Rocket fuels. The rocket fuel saga

From the book Rules for the technical operation of thermal power plants author Team of authors

How does a liquid-propellant jet engine work? Liquid-propellant jet engines are currently used as engines for heavy air defense missiles, long-range and stratospheric missiles, rocket aircraft, rocket-propelled bombs,

4. FUEL ECONOMY. SOLID, LIQUID AND GASEOUS FUELS 4.1. General provisions4.1.1. The operation of fuel equipment must ensure timely, uninterrupted preparation and supply of fuel to the boiler room. There must be a supply of basic and

6.1.1. Fuel Gasoline. To ensure reliable operation of carburetor engines in all modes, gasoline must have: high detonation resistance; optimal fractional composition; low content of resin and carbon-forming compounds and

In the general case, heating of the working fluid is present as a component of the working process of a thermal rocket engine. Moreover, the presence of a heat source - a heater is formally required (in a particular case, its thermal power may be zero). Its type can be characterized by the type of energy converted into heat. Thus, we obtain a classification sign according to which thermal rocket engines, according to the type of energy converted into thermal energy of the working fluid, are divided into electrical, nuclear (Fig. 10.1.) and chemical (Fig. 13.1, level 2).

The design, design and achievable parameters of a chemical fuel rocket engine are largely determined by the aggregate state of the rocket fuel. Chemical fuel rocket engines (sometimes called chemical rocket engines in foreign literature) based on this criterion are divided into:

liquid-propellant rocket engines - liquid-propellant rocket engines, the fuel components of which, when stored on board, are liquid (Fig. 13.1, level 3; photo, photo),

solid fuel rocket engines - solid propellant rocket motors (Fig. 1.7, 9.4, photo, photo),

hybrid rocket engines - GRDs, the fuel components of which are on board in different states of aggregation (Fig. 11.2).

An obvious feature of the classification of chemical fuel engines is the number of propellant components.

For example, liquid-propellant engines using single-component or two-component fuel, gas-propellant engines using three-component fuel (according to foreign terminology - tribrid fuel) (Fig. 13.1, level 4).

Based on design characteristics, it is possible to classify rocket engines with dozens of categories, but the main differences in the performance of the target function are determined by the scheme for supplying components to the combustion chamber. The most typical classification on this basis is liquid propellant rocket engines.

Classification of rocket fuels.

RTs are divided into solid and liquid. Solid rocket fuels have a number of advantages over liquid ones: they are stored for a long time, do not affect the rocket shell, and do not pose a danger to personnel working with them due to their low toxicity.

However, the explosive nature of their combustion creates difficulties in their use.

Solid rocket propellants include ballista and cordite propellants based on nitrocellulose.

The liquid jet engine, the idea of ​​which belongs to K.E. Tsiolkovsky, is most common in astronautics.

Liquid RT can be one-component or two-component (oxidizer and flammable).

Oxidizing agents include: nitric acid and nitrogen oxides (dioxide, tetroxide), hydrogen peroxide, liquid oxygen, fluorine and its compounds.

Kerosene, liquid hydrogen, and hydrazines are used as fuel. The most widely used are hydrazine and unsymmetrical dimethylhydrazine (UDMH).

The substances that make up liquid RT are highly aggressive and toxic to humans. Therefore, the medical service is faced with the problem of carrying out preventive measures to protect personnel from acute and chronic CRT poisoning and organizing emergency care for injuries.

In this regard, the pathogenesis and clinical picture of lesions are studied, means of providing emergency care and treatment of the affected are being developed, means of protecting the skin and respiratory organs are being created, and the maximum permissible concentrations of various CRTs and the necessary hygienic standards are being established.

Launch vehicles 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 4,500 m/s for the oxygen-hydrogen pair, for kerosene-oxygen - 3,500 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 on the rocket, the larger the size of the containers for its storage, 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-propellant 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 more high reliability, mobility and combat readiness.

In zero gravity, the components of liquid fuel move uncontrollably in the space of the tanks. To deposit them, it is necessary to take 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 :

Space flights in near-Earth space (both manned and unmanned).

Space exploration within the Solar System using automatic vehicles (Voyager, Galileo).

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

In addition to specific impulse when choosing fuel components, other indicators of fuel properties can also play a decisive role, including:

Density, which affects the size of component tanks. As follows from the table. 1, hydrogen is flammable, with the highest specific impulse (of any oxidizer), but it has an extremely low density. Therefore, the first (largest) stages of launch vehicles usually use other (less efficient, but denser) types of fuel, for example, kerosene, which makes it possible to reduce the size of the first stage to acceptable ones. Examples of such “tactics” are the Saturn 5 rocket, the first stage of which uses oxygen/kerosene components, and the 2nd and 3rd stages use oxygen/hydrogen, and the Space Shuttle system, in which solid rocket boosters are used as the first stage.

Boiling point, which can impose serious restrictions on the operating conditions of the rocket. According to this indicator, liquid fuel components are divided into cryogenic - liquefied gases cooled to extremely low temperatures, and high-boiling - liquids with a boiling point above 0 ° C.

Cryogenic components cannot be stored for a long time or transported over long distances, so they must be manufactured (at least liquefied) in special energy-intensive production facilities located in close proximity to the launch site, which makes the launcher completely immobile. In addition, cryogenic components have other physical properties that impose additional requirements for their use. For example, the presence of even a small amount of water or water vapor in containers with liquefied gases leads to the formation of very hard ice crystals, which, if they enter the rocket fuel system, act on its parts as an abrasive material and can cause a serious accident. During the many hours of preparation of the rocket for launch, a large amount of frost freezes on it, turning into ice, and the fall of its pieces from a great height poses a danger to the personnel involved in the preparation, as well as to the rocket itself and the launch equipment. After the rockets are filled with liquefied gases, they begin to evaporate, and until the moment of launch they must be continuously replenished through a special replenishment system. Excess gas formed during the evaporation of components must be removed in such a way that the oxidizer does not mix with the fuel, forming an explosive mixture.

High-boiling components are much more convenient to transport, store and handle, so in the 1950s they replaced cryogenic components from the field of military rocketry. Subsequently, this area increasingly began to focus on solid fuels. But when creating space launch vehicles, cryogenic fuels still retain their position due to their high energy efficiency, and for maneuvers in outer space, when fuel must be stored in tanks for months, or even years, high-boiling components are the most suitable. An illustration of this “division of labor” can be seen in the liquid rocket engines involved in the Apollo project: all three stages of the Saturn 5 launch vehicle use cryogenic components, and the engines of the lunar ship, intended for trajectory correction and for maneuvers in lunar orbit, use high-boiling asymmetrical dimethylhydrazine and tetroxide dinitrogen.

Chemical aggressiveness. All oxidizing agents have this quality. Therefore, the presence of even small amounts of organic substances in tanks intended for oxidizer (for example, grease stains left by human fingers) can cause a fire, which can cause the material of the tank itself to catch fire (aluminum, magnesium, titanium and iron burn very vigorously in the environment of the rocket oxidizer ). Due to their aggressiveness, oxidizers, as a rule, are not used as coolants in liquid-propellant rocket engine cooling systems, and in TNA gas generators, to reduce the thermal load on the turbine, the working fluid is oversaturated with fuel rather than oxidizer. At low temperatures, liquid oxygen is perhaps the safest oxidizer because alternative oxidizers such as dinitrogen tetroxide or concentrated nitric acid react with metals, and although they are high-boiling oxidizers that can be stored for long periods at normal temperatures, the service life The tanks in which they are located are limited.

The toxicity of fuel components and their combustion products is a serious limitation of their use. For example, fluorine, as follows from Table 1, as an oxidizing agent, is more effective than oxygen, but when paired with hydrogen it forms hydrogen fluoride - an extremely toxic and aggressive substance, and the release of several hundred, much less thousands of tons of such combustion product into atmosphere when launching a large rocket is itself a major man-made disaster, even with a successful launch. And in the event of an accident and a spill of such a quantity of this substance, the damage cannot be accounted for. Therefore, fluorine is not used as a fuel component. Nitrogen tetroxide, nitric acid and unsymmetrical dimethylhydrazine are also toxic. Currently, the preferred (from an environmental point of view) oxidizer is oxygen, and the fuel is hydrogen, followed by kerosene.

A powerful space rocket is driven by the same force as a festive fireworks display in a recreation park - the reaction force of the gases flowing from the nozzle. Escaping from the rocket engine in a column of fire, they push the engine itself and everything that is structurally connected to it in the opposite direction.

The main fundamental difference between any jet engine (rocket engines are a powerful branch of the vast family of jet engines, direct reaction engines) is that it directly generates motion, itself propelling the associated vehicle without the participation of intermediate units called propulsors. In an aircraft with piston or turboprop engines, the engine causes the propeller to rotate, which, when it hits the air, throws a mass of air back and forces the aircraft to fly forward. In this case, the propeller is the propeller. A ship's propeller works in a similar way: it throws out a mass of water. In a car or train, the propeller is the wheel. And only a jet engine does not need support in the environment, in the mass from which the vehicle would push off. The mass that a jet engine throws back and thereby gains forward motion is located within itself. It is called the working fluid, or the working substance of the engine.

Typically, hot gases operating in an engine are formed during the combustion of fuel, that is, during a chemical reaction of violent oxidation of a combustible substance. The chemical energy of the combustion substances is converted into thermal energy of combustion products. And the thermal energy of the hot gases obtained in the combustion chamber is converted, when they expand in the nozzle, into the mechanical energy of the translational motion of a rocket or jet aircraft.

The energy used in these engines is the result of a chemical reaction. Therefore, such engines are called chemical rocket engines.

This is not the only possible case. In nuclear rocket engines, the working substance must receive energy from the heat generated during the reaction of nuclear decay or fusion. In some types of electric rocket engines, the working substance is accelerated without the participation of heat at all due to the interaction of electric and magnetic forces. Nowadays, however, the basis of rocket technology is chemical, or, as they are also called, thermochemical rocket engines.

