Liquid propellant rocket engine on cryogenic fuel. Pumpless cryogenic liquid rocket engine (options) Cryogenic engine

A liquid propellant rocket engine is an engine that is fueled by liquefied gases and chemical liquids. Depending on the number of components, liquid-propellant rocket engines are divided into one-, two- and three-component ones.

Brief history of development

For the first time, the use of liquefied hydrogen and oxygen as fuel for rockets was proposed by K.E. Tsiolkovsky in 1903. The first prototype of the rocket engine was created by the American Robert Howard in 1926. Subsequently, similar developments were carried out in the USSR, USA, Germany. The greatest successes were achieved by German scientists: Thiel, Walter, von Braun. During World War II, they created a whole line of rocket engines for military purposes. There is an opinion that if they had created the V-2 Reich earlier, they would have won the war. Subsequently, the Cold War and the arms race became the catalyst for accelerating the development of liquid propellant rocket engines with a view to applying them to the space program. With the help of RD-108, the first artificial Earth satellites were put into orbit.

Today, LRE is used in space programs and heavy rocket weapons.

Scope of application

As mentioned above, LRE is used mainly as an engine spacecraft and launch vehicles. The main advantages of LRE are:

• the highest specific impulse in the class;
• the ability to perform a full stop and restart paired with traction control gives increased maneuverability;
• significantly less weight of the fuel compartment in comparison with solid fuel engines.

Among the disadvantages of LRE:

• more complex device and high cost;
• increased requirements for safe transportation;
• in a state of weightlessness, it is necessary to use additional engines to deposit fuel.

However, the main disadvantage of liquid-propellant rocket engines is the limit of the energy capabilities of the fuel, which limits space exploration with their help to the distance of Venus and Mars.

Device and principle of operation

The principle of operation of the LRE is the same, but it is achieved using different device schemes. Fuel and oxidizer are pumped from different tanks to the nozzle head, injected into the combustion chamber and mixed. After ignition under pressure, the internal energy of the fuel is converted into kinetic energy and flows out through the nozzle, creating jet thrust.

The fuel system consists of fuel tanks, pipelines and pumps with a turbine for pumping fuel from the tank into the pipeline and a control valve.

Pumping fuel supply creates a high pressure in the chamber and, as a result, a greater expansion of the working fluid, due to which the maximum value of the specific impulse is achieved.

Injector head - a block of injectors for injecting fuel components into the combustion chamber. The main requirement for the nozzle is high-quality mixing and the speed of fuel supply to the combustion chamber.

Cooling system

Although the proportion of heat transfer from the structure during the combustion process is insignificant, the problem of cooling is relevant due to the high combustion temperature (>3000 K) and threatens with thermal destruction of the engine. There are several types of chamber wall cooling:

Regenerative cooling is based on creating a cavity in the chamber walls through which fuel passes without an oxidizer, cooling the chamber wall, and the heat, together with the coolant (fuel), returns to the chamber.

The near-wall layer is a layer of gas created from combustible vapors near the walls of the chamber. This effect is achieved by installing injectors on the periphery of the head that supply only fuel. Thus, the combustible mixture lacks an oxidizing agent, and combustion near the wall is not as intense as in the center of the chamber. The temperature of the near-wall layer isolates the high temperatures in the center of the chamber from the walls of the combustion chamber.

The ablative method of cooling a liquid-propellant rocket engine is carried out by applying a special heat-shielding coating to the walls of the chamber and nozzles. The coating at high temperatures changes from a solid to a gaseous state, absorbing a large proportion of heat. This method of cooling a liquid rocket engine was used in the Apollo lunar program.

The launch of a rocket engine is a very responsible operation in terms of explosiveness in case of failures in its implementation. There are self-igniting components with which there are no difficulties, however, when using an external initiator for ignition, ideal coordination of its supply with the fuel components is necessary. The accumulation of unburned fuel in the chamber has a destructive explosive force and promises dire consequences.

The launch of large liquid rocket engines takes place in several stages, followed by reaching maximum power, while small engines are launched with an immediate output of one hundred percent power.

