSME (SSME), numerous LRE medium and low thrust on high-boiling two-component fuel. Most of the powerful ones mentioned LRE created under the leadership of S. Hoffman. Aerojet General Corporation has created a number of LRE on high-boiling two-component fuel, incl. LRE LR-87-ADzhey-5 (LR-87-AJ-5) and LR-91-ADzhey-5 (LR-91-AJ-5), series LRE medium thrust AJ-10 (AJ-10), including AJ-10-137 (AJ-10-137) and AJ-10-138 (AJ-10-138). Pratt & Whitney created the world's first oxygen-hydrogen LRE RL-10 (RL-10), Bell Aerospace Textron - numerous auxiliary LRE, and LRE medium thrust LR-81-BA-9 (LR-81-BA-9), company "TRV" - LRE medium thrust LMDE (Marquardt) - series LRE on high-boiling two-component fuel for spacecraft and interplanetary spacecraft. Several dozen types of hydrazine have been created in the USA. LRE(tested in flight LRE with thrust from 0.4 N to 2.7 kN). Among the developers LRE for interplanetary spacecraft - the company Reaction Motors, which also created a powerful LRE LR-99-RM-1 (LR-99-RM-1). The most famous of Western European LRE- AshM-7 (HM-7), “Valois”, “Vexen”, “Viking” (Viking, France), “Gamma-2” (Gamma), “Gamma-8”, RZet- 2 (RZ-2, UK). IN Western Europe are also being developed LRE low thrust on two- and one-component fuels for artificial satellites. Japan produces American products under license LRE LR-79-NA for its own version of the Delta launch vehicle (Delta). For one of the stages of this launch vehicle, the Mitsubishi company has developed a high-boiling fuel rocket engine with a thrust of 53 kN with displacement feed. Oxygen-hydrogen gases were tested on the stands LRE thrust up to 0.1 MN with pump supply. Chinese launch vehicles are used LRE thrust 0.7 MN with pump supply of high-boiling fuel.

Space LRE varied in design and characteristics. The greatest difference exists between powerful LRE, providing acceleration of the launch vehicle, and LRE rocket control systems for spacecraft. The first ones operate on two-component fuel. The pull of these LRE reaches 8 MN (with a total thrust of up to 40 MN), dimensions are several meters, and weight is several tons. They are usually designed for one-time activation (except for some LRE upper stages of the launch vehicle) and work for 2-10 minutes when changing parameters within narrow limits. To these LRE there is a requirement to provide a high specific impulse with small dimensions and weight. Therefore, they use pumping fuel supply into the chamber (the exception is LRE"Vexen" and "Valois"). To this end, in LRE provided turbopump unit(TNA) and gas generator(GG). The fuel pump contains high-pressure fuel pumps (usually axial-centrifugal) and a turbine driving them, which is rotated by the gas produced in the gas generator. IN LRE Without afterburning, the generator gas exhausted in the turbine is discharged into the exhaust pipe, steering nozzle or chamber nozzle. IN LRE with afterburning, this gas enters the afterburning chamber with the rest of the fuel.

IN LRE without afterburning, 2-3% of the total fuel can be consumed through the GG, and the expedient pressure limit in the combustion chamber is limited to ~ 10 MPa, which is associated with losses of specific impulse to the TPU drive: for LRE in general, this parameter is lower than for the camera, because the additional thrust created by the outflow of exhaust gas from the generator is small. The reason for this is the low pressure and temperature of this gas. For LRE RD-216 they are, for example, 0.12 MPa and 870 K, respectively; in this case, specific impulse losses reach 1.5% (over 40 m/s). With increasing pressure in the combustion chamber, an increase in its specific impulse is observed, but for this it is necessary to increase the flow of generator gas (to provide the required power of the fuel pumps). From a certain moment, the ever-increasing losses of specific impulse to the TNA drive balance out and then exceed the increase in the specific impulse of the chamber. IN LRE with afterburning through the GG, a portion of the total fuel is consumed (20-80%), however, the TPU drive is carried out without deterioration in efficiency LRE(chamber specific impulse values ​​and LRE match up). In the combustion chambers of these LRE it is possible to realize a pressure of 15-25 MPa (the pressure in the GG is approximately twice as high). For the powerful LRE with pump fuel supply, the specific impulse reaches 3430 m/s when using oxygen-kerosene fuel and 4500 m/s when using oxygen-hydrogen fuel; specific gravity LRE can be only 0.75-0.85 g/N.