Not all jet engines are suitable for space flight. A large class of these machines, the so-called air-breathing engines, use ambient air to oxidize fuel. Naturally, they can only work within the earth's atmosphere.

To operate in space, two types of thermochemical rocket engines are used: solid rocket engines (solid propellant rocket engines) and liquid rocket engines (LPRE). In these engines, the fuel contains everything that is needed for combustion, i.e., both fuel and oxidizer. Only the aggregate state of this fuel is different. A solid propellant rocket engine is a solid mixture of necessary substances. In a liquid rocket engine, the fuel and oxidizer are stored in liquid form, usually in separate tanks, and ignition occurs in the combustion chamber, where the fuel is mixed with the oxidizer.

The movement of the rocket occurs when the working substance is thrown away. It is far from indifferent at what speed the working fluid flows out of the jet engine nozzle. The physical law of conservation of momentum says that the momentum of a rocket (the product of its mass and the speed at which it flies) will be equal to the momentum of the working fluid. This means that the greater the mass of gases ejected from the nozzle and the speed of their expiration, the greater the engine thrust, the greater the speed that can be given to the rocket, the greater its mass and payload can be.

In a large rocket engine, within a few minutes of operation, a huge amount of fuel, the working fluid, is processed and ejected from the nozzle at high speed. To increase the speed and mass of a rocket, besides dividing it into stages, there is only one way - increasing the thrust of the engines. And you can increase thrust without increasing fuel consumption only by increasing the speed of gas flow from the nozzle.

In rocket technology there is a concept of specific thrust of a rocket engine. Specific thrust is the thrust obtained in the engine when consuming one kilogram of fuel in one second.

Specific thrust is identical to specific impulse - the impulse developed by a rocket engine for each kilogram of fuel consumed (working fluid). Specific impulse is determined by the ratio of engine thrust to the mass of fuel consumed per second. Specific impulse is the most important characteristic of a rocket engine.

The specific impulse of the engine is proportional to the speed of gas flow from the nozzle. Increasing the exhaust speed allows you to reduce fuel consumption per kilogram of thrust developed by the engine. The greater the specific thrust, the greater the exhaust velocity of the working fluid, the more economical the engine, the less fuel the rocket needs to complete the same flight.

And the flow rate directly depends on the kinetic energy of the movement of gas molecules, on its temperature and, consequently, on the calorific value (calorific value) of the fuel. Naturally, the higher the calorie content and energy productivity of the fuel, the less it is needed to perform the same work.

But the flow rate depends not only on temperature; it increases with decreasing molecular weight of the working substance. The kinetic energy of molecules at the same temperature is inversely proportional to their molecular weight. The lower the molecular weight of the fuel, the greater the volume of gases produced during its combustion. The greater the volume of gases formed during fuel combustion, the greater the speed of their flow. Therefore, hydrogen as a component of rocket fuel is doubly beneficial due to its high calorific value and low molecular weight.

A very important characteristic of a rocket engine is its specific mass, that is, the mass of the engine per unit of its thrust. A rocket engine must develop high thrust and at the same time be very light. After all, lifting every kilogram of load into space comes at a high price, and if the engine is heavy, it will mainly lift only itself. Most jet engines generally have a relatively low specific gravity, but this indicator is especially good for liquid propellant engines and solid propellant rocket engines. This is due to the simplicity of their design.

Solid propellant rocket engine and rocket engine

Solid fuel rocket engines are extremely simple in design. They essentially have two main parts: the combustion chamber and the jet nozzle. The combustion chamber itself serves as the fuel tank. True, this is not only an advantage, but also a very significant drawback. It is difficult to turn off the engine until all the fuel has burned out. Its operation is extremely difficult to regulate. The fuel must burn slowly, at a more or less constant rate, regardless of changes in pressure and temperature. The thrust value of a solid propellant rocket engine can be adjusted only within certain, predetermined limits, by selecting solid propellant charges of the appropriate geometry and structure. In a solid propellant rocket engine it is difficult to regulate not only the thrust force, but also its direction. To do this, you need to change the position of the traction chamber, and it is very large, because it contains the entire fuel supply. Solid fuel rockets with rotating nozzles have appeared; their design is quite complex, but this allows us to solve the problem of controlling the direction of thrust.

However, solid fuel rocket engines also have a number of serious advantages: constant readiness for action, reliability and ease of operation. Solid propellant rocket engines have found wide application in military affairs.

The most important element in a solid propellant rocket engine is the solid fuel charge. Engine characteristics depend on both the fuel elements and the charge structure and device. There are two main types of solid rocket fuels: dibasic, or colloidal, and mixed. Colloidal fuels are a solid homogeneous solution of organic substances, the molecules of which contain oxidizing and combustible elements. The most widely used is a solid solution of nitrocellulose and nitroglycerin.

Mixed fuels are mechanical mixtures of fuel and oxidizer. Inorganic crystalline substances such as ammonium perchlorate, potassium perchlorate, etc. are usually used as an oxidizer in these fuels. Typically, such fuel consists of three components: in addition to the oxidizer, it includes a polymer fuel that serves as a binding element, and a second fuel in the form of powdered metal additives, which significantly improve the energy characteristics of the fuel. The binding fuel can be polyester and epoxy resins, polyurethane and polybutadiene rubber, etc. The second fuel is most often powdered aluminum, sometimes beryllium or magnesium. Mixed fuels usually have a higher specific impulse than colloidal ones, higher density, greater stability, better storage, and more processability.

Solid fuel charges can be attached to the engine chamber body (they are made by pouring fuel directly into the body) and insert charges, which are made separately and inserted into the body in the form of one or more blocks.

The geometric shape of the charge is very important. By changing it and using armor coatings on charge surfaces that should not burn, we achieve the desired change in the combustion area and, accordingly, the gas pressure in the chamber and engine thrust.

There are charges that provide neutral combustion. Their combustion area remains unchanged. This happens if, for example, a stick of solid fuel burns from the end or simultaneously from the outer and inner surfaces (for this purpose, a cavity is made inside the charge). With regressive combustion, the combustion surface decreases. Thek is obtained if the cylindrical block burns from the outer surface. And finally, for progressive combustion, which ensures an increase in pressure in the combustion chamber, an increase in the combustion area is necessary. The simplest example of such a charge is a checker burning on the inner cylindrical surface.

Bonded charges with internal combustion have the most significant advantages. In them, hot combustion products do not come into contact with the walls of the housing, which makes it possible to do without special external cooling. In astronautics, solid fuel rocket engines are currently used to a limited extent. Powerful solid propellant rocket engines are used on some American launch vehicles, for example, on the Titan rocket.

Large modern solid propellant engines develop a thrust of hundreds of tons, even more powerful engines with a thrust of thousands of tons are being developed, solid fuels are being improved, and thrust control systems are being designed. And yet, liquid rocket engines undoubtedly dominate in astronautics. The main reason for this is the lower efficiency of solid rocket fuel. The best solid propellant rocket motors have a gas flow rate from the nozzle of 2500 meters per second. Liquid rocket engines have a higher specific thrust and an exhaust velocity of (for the best modern engines) 3500 meters per second, and using fuel with a very high calorific value (for example, liquid hydrogen as a fuel and liquid oxygen as an oxidizer), an exhaust velocity of four seconds can be achieved half a kilometer per second.

For the design and operation of a liquid-propellant rocket engine, the fuel on which the engine runs is of great importance.

There are known fuels that release energy during decomposition reactions, for example, hydrogen peroxide, hydrazine. They naturally consist of one component, one liquid. However, chemical fuels that release energy during combustion reactions are most widely used in rocket technology. They consist of an oxidizer and a fuel. Such fuels can also be single-component, i.e., represent one liquid. This can be a substance whose molecule contains both oxidizing and flammable elements, for example, nitromethane, or a mixture of an oxidizer and a fuel, or a solution of a fuel in an oxidizer. However, such fuels are usually prone to explosion and are rarely used. The vast majority of liquid-propellant rocket engines operate on two-propellant fuel. The oxidizer and fuel are stored in separate tanks, and their mixing occurs in the engine chamber. The oxidizer usually makes up the majority of the fuel mass - it is consumed two to four times more than the fuel. Liquid oxygen, nitrogen tetroxide, nitric acid, and hydrogen peroxide are most often used as an oxidizing agent. Kerosene, alcohol, hydrazine, ammonia, liquid hydrogen, etc. are used as fuel.

The Soviet Vostok launch vehicle ran on fuel consisting of liquid oxygen and kerosene, which ensured the launch of many of our spacecraft with astronauts on board. The same fuel powered the engines of the American Atlas and Titan rockets and the first stage of the Saturn 5 rocket, which launched the Apollo spacecraft to the Moon. The fuel, consisting of liquid oxygen and kerosene, is well developed in production and operation, reliable and cheap. It is widely used in liquid rocket engines.

Unsymmetrical dimethylhydrazine has found use as a fuel. This fuel, paired with an oxidizer - liquid oxygen - is used in the RD-119 engine, widely used in the launch of Cosmos satellites. This engine achieves the highest specific impulse for liquid propellant engines running on oxygen and high-boiling fuels.

The most effective rocket fuel currently in widespread use is liquid oxygen plus liquid hydrogen. It is used, for example, in the engines of the second and third stages of the Saturn 5 rocket.

The search for new, increasingly efficient rocket fuels continues constantly. Scientists and designers are working hard to use fluorine in liquid rocket engines, which has a stronger oxidizing effect than oxygen. Fuels produced using fluorine make it possible to obtain the highest specific impulse for liquid rocket engines and have a high density. However, its use in liquid-propellant rocket engines is complicated by the high chemical aggressiveness and toxicity of liquid fluorine, high combustion temperature (more than 4500° C) and high cost.

Nevertheless, development and bench testing of liquid propellant rocket engines using fluorine are underway in a number of countries. F.A. Tsander first proposed the use of liquid fluorine for liquid propellant engines in 1932, and in 1933 V.P. Glushzho proposed a mixture of liquid fluorine and liquid oxygen as an oxidizer.