The automatic control system of liquid-propellant rocket engines is characterized by the implementation of a safe engine start and exit to the main mode, control of stable operation, thrust adjustment according to the flight plan, adjustment of consumables, shutdown when reaching a given trajectory. Due to the moments that cannot be calculated, the liquid-propellant rocket engine is equipped with a guaranteed supply of fuel so that the rocket can enter the desired orbit in case of deviations in the program.

The propellant components and their choice during the design process are decisive in the design of a liquid rocket engine. On this basis, the conditions of storage, transportation and production technology are determined. The most important indicator of the combination of components is the specific impulse, on which the distribution of the percentage of the mass of fuel and cargo depends. The dimensions and mass of the rocket are calculated using the Tsiolkovsky formula. In addition to specific impulse, density affects the size of tanks with fuel components, boiling point can limit the operating conditions of missiles, chemical aggressiveness is characteristic of all oxidizers and, if the rules for operating tanks are not followed, can cause a tank fire, the toxicity of some fuel compounds can cause serious harm to the atmosphere and the environment . Therefore, although fluorine is a better oxidizing agent than oxygen, it is not used due to its toxicity.

Single-component liquid-propellant rocket engines use liquid as fuel, which, interacting with the catalyst, decomposes with the release of hot gas. The main advantage of single-component rocket engines is their simplicity of design, and although the specific impulse of such engines is small, they are ideally suited as low-thrust engines for orientation and stabilization of spacecraft. These engines use a displacement fuel supply system and, due to the low process temperature, do not need a cooling system. Single-component engines also include gas-jet engines, which are used in conditions where thermal and chemical emissions are unacceptable.

In the early 1970s, the United States and the USSR were developing three-component liquid-propellant rocket engines that would use hydrogen and hydrocarbon fuels as fuel. This way the engine would run on kerosene and oxygen at startup and switch to liquid hydrogen and oxygen at high altitude. An example of a three-component rocket engine in Russia is the RD-701.

Rocket control was first used in V-2 rockets using graphite gas-dynamic rudders, but this reduced engine thrust, and modern rockets use rotary chambers attached to the body with hinges that create maneuverability in one or two planes. In addition to rotary cameras, control motors are also used, which are fixed with nozzles in the opposite direction and are turned on if it is necessary to control the apparatus in space.

A closed-cycle liquid-propellant rocket engine is an engine, one of the components of which is gasified by burning at a low temperature with a small part another component, the resulting gas acts as the working fluid of the turbine, and then is fed into the combustion chamber, where it burns with the remnants of the fuel components and creates jet thrust. The main disadvantage of this scheme is the complexity of the design, but the specific impulse increases.

The prospect of increasing the power of liquid rocket engines

In the Russian school of LRE creators, headed by Academician Glushko for a long time, they strive for the maximum use of fuel energy and, as a result, the maximum possible specific impulse. Since the maximum specific impulse can be obtained only by increasing the expansion of the combustion products in the nozzle, all developments are carried out in search of the ideal fuel mixture.

Dearman in partnership with scientists, leaders industrial enterprises and specialists in cryogenic equipment specializes in the development of technologies using liquefied gases. The crowning achievement of this research is the Dearman engine, a state-of-the-art reciprocating engine powered by the expansion of liquid nitrogen or liquid air to produce environmentally friendly cold and mechanical power.

When nitrogen passes from a liquid to a gaseous state of aggregation, this gas expands 710 times. This increase in volume is used to drive the engine pistons. Dearman engines work like steam engines high pressure, but at a low boiling point of liquid nitrogen. This means that both waste heat and ambient temperature can be used as a source of thermal energy, eliminating the need for conventional fuels.

A unique feature of Dearman engines is the use of a mixture of water and glycol as the coolant. When this coolant is mixed with extremely cold nitrogen, this liquid expands quasi-isothermally, which greatly improves engine efficiency.

It is important to note that the Dearman engine only emits air or nitrogen, no oxides of nitrogen (NOx), carbon dioxide(CO2) or particulate matter.

Dearman technology has many advantages over other low carbon technologies:

• Low capital expenditures and bonded carbon – Dearman engines are made from conventional materials, using technologies common in the engine industry.
• Fast filling - liquid gas can be transferred between tanks at high speeds. The modern gas industry uses systems capable of distilling over 100 liters of liquid gas per minute.
• Large amounts of existing infrastructure – the gas industry is global. There is now a liquid nitrogen production facility sufficiently developed to power thousands of Dearman engines.
• The efficiency of the "fuel" production process - liquefaction of air is a long-established process that requires only air and electricity.