In addition to the camera, TNA and GG, powerful LRE contain fuel pipelines with bellows hoses and angular and linear movement compensators, facilitating assembly and installation LRE, as well as providing relief from thermal stresses and allowing the camera to be deflected in order to control the movement of the launch vehicle; pipelines for generator gas and fuel drainage; devices and systems rocket engine launch; automation units with electric drives, pneumatic, pyro- and hydraulic systems and devices for operating control LRE(including for him throttling); emergency protection system units; telemetry measurement system sensors; electrical cable trunks for supplying signals to automation units and receiving signals from telemetry sensors; thermal insulating covers and screens that ensure the proper temperature in the engine compartment and prevent overheating or hypothermia individual elements; system elements tank pressurization(heat exchangers, mixers, etc.); gimbal or mounting frame LRE to the launch vehicle (the frame that receives the thrust is also the element on which the engine is assembled); often - steering chambers and nozzles with systems that ensure their operation; elements of the general assembly (brackets, fasteners, seals). According to the device they distinguish block liquid rocket engines, single- and multi-chamber (with power supply to several chambers from one TNA).

LRE jet control systems belong to low-thrust engines, their weight usually does not reach 10 kg, and their height is 0.5 m; mass of many LRE does not exceed 0.5 kg, and they fit in the palm of your hand. A characteristic feature of these LRE is operation in pulse mode (over several years of operation of the spacecraft, the total number of starts LRE can reach several hundred thousand, and the operating time can reach several hours). These LRE are single-walled chambers equipped with fuel start-up and shut-off valves and are designed for displacement feed high-boiling fuel (two-component self-igniting or one-component). Pressure in the combustion chambers indicated LRE, determined mainly by the boost pressure of the remote control tanks and the hydraulic resistance of the supply lines, is in the range of 0.7-2.3 MPa. In the case when the gas for pressurizing fuel tanks is located in the tanks themselves, its pressure decreases as fuel is consumed, which leads to deterioration in performance LRE. Relatively high specific impulse LRE(up to 3050 m/s for two-component fuel and up to 2350 m/s for hydrazine) is achieved due to the relatively large size of the jet nozzle, which ensures the expansion of combustion products to very low pressure. Despite the small absolute mass LRE reactive control systems, their specific mass is high (with a decrease in thrust from 500 to 1 N, it increases from approximately 5 to 150 g/N).

LRE spacecraft occupy an intermediate position in their characteristics between powerful LRE launch vehicles and LRE reactive control systems. Their thrust covers the range from hundreds of N to tens of kN and can be either unregulated or adjustable; they can operate continuously for tenths of seconds and several thousand seconds with a number of starts from 1 to several tens. In the specified LRE the same types of fuels are used as in LRE jet control systems (single-component fuel is used only in LRE low thrust).

Plans for further space exploration LRE plays a big role. Powerful LRE, designed to cost-effectively use efficient fuels, continue to be the focus. By 1981, an oxygen-hydrogen generator was created LRE with a thrust of over 2 MN, designed to accelerate an aircraft from launch to insertion into low-Earth orbit. Thanks to advances in the field of cryogenic technology and thermal insulation materials, it becomes feasible to create LRE on low-boiling fuels developing a high specific impulse, for use in spacecraft operating in space. Progress in development LRE with thrust up to several tens of kN, operating on fuels containing fluorine and its derivatives (see, for example, RD-301), makes the use of fluorine fuels realistic LRE in the upper stages of the launch vehicle and in automatic spacecraft that will fly to the planets. During bench tests in 1977, an experimental oxygen-hydrogen LRE(thrust 0.1 MN), developed for these purposes, a specific impulse of 4690 m/s was achieved. Experimental studies are being carried out on various problems of creating LRE on metal-containing fuel.

Along with development for LRE new fuels, a search is underway for technical principles that provide a further increase in efficiency and a reduction in size and weight LRE. The improvement in parameters achieved by increasing the pressure in the chamber becomes less noticeable with increasing pressure, and the difficulties of creating LRE are increasing more and more. Increasing this parameter above 25-30 MPa is ineffective and difficult to implement. Shows interest in LRE, equipped nozzles with a central body. In order to reduce the cost of launching payloads, LRE(for reusable spacecraft), designed for several dozen flights and a service life of several hours with a small amount of inter-flight maintenance work.

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