Many fluorine-based fuels self-ignite when the oxidizer and fuel are mixed. Some fuel vapors that do not contain fluorine also ignite spontaneously. Self-ignition is a great advantage of fuel. It makes it possible to simplify the design of the rocket engine and increase its reliability. Some fuels become self-igniting when a catalyst is added. So, if a hundredth of a percent of ozone fluoride is added to the oxidizing agent—liquid oxygen—then the combination of this oxidizing agent with kerosene becomes self-igniting.

Self-ignition of the fuel (if it is not self-igniting, then pyrotechnic or electrical ignition is used, or injection of a portion of starting self-igniting fuel) occurs in the engine chamber. The chamber is the main unit of the liquid rocket engine. It is in the chamber that the fuel components are mixed, its combustion occurs, and as a result a gas is formed at a very high temperature (2000-4500 ° C) and under high pressure (tens and hundreds of atmospheres). Flowing out of the chamber, this gas creates reactive force, engine thrust. The liquid rocket engine chamber consists of a combustion chamber with a mixing head and a nozzle. Mixing of fuel components occurs in the mixing head, combustion occurs in the combustion chamber, and gases flow out through the nozzle. Typically, all chamber units are made as one whole. Most often, combustion chambers are cylindrical in shape, but they can also be conical or spherical (pear-shaped).

The mixing head is a very important part of the combustion chamber and the entire liquid propellant engine. In it, the so-called mixture formation-injection, atomization and mixing of fuel components occurs. The fuel components - oxidizer and fuel - enter the mixing head of the chamber separately. They are introduced into the chamber through the head nozzles due to the pressure difference in the fuel supply system and the chamber head. In order for the reaction in the combustion chamber to proceed as quickly as possible and be as complete as possible - and this is a very important condition for the efficiency and economy of the engine - it is necessary to ensure the most rapid and complete formation of the fuel mixture burning in the chamber, to ensure that every particle of the oxidizer meets a particle fuel.

The formation of a fuel mixture prepared for combustion consists of three processes that transform into one another - atomization of liquid components, their evaporation and mixing. When atomizing - crushing a liquid into droplets - its surface significantly increases and the evaporation process accelerates. The fineness and uniformity of the spray is very important. The subtlety of this process is characterized by the diameter of the resulting droplets: the smaller each droplet, the better. After atomization, the next stage of preparing the fuel for combustion is its evaporation. It is necessary to ensure the most complete evaporation of the oxidizer and fuel in the shortest possible time. The process of evaporation of droplets formed during spraying in the liquid-propellant rocket engine chamber takes only two to eight thousandths of a second.

As a result of atomization and evaporation of fuel components, vapors of oxidizer and fuel are formed, from which a mixture burning in the engine chamber is obtained. The mixing of components begins, essentially, immediately after the components enter the chamber and ends only as the fuel burns. With self-igniting fuels, the combustion process begins in the liquid phase, during atomization of the fuel. With non-self-igniting fuels, combustion begins in the gas phase when heat is supplied from an external source.

Liquid fuel components are supplied into the chamber through nozzles located in the head. Most often, two types of nozzles are used: jet or centrifugal. But the fuel was atomized, mixed, and ignited. When it burns in the combustion chamber, a large amount of thermal energy is released. Further energy conversion occurs in the nozzle. The successful design of the mixing head primarily determines the perfection of the engine - it ensures complete fuel combustion, combustion stability, etc.

The nozzle is a part of the combustion chamber in which the thermal energy of the compressed working fluid (gas mixture) is converted into the kinetic energy of the gas flow, i.e., it accelerates to the speed of exhaust from the engine. The nozzle usually consists of a converging and diverging part, which are connected at the critical (minimum) section.

A very difficult task is to ensure cooling of the liquid-propellant rocket engine chamber. Typically the chamber consists of two shells - an inner fire wall and an outer jacket. Liquid flows through the space between the shells, cooling the inner wall of the liquid-propellant rocket engine chamber. Typically one of the fuel components is used for this. The heated fuel or oxidizer is removed and enters the chamber head for use, so to speak, for its intended purpose. In this case, the thermal energy taken from the chamber walls is not lost, but is returned to the chamber. Such cooling (regenerative) was first proposed by K. E. Tsiolkovsky and is widely used in rocket technology.

Most modern liquid propellant engines use special turbopump units to supply fuel. To power such a powerful pump, fuel is burned in a special gas generator - usually the same fuel and the same oxidizer as in the combustion chamber of the engine. Sometimes the pump turbine is driven by steam, which is generated when the engine combustion chamber cools. There are other pump drive systems.

The creation of modern liquid-propellant rocket engines requires a high level of development of science and technology, perfection of design ideas, and advanced technology. The fact is that very high temperatures are reached in liquid rocket engines, enormous pressure develops, combustion products, and sometimes the fuel itself, are very aggressive, fuel consumption is unusually high (up to several tons per second!). With all this, the liquid-propellant rocket engine must have, especially when launching spacecraft with astronauts on board, a very high degree of reliability. It is high reliability and many other advantages that distinguish the liquid rocket engines of the famous Soviet space rocket “Vostok” - RD-107 (first stage engine) and RD-108 (second stage engine), developed in 1954-1957 under the leadership of the chief designer of rocket engines V P. Glushko. These are the world's first production engines running on high-calorie fuel; liquid oxygen and kerosene. They have a high specific thrust, which makes it possible to obtain enormous power with relatively moderate fuel consumption. In vacuum, the thrust of one RD-107 engine is 102 tons. (The first stage of the Vostok launch vehicle has four such engines.) The pressure in the combustion chamber is 60 atmospheres.

The RD-107 engine has a turbopump unit with two main centrifugal pumps; one supplies fuel, the other supplies oxidizer. Both fuel and oxidizer are supplied through a large number of nozzles into four main and two steering combustion chambers. Before entering the combustion chambers, the fuel flows around them from the outside, i.e., it is used for cooling. Reliable cooling allows you to maintain high temperatures inside the combustion chambers. Oscillating steering combustion chambers, similar in design to the main ones, were used for the first time in this engine to control the direction of thrust.

The second stage engine of the Vostok rocket RD-108 has a similar design. True, it has four steering cameras and some other differences. Its vacuum thrust is 96 tons. Interestingly, it is launched on Earth simultaneously with the first stage engines. RD-107 and RD-108 engines of various modifications have been used for many years to launch spacecraft, artificial Earth satellites, and spacecraft to the Moon, Venus and Mars.

The second stage of the two-stage Cosmos launch vehicle is equipped with the RD-119 liquid-propellant rocket engine, developed in 1958-1962 (also at the GDL-OKB), having a thrust of 11 tons; The fuel of this engine is asymmetrical dimethylhydrazine, the oxidizer is liquid oxygen. Titanium and other modern structural materials are widely used in its design. Along with high reliability, a distinctive feature of this engine is its very high efficiency. In 1965, powerful small-sized engines with very high energy characteristics were created in our country for the Proton rocket and space system. The total net power of the Proton rocket propulsion systems is three times greater than the power of the Vostok rocket engines and amounts to 60 million horsepower. These engines ensure high combustion efficiency, significant pressure in the system, and uniform and balanced flow of combustion products from the nozzles.

Currently, liquid-propellant engines have reached a high degree of perfection and their development continues. Liquid-propellant engines of various classes have been created - from micro-rocket engines for orientation and stabilization systems of aircraft with very little thrust (several kilograms or less) to huge powerful rocket engines with a thrust of hundreds of tons (for example, the American G-1 liquid-propellant rocket engine for the first stage of the Saturn-5 launch vehicle has a thrust of 690 tons. The rocket has five such engines installed).

Liquid rocket engines are being developed using highly efficient fuels - a mixture of liquid hydrogen (fuel) and liquid oxygen or liquid fluorine as oxidizers. Engines have been created using long-lasting fuel that can operate during long-term space flights.

There are projects for combined rocket engines - turbo-rocket and ramjet rocket engines, which should be an organic combination of liquid-propellant rocket engines with air-breathing ones. The creation of such engines makes it possible to use atmospheric oxygen as an oxidizer at the initial and final stages of space flight and thereby reduce the fuel supply on board the rocket. Work is also underway to create the first reusable stages. Such stages, equipped with air-breathing engines and capable of taking off and, after the separation of subsequent stages, landing like airplanes, will reduce the cost of launching spacecraft.

NUCLEAR ROCKET ENGINES

Scientists and designers have created thermochemical engines of a high degree of perfection and, no doubt, even more perfect examples will be created. However, the capabilities of thermochemical rockets are limited by the very nature of the fuel, oxidizer, and reaction products. Given the limited energy productivity of rocket fuels, which does not allow obtaining a very high speed of flow of the working fluid from the nozzle, a huge supply of fuel is required to accelerate the rocket to the required speed. Chemical rockets are unusually voracious. This is a question not only of saving, but sometimes of doing the most possible! and space flight.

Even to solve a relatively simpler problem in the field of space flights - launching artificial Earth satellites, the launch mass of a chemical rocket, due to the huge amount of fuel, must be many tens of times greater than the mass of the cargo launched into orbit. To achieve the second escape velocity, this ratio is even greater. But humanity is beginning to settle in space, people are going to build scientific stations on the Moon, they are striving for Mars and Venus, they are thinking about flights to the distant outskirts of the solar system. The rockets of tomorrow will have to carry many tons of scientific equipment and cargo in space.

For interplanetary flights, more fuel is needed to adjust the flight orbit, slow down the spacecraft before landing on the target planet, take off to return to Earth, etc. The launch mass of thermochemical rockets for such flights becomes incredibly large - several million tons!

Scientists and engineers have long been thinking about what the rocket engines of the future should be? Naturally, scientists turned their attention to nuclear energy. A tiny amount of nuclear fuel contains a very large amount of energy. The nuclear fission reaction releases millions of times more energy per unit mass than the combustion of the best chemical fuels. For example, 1 kilogram of uranium during a fission reaction can release the same amount of energy as 1,700 tons of gasoline when burned. The nuclear fusion reaction produces several times more energy.