• Air liquefaction production facilities can be used very flexibly – for example in non-working hours or during an incomplete download. Renewable energy sources can be used to further reduce costs.

How it works

The Dearman engine works like this:
1. the coolant is pumped into the engine cylinders, filling almost their entire volume;

2. then cryogenic nitrogen is introduced into the cylinder, which comes into contact with the heat exchange fluid and begins to expand;

3. heat from the coolant is absorbed by the expanding gas, resulting in an almost isothermal expansion;

4. The piston moves down, the exhaust valve opens, and the mixture of gas and liquid coolant exits the engine;

5. The coolant is recovered, heated and reused while nitrogen or air is released into the atmosphere.

A MOST IMPORTANT event in the history of the Indian space program took place: the first tests of a cryogenic engine, entirely of Indian production, designed to launch satellites into geosynchronous orbit, took place.

Tests conducted at a specialized center in Mahendragiri (South Indian state of Tamil Nadu) recorded stable operation of the engine for 15 seconds, the Indian Space Exploration Organization (IOIC) said. Thus, according to the IOIC statement, the first stage in the implementation of the entire task has been successfully overcome: the cryogenic engine was created by the Indians independently and tested.

Let me remind you the history of the issue. In July 1993, Moscow made another serious concession to Washington, assuring it that it was refusing to transfer to India the technology for creating cryogenic space engines needed to launch rockets into high and very high near-Earth orbits. At the same time, Yeltsin's Russia not only "cave in" before its overseas partner, but also lost economic benefits and damaged its international prestige. Moreover, according to the 1991 agreements, Russia not only undertook to supply Delhi with two cryogenic engines, but also to transfer to the Indians the technology for their manufacture. This was confirmed during President Yeltsin's visit to India in January 1993, and then, six months later, Moscow's word was broken. The Russian-Indian "deal of the century" was thwarted.

It is clear that cooperation in the field of creating cryogenic (based on liquid hydrogen) space engines (environmentally friendly and unsuitable for the needs of the military, since filling rockets with liquid hydrogen must be carried out extremely carefully, slowly and for a long time) was beneficial to both India and Russia, but extremely annoyed Washington. It was no secret to anyone that the United States really wanted to take Russia's place in that deal, but it didn't work out. Then, taking advantage of some favorable features of Russia's domestic political life, the United States literally forced the Kremlin to abandon most of the lucrative contract, which naturally caused great indignation in India.

The then head of the Indian space program W.R. Rao stressed that "a serious blow has been dealt, now we must rely on our own strength." And the Prime Minister of India of those years, P.V. Narasimha Rao, in one of his private conversations, remarked: "If so, then India itself will create cryogenic engines, it will not take even a few years for our scientists to achieve this." The prediction came true.

Seven years have passed, and India has created and tested a cryogenic engine, with the help of which Indian rockets will launch satellites into space from the Sriharikota cosmodrome.

The successful test is noteworthy not only from a technical but also a political point of view. First, it is important for Delhi's prestige on the eve of negotiations with US President Clinton, who is going to visit India in March. Secondly, now it will be possible to perform an act of mercy in relation to the An-26 pilots of the Latvian airline sentenced to life imprisonment, most of who took Russian citizenship while in a Calcutta prison. This was requested by the Indian authorities, who is on an official visit to Delhi, Chairman State Duma Gennady Seleznev. A possible pardon is believed in local journalistic circles if it happens just before the long-anticipated visit. Russian President to India.

We all know that one of the foundations of material life modern humanity are the well-known minerals oil and gas. Blessed hydrocarbons are present in one way or another in any area of ​​our lives, and the first thing that comes to mind for any person is fuel. These are gasoline, kerosene and natural gas used in various energy systems (including vehicle engines).

How many cars on the roads of the world and airplanes in the air burn gasoline and kerosene in their engines ... Their number is huge and the volume of fuel flying out, so to speak, into the pipe (and at the same time still striving to contribute its considerable share to the poisoning of the atmosphere) is just as huge: -)). However, this process is not endless. Oil, from which the lion's share of the world's fuel is produced (despite the fact that it is gradually losing ground to natural gas), is rapidly declining. It constantly rises in price and its shortage is felt more and more.

This situation has long been forcing researchers and scientists around the world to look for alternative sources of fuel, including for aviation. One of the directions of such activity was the development aircraft using cryogenic fuel.

Cryogenic means "born cold" and the fuel in this case is liquefied gas, which is stored at very low temperatures. The first gas that attracted the attention of developers in this regard was hydrogen. This gas is three times more calorific than kerosene and, in addition, when it is used in an engine, water and a very small amount of nitrogen oxides are released into the atmosphere. That is, it is harmless to the atmosphere.
cryogenic fuel

Aircraft TU-154B-2

In the mid-80s of the last century, the design bureau of A.N. Tupolev began to create an aircraft using liquid hydrogen as fuel. It was developed on the basis of the serial TU-154B using the NK-88 turbojet bypass engine. This engine was created in the engine-building design bureau. Kuznetsov (Samara), again based on the serial engine for the Tu-154 NK-8-2 and was intended to run on hydrogen or natural gas. It must be said that this bureau has been working on new topics since 1968.
cryogenic fuel

The same Tu-155 aircraft is in storage ... Unfortunately, disgusting storage :-(

The new aircraft powered by cryogenic fuel was named TU-155. However, everything is not so simple. The fact is that hydrogen is a dangerous fuel. It is extremely flammable and explosive. It has an exceptional penetrating ability, and can only be stored and transported in a liquefied state at very low temperatures close to absolute zero (-273 degrees Celsius). These features of hydrogen present a fairly large problem.

Therefore, the TU-155 was a flying laboratory for researching and solving existing problems, and the base aircraft underwent a radical alteration during its creation. Instead of the right engine NK-8-2, a new cryogenic NK-88 was installed (the other two remained native :-)). In the rear part of the fuselage, in place of the passenger compartment, a special tank for cryogenic fuel, liquid hydrogen, with a volume of 20 cubic meters was placed. with reinforced screen-vacuum insulation, where hydrogen could be stored at temperatures below minus 253 degrees Celsius. It was supplied to the engines by a special turbopump unit, like on a rocket.
cryogenic fuel

Engine NK-88. A massive turbopump unit is visible on top of the engine

Due to the high explosiveness, it was necessary to remove almost all electrical equipment from the compartment with the fuel tank in order to exclude the slightest possibility of sparking, and the entire compartment was constantly purged with nitrogen or air. To control the units of the power plant, a special helium control system was created. In addition, hydrogen vapor from the tank had to be vented away from the engines to avoid ignition. To do this, they made a drainage system. On the plane, its branches are clearly visible in the tail section of the fuselage (especially on the keel).
cryogenic fuel

The layout scheme of the TU-155. Blue - fuel tank. In the front compartment - supporting equipment. Red - cryogenic engine

In general, more than 30 new aircraft systems were created and implemented. In general, the work was carried out grandiose :-) . But we also needed ground equipment, no less complex, providing refueling and storage equipment. True then full swing The Buran system was being developed, on the launch vehicle of which one of the fuel components was liquid hydrogen. Therefore, it was believed that everything would be put on an industrial basis and there would be no shortage of fuel. But, I think, everyone understands that cryogenic fuel in such a system becomes simply “golden” in value. And this means that the commercial use of liquid hydrogen in the near future is hardly possible. Therefore, even then preparations were underway for the transition to another type of cryogenic fuel - liquefied natural gas (LNG).

Nevertheless, the first flight of the TU-155 on liquid hydrogen took place on April 15, 1988. In addition to this, there were 4 such flights. After that, the TU-155 was modified for flights using liquefied natural gas (LNG).

This type of fuel is much cheaper and more accessible than hydrogen, and it is also several times cheaper than kerosene. Its calorific value is 15% higher than that of kerosene. In addition, it also clogs the atmosphere little, and it can be stored at a temperature of minus 160 degrees, which is as much as 100 degrees higher than that of hydrogen. In addition, against the background of hydrogen, LNG is still less flammable (although, of course, such a danger still exists) and there is sufficient experience in maintaining it in a safe state. The organization of gas supply (LNG) to airfields, in general, is also not extremely difficult. Almost every major airport has gas pipelines. In general, there are enough advantages :-).