The use of nuclear energy can dramatically reduce the fuel supply on board a rocket, but there remains a need for a working substance that will be heated in the reactor and ejected from the engine nozzle. Upon closer examination, it turns out that the separation of fuel and working substance in a nuclear rocket has certain advantages.

The choice of working substance for a chemical rocket is very limited. After all, it also serves as fuel. This is where the advantage of separating fuel and working substance comes into play. It becomes possible to use a working substance with the lowest molecular weight—hydrogen.

A chemical rocket also uses a combination of the relatively high energy efficiency of hydrogen with low molecular weight. But there the working substance is the combustion product of hydrogen with a molecular weight of 18. And the molecular weight of pure hydrogen, which can serve as the working fluid of a nuclear rocket engine, is 2. Reducing the molecular weight of the working substance by 9 times at a constant temperature makes it possible to increase the exhaust speed by 3 times . Here it is, a tangible advantage of a nuclear rocket engine!

We are talking about atomic rocket engines that use the energy of fission of nuclei of heavy elements. The nuclear fusion reaction has so far been artificially carried out only in the hydrogen bomb, and the controlled thermonuclear fusion reaction is still a dream, despite the intensive work of many scientists in the world.

So, in a nuclear rocket engine it is possible to obtain a significant increase in the gas flow rate due to the use of a working substance with a minimum molecular weight. Theoretically, it is possible to obtain a very high temperature of the working substance. But in practice it is limited by the melting temperature of the reactor fuel elements.

In most proposed schemes for nuclear rocket engines, the working fluid is heated, washing the fuel elements of the reactor, then expands in the nozzle and is thrown out of the engine. The temperature is approximately the same as in chemical rocket engines. True, the engine itself turns out to be much more complex and heavy. Especially when you consider the need for a shield to protect astronauts from radiation on manned spacecraft. Still, a nuclear missile promises considerable gains.

In the United States, under the so-called “Rover” program, intensive work is underway to create a nuclear rocket engine. Projects for nuclear rocket engines have also emerged in which the core is in a dusty, liquid or even gaseous phase. This makes it possible to obtain a higher temperature of the working substance. The use of such reactors (they are called cavity reactors) would probably make it possible to significantly increase the flow rate of the working fluid. But the creation of such reactors is an extremely difficult matter: nuclear fuel is mixed with the working substance, and it is necessary to somehow separate it before the working fluid is ejected from the engine nozzle. Otherwise, there will be continuous losses of nuclear fuel, and a deadly trail of high radiation will follow the rocket. And the critical mass of nuclear fuel required to maintain reactions, in a gaseous state, will occupy a very large volume that is not acceptable for a rocket.
(L.A. Gilberg: Conquest of the sky)

"Buran", like its overseas brother - the reusable rocket system "Shuttle", in its characteristics leaves much to be desired.

They turned out to be not so reusable. The launch boosters can withstand the entire 3-4 flights, and the winged vehicle itself burns out and requires very expensive repairs. But the main thing is that their efficiency is not great.

And here is such a temptation - to create a manned winged vehicle capable of independently launching from Earth, going into outer space and returning back. True, the main problem remains unresolved - the engine. Air-jet engines of known types are capable of operating only up to a speed of 4-5 Mach (Mach is the speed of sound), and the first cosmic speed, as is known, is 24 Mach. But even here, it seems, the first steps towards success have already been outlined.

At the Aviadvigatele-Stroenie-92 exhibition, held in Moscow, among all kinds of exhibits - from ancient steam engines for airships to giant turbines of ultra-modern transport aircraft - a small barrel stood modestly on the stand - the world's first and only model of a hypersonic (Hypersonic - from 6M and higher) air-breathing jet engine (scramjet engine). It was created at the Central Institute of Aviation Engine Engineering (CIAM). Of course, this is the result of the work of a large team. First of all, the chief designer D. A. Ogorodnikov, his associates A. S. Rudakov, V. A. Vinogradov... Really, we should not forget those who are no longer alive - this is Doctor of Technical Sciences R. I. Kurziner and Professor E. S. Shchetinkov. The latter, several decades ago, proposed the basic principle underlying all modern scramjet engines. The engine he developed was already at that time capable of operating at hypersonic (above 5-6 M) speeds. These people created a miracle of technology, which, perhaps, in the near future will revolutionize space propulsion engineering.

But let’s not rush to “adapt” a new engine to a space plane, be it Buran or Spiral, let’s turn to theory. The fact is that each engine can operate only in a certain range, which is too narrow for space tasks, and getting it to master hypersound is far from easy. Let's figure out why.

In any WFD, three most important conditions must be met for successful operation. First of all, you need to compress the air as much as possible. Then burn the fuel in the combustion chamber without loss. Finally, with the help of a nozzle, the combustion products must expand to atmospheric pressure. Only then will the efficiency be high enough.

Look at the picture. Here is a diagram of the world's first hypersonic ramjet engine (scramjet). He solves his first problem - compressing air - in a very original way - according to the principle of... a cleaver. Imagine: a cleaver cuts into a soft, dense log, the layers of wood in front of it remain unchanged, but become compacted on the sides. Scientists call the boundary between normal and denser layers a “compression shock.” This happens in the engine as well. A pointed central body is located along its axis. Crashing into the air, it creates such a “jump” - a zone of increased pressure. There is a “reflection” of air from the central body to the walls of the housing. At the same time, it is further compressed many times. The air speed decreases and the temperature rises, kinetic energy is converted into internal, thermal energy.

Now, in order for the fuel injected into the flow to burn completely, it is desirable to obtain a speed as low as possible. But then the air temperature can reach 3-5 thousand degrees. It would seem good - the fuel will ignite like gunpowder. But even if there was real gunpowder there, there would be no flash. The thing is that at such high temperatures, along with the oxidation process, molecules also disintegrate into individual atoms. If in the first one energy is released, then in the second it is absorbed. And the paradox is that as the temperature rises, a moment may come when more is absorbed than released. In other words, the firebox will turn into... a refrigerator.

Professor Shchetinkov suggested an original way out of the situation back in 1956. He suggested compressing the air only until its supersonic speed becomes approximately the same as that of... a bullet. As is now recognized throughout the world, only under these conditions is scramjet operation possible.

But this also has its difficulties: even a mixture of hydrogen with air, known to us from chemistry courses as “explosive gas,” will barely have time to ignite under such conditions. And although liquid hydrogen was chosen as fuel for the engine, we had to resort to tricks. First, hydrogen cools the walls. Heating itself from -256° C to +700° C, it saves the metal from melting. Some of the fuel is injected through the injectors directly into the air stream. And the other part falls on the nozzles located in special rectangular niches. Powerful hydrogen torches burn here, capable of instantly burning through a sheet of steel. They ignite the hydrogen-air mixture. The same one that under normal conditions explodes from a spark dropped from a nylon shirt.

But this is, perhaps, the main task on which we and the Americans spent about 30 years. How to achieve complete combustion with a chamber of acceptable length - 3-5 m? It is known that a theory without a verification experiment is worth little. And to test the operation of such an engine, it must be placed in a hypersonic flow. There are no such aircraft, although there are wind tunnels, but they are very, very expensive. For the final check of the scramjet, the designers installed their device in the nose of the rocket and accelerated it to the required speed.

Let us clarify that the discussion here was not about creating a new type of rocket, but only about checking the quality of hydrogen combustion in the engine. It was a complete success. Now, as the Americans admit, our scientists have the secret of creating reliable combustion chambers.

Well, now let's think about what will happen if we want to enlarge this small exhibition model, making it suitable for lifting an airplane into the air. Apparently, it will take on the features of a heavy thirty-meter pipe with a huge diffuser and nozzle and a very modest combustion chamber. Who needs such an engine? Dead end? No, there is a way out and it has long been known. Many functions in its work can be assigned to... the fuselage and wing of the aircraft!

The prototype of such an aerospace aircraft (VKS) is shown in the figure. “Wedging” its nose part into the air, it creates a series of shock waves, and all of them directly fall on the entrance of the combustion chamber. The hot gases coming out of it, expanding to atmospheric pressure, slide along the surface of the rear of the aircraft, creating thrust, like in a good nozzle. At hypersonic speeds this is also possible! Surprisingly, theoretically, you can even do without a camera and “simply” inject fuel near the protrusion on the belly of the VKS! You will get an engine that seems to not exist. It's called an "external combustion" scramjet. True, its “simplicity” in research work is so expensive that so far no one has seriously studied it.

Therefore, let’s return to an aerospace aircraft with a classical scramjet engine. Its launch and acceleration to b M should occur using conventional turbojet engines. In the picture you see a unit consisting of a traditional turbojet engine and a nearby scramjet engine. At “low” speeds, the scramjet is separated by a streamlined partition and does not interfere with flight.

And on large ones, the partition blocks the air flow going to the turbojet engine, and the scramjet engine turns on.

At first everything will go well, but then, as the speed increases, the engine thrust will begin to fall, and appetites - fuel consumption - will increase. At this moment, his insatiable belly must be fed with liquid oxygen. Whether you like it or not, you still have to take it with you. True, in quantities much smaller than on a conventional rocket. Somewhere 60 kilometers from the Earth, the scramjet engine will stall from lack of air. And then a small liquid-propellant rocket engine comes into action. The speed is already high, and very little fuel and oxidizer will be consumed before entering orbit. With the same launch weight of the rocket, the aerospace aircraft is launched into orbit with a payload that is 5-10 times larger. And the cost of launching each kilogram will be tens of times lower than missiles. This is exactly what scientists and designers are striving for today.

In the general case, heating of the working fluid is present as a component of the working process of a thermal rocket engine. Moreover, the presence of a heat source - a heater is formally required (in a particular case, its thermal power may be zero). Its type can be characterized by the type of energy converted into heat. Thus, we obtain a classification sign according to which thermal rocket engines, according to the type of energy converted into thermal energy of the working fluid, are divided into electrical, nuclear (Fig. 10.1.) and chemical (Fig. 13.1, level 2).