The first flights of the TU-155, which already uses liquefied natural gas as a cryogenic fuel, took place in January 1989. (The video below shows this.) There were also about 90 such flights. All of them showed that fuel consumption is reduced by almost 15% compared to kerosene, that is, the aircraft becomes more economical and profitable.

Now a little about the prospects... In the late 90s, the main manager of Russian gas reserves, Gazprom, took the initiative to build at the beginning a cargo-passenger, and then just a passenger aircraft that could run entirely on LNG. The aircraft received the name TU-156 and was created on the basis of the existing TU-155. Three new NK-89 engines were to be installed on it. These are turbofan engines similar to the NK-88, but with two independent fuel systems: one for kerosene and the other for cryogenic fuel (LNG). This was convenient in the sense that not everywhere there was the possibility of refueling with gas, and the aircraft could, as necessary, switch from one power supply system to another. This, according to the developed technology, took only five minutes. NK-89 also had a heat exchanger in the after-turbine space, where the liquefied gas passed into a gaseous state and then entered the combustion chamber.

A lot of research and calculation work was carried out on the rearrangement of the compartments and the location of the fuel tanks. By the year 2000, three TU-156s were to be produced at the Samara Aviation Plant and their certification and trial operation began. But... Unfortunately, this was not done. And the obstacles to the implementation of the planned plans were exclusively financial.

After that, several more projects of aircraft using cryogenic fuel (CNG) were developed, such as, for example, the TU-136 with turboprop engines running on both kerosene and liquefied gas and the wide-body TU-206 with turbojet engines running on LNG . However, on this moment All these projects have so far remained projects.
cryogenic fuel

Tu-136 aircraft model

cryogenic fuel

Aircraft model TU-206 (TU-204K)

Time will tell how things will turn out in this area of ​​aviation science and technology. So far, the creation of aircraft using cryogenic fuel is hampered by various circumstances, both objective and subjective. Much remains to be done in the development of special aircraft systems, the development of ground infrastructure, fuel transportation and storage systems. But this topic is extremely promising (and, in my opinion, very interesting :-)). Hydrogen, with its enormous energy intensity and virtually inexhaustible reserves, is the fuel of the future. This can be said with complete confidence. The transitional stage to this is the use of natural gas.

And this decisive step into the future was made in Russia. I feel proud once again talking about it :-) . Nowhere in the world has there been and to this day there are no aircraft like our TU-155. I would like to quote the famous American aviation engineer Karl Brever: "The Russians have accomplished in aviation a deed commensurate with the flight of the first Earth satellite!"

This is the real truth! I really just want these things to go in a stream (and the Russians can do it :-)), and for this stream to be continuous, and not move in jerks, as is often the case with us ...

And the rivalry between the USSR and the USA for leadership in space exploration was a powerful stimulus for the development of liquid-propellant rocket engines.

In 1957, in the USSR, under the leadership of S.P. Korolev, the R-7 ICBM was created, equipped with the RD-107 and RD-108 rocket engines, 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 spacecraft of the Apollo series was launched in the USA, launched on a flight path to the Moon by a Saturn-5 launch vehicle, the first stage of which was equipped with 5 F-1 engines. F-1 is currently the most powerful among single-chamber rocket engines, yielding in thrust to the four-chamber engine RD-170, developed by Energomash Design Bureau in the Soviet Union in 1976.

Currently space programs of all countries are based on the use of LRE.

Scope of use, advantages and disadvantages

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

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

On fig. 1 schematically shows the LRE device.

Fuel system

The LRE fuel system 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, a nozzle head, and valves that regulate the fuel supply.

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

In Fig. 1, the exhaust gases from the HP turbine enter through the nozzle head into the combustion chamber along with the fuel components (11). Such an engine is called an engine closed loop(otherwise - with a closed cycle), in which the entire fuel consumption, including that used in the TNA drive, passes through the LRE combustion chamber. The pressure at the outlet of the turbine in such an engine, obviously, should be higher than in the combustion chamber of the rocket engine, and at the inlet to the gas generator (6) that feeds the turbine, it should be even higher. To meet these requirements, the same fuel components are used to drive the turbine (under high pressure), on which the LRE itself operates (with a different ratio of components, as a rule, with an excess of fuel in order 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 in environment through the outlet pipe. The implementation of an open cycle is technically simpler, since the operation of the turbine is not related to the operation of the LRE chamber, and in this case, the HP can generally have its own independent fuel system, which simplifies the procedure for starting the entire propulsion system. But closed loop systems have several best values specific impulse, and this forces designers to overcome the technical difficulties of their implementation, especially for large launch vehicle engines, which are subject to particularly high requirements for this indicator.