The design, design and achievable parameters of a chemical fuel rocket engine are largely determined by the aggregate state of the rocket fuel. Chemical fuel rocket engines (sometimes called chemical rocket engines in foreign literature) based on this criterion are divided into:

liquid-propellant rocket engines - liquid-propellant rocket engines, the fuel components of which, when stored on board, are liquid (Fig. 13.1, level 3; photo, photo),

solid fuel rocket engines - solid propellant rocket motors (Fig. 1.7, 9.4, photo, photo),

hybrid rocket engines - GRDs, the fuel components of which are on board in different states of aggregation (Fig. 11.2).

An obvious feature of the classification of chemical fuel engines is the number of propellant components.

For example, liquid-propellant engines using single-component or two-component fuel, gas-propellant engines using three-component fuel (according to foreign terminology - tribrid fuel) (Fig. 13.1, level 4).

Based on design characteristics, it is possible to classify rocket engines with dozens of categories, but the main differences in the performance of the target function are determined by the scheme for supplying components to the combustion chamber. The most typical classification on this basis is liquid propellant rocket engines.

Classification of rocket fuels.

RTs are divided into solid and liquid. Solid rocket fuels have a number of advantages over liquid ones: they are stored for a long time, do not affect the rocket shell, and do not pose a danger to personnel working with them due to their low toxicity.

However, the explosive nature of their combustion creates difficulties in their use.

Solid rocket propellants include ballista and cordite propellants based on nitrocellulose.

The liquid jet engine, the idea of ​​which belongs to K.E. Tsiolkovsky, is most common in astronautics.

Liquid RT can be one-component or two-component (oxidizer and flammable).

Oxidizing agents include: nitric acid and nitrogen oxides (dioxide, tetroxide), hydrogen peroxide, liquid oxygen, fluorine and its compounds.

Kerosene, liquid hydrogen, and hydrazines are used as fuel. The most widely used are hydrazine and unsymmetrical dimethylhydrazine (UDMH).

The substances that make up liquid RT are highly aggressive and toxic to humans. Therefore, the medical service is faced with the problem of carrying out preventive measures to protect personnel from acute and chronic CRT poisoning and organizing emergency care for injuries.

In this regard, the pathogenesis and clinical picture of lesions are studied, means of providing emergency care and treatment of the affected are being developed, means of protecting the skin and respiratory organs are being created, and the maximum permissible concentrations of various CRTs and the necessary hygienic standards are being established.

Launch vehicles 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 4,500 m/s for the oxygen-hydrogen pair, for kerosene-oxygen - 3,500 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 on the rocket, the larger the size of the containers for its storage, 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-propellant 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 more high reliability, mobility and combat readiness.

In zero gravity, the components of liquid fuel move uncontrollably in the space of the tanks. To deposit them, it is necessary to take 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 :

Space flights in near-Earth space (both manned and unmanned).

Space exploration within the Solar System using automatic vehicles (Voyager, Galileo).

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

In addition to specific impulse when choosing fuel components, other indicators of fuel properties can also play a decisive role, including:

Density, which affects the size of component tanks. As follows from the table. 1, hydrogen is flammable, with the highest specific impulse (of any oxidizer), but it has an extremely low density. Therefore, the first (largest) stages of launch vehicles usually use other (less efficient, but denser) types of fuel, for example, kerosene, which makes it possible to reduce the size of the first stage to acceptable ones. Examples of such “tactics” are the Saturn 5 rocket, the first stage of which uses oxygen/kerosene components, and the 2nd and 3rd stages use oxygen/hydrogen, and the Space Shuttle system, in which solid rocket boosters are used as the first stage.

Boiling point, which can impose serious restrictions on the operating conditions of the rocket. According to this indicator, liquid fuel components are divided into cryogenic - liquefied gases cooled to extremely low temperatures, and high-boiling - liquids with a boiling point above 0 ° C.

Cryogenic components cannot be stored for a long time or transported over long distances, so they must be manufactured (at least liquefied) in special energy-intensive production facilities located in close proximity to the launch site, which makes the launcher completely immobile. In addition, cryogenic components have other physical properties that impose additional requirements for their use. For example, the presence of even a small amount of water or water vapor in containers with liquefied gases leads to the formation of very hard ice crystals, which, if they enter the rocket fuel system, act on its parts as an abrasive material and can cause a serious accident. During the many hours of preparation of the rocket for launch, a large amount of frost freezes on it, turning into ice, and the fall of its pieces from a great height poses a danger to the personnel involved in the preparation, as well as to the rocket itself and the launch equipment. After the rockets are filled with liquefied gases, they begin to evaporate, and until the moment of launch they must be continuously replenished through a special replenishment system. Excess gas formed during the evaporation of components must be removed in such a way that the oxidizer does not mix with the fuel, forming an explosive mixture.

High-boiling components are much more convenient to transport, store and handle, so in the 1950s they replaced cryogenic components from the field of military rocketry. Subsequently, this area increasingly began to focus on solid fuels. But when creating space launch vehicles, cryogenic fuels still retain their position due to their high energy efficiency, and for maneuvers in outer space, when fuel must be stored in tanks for months, or even years, high-boiling components are the most suitable. An illustration of this “division of labor” can be seen in the liquid rocket engines involved in the Apollo project: all three stages of the Saturn 5 launch vehicle use cryogenic components, and the engines of the lunar ship, intended for trajectory correction and for maneuvers in lunar orbit, use high-boiling asymmetrical dimethylhydrazine and tetroxide dinitrogen.

Chemical aggressiveness. All oxidizing agents have this quality. Therefore, the presence of even small amounts of organic substances in tanks intended for oxidizer (for example, grease stains left by human fingers) can cause a fire, which can cause the material of the tank itself to catch fire (aluminum, magnesium, titanium and iron burn very vigorously in the environment of the rocket oxidizer ). Due to their aggressiveness, oxidizers, as a rule, are not used as coolants in liquid-propellant rocket engine cooling systems, and in TNA gas generators, to reduce the thermal load on the turbine, the working fluid is oversaturated with fuel rather than oxidizer. At low temperatures, liquid oxygen is perhaps the safest oxidizer because alternative oxidizers such as dinitrogen tetroxide or concentrated nitric acid react with metals, and although they are high-boiling oxidizers that can be stored for long periods at normal temperatures, the service life The tanks in which they are located are limited.

The toxicity of fuel components and their combustion products is a serious limitation of their use. For example, fluorine, as follows from Table 1, as an oxidizing agent, is more effective than oxygen, but when paired with hydrogen it forms hydrogen fluoride - an extremely toxic and aggressive substance, and the release of several hundred, much less thousands of tons of such combustion product into atmosphere when launching a large rocket is itself a major man-made disaster, even with a successful launch. And in the event of an accident and a spill of such a quantity of this substance, the damage cannot be accounted for. Therefore, fluorine is not used as a fuel component. Nitrogen tetroxide, nitric acid and unsymmetrical dimethylhydrazine are also toxic. Currently, the preferred (from an environmental point of view) oxidizer is oxygen, and the fuel is hydrogen, followed by kerosene.

"...And there is nothing new under the sun"

One of the first works on liquid rocket engine fuels can be considered the book by V.P. Glushko "Liquid fuel for jet engines", published in 1936.

For me, the topic seemed interesting, related to my former specialty and studies at the university, especially since my youngest son “dragged” it: “Chief, let’s knead what the thread is and run it, and if you’re lazy, then we ourselves"Let's figure it out." Apparently they don't give me peace.

I really want to blow up my rocket engine properly.

An important parameter is the oxidizer excess coefficient (denoted by the Greek “α” with the subscript “ok.”) and the mass ratio of the components Km.

Km=(dmok./dt)/(dmg../dt), i.e. the ratio of the mass flow rate of the oxidizer to the mass flow rate of the fuel. It is specific for each fuel. Ideally, it is a stoichiometric ratio of oxidizer and fuel, i.e. shows how many kg of oxidizer is needed to oxidize 1 kg of fuel. However, real values ​​differ from ideal ones. The ratio of real Km to ideal is the oxidizer excess coefficient.

As a rule, α is approx.<=1. И вот почему. Зависимости Tk(αок.) и Iуд.(αок.) нелинейны и для многих топлив последняя имеет максимум при αок. не при стехиометрическом соотношении компонентов, т.е макс. значения Iуд. получаются при некотором снижении количества окислителя по отношению к стехиометрическому. Ещё немного терпения, т.к. не могу обойти понятие: . Это пригодится и в статье, и в повседневной жизни.

In short, enthalpy is energy. Two aspects of this article are important:
Thermodynamic enthalpy- the amount of energy spent on the formation of a substance from the initial chemical elements. For substances consisting of identical molecules (H 2, O 2, etc.), it is equal to zero.
Enthalpy of combustion- makes sense only if a chemical reaction occurs. In reference books you can find values ​​of this quantity experimentally obtained under normal conditions. Most often, for combustibles this is complete oxidation in an oxygen environment, for oxidizers it is the oxidation of hydrogen with a given oxidizer. Moreover, the values ​​can be both positive and negative depending on the type of reaction.

“The sum of the thermodynamic enthalpy and the enthalpy of combustion is called the total enthalpy of the substance. Actually, this value is used in the thermal calculation of liquid-propellant rocket engine chambers.”

Requirements for ZhRT:
-as a source of energy;
-as a substance that has to be used (at this level of technology development) for cooling the rocket engines and pumping pumps, sometimes for pressurizing tanks with rocket engines, providing it with volume (stage rocket tanks), etc.;
-as to a substance outside the rocket engine, i.e. during storage, transportation, refueling, testing, environmental safety, etc.