In the diagram in fig. 1 one HP pumps both components, which is acceptable in cases where the components have comparable densities. For most liquids used as propellant components, the density ranges from 1 ± 0.5 g/cm³, which allows one turbo drive to be used 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 rotational speed. Therefore, in the case of using hydrogen as a fuel, an independent THA is provided for each component.

With a small engine thrust (and, consequently, low fuel consumption), the turbopump unit becomes too “heavy” an element that worsens the weight characteristics of the propulsion system. An alternative to a pumped fuel system is displacement, at which the flow of fuel into the combustion chamber is provided 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 manufacture. Helium is used to pressurize tanks with liquid hydrogen, since other gases condense and turn into liquids at the temperature of liquid hydrogen.

When considering the operation of an engine with a displacement fuel supply system from the diagram in fig. 1, the THA is excluded, and the fuel components come from the tanks directly to the main LRE valves (9) and (10). The pressure in the fuel tanks during displacement supply must be higher than in the combustion chamber, the tanks are stronger (and heavier) than in the case of a pumped 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 small thrust (within 10 tons). The advantages of the displacement system are the simplicity of design and the speed of the engine's reaction to the start command, especially in the case of using self-igniting fuel components. Such engines are used to perform spacecraft maneuvers 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 node in which they are mounted nozzles designed to inject 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 must burn as completely as possible, since unburned fuel is a loss of thrust and specific impulse. The solution to this problem is achieved by a number of measures:

• The maximum increase in the number of nozzles in the head, with a proportional minimization of the flow rate through one nozzle. (There are 2,600 oxygen nozzles and 3,700 kerosene nozzles in the nozzle head of the engine).
• The special geometry of the injectors in the head and the alternation of the fuel and oxidizer injectors.
• The special shape of the nozzle channel, due to which, when moving through the channel, the liquid is rotated, 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 LRE combustion chamber, only an insignificant part (fractions of a percent) of all the 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 the thermal destruction of the engine, so the problem of LRE cooling is very relevant.

For LRE with pumped fuel supply, two methods of cooling the walls of the LRE chamber are mainly used: regenerative cooling and wall layer, which are often used together. For small engines with a positive displacement fuel system, it is often used 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, in one way or another, a cavity is created (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 LRE chamber of the V-2 rocket, for example, consisted of two steel shells, inner and outer, repeating the shape of each other. A cooling component (ethanol) passed through the gap between these shells. Due to technological deviations in the thickness of the gap, uneven fluid flow occurred, as a result, local overheating zones of the inner shell were created, which often “burned out” in these zones, with catastrophic consequences.

In modern engines, the inner part of the chamber wall is made of highly heat-conducting bronze alloys. Narrow thin-walled channels are created in it by milling (15D520 RN 11K77 Zenith, RN 11K25 Energia), or acid etching (SSME Space Shuttle). From the outside, this structure is tightly wrapped around a load-bearing steel or titanium sheet shell, which perceives the power 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 soldered with a bronze alloy for tightness, but such chambers are designed for lower pressure.

Wall layer(boundary layer, the Americans also use the term “curtain” - 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 injectors are installed along the periphery of the mixing head. Due to the excess of fuel and the lack of an oxidizer, the chemical reaction of combustion in the near-wall layer occurs much less intensively than in the central zone of the chamber. As a result, the temperature of the near-wall layer is much lower than the temperature in the central zone of the chamber, and it isolates 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, which bring part of the fuel into the chamber directly from the cooling jacket, also in order to create a near-wall layer.

LRE launch

The launch of an LRE 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 igniter is needed, the action of which must be precisely coordinated with the supply of fuel components to the combustion chamber. The unburned fuel mixture is an explosive of great destructive power, and its accumulation in the chamber threatens a severe accident.

After the ignition of the fuel, the maintenance of a continuous process of its combustion occurs by itself: the fuel re-entering the combustion chamber ignites due to the high temperature created during the combustion of previously introduced portions.

For the initial ignition of the fuel in the combustion chamber during the launch of the LRE, different methods are used:

• The use of self-igniting components (as a rule, based on phosphorus-containing starting fuels, self-igniting when interacting with oxygen), which are introduced into the chamber at the very beginning of the engine start process 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 its repeated restart.
• An electric igniter placed in the combustion chamber near the mixing head, which, when switched on, creates an electric arc or a series of high voltage spark discharges. This igniter is disposable. After the fuel is ignited, it burns.
• Pyrotechnic igniter. Near the mixing head in the chamber is placed a small incendiary pyrotechnic checker, which is ignited by an electric fuse.

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

The launch of large LRE with a pumped fuel system consists of several stages: first, the HP is launched and gains momentum (this process can also consist of several phases), then the main valves of the LRE are turned on, as a rule, in two or more stages with a gradual increase in thrust from stage to stage. steps to normal.

For relatively small engines, it is practiced to start with the output of 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 start of the engine and bringing it to the main mode.
• Maintaining stable operation.
• Thrust change in accordance with the flight program or at the command of external control systems.
• Shutdown of the engine when the rocket reaches a given orbit (trajectory).
• Regulation of the ratio of the consumption of components.

Due to technological dispersion hydraulic resistance fuel and oxidizer paths, the ratio of component costs 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 not fulfill its task, having completely consumed one of the fuel components. At the dawn of rocket science, this was fought 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 the actual flight conditions from the calculated ones). The guaranteed fuel supply is created at the expense of the payload. At present, 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 executive bodies it is the main valves of the rocket 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 the most important decisions in the design of a rocket engine, which predetermines many details of the engine design and subsequent technical solutions. Therefore, the choice of fuel for LRE is carried out with a comprehensive consideration of the purpose of the engine and the rocket on which it is installed, the conditions for their operation, the technology of production, storage, transportation to the launch site, etc.

One of the most important indicators characterizing the combination of components is specific impulse, which has especially importance when designing launch vehicles for spacecraft, since the ratio of the mass of fuel and payload, and, consequently, the dimensions and mass of the entire rocket (see Tsiolkovsky's formula), which, if the specific impulse is not high enough, 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 momentum, 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
dinitrogen 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

One-component are also jet engines running on compressed cold gas (for example, air or nitrogen). 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 the simplicity of the design. These requirements must be met, for example, by individual cosmonaut movement and maneuvering devices (UPMK) located in a knapsack behind the back and intended for movement during work outside the spacecraft. UPMK operate from two cylinders with compressed nitrogen, which is supplied through solenoid valves to the propulsion system, consisting of 16 engines.

Three-component rocket engines

Since the beginning of the 1970s, the concept of three-component engines has been studied in the USSR and the USA, which would combine a high specific impulse when used as a combustible hydrogen, and a higher average fuel density (and, consequently, a smaller volume and weight of fuel tanks), characteristic of hydrocarbon fuels. At start-up, such an engine would run on oxygen and kerosene, and at high altitudes it would switch to using liquid oxygen and hydrogen. Such an approach may make it possible to create a single-stage space carrier. Russian example The three-component engine is the RD-701 liquid-propellant rocket engine, 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.

Missile control

In liquid-propellant rockets, engines often, in addition to their main function - creating thrust, also play the role of flight controls. Already the first V-2 guided ballistic missile was controlled using 4 graphite gas-dynamic rudders placed in the jet stream of the engine along the periphery of the nozzle. Deviating, 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, besides, graphite rudders in a jet stream are subject to severe erosion and have a very short time resource.
Modern missile control systems use PTZ cameras LRE, which are attached to the bearing elements of the rocket body with the help of hinges that allow you to rotate the camera in one or two planes. The fuel components are brought to the chamber with the help of 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 moment of force. The cameras are rotated by hydraulic or pneumatic steering machines, which execute commands generated by the rocket control system.
In the domestic space carrier Soyuz (see photo in the title of the article), in addition to 20 main, fixed cameras of the propulsion system, there are 12 rotary (each in its own plane), smaller control cameras. Steering chambers have a common fuel system with the main engines.
Of the 11 sustainer 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 the camera rotation is no more than ±5 °: due to the large 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 camera deviation creates a significant control