This gradation is relatively arbitrary, but in principle it reflects the essence. I will name these requirements as follows: No. 1, No. 2, No. 3. Someone can add to the list in the comments.
These requirements are a classic example that “pull” the creators of RD in different directions:

# From the point of view of the LRE energy source (No. 1)

Those. you need to get max. Iud. I won’t bother everyone further, in general:

With other important parameters for No. 1, we are interested in R and T (with all indices).
Need to: the molecular weight of combustion products was minimal, and the specific heat content was maximum.

# From the point of view of the launch vehicle designer (No. 2):

TCs must have maximum density, especially in the first stages of rockets, because they are the most voluminous and have the most powerful thrusters, with a high per second flow rate. Obviously, this is not consistent with requirement No. 1.

# From operational tasks important (No. 3):

Chemical stability of TC;
- ease of refueling, storage, transportation and manufacturing;
- environmental safety (in the entire “field” of application), namely toxicity, cost of production and transportation, etc. and safety during RD operation (explosion hazard).

For more details, see "The saga of rocket fuels - the other side of the coin."

I hope no one has fallen asleep yet? I feel like I'm talking to myself. Coming soon about alcohol, stay tuned!

Of course, this is just the tip of the iceberg. There are also additional requirements here, due to which one should look for CONSENSUSES and COMPROMISES. One of the components must have satisfactory (preferably excellent) coolant properties, because at this level of technology, it is necessary to cool the combustor and the nozzle, as well as protect the critical section of the jet engine:

The photograph shows the nozzle of the XLR-99 liquid-propellant rocket engine: a characteristic feature of the design of American liquid-propellant rocket engines of the 50-60s is clearly visible - a tubular chamber:

It is also required (as a rule) to use one of the components as a working fluid for the turbocharger turbine:

For fuel components, “saturated vapor pressure is of great importance (roughly speaking, the pressure at which a liquid begins to boil at a given temperature). This parameter greatly influences the design of pumps and the weight of tanks.”/ S.S. Fakas/

An important factor is the aggressiveness of the TC towards the materials (CM) of the liquid propellant rocket engine and the tanks for their storage.
If fuel oils are very “harmful” (as some people are), then engineers have to spend money on a number of special measures to protect their structures from fuel.

Classification of liquid gas is most often based on saturated vapor pressure or, more simply put, boiling point at normal pressure.

High-boiling components of liquid fuel.

Such liquid rocket engines can be classified as multi-fuel.
A liquid-propellant rocket engine using three-component fuel (fluorine+hydrogen+lithium) was developed in.

Binary fuels consist of an oxidizer and a fuel.
Liquid propellant engine Bristol Siddeley BSSt.1 Stentor: two-component liquid propellant engine (H2O2 + kerosene)

Oxidizing agents

Oxygen

Chemical formula-O 2 (dioxygen, American designation Oxygen-OX).
Liquid-propellant rocket engines use liquid oxygen, not gaseous oxygen - Liquid oxygen (LOX - briefly and everything is clear).
Molecular weight (for a molecule) is 32 g/mol. For lovers of precision: atomic mass (molar mass) = 15.99903;
Density=1.141 g/cm³
Boiling point=90.188K (−182.96°C)

From a chemical point of view, it is an ideal oxidizing agent. It was used in the FAA's first ballistic missiles and its American and Soviet copies. But its boiling point did not suit the military. The required operating temperature range is from –55°C to +55°C (long preparation time for launch, short time spent on combat duty).

Very low corrosiveness. Production has been mastered for a long time, the cost is low: less than $0.1 (in my opinion, several times cheaper than a liter of milk).
Flaws:

Cryogenic - cooling and constant refueling are required to compensate for losses before launch. It can also spoil other TCs (kerosene):

In the photo: the doors of the protective devices of the kerosene refueling automatic docking station (ZU-2), 2 minutes before the end of the cyclogram when performing the operation CLOSE CHECKER did not close completely due to icing. At the same time, due to icing, the signal about the TUA leaving the launcher did not go through. The launch took place the next day.

The RB liquid oxygen filling unit was removed from the wheels and installed on the foundation.

It is difficult to use the CS and liquid rocket engine nozzle as a cooler.

"ANALYSIS OF THE EFFICIENCY OF USE OF OXYGEN AS A COOLER FOR A LIQUID ROCKET ENGINE CHAMBER" SAMOSHKIN V.M., VASYANINA P.YU., Siberian State Aerospace University named after Academician M.F. Reshetnyova

Now everyone is studying the possibility of using supercooled oxygen or oxygen in a sludge-like state, in the form of a mixture of solid and liquid phases of this component. The view will be approximately the same as this beautiful ice slush in the bay to the right of Shamora:

Sugaring will increase the overall density of the oxidizer.

An example of cooling (supercooling) of the R-9A ballistic missile: for the first time, it was decided to use supercooled liquid oxygen as an oxidizer in a rocket, which made it possible to reduce the total time to prepare the rocket for launch and increase the degree of its combat readiness.

Note: For some reason, the famous writer Dmitry Konanykhin bent over (almost “chocked”) Elon Musk for this same procedure.
Cm:

Ozone-O 3

Molecular mass=48 amu, molar mass=47.998 g/mol
The density of the liquid at -188 °C (85.2 K) is 1.59 (7) g/cm³
The density of solid ozone at −195.7 °C (77.4 K) is 1.73(2) g/cm³
Melting point −197.2(2) °C (75.9 K)

Engineers have long struggled with it, trying to use it as a high-energy and at the same time environmentally friendly oxidizer in rocket technology.

The total chemical energy released during the combustion reaction involving ozone is approximately one quarter greater than for simple oxygen (719 kcal/kg). Accordingly, Iud will be greater. Liquid ozone has a higher density than liquid oxygen (1.35 versus 1.14 g/cm³, respectively), and its boiling point is higher (−112 °C and −183 °C, respectively).

So far, an insurmountable obstacle is the chemical instability and explosiveness of liquid ozone with its decomposition into O and O2, in which a detonation wave appears moving at a speed of about 2 km/s and a destructive detonation pressure of more than 3 107 dyne/cm2 (3 MPa) develops, which makes the use of liquid ozone is impossible with the current level of technology, with the exception of the use of stable oxygen-ozone mixtures (up to 24% ozone). The advantage of such a mixture is also a higher specific impulse for hydrogen engines compared to ozone-hydrogen engines. Today, such highly efficient engines as RD-170, RD-180, RD-191, as well as accelerating vacuum engines, have reached Isp parameters close to the maximum values, and to increase the efficiency there is only one option left, related to the transition to new types of fuel .

Nitric acid-HNO3

Condition - liquid at no.
Molar mass 63.012 g/mol (doesn't matter what I use or molecular mass - it doesn't change the point)
Density=1.513 g/cm³
T. melt.=-41.59 °C, T. boil.=82.6 °C

HNO3 has a high density, low cost, is produced in large quantities, is quite stable, including at high temperatures, and is fire and explosion-proof. Its main advantage over liquid oxygen is its high boiling point, and, therefore, the ability to be stored indefinitely without any thermal insulation. The nitric acid molecule HNO 3 is an almost ideal oxidizing agent. It contains a nitrogen atom and a “half” water molecule as “ballast”, and two and a half oxygen atoms can be used to oxidize the fuel. But it was not there! Nitric acid is such an aggressive substance that it continuously reacts with itself—hydrogen atoms are split off from one acid molecule and join neighboring ones, forming fragile but extremely chemically active aggregates. Even the most resistant grades of stainless steel are slowly destroyed by concentrated nitric acid (as a result, a thick greenish “jelly”, a mixture of metal salts, is formed at the bottom of the tank). To reduce the corrosive activity, various substances began to be added to nitric acid; just 0.5% hydrofluoric acid reduces the corrosion rate of stainless steel tenfold.

To increase the shock pulse, nitrogen dioxide (NO 2) is added to the acid. The addition of nitrogen dioxide to the acid binds the water entering the oxidizer, which reduces the corrosive activity of the acid, increases the density of the solution, reaching a maximum at 14% dissolved NO 2. The Americans used this concentration for their military missiles.

We have been looking for suitable containers for nitric acid for almost 20 years. It is very difficult to select construction materials for tanks, pipes, and combustion chambers of liquid propellant rocket engines.

The oxidizer option that was chosen in the USA is with 14% nitrogen dioxide. But our rocket scientists acted differently. It was necessary to catch up with the United States at any cost, so Soviet brand oxidizers - AK-20 and AK-27 - contained 20 and 27% tetroxide.

Interesting fact: In the first Soviet rocket fighter BI-1, nitric acid and kerosene were used for flight.

The tanks and pipes had to be made of Monel metal: an alloy of nickel and copper, which became a very popular structural material among rocket scientists. Soviet rubles were almost 95% made from this alloy.

Disadvantages: tolerable "muck". Corrosive active. The specific impulse is not high enough. Currently, it is almost never used in its pure form.

Nitrogen tetroxide-AT (N 2 O 4)

Molar mass=92.011 g/mol
Density=1.443 g/cm³

Disadvantages: the same nasty thing as HNO 3, but with its own quirks. May decompose into nitric oxide. Toxic. Low specific impulse. The oxidizing agent AK-NN was and is often used. It is a mixture of nitric acid and nitric tetroxide, sometimes called "red fuming nitric acid." The numbers indicate the percentage of N 2 O 4.

These oxidizers are mainly used in military rocket engines and spacecraft rocket engines due to their properties: durability and self-ignition. Typical fuels for AT are UDMH and hydrazine.

Fluorine-F 2

Atomic mass = 18.998403163 a. e.m. (g/mol)
Molar mass of F2, 37.997 g/mol
Melting point=53.53 K (−219.70 °C)
Boiling point = 85.03 K (−188.12 °C)
Density (for liquid phase), ρ=1.5127 g/cm³

Fluorine chemistry began to develop in the 1930s, especially quickly during the Second World War of 1939-45 and after it in connection with the needs of the nuclear industry and rocket technology. The name "Fluorine" (from the Greek phthoros - destruction, death), proposed by A. Ampere in 1810, is used only in Russian; in many countries the name is accepted "fluor". It is an excellent oxidizing agent from a chemical point of view. It oxidizes oxygen, water, and practically everything. Calculations show that the maximum theoretical Isp can be obtained on the F2-Be (beryllium) pair - about 6000 m/s!

Super? Bummer, not "super"...

Safety precautions. Fluorine is toxic, its maximum permissible concentration in the air is approximately 2·10-4 mg/l, and the maximum permissible concentration with exposure for no more than 1 hour is 1.5·10-3 mg/l.

The 8D21 liquid-propellant rocket engine using the fluorine + ammonia pair gave a specific impulse at the level of 4000 m/s.
For the pair F 2 +H 2 it turns out Isp = 4020 m/s!
Trouble: HF hydrogen fluoride in the exhaust.

Starting position after launching such an “energetic engine”?
A puddle of liquid metals and other chemical and organic objects dissolved in hydrofluoric acid!
H 2 +2F=2HF, at room temperature exists in the form of a dimer H 2 F 2.

Mixes with water in any ratio to form hydrofluoric acid. And its use in rocket engines of spacecraft is not realistic due to the deadly complexity of storage and the destructive effect of combustion products.

The same applies to other liquid halogens, for example, chlorine.

A hydrogen fluorine liquid-propellant rocket engine with a thrust of 25 tons to equip both stages of the rocket accelerator was supposed to be developed in V.P. Glushko based on a spent liquid-propellant rocket engine with a thrust of 10 tons using fluoroammonia (F 2 + NH 3) fuel.

Hydrogen peroxide-H 2 O 2 .

I mentioned it above in single-component fuels.

Walter HWK 109-507: advantages in the simplicity of the rocket engine design. A striking example of such a fuel is hydrogen peroxide.

Alles: the list of more or less real oxidizing agents is complete. I focus on HCl O 4. As independent oxidizing agents based on perchloric acid, the only ones of interest are: monohydrate (H 2 O + ClO 4) - a solid crystalline substance and dihydrate (2HO + HClO 4) - a dense viscous liquid. Perchloric acid (which, due to Isp, in itself is unpromising), is of interest as an additive to oxidizers, guaranteeing the reliability of self-ignition of the fuel.

Oxidizing agents can be classified as follows:

The final (most often used) list of oxidizers in conjunction with real combustibles:

Note: if you want to convert one specific impulse option to another, you can use a simple formula: 1 m/s = 9.81 s.

Flammable

Based on their physical and chemical composition, they can be divided into several groups:

Hydrocarbon fuels.
Low molecular weight hydrocarbons.
Simple substances: atomic and molecular.

For this topic, so far only hydrogen (Hydrogenium) is of practical interest.
I will not consider Na, Mg, Al, Bi, He, Ar, N 2, Br 2, Si, Cl 2, I 2, etc. in this article.
Hydrazine fuels ("stinkers").

Wake up, sleepyheads - we have already reached alcohol (C2H5OH).

The search for the optimal fuel began with the development of liquid propellant rocket engines by enthusiasts. The first widely used fuel was ethanol), used in the first
Soviet missiles R-1, R-2, R-5 ("legacy" of the FAU-2) and on the Vergeltungswaffe-2 itself.

More precisely, a solution of 75% ethyl alcohol (ethanol, ethyl alcohol, methyl carbinol, wine alcohol or alcohol, often colloquially simply “alcohol”) - monohydric alcohol with the formula C 2 H 5 OH (empirical formula C 2 H 6 O), another option: CH 3 -CH 2 -OH
This fuel two serious shortcomings, which obviously did not suit the military: low energy performance and.

Proponents of a healthy lifestyle (alcohol phobes) tried to solve the second problem with the help of furfuryl alcohol. It is a poisonous, mobile, transparent, sometimes yellowish (to dark brown) liquid that turns red over time when exposed to air. BARBARIANS!

Chem. formula: C 4 H 3 OCH 2 OH, Rat. formula:C 5 H 6 O 2. Disgusting slurry. Not suitable for drinking.

Hydrocarbon group.

Kerosene

Conditional formula C 7.2107 H 13.2936
A flammable mixture of liquid hydrocarbons (from C 8 to C 15) with a boiling point in the range of 150-250 ° C, transparent, colorless (or slightly yellowish), slightly oily to the touch
density - from 0.78 to 0.85 g/cm³ (at a temperature of 20°C);
viscosity - from 1.2 – 4.5 mm²/s (at a temperature of 20°C);
flash point - from 28°C to 72°C;
calorific value - 43 MJ/kg.

My opinion: it is pointless to write about the exact molar mass

Kerosene is a mixture of various hydrocarbons, which is why scary fractions appear (in the chemical formula) and a “smeared” boiling point. Convenient high-boiling fuel. It has been used for a long time and successfully all over the world in engines and aviation. This is what Soyuz aircraft still fly on. Low toxicity (I strongly do not recommend drinking), stable. Still, kerosene is dangerous and harmful to health (oral consumption).
The Ministry of Health is categorically against it!
Soldier's tales: good for getting rid of nasty ones.

However, it also requires careful handling during operation:

Significant advantages: relatively inexpensive, mastered in production. The kerosene-oxygen pair is ideal for the first stage. Its specific impulse on the ground is 3283 m/s, void 3475 m/s. Flaws. Relatively low density.

American rocket kerosene Rocket Propellant-1 or Refined Petroleum-1

Kerosene has a tendency to deposit tarry deposits in the lines and cooling path, which negatively affects cooling. They focus on this bad quality of his.
Kerosene engines were most developed in the USSR.

A masterpiece of human intelligence and engineering, our “pearl” RD-170/171:

Now the term “hydrocarbon fuel” has become a more correct name for kerosene-based fuels, because from kerosene, which was burned in safe kerosene lamps by I. Lukasiewicz and J. Zech, the used UVG “went away” very much.

In fact, Roscosmos gives out misinformation:

After fuel components are pumped into its tanks - naphthyl (rocket fuel)), liquefied oxygen and hydrogen peroxide, the space transport system will weigh more than 300 tons (depending on the modification of the launch vehicle.

Low molecular weight hydrocarbons

Methane-CH4

Molar mass: 16.04 g/mol
Density gas (0 °C) 0.7168 kg/m³;
liquid (−164.6 °C) 415 kg/m³
Melting temperature=-182.49 °C
Bp = -161.58 °C

It is now considered by everyone as a promising and cheap fuel, as an alternative to kerosene and hydrogen.
Chief designer Vladimir Chvanov:

The specific impulse of an LNG engine is high, but this advantage is offset by the fact that methane fuel has a lower density, so the total energy advantage is insignificant. From a design point of view, methane is attractive. To free the engine cavities, you only need to go through an evaporation cycle - that is, the engine is more easily freed from product residues. Due to this, methane fuel is more acceptable from the point of view of creating a reusable engine and a reusable aircraft.

Inexpensive, common, stable, low-toxic. Compared to hydrogen, it has a higher boiling point, and the specific impulse paired with oxygen is higher than that of kerosene: about 3250-3300 m/s on earth. Not a bad cooler.

Flaws. Low density (half that of kerosene). In some combustion modes, it can decompose with the release of carbon in the solid phase, which can lead to a drop in momentum due to the two-phase flow and a sharp deterioration in the cooling mode in the chamber due to soot deposition on the walls of the combustion chamber. Recently, active research and development activities have been carried out in the field of its use (along with propane and natural gas), even in the direction of modifying existing gas. LRE (in particular, such work was carried out on).

Already in 2016, Roscosmos began developing a power plant using liquefied natural gas.

Or "Kinder Surpeis", as an example: American Raptor engine from Space X:

These fuels include propane and natural gas. Their main characteristics as combustibles are close (with the exception of higher density and higher boiling point) to hydrocarbons. And there are the same problems when using them.

-H 2 (Liquid: LH 2) stands out among flammables.

The molar mass of hydrogen is 2016 g/mol or approximately 2 g/mol.
Density (at no.)=0.0000899 (at 273 K (0 °C)) g/cm³
Melting point=14.01K (-259.14 °C);
Boiling point=20.28K (-252.87 °C);

From the point of view of thermodynamics, H 2 is an ideal working fluid for both the liquid propellant engine itself and the TNA turbine. An excellent coolant, both in liquid and gaseous states. The latter fact makes it possible not to be particularly afraid of the boiling of hydrogen in the cooling path and to use hydrogen gasified in this way to drive the pump.

This scheme is implemented in the Aerojet Rocketdyne RL-10 - simply a gorgeous (from an engineering point of view) engine:

Our analogue ( even better, because younger): RD-0146 (D, DM) - a gas-free liquid-propellant rocket engine developed by the Chemical Automatics Design Bureau in Voronezh.

Particularly effective with a nozzle made of Grauris material. But it doesn't fly yet

This TC provides a high specific impulse - when paired with oxygen, 3835 m/s.

This is the highest figure among those actually used. These factors determine the keen interest in this fuel. Environmentally friendly, at the “output” in contact with O 2: water (water vapor). Common, virtually unlimited supplies. Mastered in production. Non-toxic. However, there are a lot of fly in the ointment in this barrel of honey.

1. Extremely low density. Everyone has seen the huge hydrogen tanks of the Energia launch vehicle and the Space Shuttle. Due to the low density, it is applicable (as a rule) at the upper stages of the launch vehicle.

In addition, low density poses a difficult challenge for pumps: hydrogen pumps are multistage in order to provide the required mass flow without cavitating.

For the same reason it is necessary to install the so-called fuel booster pumping units (FPU) immediately behind the intake device in the tanks, in order to make life easier for the main fuel pump.

Hydrogen pumps also require a significantly higher rotation speed of the pump for optimal operation.

2. Low temperature. Cryogenic fuel. Before refueling, it is necessary to cool (and/or supercool) the tanks and the entire tract for many hours. LV tanks "Falocn 9FT" - a look from the inside:

More about "surprises":
"MATHEMATICAL MODELING OF HEAT AND MASS TRANSFER PROCESSES IN HYDROGEN SYSTEMS" N0R V.A. Gordeev V.P. Firsov, A.P. Gnevashev, E.I. Postoyuk
FSUE "GKNPTs im. M.V. Khrunicheva, KB "Salyut"; "Moscow Aviation Institute (State Technical University)

The paper describes the main mathematical models of heat and mass transfer processes in the tank and hydrogen lines of the oxygen-hydrogen upper stage 12KRB. Anomalies in the supply of hydrogen to liquid propellant engines were identified and their mathematical description was proposed. The models were tested during bench and flight tests, which made it possible to use them to predict the parameters of serial upper stages of various modifications and make the necessary technical decisions to improve pneumohydraulic systems.

3. Liquid hydrogen has some properties of a gas:

Compressibility coefficient (pv/RT) at 273.15 K: 1.0006 (0.1013 MPa), 1.0124 (2.0266 MPa), 1.0644 (10.133 MPa), 1.134 (20.266 MPa), 1.277 (40.532 MPa);
Hydrogen can be in ortho and para states. Orthohydrogen (o-H2) has a parallel (same sign) orientation of nuclear spins. Para-hydrogen (p-H2)-antiparallel.

At normal and high temperatures, H2 (normal hydrogen, n-H2) is a mixture of 75% ortho and 25% para modifications, which can mutually convert into each other (ortho-para transformation). When o-H 2 is converted to p-H 2, heat is released (1418 J/mol).

4. Hydrogen gas spreads faster than other gases in space, passes through small pores, and at high temperatures penetrates steel and other materials relatively easily. H 2g has high thermal conductivity, equal to 0.1717 W/(m*K) at 273.15 K and 1013 hPa (7.3 relative to air).

Hydrogen in its normal state at low temperatures is inactive; without heating it reacts only with F 2 and in the light with Cl 2. Hydrogen reacts more actively with nonmetals than with metals. Reacts with oxygen almost irreversibly, forming water with the release of 285.75 MJ/mol of heat;

5. Hydrogen forms hydrides with alkali and alkaline earth metals, elements of groups III, IV, V and VI of the periodic table, as well as with intermetallic compounds. Hydrogen reduces the oxides and halides of many metals to metals, and unsaturated hydrocarbons to saturated ones (see).
Hydrogen gives up its electron very easily. In solution, it is detached in the form of a proton from many compounds, causing their acidic properties. In aqueous solutions, H+ forms a hydronium ion H 3 O with a water molecule. Being part of the molecules of various compounds, hydrogen tends to form a hydrogen bond with many electronegative elements (F, O, N, C, B, Cl, S, P).

6. Fire and explosion hazard. There is no need to pickle it: everyone knows the explosive mixture.
A mixture of hydrogen and air explodes from the slightest spark in any concentration - from 5 to 95 percent.

Is the Space Shuttle Main Engine (SSME) impressive?

Now estimate its cost!
Probably, having seen this and calculating the costs (the cost of putting 1 kg of payload into orbit), legislators and those who rule the budget of the United States and NASA in particular... decided “well, screw it.”
And I understand them - the Soyuz launch vehicle is both cheaper and safer, and the use of the RD-180/181 eliminates many of the problems of American launch vehicles and significantly saves taxpayers’ money in the richest country in the world.

The best rocket engine is one that you can make/buy that will have the thrust you want (not too much or too little) and be efficient enough (specific impulse, combustion chamber pressure) to cost will not become too heavy for you. /Philip Terekhov@lozga

Hydrogen engines are the most developed in the USA.
Now we are positioned in 3-4 place in the “Hydrogen Club” (after Europe, Japan and China/India).

I will separately mention solid and metallic hydrogen.

Solid hydrogen crystallizes in a hexagonal lattice (a = 0.378 nm, c = 0.6167 nm), at the nodes of which there are H 2 molecules connected to each other by weak intermolecular forces; density 86.67 kg/m³; С° 4.618 J/(mol*K) at 13 K; dielectric. At pressures above 10,000 MPa, a phase transition is expected with the formation of a structure built from atoms and possessing metallic properties. The possibility of “metallic hydrogen” superconductivity has been theoretically predicted.

Solid hydrogen is the solid state of aggregation of hydrogen.
Melting point −259.2 °C (14.16 K).
Density 0.08667 g/cm³ (at −262 °C).
White snow-like mass, crystals of hexagonal system.

The trouble is with him. He constantly gets lost: . This is understandable: a cube of molecules is obtained: 6x6x6. Just “giant” volumes - just “refuel” the rocket right now. For some reason this reminded me. This nano-miracle has not been found for 7 years or more.

I’ll leave anameson, antimatter, and metastable helium behind the scenes for now.

State at zero - colorless liquid
Molar mass=32.05 g/mol
Density=1.01 g/cm³

For those who are confused by Iud in the dimension N*s/kg, I answer: this designation is “loved” by the military.
Newton is a derived unit, based on which it is defined as a force that changes the speed of a body weighing 1 kg by 1 m/s in 1 second in the direction of the force. Thus, 1 N = 1 kg m/s 2.
Accordingly: 1 N*s/kg =1 kg m/s 2 *s/kg=m/s.
Mastered in production.

Disadvantages: toxic, smelly.

The toxicity of hydrazine to humans has not been determined. According to calculations by S. Krop, a dangerous concentration should be considered 0.4 mg/l. Ch. Comstock and co-workers believe that the maximum permissible concentration should not exceed 0.006 mg/l. According to more recent American data, this concentration at 8-hour exposure is reduced to 0.0013 mg/l. It is important to note that the threshold for the olfactory sensation of hydrazine in humans significantly exceeds the indicated numbers and is equal to 0.014-0.030 mg/l. Significant in this regard is the fact that the characteristic odor of a number of hydrazine derivatives is felt only in the first minutes of contact with them. Subsequently, due to the adaptation of the olfactory organs, this sensation disappears, and a person, without noticing it, can remain for a long time in a contaminated atmosphere containing toxic concentrations of the said substance.

Hydrazine vapors explode under adiabatic compression. It is prone to decomposition, which, however, allows it to be used as a monopropellant for low-thrust liquid rocket engines (LPRE). Due to the development of production, it is more common in the USA.

Unsymmetrical dimethylhydrazine (UDMH)-H 2 N-N(CH 3) 2

Chem. formula: C2H8N2, Rat. formula:(CH3)2NNH2
State at zero - liquid
Molar mass=60.1 g/mol
Density=0.79±0.01 g/cm³

Has higher impulse compared to hydrazine.

Density and specific impulse with basic oxidizers are lower than kerosene with the same oxidizers. Will spontaneously ignite with nitrogen oxidizers. Mastered in production in the USSR.
More common in the USSR.
And in the jet engine of a French fighter-bomber (good video, I recommend) UDMH is used as an activating additive to traditional fuel.

Regarding hydrazine fuels.

Specific thrust is equal to the ratio of thrust to weight fuel consumption; in this case it is measured in seconds (s = N s/N = kgf s/kgf). To convert weight specific thrust into mass thrust, it must be multiplied by the acceleration of gravity (approximately equal to 9.81 m/s²)

Left behind the scenes:
Aniline, methyl-, dimethyl- and trimethylamines and CH 3 NHNH 2 -Methylhydrazine (aka monomethylhydrazine or heptyl), etc.

They are not that common. The main advantage of flammable hydrazine group is its long shelf life when using high-boiling oxidizers. Working with them is very unpleasant - toxic flammable, aggressive oxidizing agents, toxic combustion products.

We can say with a high degree of confidence that if the launch vehicle has “smelly” engines, then “before marriage” it was a combat missile (ICBM, SLBM or missile defense system - which is already a rarity). Chemistry in the service of both the army and civilians.

The only exception, perhaps, is the Ariane launch vehicle - the creation of a cooperative: Aérospatiale, Matra Marconi Space, Alenia, Spazio, DASA, etc. It suffered a similar military fate in its “girlhood”.

Almost all of the military switched to solid propellant rocket engines, as they were more convenient to use. The niche for “smelly” fuels in astronautics has narrowed to use in spacecraft propulsion systems, where long-term storage is required without special material or energy costs.
Perhaps the overview can be briefly expressed graphically:

Rocket scientists are also actively working with methane. There are no particular operational difficulties: it allows you to raise the pressure in the chamber quite well (up to 40 M Pa) and get good performance.
() and other natural gases (LNG).

I will write about other directions for improving the performance of liquid-propellant rocket engines (metallization of fuels, use of He 2, acetam, etc.) later. If there is interest.

Using the effect of free radicals is a good prospect.
Detonation combustion is an opportunity for the long-awaited jump to Mars.

Afterword:

All home (garage) manipulations with such chemical components are extremely dangerous and sometimes illegal. It is BETTER not to approach the places where they spill without a protective equipment and a gas mask:

Just like with spilled mercury: call the Ministry of Emergency Situations, they will quickly come and professionally pick up everything.

Thank you to everyone who was able to endure it all to the end.

Primary sources:
Kachur P. I., Glushko A. V. "Valentin Glushko. Designer of rocket engines and space systems", 2008.
G.G. Gahun "Design and design of liquid rocket engines", Moscow, "Mechanical Engineering", 1989.
Possibility of increasing the specific impulse of a liquid-propellant rocket engine
when adding helium to the combustion chamber S.A. Orlin MSTU named after. N.E. Bauman, Moscow
M.S. Shekhter. "Fuels and working fluids of rocket engines", Mechanical Engineering" 1976
Zavistovsky D.I. "Conversations about rocket engines."
Philip Terekhov @lozga (www.geektimes.ru).
"Types of fuel and their characteristics. Fuel is a flammable substance used to produce heat. Composition of the fuel. Combustible part - carbon C-hydrogen H-sulfur." ​​- presentation by Oksana Kaseeva
Fakas S.S. "Fundamentals of liquid propellant engines. Working fluids"
Photos and video materials were used from the sites:

http://technomag.bmstu.ru
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